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SpaceX Vehicles and Missions => SpaceX Starship Program => Topic started by: Chris Bergin on 10/03/2016 02:28 pm

Title: SpaceX Raptor engine (Super Heavy/Starship Propulsion) - General Thread 1
Post by: Chris Bergin on 10/03/2016 02:28 pm
https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/ - by Alejandro G. Belluscio.

Follows on from his previous Raptor overview two years ago:
https://forum.nasaspaceflight.com/index.php?topic=34197.0

And because this is now updated to what has been revealed, this is the continuation thread.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AndyX on 10/03/2016 02:39 pm
Fascinating read into the challenges of a full flow engine unit. Didn't realize it was that unique and that it was more unique to the west.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/03/2016 02:44 pm
That was a terrific article, many thanks for that!!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/03/2016 03:07 pm
Also, this makes the thing more intriguing. It might be a big coincidence, but a 1MN dev model with a nozzle area ratio of 150:1 might be very close/exactly what is needed for a Falcon9/FH Mvac methalox replacement.
Which is what incidentally the USAF paid for when entering a contract with SpaceX for this.
Too many coincidences?...XD
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cro-magnon gramps on 10/03/2016 03:07 pm
That was an excellent article, that even a novice like myself could follow...
one question popped up: will the Raptor be more difficult to mass produce than the present Merlin engines?

Thanks...

Gramps...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Mongo62 on 10/03/2016 03:08 pm
"Mr. Musk has since confirmed that the development engine will eventually have a nozzle with an expansion ratio of 150, the maximum possible within Earth’s atmosphere."

Is this correct? I thought the SL Raptor had an expansion ratio of around 50? This seems supported by the difference in the nozzle diameters, ~2m vs ~4m for the Vac nozzle with an expansion ratio of ~200.

On the other hand, with three times the chamber pressure of the M1D it seems reasonable that the SL expansion ratio could be three times as great as well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/03/2016 03:11 pm
"Mr. Musk has since confirmed that the development engine will eventually have a nozzle with an expansion ratio of 150, the maximum possible within Earth’s atmosphere."

This is for the 1MN dev article.

Btw...I think we can get a mass estimate for the Raptors too. We don't have any concrete info yet, though Musk has hinted that it would probably unseat the M1-D as a TWR champion. 

If we assume that to be true, it potentially gives us a max weight for the engine.

Merlin SL TWR = 183.3
Merlin Vac TWR = 198.5
Merlin weight = 470 kg

Raptor SL TWR = 183.3+
Raptor Vac TWR = 198.5+
Raptor maximum speculated Weight = (311,013 / 183.3)+(334,976/198.5) / 2 = (1696+1687)/2 = ~ 1690 kg

In other words, to beat Merlin in TWR Raptor would have to be less than 1690kg.
 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: john smith 19 on 10/03/2016 03:16 pm
An impressive article.  I did not realize the engine Musk showed on the video was a 1/3 full scale unit.

The real surprise is the speed with which this engine has been built given the very limited prior art in the West on such designs.  IIRC Aerojet regularly put them into their design proposals but I don't know if many (any?) of them got to development

I would guess they studied the SSME development history very carefully and started trying to take the engine through simulated start ups and downs much earlier in the timeline than the SSME developers were able.

An interesting question would be wheather SX were able to avoid putting an oxidation resistant coating on the O2 rich pre burner turbine blades. IIRC the Russians could not quite guarantee the blades would survive without it and it's one of the issues that have made making the RD180 in the US difficult.

For a single use engine this is not an issue but for a reusable engine it becomes a critical  inspection issue. SSME had it with their gold plating of the turbine blades to resist attack by the high temperature GH2/Steam stream from the pre burners.

Fortunately Methane is not Hydrogen so a resistant alloy should be possible but time will tell how robust the engine is.

For those worried about the size of the SL nozzle keep in mind how much above the SSME main chamber pressure Raptor is.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Norm38 on 10/03/2016 03:19 pm
In other words, to beat Merlin in TWR Raptor would have to be less than 1690kg.

If this image is close to accurate, that doesn't seem a hard target to reach.  About 4x mass to work with, and it's not 4x the size.

http://forum.nasaspaceflight.com/index.php?action=dlattach;topic=34197.0;attach=1373555;sess=20788
(Tried to quote the image, but can't quote from locked threads)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: F9man on 10/03/2016 03:44 pm
Very exciting. Can't wait to meet a raptor in person
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 03:48 pm
Also, this makes the thing more intriguing. It might be a big coincidence, but a 1MN dev model with a nozzle area ratio of 150:1 might be very close/exactly what is needed for a Falcon9/FH Mvac methalox replacement.
Which is what incidentally the USAF paid for when entering a contract with SpaceX for this.
Too many coincidences?...XD
I understand that articles are not places to speculate. But yes, now that the size is known, it is, in fact, the perfect size for a Falcon Heavy upper stage. In fact, it might enable SpaceX to make a reusable upper stage for FH. Only issue I see, is that it would seem that the ITS upper stage has 9 engines, and they would only use the inner 3 for landing. At 20% of thrust, that would be 6,67% of thrust. Using a single Raptor would mean 3 times that thrust and thus quite an hoverslam.
But in expendable mode, Dimitry could probably surprise us.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: clongton on 10/03/2016 03:49 pm
Awesome write-up. Thank you
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 03:54 pm
That was an excellent article, that even a novice like myself could follow...
one question popped up: will the Raptor be more difficult to mass produce than the present Merlin engines?

Thanks...

Gramps...
It will probably cost more to produce, since it will probably need higher tolerances and a lot more material. Which, when 3D printed, means a lot more print time. Also, things like valves, integration, certification and such will also cost more.
But if you look at the previous thread, they appear to have used the 3D printing capabilities in very exiting ways. For example, the LOX TP appears to be integrated straight over the injector. If they can arbitraty passages, they will simplify basically everything because the oxidizer rich gases only need to travel through the preburner/turbine/injector without needed connecting piping.
And the fuel TP case is also integrated to the side, but all the cooling passages also appear to be 3D printed. We will see how the production engines are, but this engine looks a lot like a Tesla, it looks like a conventional car, but the construction and internal layout are completely different.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/03/2016 04:07 pm
An impressive article.  I did not realize the engine Musk showed on the video was a 1/3 full scale unit.

The real surprise is the speed with which this engine has been built given the very limited prior art in the West on such designs.  IIRC Aerojet regularly put them into their design proposals but I don't know if many (any?) of them got to development

I would guess they studied the SSME development history very carefully and started trying to take the engine through simulated start ups and downs much earlier in the timeline than the SSME developers were able.

I think two factors are the most important ones for how for accelerating development and avoiding some SSME pitfalls:
 - CFD analysis has improved to the point that you can use it for combustion chamber simulation
 - 3D/additive printing

They are clearly aware of past engine development history and some of the pitfalls (SSME, J-2X), which helps a lot.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 10/03/2016 04:14 pm
One thing about 3-D printing the innards, I believe it limits what can be coated or left uncoated. Not sure what secret sauce is required, but previous engines of this type relied on some sort of covering to protect engine structures.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 04:27 pm
The real surprise is the speed with which this engine has been built given the very limited prior art in the West on such designs.  IIRC Aerojet regularly put them into their design proposals but I don't know if many (any?) of them got to development

I would guess they studied the SSME development history very carefully and started trying to take the engine through simulated start ups and downs much earlier in the timeline than the SSME developers were able.

An interesting question would be wheather SX were able to avoid putting an oxidation resistant coating on the O2 rich pre burner turbine blades. IIRC the Russians could not quite guarantee the blades would survive without it and it's one of the issues that have made making the RD180 in the US difficult.

For a single use engine this is not an issue but for a reusable engine it becomes a critical  inspection issue. SSME had it with their gold plating of the turbine blades to resist attack by the high temperature GH2/Steam stream from the pre burners.

Fortunately Methane is not Hydrogen so a resistant alloy should be possible but time will tell how robust the engine is.

For those worried about the size of the SL nozzle keep in mind how much above the SSME main chamber pressure Raptor is.
Well, you Aerojet's proposals were mostly for a dual expander. And they had did the fuel rich preburner of the IPD. Yet, they like the use of dual expander, where they use the Hydrogen to absorb all possible heat and then a closed Bayrton heat exchanger to transfer some of that heat to the LOX to drive the LOX turbine.

With the absorption of Rocketdyne, they had all gas-gas experience out of SpaceX. But there had been other proposals to make the SSME full flow. But NASA apparently didn't wanted to mess with their most expansive and crew rated engine.

SpaceX, definitely needed an oxidizer rich resistant coating for the preburner, turbine and injectors. But now a days, Russia, China, Ukraine, India and the US have the material technology. And the truth is that any country that have to process uranium, have to develop Fluorine resistant coatings, which are actually a lot harder than just O2 resistant.

But SpaceX had a series of critical developments. For examples, they went and developed a software that used a wavelet abstraction to be able to simulate only the boundary of the gas mixture with ns and nm detail and less demanding time slices and volume matrix for the rest of the flow. This enabled very high simulation fidelity with reasonable computing power. Then they went forward and use Stennis E2 to simulate and adjust.

But I believe that the actual breakthrough was just daring to the the full flow design. That probably made all the difference.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kansan52 on 10/03/2016 04:41 pm
Wonderful article and very informative to a lay person (like me). The article presents how difficult this engine is, what they have done to manage the development, and shows the path ahead.

Thanks!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/03/2016 05:02 pm
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 10/03/2016 05:03 pm
How was it determined that this was a 1MN 1/3 scale engine?
I didn't see it any forum posts.
Didn't see it in any an announcement.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 05:13 pm
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?
We don't know the details, but Elon said theyu usd heat exchangers. Also, expanded methane is not only hot, it is very high pressure, well past its critical point, in fact. So I guess they could use tap off, but I can't see one from the pictures and it would be quite safer to use a heat exchanger.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/03/2016 05:15 pm
One thing about 3-D printing the innards, I believe it limits what can be coated or left uncoated. Not sure what secret sauce is required, but previous engines of this type relied on some sort of covering to protect engine structures.

Matthew

I would think they would make use of a Mondaloy (or similar) oxidation resistant material instead of (or in conjunction with) coatings.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ellindsey on 10/03/2016 05:16 pm
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?

From looking at the engine, it appears that the methane is tapped right after it comes out of the regenerative cooling circuit of the main combustion chamber and nozzle.  Only the oxygen feed has a separate heat exchanger for pressurization gas heating.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 05:35 pm
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?

From looking at the engine, it appears that the methane is tapped right after it comes out of the regenerative cooling circuit of the main combustion chamber and nozzle.  Only the oxygen feed has a separate heat exchanger for pressurization gas heating.

I see more a tap for the LOX preburner. It is not quite clear now what's the exact schematic.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/03/2016 05:55 pm
The 1MN dev. model of Raptor should be mass produced to replace Merlin to do away with the He system on F9 and FH.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/03/2016 05:58 pm
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?

From looking at the engine, it appears that the methane is tapped right after it comes out of the regenerative cooling circuit of the main combustion chamber and nozzle.  Only the oxygen feed has a separate heat exchanger for pressurization gas heating.

I see more a tap for the LOX preburner. It is not quite clear now what's the exact schematic.

Speculation...
The heat exchanger is 3D printed into the pump housing between the pump output and the preburner inlet...
The hot gases going to tank pressurization would be cooled by the cold fluids chilling the housing...
Would have to see a print of the housing to know it's there...  ;)
It's amazing what 3D printing lets you do...  8)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MAC74 on 10/03/2016 06:01 pm
One thing about 3-D printing the innards, I believe it limits what can be coated or left uncoated. Not sure what secret sauce is required, but previous engines of this type relied on some sort of covering to protect engine structures.

Matthew

My guess is that they are 3D printing or casting the parts that are exposed to oxygen rich hot gas from Mondaloy 200.  The parts that are on the fuel rich side will probably be Inconel.  Mondaloy is the new US equivalent to the exotic Russian metallurgy.  It is a zinc rich superalloy that can resist high temperature oxidation without a protective coating.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 10/03/2016 06:08 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/03/2016 06:23 pm
I think that Musk will be doing and AMA this week or the next. It would be pretty cool to get some more answers about Raptor, especially after the added info we got from this great article.

1. Was the test firing using the full engines' powerpack, or was it only a chamber test?
2. Was TEA-TEB used, or a spark igniter (the video I think is inconclusive on that)?
3. Will this dev article reach during development the high pressures intended for the ITS Raptor?
4. Will the end of development for this 1MN variant involve an acceptance test at Stennis (as per the USAF contract)?

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 10/03/2016 07:05 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/

I have been wondering about this. Since SpaceX keeps so much of the details of its tech secret, what other than honor stops them from copying all sorts of proprietary things.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/03/2016 07:23 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/

I have been wondering about this. Since SpaceX keeps so much of the details of its tech secret, what other than honor stops them from copying all sorts of proprietary things.

Matthew

The reason is ITAR why SpaceX have to keep details of it's tech. including Raptor secret.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 07:49 pm
I think that Musk will be doing and AMA this week or the next. It would be pretty cool to get some more answers about Raptor, especially after the added info we got from this great article.

1. Was the test firing using the full engines' powerpack, or was it only a chamber test?
2. Was TEA-TEB used, or a spark igniter (the video I think is inconclusive on that)?
3. Will this dev article reach during development the high pressures intended for the ITS Raptor?
4. Will the end of development for this 1MN variant involve an acceptance test at Stennis (as per the USAF contract)?

1) It was a complete rocket, it included a 27MW turbo machinery. It's in the article.
2) I don't know if it included the spark ignition. Somebody should include that question in the AMA.
3) I would guess that it has the capability of reaching full Pc, because 27MW is more MW/kN of any non hydrogen rocket.
4) I think it is a possibility. I don't have information but I would be surprised if two things were not true:
a) this won't be the only demonstrator.
b) this prototype or the next one isn't used to complete the USAF contract.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/03/2016 07:51 pm
Nice article, Baldusi (by the way)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mheney on 10/03/2016 08:37 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/

I have been wondering about this. Since SpaceX keeps so much of the details of its tech secret, what other than honor stops them from copying all sorts of proprietary things.

Matthew



Lawsuits.  People move around, and you couldn't keep stealing other people's work secret for long.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/03/2016 08:42 pm
Mr. Belluscio, a very nice article – thank you.

One note of correction: the 361s ISP you cite for the first stage's Raptors in vacuum is actually the sea level value for the three inner Raptors of the second stage. See pp. 36 of SpaceX's published PDF (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf). 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MAC74 on 10/03/2016 08:49 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/

Mondaloy is an Air Force Research Laboratory program.  It says right on the program that the information is to be shared with the entire US Rocket Community.  Here are the exact words.

"The improved knowledge base, test results, and lessons learned in the HCB program and other BPTM activities are shared with the entire U.S. rocket propulsion community."
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Science on 10/03/2016 08:50 pm
Great work on the article Alejandro, thank you! :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 09:09 pm
Mr. Belluscio, a very nice article – thank you.

One note of correction: the 361s ISP you cite for the first stage's Raptors in vacuum is actually the sea level value for the three inner Raptors of the second stage. See pp. 36 of SpaceX's published PDF (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).
I believe that you are misreading the information. Vacuum optimized nozzle can't be used at sea level since they would get into flow separation issues. When they say Sea Level and Vacuum they refer to the two different Raptor versions.
There is no way you can get 361 seconds of isp with methane/LOX at sea level. Best I could get was 355 theoretical, without losses, and that was with a Pc of 70MPa. At 30MPa you can't get past 337s.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/03/2016 09:24 pm
Mr. Belluscio, a very nice article – thank you.

One note of correction: the 361s ISP you cite for the first stage's Raptors in vacuum is actually the sea level value for the three inner Raptors of the second stage. See pp. 36 of SpaceX's published PDF (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).
I believe that you are misreading the information. Vacuum optimized nozzle can't be used at sea level since they would get into flow separation issues. When they say Sea Level and Vacuum they refer to the two different Raptor versions.
There is no way you can get 361 seconds of isp with methane/LOX at sea level. Best I could get was 355 theoretical, without losses, and that was with a Pc of 70MPa. At 30MPa you can't get past 337s.

Understood. I'm simply pointing out the information SpaceX has and has not provided us with.

For the first stage SX provides thrust value only, not ISP. On the second stage they provide vacuum thrust only, then separate sea-level and vacuum ISP values.

The 361s value is interesting. Perhaps the second stage's three inner Raptors are configured differently than those on the first stage given they are used for Earth landing but not Earth lift-off.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MATTBLAK on 10/03/2016 09:27 pm
Has a version of the Merlin ever seriously been considered that runs on LOX/CH4? Even without all the full flow, staged combustion features of the Raptor; with subcooled propellants, what kind of performance could be squeezed out of them?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeAtkinson on 10/03/2016 09:42 pm
Mr. Belluscio, a very nice article – thank you.

One note of correction: the 361s ISP you cite for the first stage's Raptors in vacuum is actually the sea level value for the three inner Raptors of the second stage. See pp. 36 of SpaceX's published PDF (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).

It says

Raptor Engines
   3 Sea-Level - 361 Isp
   6 Vacuum - 382 Isp

Meaning 3 Sea-Level engines and 6 Vacuum engines, with Isp 361 and 382 seconds in vacuum respectively.

It is easy to see that they mean the vacuum Isp for the Sea-Level engines as page 31 gives the sea-level Isp as 334 and the main use of the Sea-Level engines in the Ship will be for Earth ascent, Mars landing and Mars descent all of which are in near vacuum.

Edit: the Ship total thrust of 31 MN allows us to estimate the Raptor (SL) thrust in vacuum. As

(31- 6 x 3.5) / 3 = 3.33 MN
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/03/2016 09:53 pm
Mr. Belluscio, a very nice article – thank you.

One note of correction: the 361s ISP you cite for the first stage's Raptors in vacuum is actually the sea level value for the three inner Raptors of the second stage. See pp. 36 of SpaceX's published PDF (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).

It says

Raptor Engines
   3 Sea-Level - 361 Isp
   6 Vacuum - 382 Isp

Meaning 3 Sea-Level engines and 6 Vacuum engines, with Isp 361 and 382 seconds in vacuum respectively.

It is easy to see that they mean the vacuum Isp for the Sea-Level engines as page 31 gives the sea-level Isp as 334 and the main use of the Sea-Level engines in the Ship will be for Earth ascent, Mars landing and Mars descent all of which are in near vacuum.

That seems a stretch of interpretation to me. If you state 'Sea-Level' and follow with an ISP value then... what might one suppose you are trying to communicate?

Is it possible that, for the three inner Raptors of the second stage, they have a third variant? After all, these engines need never fight Earth's gravity when velocity=0.

EDIT:
An exercise: go to the PDF (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf) and measure nozzle lengths. I'm working from the ITS cutaway view on page 26, and find the Raptors' nozzles on the first stage to be approximately 80% the length of those on second stage's inner three engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/03/2016 10:04 pm
If you are not convinced, do 138MN/128MN*334 seconds=360.3seconds. Given the rounding on the MN, it is totally consistent with the 361s vacuum performance for Sea Level optimized Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: SirKeplan on 10/03/2016 10:09 pm
That seems a stretch of interpretation to me. If you state 'Sea-Level' and follow with an ISP value then... what might one suppose you are trying to communicate?

Is it possible that, for the three inner Raptors of the second stage, they have a third variant? After all, these engines need never fight Earth's gravity when velocity=0.
I can see where the confusion comes in, but if you compare with page 31 you see ISP is given as vacuum ISP, unless qualified with "(SL)"

on page 34 for the Spaceship it only makes sense to quote vacuum ISPs. for the sea level optimised engine we already know it's ISP at sea level, as it was stated earlier.


However, it is entirely possible the Sea-Level Raptors on the second stage are slightly different to on the first stage. the second stage does not have the same space constraints as the booster, and indeed if you measure the pixel sizes the second stage has wider nozzles in the images. this would allow the engine expansion to be slightly more optimal than if it used booster engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nilof on 10/03/2016 10:19 pm
The wikipedia edits are getting annoying. A few days ago I saw 382 indicated as the vaccum ISP of the ITS first stage and corrected it to ~360s . Apparently some confused soul changed it back to 382 seconds, looking back at the edit history I saw an edit war between a few other editors between the two values, and then at some point the vaccum isp was deleted outright.

The wikipedia article on the ITS seems to be Encyclopedia Astronautica-tier unreliable right now.

It would be so much nicer if anyone who edited rocket engine ISP's on any wiki was forced to sanity test said ISP's in RPA before making the edits...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/03/2016 10:23 pm
If you are not convinced, do 138MN/128MN*334 seconds=360.3seconds. Given the rounding on the MN, it is totally consistent with the 361s vacuum performance for Sea Level optimized Raptor.

(138*334)/128... yes, that is convincing.

When would we expect to see those three engines firing in a vacuum?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/03/2016 10:26 pm
The wikipedia edits are getting annoying. A few days ago I saw 382 indicated as the vaccum ISP of the ITS first stage and corrected it to ~360s . Apparently some confused soul changed it back to 382 seconds, looking back at the edit history I saw an edit war between a few other editors between the two values, and then at some point the vaccum isp was deleted outright.

The wikipedia article on the ITS seems to be Encyclopedia Astronautica-tier unreliable right now.

It would be so much nicer if anyone who edited rocket engine ISP's on any wiki was forced to sanity test said ISP's in RPA before making the edits...

SpaceX have not provided a formal ISP value for the first stage Raptors in vacuum, though Baldusi's math seems fair enough.

And yes, Wikipedia changes. Tragic, isn't it?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/03/2016 10:38 pm
If you are not convinced, do 138MN/128MN*334 seconds=360.3seconds. Given the rounding on the MN, it is totally consistent with the 361s vacuum performance for Sea Level optimized Raptor.

(138*334)/128... yes, that is convincing.

When would we expect to see those three engines firing in a vacuum?

After staging from the ITS booster, when climbing to LEO. (see the video) Also the martian atmosphere is practically a vacuum.  :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: kch on 10/03/2016 10:44 pm
The wikipedia edits are getting annoying. A few days ago I saw 382 indicated as the vaccum ISP of the ITS first stage and corrected it to ~360s . Apparently some confused soul changed it back to 382 seconds, looking back at the edit history I saw an edit war between a few other editors between the two values, and then at some point the vaccum isp was deleted outright.

The wikipedia article on the ITS seems to be Encyclopedia Astronautica-tier unreliable right now.

It would be so much nicer if anyone who edited rocket engine ISP's on any wiki was forced to sanity test said ISP's in RPA before making the edits...

SpaceX have not provided a formal ISP value for the first stage Raptors in vacuum, though Baldusi's math seems fair enough.

And yes, Wikipedia changes. Tragic, isn't it?

More amusing than tragic, though it does make it not-much-of-a-source as regards accurate information.  Useful mostly for the links to other sites.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/03/2016 10:51 pm
If you are not convinced, do 138MN/128MN*334 seconds=360.3seconds. Given the rounding on the MN, it is totally consistent with the 361s vacuum performance for Sea Level optimized Raptor.

(138*334)/128... yes, that is convincing.

When would we expect to see those three engines firing in a vacuum?

After staging from the ITS booster, when climbing to LEO. (see the video) Also the martian atmosphere is practically a vacuum.  :)

Thank you! You're right, they're all firing at that point. It's on Mars departure when we see only the outside engines firing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Toast on 10/03/2016 10:53 pm
The 1MN dev. model of Raptor should be mass produced to replace Merlin to do away with the He system on F9 and FH.

That would be a massive change, a lot of the Falcon 9 design would have to go back to the drawing board. Plus, the Merlin is an extremely reliable engine, they've only had one failure out of almost three hundred engines that have launched. The helium system is problematic, but fixable. On the other hand, Raptor is a cutting-edge engine that's not fully developed yet, and that has unknown reliability. Switching to it now would result in an extremely protracted return to flight period, and might not improve reliability overall.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wardy89 on 10/03/2016 11:05 pm
This might be a stupid question but that does 1MN mean? some people have said that makes it about 1/3 size i would just like to understand the scaling ect.

Edit: please ignore this i have since answered my own question! MN=Meganewton which is 1000 Kilonewtons so roughly 1/3 thrust!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AS-503 on 10/03/2016 11:23 pm
This might be a stupid question but that does 1MN mean? some people have said that makes it about 1/3 size i would just like to understand the scaling ect.

It means 1 Mega Newtons. Or 1,000,000 Newtons. Or 1,000,000 X 0.224 pounds (224,000 pounds of thrust).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 10/04/2016 01:01 am
SpaceX have not provided a formal ISP value for the first stage Raptors in vacuum, though Baldusi's math seems fair enough.
http://www.spacex.com/sites/spacex/files/mars_presentation.pdf
Page 36 gives the vacuum Isp for the SL Raptors.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/04/2016 01:15 am
SpaceX have not provided a formal ISP value for the first stage Raptors in vacuum, though Baldusi's math seems fair enough.
http://www.spacex.com/sites/spacex/files/mars_presentation.pdf
Page 36 gives the vacuum Isp for the SL Raptors.

Actually, that page purports to give the Isp for three sea level Raptors, then the Isp for six vacuum Raptors, all of which belong to the second stage. What exactly this means is the discussion at hand.

Moreover, it appears none of these Raptors (on the second stage) are the same as those on the first – smaller nozzles all around on stage one.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/04/2016 01:20 am
SpaceX have not provided a formal ISP value for the first stage Raptors in vacuum, though Baldusi's math seems fair enough.
http://www.spacex.com/sites/spacex/files/mars_presentation.pdf
Page 36 gives the vacuum Isp for the SL Raptors.

Actually, that page purports to give the Isp for three sea level Raptors, then the Isp for six vacuum Raptors, all of which belong to the second stage. What exactly this means is the discussion at hand.

Moreover, it appears none of these Raptors (on the second stage) are the same as those on the first – smaller nozzles all around on stage one.
Honestly, the discussion is silly. Try running RPA Lite, and the only way to make any sense of what was given is the simplest explanation:
382s is for vac-optimized Raptor at vacuum.
~360s is for sl-optimized Raptor at vacuum.
332s is for sl-optimized Raptor at sea level.

Let's not over-complicate it because the diagram may show slight /apparent differences in nozzle size. Occam's Razor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/04/2016 01:28 am
SpaceX have not provided a formal ISP value for the first stage Raptors in vacuum, though Baldusi's math seems fair enough.
http://www.spacex.com/sites/spacex/files/mars_presentation.pdf
Page 36 gives the vacuum Isp for the SL Raptors.

Actually, that page purports to give the Isp for three sea level Raptors, then the Isp for six vacuum Raptors, all of which belong to the second stage. What exactly this means is the discussion at hand.

Moreover, it appears none of these Raptors (on the second stage) are the same as those on the first – smaller nozzles all around on stage one.
Honestly, the discussion is silly. Try running RPA Lite, and the only way to make any sense of what was given is the simplest explanation:
382s is for vac-optimized Raptor at vacuum.
~360s is for sl-optimized Raptor at vacuum.
332s is for sl-optimized Raptor at sea level.

Let's not over-complicate it because the diagram may show slight /apparent differences in nozzle size. Occam's Razor.

CAD files, according to Musk... > 'a diagram'.
We're working with what we've been given.
Goodness knows many on this board have worked with less.

I don't care about the first stage vacuum Isp value; Baldusi convinced me on that.
But SpaceX have shown us three different nozzle sizes, a detail I hope you'll agree is relevant here.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/04/2016 01:30 am
Here is how I view this.

1. There is one Raptor engine.
2. It has three different nozzles. 40:1, 50:1 and 200:1
3. The smallest 40:1 nozzle is for booster engines (so as to fit). The SL Isp is 334s and the Vac Isp is unknown (around 360s would be a good guess).
4. The 50:1 nozzle is for the spaceship/tanker landing engines. The Vac Isp is 361s, and the SL Isp is unknown (around 335s would be a good bet).
5. The 200:1 nozzle is for the spaceship/tanker vacuum engines. The Vac Isp is 382s and the SL Isp (if those engines are used for abort) is unknown.
6. The CAD Raptor image that SpaceX gave us was for the booster 40:1 sea level Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 10/04/2016 01:47 am
382s is for vac-optimized Raptor at vacuum.
~360s is for sl-optimized Raptor at vacuum.
332s is for sl-optimized Raptor at sea level.
I agree! Giving anything but the vacuum Isp for a second stage engine makes no sense.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/04/2016 01:53 am
382s is for vac-optimized Raptor at vacuum.
~360s is for sl-optimized Raptor at vacuum.
332s is for sl-optimized Raptor at sea level.
I agree! Giving anything but the vacuum Isp for a second stage engine makes no sense.

...correct. Except this second stage returns to and lands on Earth.  :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 10/04/2016 03:01 am
Let me add my congratulations and thanks for a great article. Very educational for an engine tech novice like me!

How was it determined that this was a 1MN 1/3 scale engine?
I didn't see it any forum posts.
Didn't see it in any an announcement.

I was wondering about this too and haven't seen any posts (including in L2), although the forum has been a bit busy of late!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 10/04/2016 03:35 am
...correct. Except this second stage returns to and lands on Earth.  :)
And the landing burn which lasts a few seconds is the only time you have a significant burn time in dense atmosphere. Dont think the Isp is that important for that one.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/04/2016 05:49 am
I redid my Raptor engine model with MR = 3.8. Didn't change much. I also compared it with the Raptor CAD drawing to try and get a scale on it. It appears that the drawing was a 40:1 booster engine. dia ~ 1.7 m, ht ~ 3.07 m. For the vacuum engine: dia ~ 3.79 m, ht ~6.2 m.


Raptor engine model corrections and sized to ~3.5 MN VAC:

Common:
    - Chamber Pressure = 296 atmospheres (4350 psi, 30 MPa, 300 bar)
    - Mixture Ratio = 3.8
    - Diameter Throat  = .268 m
Vacuum Engine:
    - Expansion Ratio = 200
    - Isp vacuum = 382
    - Thrust Vac = 3.5 MN
    - Diameter Exit = 3.79 m
Booster Engine:
    - Expansion Ratio = 40  (I believe this is constrained by the booster base area, it should be a little higher)
    - Isp Vac = 359
    - Thrust Vac = 3.28 MN
    - Isp SL = 334
    - Thrust SL  = 3.06 MN
    - Diameter Exit = 1.7 m

OK I resized properly 

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/04/2016 06:10 am
...correct. Except this second stage returns to and lands on Earth.  :)
And the landing burn which lasts a few seconds is the only time you have a significant burn time in dense atmosphere. Dont think the Isp is that important for that one.

It's not so much about isp. It's about stability and reliability. Flow separation can have really nasty effects.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 10/04/2016 06:18 am
...correct. Except this second stage returns to and lands on Earth.  :)
And the landing burn which lasts a few seconds is the only time you have a significant burn time in dense atmosphere. Dont think the Isp is that important for that one.

It's not so much about isp. It's about stability and reliability. Flow separation can have really nasty effects.

In the octoweb arrangement the engine bells of 3 Merlins stick out a bit further than the rest. IIRC the engines themselves are identical, it's their mounting that is offset. I recall some speculation at the time, but diid we ever learn the definitive purpose for this?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ArbitraryConstant on 10/04/2016 07:50 am
I think that Musk will be doing and AMA this week or the next. It would be pretty cool to get some more answers about Raptor, especially after the added info we got from this great article.

1. Was the test firing using the full engines' powerpack, or was it only a chamber test?
2. Was TEA-TEB used, or a spark igniter (the video I think is inconclusive on that)?
3. Will this dev article reach during development the high pressures intended for the ITS Raptor?
4. Will the end of development for this 1MN variant involve an acceptance test at Stennis (as per the USAF contract)?

1) It was a complete rocket, it included a 27MW turbo machinery. It's in the article.
2) I don't know if it included the spark ignition. Somebody should include that question in the AMA.
3) I would guess that it has the capability of reaching full Pc, because 27MW is more MW/kN of any non hydrogen rocket.
4) I think it is a possibility. I don't have information but I would be surprised if two things were not true:
a) this won't be the only demonstrator.
b) this prototype or the next one isn't used to complete the USAF contract.
Am I reading this right? This sounds like it couldn't possibly be more perfect for an enhanced upper stage for Falcon 9.


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomic on 10/04/2016 10:19 am
Great article.

There's (understandably) very little information on the materials actually used in oxygen rich preburners, mondaoly is one of the better sources. Lpre.de suggests the RD-253 uses zirconium thermal barrier coatings used on , NK-33 used ceramic coatings, while the RD-170 series supposedly use multiple layers (ceramic over zirconium over nickel based material?) and some film cooling by cold LOX.

However with one of the big advantages of the FFSC cycle is the lower turbine inlet temp for a given chamber pressure, so might not need such extreme measures. 

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 10/04/2016 11:35 am
So the engine tested so far is sub-scale after all- news to me (perhaps not to those on L2).
At first this is a little disappointing. But on the up side, it opens up the possibility of a production version which would be a very useful engine indeed.

Do we have any indication that the 1MN scale engine will be taken all the way to a flight-ready production version? I would presume that a demonstrator can be built extremely conservatively, especially around mass requirements, just to prove the concept of the cycle and materials etc.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/04/2016 11:41 am
Has a version of the Merlin ever seriously been considered that runs on LOX/CH4? Even without all the full flow, staged combustion features of the Raptor; with subcooled propellants, what kind of performance could be squeezed out of them?

Something like 15-20 second(<10%) increase in isp over Merlin, but T/W would be worse due methane needing bigger pipes and bigger pumps.

Would require redesigning too many parts of the engine, that not worth doing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/04/2016 11:52 am
Here is a strictly hypothetical question.

Assuming this 1,000 kN demonstrator reaches a 30 MPa operating chamber pressure, how big/wide would a 50:1 ratio nozzle be for it? Moreover, what would be the most effective/efficient nozzle ratio that it could have, assuming it is used for first stage propulsion (among 8 other engines) and slow/low S1 separation for RTLS duties?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Silversheep2011 on 10/04/2016 12:20 pm
Question: Does the placement of the 3 sea level raptors play a further  important role by being at the conical base of the spaceship and by being in  the center section of the 6 vacuum rated Raptors on that are on the outer edge  rim  [presumably with somewhat lower exhaust pressures and exhaust velocities]

Or put another way, is there some  hidden benefits for example based in the same way the principle of an Aerospike engine works in transitioning atmospheric to vacuum environments?

https://www.youtube.com/watch?v=EWf4iOMSPNc
see 1:37 to 2:31 that makes the S.L. raptors that little bit more efficient in the vacuum of outer space?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/04/2016 12:27 pm
I don't think there are any "hidden" benefits. The SL Raptors in the spaceship and tanker will be mainly used for retro-propulsion and landing. It wouldn't make much sense to use them for vacuum propulsion (other than possibly as part of the S2 ascent), since the proper Vacuum engines are a lot more efficient.

One possible benefit I can think of for the arrangement is clearing up debris and reducing blowback when landing on unprepared Mars surfaces, if you have each SL raptor gimbaling towards the corresponding leg during the final stages of landing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/04/2016 02:08 pm
Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/04/2016 02:48 pm
Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.

Taken the liberty to arrange them with the throat as the common line. That way I think we can get a better comparative look on the powerpack, chamber and nozzle respective sizes.

btw..if you do have cad drawings like these for other engines, I would love to put them in too...;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/04/2016 03:08 pm
Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.

Taken the liberty to arrange them with the throat as the common line. That way I think we can get a better comparative look on the powerpack, chamber and nozzle respective sizes.

btw..if you do have cad drawings like these for other engines, I would love to put them in too...;)

One of the most interesting aspects from the CAD, at least from my perspective, is to see how much piping and volume is saved by the way Raptor integrates the LOX turbopump, preburner and straight to the injector. And also, how the higher pressure does means smaller pipings for the gaseous methane. Just look at the turbine outlet to the fuel ring around the LOX TP. Just look at the size of the turbine outlet as it goes straight to the fuel dome.
Look at the huge pipe from the BE-4 turbine outlet, how it has to make a U-turn, go all the way up from below the throat, and make a second U-turn. Raptor gets getting prettier the more I look at it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/04/2016 05:48 pm
Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.

Taken the liberty to arrange them with the throat as the common line. That way I think we can get a better comparative look on the powerpack, chamber and nozzle respective sizes.

btw..if you do have cad drawings like these for other engines, I would love to put them in too...;)

One of the most interesting aspects from the CAD, at least from my perspective, is to see how much piping and volume is saved by the way Raptor integrates the LOX turbopump, preburner and straight to the injector. And also, how the higher pressure does means smaller pipings for the gaseous methane. Just look at the turbine outlet to the fuel ring around the LOX TP.
Look at the huge pipe from the BE-4 turbine outlet, how it has to make a U-turn, go all the way up from below the throat, and make a second U-turn. Raptor gets getting prettier the more I look at it.

The fuel turbine outlet does not go to the fuel ring around the LOX TP. That is liquid CH4 coming out of the regen exhaust. It is also only a small portion of the total CH4 flow. Only enough to gasify the LOX sufficient to power its pump. The majority of the CH4 goes into its preburner and exits perpendicular to the preburner straight into the main chamber in what I believe is a short wide shallow duct shaped to match the depth of the fuel injector gallery below the Lox preburner's turbine. See my labled CAD drawing.

The Raptors ducting still looks too small to me.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 10/04/2016 06:02 pm
You are right, this happens when I write from memory instead of actually looking at the image again. And it still looks amazingly small to me, too.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/04/2016 06:52 pm
Examples of a 3D metal printing and 5-axis machining center in action...

I found these helped me understand how a complex thing like SpaceX Raptor can be made...  8)

https://www.youtube.com/watch?v=g8sT8ESfjrg

https://www.youtube.com/watch?v=Fr_PneeyO34

On edit... another example...
In short... by laying up some metal... then shaping it... then laying up more... back and forth...
Working from the combustion chamber out... making features in layers and shells of sorts...
You could make a very complex part with many features and passages buried in the metal...  :o  8)

https://www.youtube.com/watch?v=oaIOrQi2HLM
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: john smith 19 on 10/04/2016 11:45 pm
Quote
From looking at the engine, it appears that the methane is tapped right after it comes out of the regenerative cooling circuit of the main combustion chamber and nozzle.  Only the oxygen feed has a separate heat exchanger for pressurization gas heating.
Logical. Getting a supply of warm (hot?) fuel is rarely a problem in regeneratively cooled engines but getting the same for the oxidizer is more complex.

Note the size of the LOX HX is not that big. IIRC the SSME LOX HX was basically a half turn pipe around the the main combustion chamber. Given the Raptors higher chamber pressure I'd guess it runs a hotter chamber as well.

Obviously both gas streams will cool down a bit on their way to the tank outlets but I strongly doubt either pipe is insulated, except on the tank side, to stop boiling the tank contents.

Great article.

There's (understandably) very little information on the materials actually used in oxygen rich preburners, mondaoly is one of the better sources. Lpre.de suggests the RD-253 uses zirconium thermal barrier coatings used on , NK-33 used ceramic coatings, while the RD-170 series supposedly use multiple layers (ceramic over zirconium over nickel based material?) and some film cooling by cold LOX.

However with one of the big advantages of the FFSC cycle is the lower turbine inlet temp for a given chamber pressure, so might not need such extreme measures.
My impression is the Russians were much less inclined to treat rocket engines as "special" relative to jet engines and were quite OK with adapting jet engine practice to rocket engines.

Engine mfg have been depositing 2 layer "thermal barrier coatings" on turbine blades for decades. The inner layer is a thermal expansion matching layer while the outer is normally a metal oxide to handle high temperatures.

The issue remains that once you start relying on such coatings to deliver the necessary performance their integrity becomes critical to functioning.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 10/05/2016 12:03 am
...correct. Except this second stage returns to and lands on Earth.  :)
And the landing burn which lasts a few seconds is the only time you have a significant burn time in dense atmosphere. Dont think the Isp is that important for that one.

It's not so much about isp. It's about stability and reliability. Flow separation can have really nasty effects.
Which was not the topic of the discussion. My point was that it makes no sense to list anything but the vacuum Isp for a second stage, (even for the sealevel engines) because the sea level Isp is completely irrelevant except for a few seconds during landing. Clear now?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nathan2go on 10/05/2016 02:35 am
Has a version of the Merlin ever seriously been considered that runs on LOX/CH4? Even without all the full flow, staged combustion features of the Raptor; with subcooled propellants, what kind of performance could be squeezed out of them?

Something like 15-20 second(<10%) increase in isp over Merlin, but T/W would be worse due methane needing bigger pipes and bigger pumps.

Would require redesigning too many parts of the engine, that not worth doing.
Well, the Airforce is paying for 1/3rd of the development cost, so they apparently hope it will be used to carry their payloads.

That 10% boost in Isp (348->382 sec) on the F9 second stage will give a 23% boost in LEO payload, and a 64% boost for GTO payloads (assuming the wet&dry weights are the same, according to my calculations).  This would let the F9 match the Atlas 551, even with booster RTLS. 

For the first stage though, switching to a methalox engine would not have as big a benefit: if the tank volume stays the same, the lower fuel density (therefore lower gross weight) will offset some of the Isp advantage.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TrueBlueWitt on 10/05/2016 02:42 am
Has a version of the Merlin ever seriously been considered that runs on LOX/CH4? Even without all the full flow, staged combustion features of the Raptor; with subcooled propellants, what kind of performance could be squeezed out of them?

Something like 15-20 second(<10%) increase in isp over Merlin, but T/W would be worse due methane needing bigger pipes and bigger pumps.

Would require redesigning too many parts of the engine, that not worth doing.
Well, the Airforce is paying for 1/3rd of the development cost, so they apparently hope it will be used to carry their payloads.

That 10% boost in Isp (348->382 sec) on the second stage will give a 23% boost in LEO payload, and a 64% boost for GTO payloads (assuming the wet&dry weights are the same, according to my calculations).  This would let the F9 match the Atlas 551, even with booster RTLS. 

For the first stage though, if the tank volume stays the same, the lower fuel density (therefore lower gross weight) will offset some of the Isp advantage.

I'm thinking to do this optimally you'd readjust stage lengths..
Keep S1 KeroLox and go back to shorter tank. Makes RTLS easier. Then stretch the 1MN Raptor S2.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/05/2016 06:29 am
We are getting a little off topic here, but being able to stretch the S2 (instead of making it wider) will have three more advantages.

1. Road-transportability with the same hardware.
2. No need to change your tooling for the tanks.
3. No need to develop Dragon/Dragon2 stage adapters, payload adapters and fairings.

It could work. Changing the GSE though, as well as the engine for the stage is not going to be cheap (helium system, thrusters etc). Same goes for changing the S1 length. 

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/05/2016 12:39 pm
Here is a strictly hypothetical question.

Assuming this 1,000 kN demonstrator reaches a 30 MPa operating chamber pressure, how big/wide would a 50:1 ratio nozzle be for it? Moreover, what would be the most effective/efficient nozzle ratio that it could have, assuming it is used for first stage propulsion (among 8 other engines) and slow/low S1 separation for RTLS duties?

For a booster engine 50:1 would be about right. Cycle is the same so the nozzle scales with area so its diameter scales with the square root of the thrust ratio:
                                          40:1 diam = 1.7 x sqrt(1 / 3.05)         = . 97 m  (~38 inches).
                                          50:1 diam = 1.7 x sqrt(50 / 40 /3.05) = 1.09 m (~ 43 inches)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/06/2016 04:11 pm
I updated reply #61 on this thread to the correct my estimated size of the engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 10/07/2016 06:34 am
Here is a strictly hypothetical question.

Assuming this 1,000 kN demonstrator reaches a 30 MPa operating chamber pressure, how big/wide would a 50:1 ratio nozzle be for it? Moreover, what would be the most effective/efficient nozzle ratio that it could have, assuming it is used for first stage propulsion (among 8 other engines) and slow/low S1 separation for RTLS duties?

For a booster engine 50:1 would be about right. Cycle is the same so the nozzle scales with area so its diameter scales with the square root of the thrust ratio:
                                          40:1 diam = 1.7 x sqrt(1 / 3.05)         = . 97 m  (~38 inches).
                                          50:1 diam = 1.7 x sqrt(50 / 40 /3.05) = 1.09 n (~ 43 inches)

Many thanks for that. A couple more questions to anyone interested to answer (again, this is a hypothetical scenario).

What is the diameter of the current M9 nozzle?
If we assume that the material, width and height of the current F9 S1 remains constant, and that the common bulkhead is moved to adjust, given:

1. The known propellant ratio for the Raptor Demonstrator.
2. An SL thrust of 870kN and Vac thrust of 930kN.
3. An SL Isp of 330s and Vac Isp of 358s
4. A dry stage weight of 27 metric tons.

What would the performance delta be against the current F9 S1?

I'm not asking whether something like this is possible, probable, practicable or wanted/needed. Just want to understand the comparative difference between one engine and the other in a hypothetical scenario. I assume that the difference would be rather small, both due to having less propellant on the stage and Isp not being the most important factor in the two re-usable scenarios that F9 S1 covers (RTLS and DPL S1-S2 separations).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AP3 on 10/07/2016 07:21 pm
I redid my Raptor engine model with MR = 3.8. Didn't change much. I also compared it with the Raptor CAD drawing to try and get a scale on it. It appears that the drawing was a 40:1 booster engine. dia ~ 1.7 m, ht ~ 3.07 m. For the vacuum engine: dia ~ 3.79 m, ht ~6.2 m.

Raptor engine model corrections and sized to ~3.5 MN VAC:

Common:
    - Chamber Pressure = 296 atmospheres (4350 psi, 30 MPa, 300 bar)
    - Mixture Ratio = 3.8
    - Diameter Throat  = .268 m
Vacuum Engine:
    - Expansion Ratio = 200
    - Isp vacuum = 382
    - Thrust Vac = 3.5 MN
    - Diameter Exit = 3.79 m
Booster Engine:
    - Expansion Ratio = 40  (I believe this is constrained by the booster base area, it should be a little higher)
    - Isp Vac = 359
    - Thrust Vac = 3.28 MN
    - Isp SL = 334
    - Thrust SL  = 3.06 MN
    - Diameter Exit = 1.7 m
For comparison here are my models of Raptor (prepared for RPA 2 SE) with all results.

Model for engine with vacuum nozzle:
https://github.com/lpre/RPA-Examples/blob/master/Configs/Cycle%20Analysis/Raptor.cfg
Engine size is defined by required thrust in vacuum.

Results:
http://lpre.de/upload/Raptor_performance.txt
http://lpre.de/upload/Raptor_nozzle.txt
http://lpre.de/upload/Raptor_cycle.txt

O/F = 3.8
Ae/At = 200
Isp vac = 383 s
Thrust vac = 3.50 MN
De = 3.8 m

Model for  engine with sea-level nozzle:
https://github.com/lpre/RPA-Examples/blob/master/Configs/Cycle%20Analysis/Raptor_SL.cfg
Engine size is defined by throat diameter obtained from analysis of engine with vacuum nozzle.

Results:
http://lpre.de/upload/Raptor_SL_performance.txt
http://lpre.de/upload/Raptor_SL_nozzle.txt
http://lpre.de/upload/Raptor_SL_cycle.txt

O/F = 3.8
Ae/At = 40
Isp vac = 356 s
Thrust vac = 3.26 MN
Isp SL = 330 s
Thrust SL  = 3.02 MN
De = 1.7 m

("e" - nozzle exit, "t" - nozzle throat)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: SirKeplan on 10/07/2016 09:48 pm
For comparison here are my models of Raptor (prepared for RPA 2 SE) with all results.

Model for engine with vacuum nozzle:
https://github.com/lpre/RPA-Examples/blob/master/Configs/Cycle%20Analysis/Raptor.cfg
Engine size is defined by required thrust in vacuum.

Results:
http://lpre.de/upload/Raptor_performance.txt
http://lpre.de/upload/Raptor_nozzle.txt
http://lpre.de/upload/Raptor_cycle.txt

O/F = 3.8
Ae/At = 200
Isp vac = 383 s
Thrust vac = 3.50 MN
De = 3.8 m

Model for  engine with sea-level nozzle:
https://github.com/lpre/RPA-Examples/blob/master/Configs/Cycle%20Analysis/Raptor_SL.cfg
Engine size is defined by throat diameter obtained from analysis of engine with vacuum nozzle.

Results:
http://lpre.de/upload/Raptor_SL_performance.txt
http://lpre.de/upload/Raptor_SL_nozzle.txt
http://lpre.de/upload/Raptor_SL_cycle.txt

O/F = 3.8
Ae/At = 40
Isp vac = 356 s
Thrust vac = 3.26 MN
Isp SL = 330 s
Thrust SL  = 3.02 MN
De = 1.7 m

("e" - nozzle exit, "t" - nozzle throat)

I used the free version of RPA, and reached performance values that are very similar. What I also noticed was that setting the mix ratio to 3.4 for the Sea Level Raptor gets 334s and 359s of ISP, which is very close to the stated values.

Does running the booster and vacuum engines at different mix ratios seem likely to be what SpaceX could be doing?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/08/2016 02:46 am
We are all getting consistent numbers. Until engine is developed this is probably as close as anyone can get. And yes mixture ratios can be made to vary if need be. I'm sizing turbo pumps now. I haven't done that before.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 10/18/2016 07:07 pm
We are all getting consistent numbers. Until engine is developed this is probably as close as anyone can get. And yes mixture ratios can be made to vary if need be. I'm sizing turbo pumps now. I haven't done that before.

John

Question, are mixing ratios variable because there are 2 separate pumps?

Edit: Not a common shaft between fuel and oxidizer.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/18/2016 08:45 pm
We are all getting consistent numbers. Until engine is developed this is probably as close as anyone can get. And yes mixture ratios can be made to vary if need be. I'm sizing turbo pumps now. I haven't done that before.

John

Question, are mixing ratios variable because there are 2 separate pumps?

Edit: Not a common shaft between fuel and oxidizer.

No, mixture ratios can be controlled with variation in pressure drops between fuel and oxidizer lines, Valves can do this.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 10/24/2016 03:54 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/

Sounds like they have developed the necessary alloy and put a few more seconds on the test engine.

Quote from: Reddit User MINDMOLESTER
Hi Elon,
ITS question:
What SpaceX technology/material still requires the most development for ITS to be a success?
Thank you!
Quote from: Elon Musk
It used to be developing a new metal alloy that is extremely resistant to oxidation for the hot oxygen-rich turbopump, which is operating at insane pressure to feed a 300 bar main chamber. Anything that can burn, will burn. We seem to have that under control, as the Raptor turbopump didn't show erosion in the test firings, but there is still room for optimization.
Biggest question right now is sealing the carbon fiber tanks against cryo propellant with hot autogenous pressurization. The oxygen tank also has an oxidation risk problem as it is pressurized with pure, hot oxygen. Will almost certainly need to apply an inert layer of some kind. Hopefully, something that can be sprayed. If need be, will use thin sheets of invar welded together on the inside.
https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94tbej/?context=3&st=iuo8s2ur&sh=8d4dc7b8
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/24/2016 04:29 pm
It appears that Mondaloy is an Aerojet product, so I can imagine that SpaceX would not have access to it.

SpaceX and Tesla have hired Charles Kuehmann to lead materials development, so SpaceX probably has its own solution.

https://electrek.co/2016/02/24/apple-alloy-expert-tesla-spacex/

Sounds like they have developed the necessary alloy and put a few more seconds on the test engine.

Quote from: Reddit User MINDMOLESTER

Hi Elon,
ITS question:
What SpaceX technology/material still requires the most development for ITS to be a success?
Thank you!
Quote from: Elon Musk
It used to be developing a new metal alloy that is extremely resistant to oxidation for the hot oxygen-rich turbopump, which is operating at insane pressure to feed a 300 bar main chamber. Anything that can burn, will burn. We seem to have that under control, as the Raptor turbopump didn't show erosion in the test firings, but there is still room for optimization.
Biggest question right now is sealing the carbon fiber tanks against cryo propellant with hot autogenous pressurization. The oxygen tank also has an oxidation risk problem as it is pressurized with pure, hot oxygen. Will almost certainly need to apply an inert layer of some kind. Hopefully, something that can be sprayed. If need be, will use thin sheets of invar welded together on the inside.
https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94tbej/?context=3&st=iuo8s2ur&sh=8d4dc7b8

They could use 200 - 300 deg F nitrogen instead.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Prettz on 10/24/2016 04:40 pm
They could use 200 - 300 deg F nitrogen instead.
Then where do you store it? And how do you refill on Mars? That sounds like it introduces more problems than it solves.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 10/24/2016 05:01 pm
They could use 200 - 300 deg F nitrogen instead.
Then where do you store it? And how do you refill on Mars? That sounds like it introduces more problems than it solves.

and it condenses into subcooled lox.
77k LN2 boiling point
66K subcooled lox
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Manabu on 10/24/2016 05:06 pm
Anyone here wants to speculate/estimate/simulate the performance of Raptor if SpaceX decided to continue in the path of making it a Hydrolox engine, given their current performance goals for methane? If they could get 30Mpa chamber pressure with Hydrolox, it would surpass the SSME engine in ISP (that many call the pinnacle in rocket science), not to mention TWR, right?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Llian Rhydderch on 10/24/2016 06:40 pm
Wrong thread.  This one is about the actual as-unveiled-by-SpaceX methalox FFSC Raptor engine.

There are hundreds of other threads where speculation would fit about "What if ... " some other design decision were to be made.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 11/03/2016 05:35 pm
I'm starting to get surprised at how little interest this thread is getting since this engine family is so interesting and down right sexy.

Also, I thought we'd hear about more test firing by now.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 11/03/2016 06:42 pm
I'm super excited. But as you said no info to work with.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 11/08/2016 03:33 pm
I'm super excited. But as you said no info to work with.
I think its not only the dearth of info, but the high inconsistency on the available one. I had more than 10 questions regarding Raptor in the Reddit AMA for Elon, but obviously none was answered. Apparently "how do you feel ..." are a lot more interesting than the use of expander cycle for the low pressure turbopump.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/08/2016 04:02 pm
I'm super excited. But as you said no info to work with.
I think its not only the dearth of info, but the high inconsistency on the available one. I had more than 10 questions regarding Raptor in the Reddit AMA for Elon, but obviously none was answered. Apparently "how do you feel ..." are a lot more interesting than the use of expander cycle for the low pressure turbopump.

Expander cycle for the low pressure turbopump???
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: baldusi on 11/08/2016 08:30 pm
I'm super excited. But as you said no info to work with.
I think its not only the dearth of info, but the high inconsistency on the available one. I had more than 10 questions regarding Raptor in the Reddit AMA for Elon, but obviously none was answered. Apparently "how do you feel ..." are a lot more interesting than the use of expander cycle for the low pressure turbopump.

Expander cycle for the low pressure turbopump???
Its a relatively common trick. I didn't saw anything like that in the picture, just a speculative question. But it is a trick used by the SSME. They use a low pressure pump to avoid cavitation. And run it from the supercritical fuel that's output by the regen cooling loop.
KBKhA RD-0162/SD use the expander cycle to run the mail fuel pump. And that was a 2MN engine. So there is some significant power availability from the expander cycle for a 3MN rocket. A pity not to use it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 01/09/2017 02:10 pm
Are there any updates about Raptor development after the September test?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 01/10/2017 07:30 pm
Are there any updates about Raptor development after the September test?

None about development or test
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 01/13/2017 10:37 am
Just noticed that SpaceX posted higher resolution photos of the raptor test fire on flickr than were attached to Elon's original tweets (as originally posted below). I can't see these higher resolutions posted earlier in this, or the previous ITS propulsion thread.

Quote from: Elmar Moelzer link=topic=34197.msg1588736#msg1588736
Elon Musk on Twitter:
SpaceX propulsion just achieved first firing of the Raptor interplanetary transport engine
https://twitter.com/elonmusk/status/780280440401764353

Production Raptor goal is specific impulse of 382 seconds and thrust of 3 MN (~310 metric tons) at 300 bar
https://twitter.com/elonmusk/status/780275236922994688
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rockets4life97 on 01/27/2017 02:56 am
Any word on more tests? Anybody have a guess about how long they will test this initial engine before moving to an upgraded version?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 03/10/2017 08:21 pm
Here is a strictly hypothetical question.

Assuming this 1,000 kN demonstrator reaches a 30 MPa operating chamber pressure, how big/wide would a 50:1 ratio nozzle be for it? Moreover, what would be the most effective/efficient nozzle ratio that it could have, assuming it is used for first stage propulsion (among 8 other engines) and slow/low S1 separation for RTLS duties?

For a booster engine 50:1 would be about right. Cycle is the same so the nozzle scales with area so its diameter scales with the square root of the thrust ratio:
                                          40:1 diam = 1.7 x sqrt(1 / 3.05)         = . 97 m  (~38 inches).
                                          50:1 diam = 1.7 x sqrt(50 / 40 /3.05) = 1.09 n (~ 43 inches)

Many thanks for that. A couple more questions to anyone interested to answer (again, this is a hypothetical scenario).

What is the diameter of the current M9 nozzle?
If we assume that the material, width and height of the current F9 S1 remains constant, and that the common bulkhead is moved to adjust, given:

1. The known propellant ratio for the Raptor Demonstrator.
2. An SL thrust of 870kN and Vac thrust of 930kN.
3. An SL Isp of 330s and Vac Isp of 358s
4. A dry stage weight of 27 metric tons.

What would the performance delta be against the current F9 S1?

I'm not asking whether something like this is possible, probable, practicable or wanted/needed. Just want to understand the comparative difference between one engine and the other in a hypothetical scenario. I assume that the difference would be rather small, both due to having less propellant on the stage and Isp not being the most important factor in the two re-usable scenarios that F9 S1 covers (RTLS and DPL S1-S2 separations).

Adjusting for lower propellant density and assuming similar engine TWR, the high pressure FFSC methalox still gets 31% more payload to LEO and 38% more payload to GTO compared to low pressure GG kerolox:

Using http://www.silverbirdastronautics.com/LVperform.html
To 185 km x 28.5 deg circular LEO with no fairing and 0.5% residuals:

21162 kg for kerolox S1: 24000 kg dry, 430000 kg prop, 8000 kN avg, 297 sec avg; S2: 4500, 115000, 934, 348.

25743 kg for methalox S1: 24000 kg dry, 360000 kg prop, 8100 kN avg, 348 sec avg; S2: 4500, 96000, 1000, 374.

To 185 x 38500 km x 28.5 deg GTO with 4000 kg fairing discarded at 220 sec, and 0.5% residuals:

7006 kg for kerolox S1: 24000 kg dry, 430000 kg prop, 8000 kN avg, 297 sec avg; S2: 4500, 115000, 934, 348.

9680 kg for methalox S1: 24000 kg dry, 360000 kg prop, 8100 kN avg, 348 sec avg; S2: 4500, 96000, 1000, 374.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Manabu on 03/14/2017 06:17 pm
I've run numbers on RPA-lite to calculate a family of Raptor engines differing only in Expansion Ratio (ER). I have calibrated my model considering only two authoritative sources: The vacuum numbers for Raptor in the IAC lecture and the stage drawings that are supposedly directly from CAD. This means I'm using 3.7 O/F from spaceflight101 tank measurements (http://spaceflight101.com/spx/spacex-raptor/) instead of the 3.8 O/F that Elon said. I attached my RPA-lite configuration file for the Raptor 200 (just change the extension to .cfg). For the others I only variated the ER.

I used the 'freezing at area ratio' to aim precisely at 382 s isp for the Raptor 200. It gave an pretty high number of 12 and is still undershooting the SL variants of the engine. The RPA guys use 6 (https://www.slideshare.net/AlexanderPonomarenko/rpa-presentation-jun2013) for their RD-253 (N2O4/UDMH) performance validation and still undershot the ISP too, especially at SL. So maybe more is adequate for a methane raptor, I don't know. It is set lower for other fuel types and R7 (https://forum.nasaspaceflight.com/index.php?topic=35655.msg1262643#msg1262643) found that 3 is adequate to simulate a Russian methane rocket engine.

Leaving the throat diameter fixed at 0.2685m and using the measurements from OneSpeed (https://forum.nasaspaceflight.com/index.php?topic=42003.msg1629656#msg1629656), by simple scaling I get an ER for booster engines of 32:1 and 44:1 for the BFS SL engines. I assume that the 3050 kN 334s at 40 ER SL engine described in the IAC slides is in fact the Raptor 32 while the 361 s vacuum isp is the Raptor 44. RPA has undershot both slightly. The Raptor 40 is as far as I understand just a middle of the way designation to talk about the performance of an average SL Raptor, but I ran numbers for it too anyway, as well as the usually discussed Raptor 50.

I also ran numbers for other intermediate ER, for the benefit of those who are dreaming with a BFS SSTO (me included). 116:1 being one that fits 9 in the perimeter of BFS and 130:1 being the maximum ER that RPA-lite doesn't warns me against flow separation at SL. Some altitude performance analysis graphs are attached too. They seem based on Theoretical performance, not the estimated delivered performance.


    Nozzle size    |        Sea Level      |          Vacuum       | Optimal Expansion |
 ER | Diameter (m) | Thrust (kN) | Isp (s) | Thrust (kN) | Isp (s) |  H (km) |  P (atm)|
----|-------------:|------------:|--------:|------------:|--------:|--------:|--------:|
 32 |     1.52     |    3044     |  332.0  |    3234     |  353.0  |   0.00  |   1.002 |
 40 |     1.70     |    3037     |  331.5  |    3274     |  357.4  |   2.33  |   0.753 |
 44 |     1.78     |    3029     |  330.7  |    3290     |  359.1  |   3.29  |   0.667 |
 50 |     1.90     |    3015     |  329.1  |    3311     |  361.4  |   4.55  |   0.566 |
 57 |     2.03     |    2994     |  326.8  |    3332     |  363.7  |   5.79  |   0.479 |
 80 |     2.40     |    2908     |  317.4  |    3382     |  369.2  |   8.82  |   0.312 |
116 |     2.89     |    2746     |  299.8  |    3434     |  374.8  |  11.88  |   0.195 |
130 |     3.06     |    2678     |  292.3  |    3448     |  376.4  |  12.80  |   0.169 |
200 |     3.80     |    2315     |  253.0  |    3500     |  382.0  |  16.27  |   0.098 |



I don't know how to force a fixed width font in this forum (edit: now I know, thanks). The results are inside RPA error margin, especially considering that it should not be as tuned for methane because the lack of real engine data to check against.

I'm ignoring completely this latest information on Raptor (https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94u6zk/?context=3), as it suggests a smaller engine with a vacuum thrust at 200:1 ER in the 3125 kN range, while the IAC slides said 3500 kN.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Manabu on 03/14/2017 06:24 pm
I also did a throttling analysis on the same basis. The Raptor 40 isn't quite capable of throttling down to 20% before flow separation at SL, according to RPA-lite. But with 32:1 ER it can, and with 44:1 it can throttle down to about 30%. Maybe some nozzle tricks may prove those numbers too conservative.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 03/14/2017 06:46 pm
I also did a throttling analysis on the same basis. The Raptor 40 isn't quite capable of throttling down to 20% before flow separation at SL, according to RPA-lite. But with 32:1 ER it can, and with 44:1 it can throttle down to about 30%. Maybe some nozzle tricks may prove those numbers too conservative.

Nice work!

For the throttled engines, are you plotting chamber pressure ratios or thrust ratios? Because of atmospheric back-pressure at sea level slowing the exhaust, throttling the chamber pressure to 20% will produce less than 20% thrust.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Manabu on 03/14/2017 07:47 pm
Nice work!

For the throttled engines, are you plotting chamber pressure ratios or thrust ratios? Because of atmospheric back-pressure at sea level slowing the exhaust, throttling the chamber pressure to 20% will produce less than 20% thrust.
I specified the interval as thrust ratios, where 1.0 corresponds to the nominal thrust. RPA-lite did the rest for me. But good observation, I haven't thought about that.

EDIT: Another thing to have in mind is that those numbers use the SL performance that I estimated with RPA-lite, that is a bit lower than the ones confusingly said by SpaceX. I'm also using the 3.7 O/F that gives a little less thrust for a given ISP.

I redid the Throttled chamber performance analysis with a more orthodox 3.8 O/F, pure shifting equilibrium model for the nozzle and reaction efficiency manually raised to 99.4 to match the Raptor 40 IAC numbers. Graph in the attachment and here the engine parameters compared to the ones in the other table:


      Nozzle size     |        Sea Level      |          Vacuum       | Optimal Expansion |
  ER   | Diameter (m) | Thrust (kN) | Isp (s) | Thrust (kN) | Isp (s) |  H (km) |  P (atm)|
-------|-------------:|------------:|--------:|------------:|--------:|--------:|--------:|
40     |     1.70     |    3037     |  331.5  |    3274     |  357.4  |   2.33  |   0.753 |
40  V2 |     1.70     |    3052     |  334.1  |    3287     |  359.8  |   1.69  |   0.815 |
200    |     3.80     |    2315     |  253.0  |    3500     |  382.0  |  16.27  |   0.098 |
200 V2 |     3.80     |    2361     |  258.4  |    3536     |  387.0  |  15.55  |   0.110 |


In the end the only thing that seems to have affected the plot is the O/F ratio, and then only a little, as RPA-lite seems to also use Theoretical performance instead of the estimated delivered performance in this plot.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: OneSpeed on 03/15/2017 07:58 pm
I don't know how to force a fixed width font in this forum.

You can create a table in the reply editor, using the table tags, but it is a bit laborious:

Nozzle sizeSea LevelVacuum-OptimalExpansion
ERDiameter (m)Thrust (kN)Isp (s)Thrust (kN)Isp (s)H (km)P (atm)
401.703037331.53274357.42.330.753
40  V21.703052334.13287359.81.690.815
2003.802315253.03500382.016.270.098
200 V23.802361258.43536387.015.550.110

Is that what you are after?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: nacnud on 03/15/2017 08:03 pm
This may help in the future, but test it first!

http://www.teamopolis.com/tools/bbcode-table-generator.aspx
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AnalogMan on 03/15/2017 08:42 pm
I don't know how to force a fixed width font in this forum.

You can create a table in the reply editor, using the table tags, but it is a bit laborious:

Nozzle sizeSea LevelVacuum-OptimalExpansion
ERDiameter (m)Thrust (kN)Isp (s)Thrust (kN)Isp (s)H (km)P (atm)
401.703037331.53274357.42.330.753
40  V21.703052334.13287359.81.690.815
2003.802315253.03500382.016.270.098
200 V23.802361258.43536387.015.550.110

Is that what you are after?

You can force a fixed pitch font using the
[tt] and [/tt]
tags.  If using the simple forum editor in preview mode then you can also highlight the relevant text and click the "Tt" button - this inserts the tags for you.

This produces a monospaced teletype font - this is what it looks like applied to the text your table:

    Nozzle size    |        Sea Level      |          Vacuum       | Optimal Expansion |
 ER | Diameter (m) | Thrust (kN) | Isp (s) | Thrust (kN) | Isp (s) |  H (km) |  P (atm)|
----|-------------:|------------:|--------:|------------:|--------:|--------:|--------:|
 40 |     1.70     |    3037     |  331.5  |    3274     |  357.4  |   2.33  |   0.753 |
40  V2 |  1.70     |    3052     |  334.1  |    3287     |  359.8  |   1.69  |   0.815 |
200 |     3.80     |    2315     |  253.0  |    3500     |  382.0  |  16.27  |   0.098 |
200 V2 |  3.80     |    2361     |  258.4  |    3536     |  387.0  |  15.55  |   0.110 |
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Manabu on 03/16/2017 12:33 am
Thanks all above, I fixed the tables using the AnalogMan advice. The bbcode table is laborious to make, even with that website, while I already have a workflow for those fixed width tables and they are more "portable". But maybe I can use for some future tables to make them a little prettier.

I found a small problem in my simulation. When I went to look the logs by curiosity, I found this silent warning:
Quote
WARNING: Temperature T=93.00 K could not be assigned to the species "CH4(L)". Using T=298.15 K instead.
The minimum temperature supported for CH4 is 100 K, and that reduces the isp compared to 298.15 K by about 2 s, all else the same. When increasing the freezing area ratio to match the 382 Raptor 200 isp, the Raptor 32 isp drop up to 2 s compared to the previous simulation.

But I'm right in using those sub-cooled temperatures as they are in the tanks? Or should I use high temperatures and pressures for the fuel (and maybe the oxidizer too) because the engine is regenerative cooled? This would reduce a little, but not eliminate, the gap between SpaceX stated SL performance and my RPA-lite simulations.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 03/21/2017 09:09 pm
How far along is the Raptor engine?  Any word as to when the Raptor and the Raptor vacuum will be ready for full testing?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: macpacheco on 03/21/2017 10:18 pm
How far along is the Raptor engine?  Any word as to when the Raptor and the Raptor vacuum will be ready for full testing?
I'm no rocket scientist/engineer but it seems clear enough there will be a full year minimum testing before proper sea level / vacuum engines are produced for actual full thrust testing/qualification. The real for flight engines might not even be built in 2017.
This is still very early testing on a complete engine.
They will have to slowly increase thrust/change mixtures until the sub scale engine is running at its optimal (and more dangerous) parameters.
We don't know how much the engine components are finalized with margins to tolerate full power operations or a normal size engine.
I would wait at least until late summer/2017 to repeat such questions and hope for an actual answer.
Raptor is a crazy ambitious project. It not only intends to be one of the most efficient rocket engines in the world but also capable of 1000 mission firings (with at least 100 firings without any engine refurb). That and M1D are already good enough for current missions. They will take their time to do it right, much like M1C/M1D development progressed much slower than some people wanted, because Musk demanded the engine had crazy margins which are now paying off with Block IV/V thrust upgrades.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: dglow on 03/21/2017 11:31 pm
A different angle: SpaceX isn't the only company building a methane SC engine. And they increasingly find themselves in direct competition, on multiple levels, with the other company doing so.

So not only do we not know, to any level of precision, the progress of Raptor development; I suspect we are unlikely to ever know much detail until the rocket is finished, or very nearly so. Blue is famously tight-lipped, and we've seen SX increasingly adopt a similar approach.

Sorry, that sucks as an answer. We can scout McGregor until the cows come home – or run away! – but we won't know Raptor is ready until either, a big Elon reveal (which won't necessarily coincide with 'finished'), or when we see it fly.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/22/2017 01:08 am
Heck, SpaceX was more tight lipped than Blue Origin. Blue Origin did a press release with pictures and articles when the first BE-4 was finished, before even the first actual BE-4 test firing. SpaceX only showed the Raptor test firing. I think this may be because Blue has a customer that hasn't 100% decided on what engine to pick yet, so Blue has to make a big deal about any progress so it's obvious to all stakeholders. SpaceX just has themselves, in reality (other than some Air Force funding for development, which doesnt need public press releases).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Okie_Steve on 04/01/2017 11:09 pm
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Apollo100 on 04/03/2017 10:54 pm
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Stan-1967 on 04/03/2017 11:17 pm
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Off topic, but interesting.  Start a thread to compare this idea to ULA's IVF technology?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: kendalla59 on 04/06/2017 08:26 pm
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

The Integrated Powerhead Demonstrator used liquid hydrogen propellant, so yes of course SpaceX must have changed the design for Raptor. Thanks for the pointer -- it was fascinating to read about IPD. I wonder how many Aerojet-Rocketdyne engineers are working at SpaceX now?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/06/2017 10:31 pm
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Off topic, but interesting.  Start a thread to compare this idea to ULA's IVF technology?
SpaceX will just use batteries and a solar array. Or just batteries, most likely.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: brickmack on 04/06/2017 11:03 pm
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

The Integrated Powerhead Demonstrator used liquid hydrogen propellant, so yes of course SpaceX must have changed the design for Raptor. Thanks for the pointer -- it was fascinating to read about IPD. I wonder how many Aerojet-Rocketdyne engineers are working at SpaceX now?

Hydrogen seems to behave pretty similarly to methane with regards to engine operation. Most of the methalox engines fired to date have been lightly modified hydrolox engines, not purpose-built designs.

Though I doubt there is much IPD heritage in Raptor
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Stan-1967 on 04/07/2017 08:25 pm
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Off topic, but interesting.  Start a thread to compare this idea to ULA's IVF technology?
SpaceX will just use batteries and a solar array. Or just batteries, most likely.
 

ITS is likely going to need more than solar arrays for thermal management of propellant as well as life support.  My understanding ( which may be incorrect) of IVF was that it solved multiple problems for keeping the upper stage "alive":

1.  Thermal management of propellant ( autogenous pressurization).  The choice of an ICE was made because they needed waste heat ( entropy) to keep the stage functional.
2.  Ullage system using combustion products expanded through a rocket nozzle
3.  electricity generation. 

Solar cells only perform the electricity generating function well.  The efficiency of using solar cells, even high efficiency triple junction ones, to product the needed heat for prop management & ECLSS during cruise does not seem like a winning proposition for efficiency & mass tradeoffs.

Autogenous pressurization may work well on the ground when GSE equipment can provide the heat, & during limited loiter times in orbit around earth, but what about the 3-6 month cruise to Mars?  It would vastly increase the mass of the PV array if it had to be sized to generate electricity to power heaters for the needed thermal budget of all ITS systems vs. just electrical power for GNC.

IVF solves this for an unmanned upper stage in the vincinity of the Earth, ITS is going to have much more complex & demanding thermal requirements.  It may end up being a combination of PV, ICE/Fuel cell/solar thermal.

I also question how thermal management of deep cryogenic propellant will affect the design of Raptor.  In other F9 threads, those in the know insisted that it was non-trivial to characterize the performance of the GG turbo machinery for different temperature of prop.  Basically if they went back to non deep chilled prop, they would have to change the turbomachinery.  Does FFCS bypass this issue?  Keeping LOX/CH4 superchilled for the cruise to Mars seems demanding for power & mass requirements, so the ability to start & operate Raptor under a wide range of propellant temperatures seems necessary.


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 04/07/2017 08:44 pm


1.  Thermal management of propellant ( autogenous pressurization).  The choice of an ICE was made because they needed waste heat ( entropy) to keep the stage functional.

2.  Ullage system using combustion products expanded through a rocket nozzle


I would think the combustion products of an ICE (water and CO2) would not be a good pressurization gas.
Water with cryogenics doesn't work.

If you meant just to power the coolers I guess that works. Still think solar cells are a better choice. Maybe insulation too.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/07/2017 11:53 pm
Dumping the vehicle's internal heat all into the propellant sounds like a really good way to end up at Mars without any propellant left for landing.

ITS will have radiators. Perhaps body-mounted or something (or even just carefully using the whole side of the vehicle passively, but either way it'll dump heat radiatively
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/08/2017 01:38 am
If you wanted to use something like IVF to produce power instead of that 100-200kW of solar for 100 days, it'd consume about 250 tons of propellant. If you want to dump all that produced heat into the propellant, you'd run out of propellant before arriving at Mars, even if the ship somehow was 1950 tons full of propellant after trans-Mars-insertion burn.

People are all "solar is wimpy, use a combustion engine, ha!" but solar actually kicks butt in orbit. For a given mass in orbit (including consumables) you can produce about 200-400 times as much energy with solar as with IVF over 100 days.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 04/08/2017 04:05 am
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Off topic, but interesting.  Start a thread to compare this idea to ULA's IVF technology?
SpaceX will just use batteries and a solar array. Or just batteries, most likely.

Yup - I've always wondered about that ICE.  Batteries today can almost compete with fuel BEFORE you carry the oxygen with you, not to mention that in space heat rejection (for a heat engine) means even more mass, even beyond just the dead mass of the engine.

... and batteries can recharge.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MP99 on 04/08/2017 07:08 pm
Dumping the vehicle's internal heat all into the propellant sounds like a really good way to end up at Mars without any propellant left for landing.

ITS will have radiators. Perhaps body-mounted or something (or even just carefully using the whole side of the vehicle passively, but either way it'll dump heat radiatively
If ITS is oriented with the crew section pointing to the sun, and the prop tanks in shade, the engine bells will get very cold. Prop could be circulated through the regen channels in the engines to provide cooling to ensure ZBO.

Cheers, Martin
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeAtkinson on 04/08/2017 08:30 pm
Dumping the vehicle's internal heat all into the propellant sounds like a really good way to end up at Mars without any propellant left for landing.

ITS will have radiators. Perhaps body-mounted or something (or even just carefully using the whole side of the vehicle passively, but either way it'll dump heat radiatively
If ITS is oriented with the crew section pointing to the sun, and the prop tanks in shade, the engine bells will get very cold. Prop could be circulated through the regen channels in the engines to provide cooling to ensure ZBO.

Cheers, Martin

They may not even need to do that, the propellant tanks would receive very little direct sunlight and conduction through a carbon fibre composite should be low. SpaceX may even have to take measures to stop the propellants getting too cold!

Orbiting Earth the heat load will be higher and cannot be easily controlled by orientation, so if the engine bells were shaded from both the Sun and Earth then your idea would I think be necessary, particularly as the "fleet" might spend months in LEO waiting for the TMI window. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 04/08/2017 09:06 pm
Orbiting Earth the heat load will be higher and cannot be easily controlled by orientation, so if the engine bells were shaded from both the Sun and Earth then your idea would I think be necessary, particularly as the "fleet" might spend months in LEO waiting for the TMI window.

Maybe pointing the heatshield towards earth. It should have reasonable insulation capability with its low weight. But zero boil off is probably not achievable in LEO.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 04/08/2017 09:27 pm
If you wanted to use something like IVF to produce power instead of that 100-200kW of solar for 100 days, it'd consume about 250 tons of propellant. If you want to dump all that produced heat into the propellant, you'd run out of propellant before arriving at Mars, even if the ship somehow was 1950 tons full of propellant after trans-Mars-insertion burn.

People are all "solar is wimpy, use a combustion engine, ha!" but solar actually kicks butt in orbit. For a given mass in orbit (including consumables) you can produce about 200-400 times as much energy with solar as with IVF over 100 days.
I get about 60 tons a month fuel + LOX for 100kw methane turbine using earthbound generator specs and guessing 3.8 LOX for every 1 fuel. Doesn't really make a case for an engine over solar though.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/08/2017 10:16 pm
Stoichiometric is CH4+2*O2, and CH4 is 16 molar mass, and 2*O2 is 64 molar mass. So even using the very optimistic stoichiometric case, you're looking at 4:1. Anyway, I think we're basically in agreement. 100 days is 3 and a third months, so 60 tons per month is 200 tons by your measure. But stoichiometric would likely be way too hot and would burn out the motor. IVF runs very fuel-rich, for instance (while on Earth, 80% nitrogen in air naturally will keep you cool enough). So add at least a bit of methane to keep it cool, and you're at 250 tons for 100 days.

Anyway, whether 200 tons or 250 tons is immaterial. It's way heavier than the ~1-2 tons of solar array that is needed for 100kW at Mars.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Surgeon on 04/10/2017 04:43 am
Quick question:
Is there any way to find out what kind of ISP the 1/3 Demonstrator would have with a 150:1 expansion ratio and a thrust of 1MN?

All the estimates point towards the Raptor being of a similar physical size to the Merlin, and some back of the envelope calculations that I've done make it look like that if it can do an ISP of ~375, you could get an extra 1.25 tonne of payload to GTO using a second stage with the same volume as the existing one (using densified methalox fuel) and an extra tonne or so of mass to account for the new engine. This, even when accounting for the reduction in fuel mass due to methalox being less dense.

If the 1/3 Demonstrator could do an ISP of ~375s, a thrust of 1MN (just bigger than the current Merlin 1D+ Vac’s 0.934 MN) and still fit inside the Interstage WITHOUT having to change the dimensions of the second stage, there could be quite the argument for changing over to methalox on the second stage.

They could also save further on weight by using an IVF type system. This would be the closest they could possibly get to a ‘drop-in’ replacement to change the stage over to methalox and boost their payload, or have enough for efficient second stage reusablility. Needless to say, it WOULD NOT be a drop in replacement, but they have all the parts to redesign the second stage and use their existing infrastrauce, all they would have to do is add densified methane storage and piping to things (TEL, test stands etc.) Heck, they could even use the same tanks for the subcooled Methane as they do for the subcooled O2.

Now I’m not suggesting this is the plan, I’m just curious to see whether or not it’s possible to replace the Merlin on the second stage with a Raptor and still get a payload boost without having to make the second stage wider/longer/bigger… then it can be argued whether or not it is worth it. Such a new stage would not be optimised for its mission, but it’s already not, and gaining experience with the Raptor, IVF and reusability could be worth the inefficiencies of having to produce 2 kinds of engines.

TL;DR – How much more/less extra payload could you get by putting the 1/3 demonstrator Raptor on a stage the same size/volume as the existing second stage? What ISP would it need to make that meaningful?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/10/2017 04:45 am
So now it has become 1/3rd scale?

We need better sources, here. Throat size is the best determinant of scale, and the throat looks big enough for full scale.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Surgeon on 04/10/2017 04:47 am
So now it has become 1/3rd scale?

We need better sources, here. Throat size is the best determinant of scale, and the throat looks big enough for full scale.

I thought it was pretty much confirmed it was scaled down, is the stand it was on big enough to take 3MN?

EDIT: The article that came out last year October 3 says it was a 1/3 Demonstrator

https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/10/2017 04:50 am
So now it has become 1/3rd scale?

We need better sources, here. Throat size is the best determinant of scale, and the throat looks big enough for full scale.

I thought it was pretty much confirmed it was scaled down, is the stand it was on big enough to take 3MN?
Irrelevant. The THRUST may be scaled down, not necessarily the chamber size. After all, the most challenging part of Raptor is the insane chamber pressures, not the physical size. And even if you had a full-power capability, you'd first run it at lower pressures.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 04/10/2017 05:13 am
Quick question:
Is there any way to find out what kind of ISP the 1/3 Demonstrator would have with a 150:1 expansion ratio and a thrust of 1MN?

All the estimates point towards the Raptor being of a similar physical size to the Merlin, and some back of the envelope calculations that I've done make it look like that if it can do an ISP of ~375, you could get an extra 1.25 tonne of payload to GTO using a second stage with the same volume as the existing one (using densified methalox fuel) and an extra tonne or so of mass to account for the new engine. This, even when accounting for the reduction in fuel mass due to methalox being less dense.

If the 1/3 Demonstrator could do an ISP of ~375s, a thrust of 1MN (just bigger than the current Merlin 1D+ Vac’s 0.934 MN) and still fit inside the Interstage WITHOUT having to change the dimensions of the second stage, there could be quite the argument for changing over to methalox on the second stage.

They could also save further on weight by using an IVF type system. This would be the closest they could possibly get to a ‘drop-in’ replacement to change the stage over to methalox and boost their payload, or have enough for efficient second stage reusablility. Needless to say, it WOULD NOT be a drop in replacement, but they have all the parts to redesign the second stage and use their existing infrastrauce, all they would have to do is add densified methane storage and piping to things (TEL, test stands etc.) Heck, they could even use the same tanks for the subcooled Methane as they do for the subcooled O2.

Now I’m not suggesting this is the plan, I’m just curious to see whether or not it’s possible to replace the Merlin on the second stage with a Raptor and still get a payload boost without having to make the second stage wider/longer/bigger… then it can be argued whether or not it is worth it. Such a new stage would not be optimised for its mission, but it’s already not, and gaining experience with the Raptor, IVF and reusability could be worth the inefficiencies of having to produce 2 kinds of engines.

TL;DR – How much more/less extra payload could you get by putting the 1/3 demonstrator Raptor on a stage the same size/volume as the existing second stage? What ISP would it need to make that meaningful?

Economically not sensible until they get second stage reuse working, Raptor is MUCH more expensive than Merlin, even at 1/3 of the size. The cost increase would be much greater than the payload increase, and AFAIK they have no payloads too big for first-stage reusable FH, so this is simply not needed.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Surgeon on 04/10/2017 05:49 am
Quick question:
Is there any way to find out what kind of ISP the 1/3 Demonstrator would have with a 150:1 expansion ratio and a thrust of 1MN?

All the estimates point towards the Raptor being of a similar physical size to the Merlin, and some back of the envelope calculations that I've done make it look like that if it can do an ISP of ~375, you could get an extra 1.25 tonne of payload to GTO using a second stage with the same volume as the existing one (using densified methalox fuel) and an extra tonne or so of mass to account for the new engine. This, even when accounting for the reduction in fuel mass due to methalox being less dense.

If the 1/3 Demonstrator could do an ISP of ~375s, a thrust of 1MN (just bigger than the current Merlin 1D+ Vac’s 0.934 MN) and still fit inside the Interstage WITHOUT having to change the dimensions of the second stage, there could be quite the argument for changing over to methalox on the second stage.

They could also save further on weight by using an IVF type system. This would be the closest they could possibly get to a ‘drop-in’ replacement to change the stage over to methalox and boost their payload, or have enough for efficient second stage reusablility. Needless to say, it WOULD NOT be a drop in replacement, but they have all the parts to redesign the second stage and use their existing infrastrauce, all they would have to do is add densified methane storage and piping to things (TEL, test stands etc.) Heck, they could even use the same tanks for the subcooled Methane as they do for the subcooled O2.

Now I’m not suggesting this is the plan, I’m just curious to see whether or not it’s possible to replace the Merlin on the second stage with a Raptor and still get a payload boost without having to make the second stage wider/longer/bigger… then it can be argued whether or not it is worth it. Such a new stage would not be optimised for its mission, but it’s already not, and gaining experience with the Raptor, IVF and reusability could be worth the inefficiencies of having to produce 2 kinds of engines.

TL;DR – How much more/less extra payload could you get by putting the 1/3 demonstrator Raptor on a stage the same size/volume as the existing second stage? What ISP would it need to make that meaningful?

Economically not sensible until they get second stage reuse working, Raptor is MUCH more expensive than Merlin, even at 1/3 of the size. The cost increase would be much greater than the payload increase, and AFAIK they have no payloads too big for first-stage reusable FH, so this is simply not needed.

Fair enough, I understand that it does not make economic sense, and is certainly not part of the plan, I'm more curious as to what effect it would have and what boost/loss it would make to the payload, not so much the economic sense, and I don't have the know how to properly work this stuff out myself.

One could make the argument that switching to a Raptor upper stage would give them the margin to make second stage reuse more effective/efficient. Maybe once they've nailed down a simple second stage reuse plan, they could redesign the second stage to make reuse more stream line..I.E. actual powered landing over parachutes.

But this is getting off topic, the question is what sort of stats would a sub-scale Raptor (or as I like to think of it, a Raptor 1C :P) have and could that be used to improve GTO payload mass of the Falcon 9 without changing the dimensions of the second stage.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 04/10/2017 08:20 am
Could it potentially make F9 so much more capable that they can do many flights witout using FH? Could they get away with maybe 50cm more stage diameter without changing the TE? As I expect the new carbon fiber body for a Raptor upper stage they would not be fixed to the same diameter except for TE-restrictions.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 04/10/2017 02:05 pm
There have been huge threads on raptor-based upper stages already. Let's not turn this one into a rehash, please.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 04/10/2017 09:57 pm
A scaled Raptor should e able to hit Isp of 375 sec as long as the expansion ratio is near 150 and the pressures are in the neighborhood of the full scale design.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Surgeon on 04/10/2017 11:08 pm
A scaled Raptor should e able to hit Isp of 375 sec as long as the expansion ratio is near 150 and the pressures are in the neighborhood of the full scale design.

John

Brilliant! Thanks John!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/11/2017 01:02 am
Scaled up or down? :P
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 04/11/2017 01:26 am
Scaled up or down? :P

Seriously, they have referred to Raptor as "scalable" on several occasions. And, it's largely 3-d printed. And we know they were doing massively parallel GPU combustion modeling.

What if they have made a parametric engine design that can be scaled to any size and would just need an extended acceptance test for the first example of a new size? This seems like the sort of thing that Elon "first principles" Musk would figure out how to do.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/11/2017 02:36 am
No.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 04/11/2017 06:40 am
Performance in rocket engines does not scale linearly in proportion to size. That has been known since liquid engine production first began. Musk said they did extensive modeling and are building at the size which provides optimum performance. Can you scale it? Yes. If you double the mass of the full sized Raptor with all proportions the same, will you get double the performance? No.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 04/11/2017 07:42 am
Let me try to explain with an analogy. I am not a rocket scientist. One of my four teaching credentials is in Physics, though. Let me try to explain this in a manner that a HS physics student would understand.

Let's say you were in deep space where the temperature is essentially absolute zero. You have a cube of pure titanium that is exactly one cubic centimeter in size and has a temperature of 300 °C. I'm not going to calculate the joules or BTU that equals; let's just say it is x units of energy. Heat is going to radiate into space according to a given formula from the 6 cm2 surface area. If you had 1000 of these cubes at a substantial distance from each other, you would have 1000 times the mass, 1000 times the energy, 1000 times the surface area, and 1000 times the energy output at the same rate of radiation.

But let's say that instead of 1000 single cubic centimeter objects, you arranged them into one single cube that is 10 cm wide, 10 cm high, and 10 cm deep. You still have 1000 times the mass of the original cube, one thousand times the total energy, but not 1000 times the surface area. Rather than 6000 square cm of surface area, you now have only 600 square centimeters of surface area. You now have only 10% of the surface area from which to radiate the heat. You have scaled the mass, dimensions, and energy with perfect proportion, but the surface area differs dramatically. Those first three things have increased by 1 x 103 while the surface area has only increased by 1 x 102.

A rocket engine is not a set of cubes sitting in space; it is profoundly more complex. The principle we have seen in relation to scalability of size and mass vs. surface area comes into play with combustion chambers and expansion nozzles. And rather than plane geometry, we are dealing with complex calculus. Your scaled up Raptor is not going to have proportional surface area against which the expanding gasses push. The temperature of the oxidizing prop is not scaled. The manner in which the prop fluids mix, oxidize, and expand are not the same either. There are similarities, but swirling oxidizing gasses will behave differently according to a fractal equation. Watch a Youtube video of a 2D Mandlebrot set to see an analogy.

When scaling proportionally, the larger you make something, the lower the ratio to surface area. Conversely, the smaller you make it, the greater the ratio for surface area. (Just think about dicing a cube of cheese into ever smaller cubes. You keep the same mass but keep increasing the surface area.) With rocket engines, there comes a point, however, at which you have too many small engines and too much difficulty mounting them on a thrust plate, Your prop lines have ever smaller diameter and the coefficient of friction in relation to surface area inside the lines becomes problematic. Simply put, there is a sweet spot in size for any given rocket engine design. Scale it up or down proportionally in size and mass, it just ain't gonna work the same. Elon has said the optimal size for this engine produces a bit over 500k lb thrust. They are experimenting with a smaller prototype because that is simpler, but it is not as efficient, in mathematical theory, as the full size engine.

So, again, can you scale the thing up or down proportionally in mass and dimension? Yes. Will you get proportional performance? No. There's nothing simple about any of this stuff. That's why it's called Rocket Science.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeAtkinson on 04/11/2017 08:02 am
Yes, but a parametric engine design could take those factors into account, it just needs more (and more complex) parameters.

The difficulty for something like Raptor is that simulations would only get them so far, they would actually need to build and test various scales of engines, and then tweek the parameters to agree with reality.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 04/11/2017 08:21 am
And now you have a different engine. You can't just tell a computer controlled 3D printer, Make me a Raptor engine that is scaled proportionally at 1.5 x the size and expect to get 1.5 x the performance. What you are doing is creating a new and different engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 04/11/2017 01:03 pm
Scaling an engine (within reason 1/2 - 2 times) will not have a significant impact on PERFORMANCE. Square cube scaling relations do add to cooling requirements for smaller engines, but as long as there is sufficient cooling capacity in the propellants to handle it it will not effect performance.

Having said that, DEVELOPMENT of a scaled engine is complex, time consuming and expensive. The more you scale up or down away, from a fully developed baseline design, the more complex and expensive it is. Some reasons for this have been outlined above.

Bottom line, different physical phenomenons scale differently, hence many things need to be changed when scaling. Pumps, combustion, cooling, etc. all require non photographic scaling of parts. Then the development testing and design iteration must be redone. It won't be as hard as starting from scratch, but hard non the less.

I am a rocket engineer.  ;^)

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 04/11/2017 02:28 pm
OK, so this was precisely the discussion I've been trying to provoke about this for some time.

TomH ... I should have saved you some typing of your excellent scaling example, I have a degree in physics and two in EE, have worked on computational fluid models (ocean, atmosphere, supersonic flow), and I grok scaling laws :-)

The expert consensus seems to be that even with knowledge of the physical scaling laws that govern engine design and the advances in computer modeling of structures and combustion, it is still a lot of work.

Is combustion instability the biggest issue? It seems like structures, pressure vessels, piping, turbines, and pumps are pretty amenable to evaluation by computational methods.

Once you have enough modeling to get somewhat close to reality, you can start to thing of a parametric model that generates an engine design. It would have a lot of parameters to cover even things like component placement. Then you can automate the initial design process with a genetic algorithm system. Fitness of a particular design means it passes basic structural tests and has top scores for CFD flow and weight of materials used. That process can examine a large state space and come up with potential starting points for the designer to use. Could be more trouble than it's worth, but I think nearly all the pieces have to be present already. GA driver to control the search is not hard to do (I've built two different ones). Parametric generation of the piping and structures is the missing piece but you could have a crew of interns doing that in CAD :-)

Could be one explanation of how they came up with the Raptor design layout and the lox-side turbopump integrated into the combustion chamber.

OK, beat me up, I am a compiler guy, parallel programmer, chip designer but NOT a rocket engineer.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 04/11/2017 02:46 pm
I think the CFD and parametric modeling is the "easy" part.  What is hard (and what takes all the time) is taking the theoretically perfect design, putting it on a test stand, and seeing what breaks.  And this has to be done methodically and slowly because you're talking about real physical objects and test stands that have to be rebuild from scratch if you goof and blow everything up.

So the "parameterized" guys are right: you could totally design a parameterized engine---the *theory* is understood well enough.  It wouldn't be linear scaling, you'd adjust everything together to take into account scaling laws and what's known about all the processes.

 But the actual rocket engineers are *more right*.  Once you've done that and sent it to the 3d printer you're just on "day 2" of your multi-year development effort.  That's when you start finding out, not just the places where your parameterization was off, but all the places where "unknown unknowns" start to get you.

Although some small amount of testing might be shared among your different parameterized designs (say, you've characterized the actual material characteristics of your 3d-printed parts well enough that you can feed the results back into your parameterization), the vast majority of the work needs to be repeated for each set of parameters.  Practically speaking, it's not worth it: the amount of effort to make a parameterization that yields an actual production-quality engine at two different parameter values is greater than the effort to just design two distinct engines from scratch.  And by forcing the parameterization you're missing out on opportunities for specializing the designs.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 04/11/2017 05:46 pm
I would have expected there was already a "paramaterized" version of the Raptor, and then they did a "best fit" function on all the theoretical results, to get elon quotes like,

" Raptor TWR Optimization is settling on a surprisingly low thrust, even including mounts for additional engines"

(hopefully I didnt mangle that too bad, it's from memory)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Oersted on 04/11/2017 06:11 pm
You can scale the engine but you cannot scale the molecular and heat-transfer properties of the fuel and oxidiser.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 04/11/2017 09:16 pm
Are we taking into account SpaceX's advanced CFD system wrt to how fast they can scale a design? Combustion instability is addressed starting about 05:30 in this 2015 video. I assume today they're using NVIDIA's supercomputer since Tesla is using an automotive variant of it.

https://www.youtube.com/watch?v=txk-VO1hzBY
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 04/11/2017 10:22 pm
OK, so this was precisely the discussion I've been trying to provoke about this for some time.

TomH ... I should have saved you some typing of your excellent scaling example, I have a degree in physics and two in EE, have worked on computational fluid models (ocean, atmosphere, supersonic flow), and I grok scaling laws :-)

The expert consensus seems to be that even with knowledge of the physical scaling laws that govern engine design and the advances in computer modeling of structures and combustion, it is still a lot of work.

Is combustion instability the biggest issue? It seems like structures, pressure vessels, piping, turbines, and pumps are pretty amenable to evaluation by computational methods.

Once you have enough modeling to get somewhat close to reality, you can start to thing of a parametric model that generates an engine design. It would have a lot of parameters to cover even things like component placement. Then you can automate the initial design process with a genetic algorithm system. Fitness of a particular design means it passes basic structural tests and has top scores for CFD flow and weight of materials used. That process can examine a large state space and come up with potential starting points for the designer to use. Could be more trouble than it's worth, but I think nearly all the pieces have to be present already. GA driver to control the search is not hard to do (I've built two different ones). Parametric generation of the piping and structures is the missing piece but you could have a crew of interns doing that in CAD :-)

Could be one explanation of how they came up with the Raptor design layout and the lox-side turbopump integrated into the combustion chamber.

OK, beat me up, I am a compiler guy, parallel programmer, chip designer but NOT a rocket engineer.

I have spent my career generating parametric models of aircraft and rockets. Most of the rocket engine physics models exist, but it takes considerable work to set up the functional dependencies and integrate them together. Once you have a model then, you explore your design space. GA can certainly work.  I use Pareto Fronts a lot since the engine has multi-objectives. As you explore your design space, you always find shortcomings, so you are always tinkering with the model.

I have a very basic Raptor model which currently contains the following:
 - CEA chemistry
 - Combustor sized by combustion characteristic length
 - Rao Nozzle model with viscous losses
I am working on integrating:
 - turbopump models
This leaves:
 - pre-burners
 - injectors
 - coolant model
 - valves
 - engine controls
 - transient models
 - failure mode models
... and a whole bunch more..... I'll never get to these.

I estimate (ROM) that a complete, multi fidelity, Raptor parametric model (which I am sure SpaceX has) would take somewhere around 10 to 20 man-years to develop and would cost $3-6 million. This assumes that all needed component models are available (most are) and just need to be integrated together, and developed. Developing the model and validating it will take time.

 Now, this is just the model. We haven't even started the component and engine hardware development. Once the testing is started, the model will be updated to stay synced up with the  test data.  ROM staged combustion engine development costs of $250 - 500 million?

The model will be very useful for any subsequent engine development, but most of the testing has to be done again.

Combustion stability and vibrations in general are common problems. These also are often interrelated. Also minor manufacturing details can have large impacts on component life. By large, I mean as much as an order of magnitude. We know these sorts of things are going to happen, but cannot predict them very well. So we test and test and test.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 04/12/2017 01:14 am
And, it's largely 3-d printed.

The only information I've seen for the 3D printing is 40% of the demonstrator engine by mass, and the notes from Elon's NRO talks that 3D printing doesn't work nearly as well for Raptor as it does for SuperDraco.  They can't just scale it up a little and hit the Print button.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/12/2017 03:08 am
And, it's largely 3-d printed.

The only information I've seen for the 3D printing is 40% of the demonstrator engine by mass, and the notes from Elon's NRO talks that 3D printing doesn't work nearly as well for Raptor as it does for SuperDraco.  They can't just scale it up a little and hit the Print button.
Yup. And 3D printing produces worse strength than forging (for example). And you're limited in your alloy selection.

And just in general: parametric design works fantastic with simple objects. But as soon as you get to a certain level of complexity, a fully parametric design simply isn't feasible. It gets super complicated, and you get constraints that screw up under certain conditions, and at some point you'll get tired of fighting your model and just redo parts of it from scratch.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rei on 04/13/2017 01:39 pm
And, it's largely 3-d printed.

The only information I've seen for the 3D printing is 40% of the demonstrator engine by mass, and the notes from Elon's NRO talks that 3D printing doesn't work nearly as well for Raptor as it does for SuperDraco.  They can't just scale it up a little and hit the Print button.
Yup. And 3D printing produces worse strength than forging (for example). And you're limited in your alloy selection.

We should not make the mistake of assuming that all 3d printing technologies are the same. ;) For example, the DMG Mori Lasertec 65:

http://be-nl.dmgmori.com/blob/334060/67241acc5e196393c59bb68002da7c56/pl1uk15-lasertec-65-3d-pdf-data.pdf
http://en.dmgmori.com/blob/120872/cc1b707f03ee3c2b0bfc81d22c3442ca/pl0uk13-lasertec-series-pdf-data.pdf
https://www.youtube.com/watch?v=L3CkzQQFZXs

First off, it's both CNC and printing on the same system, so you can start out with an existing shape and mill elements down, then add onto it.  Secondly, it's laser spraying, not powder bed.  So you don't have to lay down layers across a build, it has a continuous, rapid stream of powder which it melts with a laser as it impacts.  The high speed of the particles means that they also compact as they impact, yielding excellent material properties. The CNC side can mill off all 3d print marks, while the laser can engrave tiny details (holes, etc). The combination of CNC with additive manufacturing means that you can even machine internal areas that normally would be inaccessible. The potential range of materials you can print from is basically unlimited, anything that you can suspend in a dust and which will attach with some combination of impact force and heat. They've validated it with among other things stainless, inconel, bronze, brass, chrome-cobalt-molybdenum alloys, tool steel, stellite, and even tungsten carbide.  Multiple materials printed onto the same part. And part sizes up to half a meter diameter.

We're not talking Makerbots here  ;)

Even if for some reason the quality wasn't right, or you wanted to focus on mass production, you can always use the 3d printer to make molds / die heads / etc for parts. 

It doesn't state, but I wonder if you can "resume" a previous build.  If so, you could take your previously-built engine and tweak its geometry without having to print a new one from scratch (since, again, it can both add and subtract).  Now that would be some fast manufacturing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/13/2017 05:26 pm
Um, yeah, I know all about those types of additive manufacturing. It is, in fact, my job.

Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

Again, additively manufactured metal parts are significantly inferior to forged metal parts.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 04/13/2017 05:35 pm
Um, yeah, I know all about those types of additive manufacturing. It is, in fact, my job.

Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

Again, additively manufactured metal parts are significantly inferior to forged metal parts.

The material as manufactured with additive is inferior to forging, but part properties are a function of both material and geometry. AM allows geometries that are infeasible or completely impossible with forging. So it's possible to make a part with AM that is far superior to a forging serving the same purpose - especially for extremely complex integrated parts, like Raptor appears to use.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RonM on 04/13/2017 05:44 pm
As the old saying goes, use the right tool for the job.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 04/13/2017 06:07 pm
Um, yeah, I know all about those types of additive manufacturing. It is, in fact, my job.

Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

Again, additively manufactured metal parts are significantly inferior to forged metal parts.

The material as manufactured with additive is inferior to forging, but part properties are a function of both material and geometry. AM allows geometries that are infeasible or completely impossible with forging. So it's possible to make a part with AM that is far superior to a forging serving the same purpose - especially for extremely complex integrated parts, like Raptor appears to use.

I think a more salient comparison is between casting and AM. You can cast just about anything, but the cost/difficulty really shoots up when casting more complex parts. I think AM and casting produce parts with similar properties these days, though the state of the art is a rapidly moving target for AM.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/13/2017 06:14 pm
Yeah, for small cast parts, additive is a big threat. Strength can be greater for additive. Really expensive, but not a problem for low part count runs.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rei on 04/17/2017 09:12 am
Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

Again, additively manufactured metal parts are significantly inferior to forged metal parts.

https://link.springer.com/article/10.1007%2Fs00170-011-3423-2?LI=true

Quote
Tensile mechanical properties of selective laser-melted Hastelloy® X alloy in as-deposited condition and after hot isostatic pressing (HIP) have been studied at ambient and elevated temperatures. Room temperature four-point bending and tension–tension fatigue properties have also been investigated in as-deposited condition and after HIP. The yield strength of the as-deposited selective laser-melted Hastelloy® X specimen is higher than the heat-treated (hot forged) samples. The ultimate strength is also higher than that of the hot forged samples while the elongation property is lower. This can be attributed to its ultrafine microstructure caused by rapid solidification, which is characteristic of the selective laser melting process. It is also found that the mechanical properties (tensile and fatigue) do not vary with samples built in different bed locations.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/17/2017 03:41 pm
Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

Again, additively manufactured metal parts are significantly inferior to forged metal parts.

https://link.springer.com/article/10.1007%2Fs00170-011-3423-2?LI=true

Quote
Tensile mechanical properties of selective laser-melted Hastelloy® X alloy in as-deposited condition and after hot isostatic pressing (HIP) have been studied at ambient and elevated temperatures. Room temperature four-point bending and tension–tension fatigue properties have also been investigated in as-deposited condition and after HIP. The yield strength of the as-deposited selective laser-melted Hastelloy® X specimen is higher than the heat-treated (hot forged) samples. The ultimate strength is also higher than that of the hot forged samples while the elongation property is lower. This can be attributed to its ultrafine microstructure caused by rapid solidification, which is characteristic of the selective laser melting process. It is also found that the mechanical properties (tensile and fatigue) do not vary with samples built in different bed locations.
What's the ultimate tensile strength in MPa of this printed sample?

I'm distrustful when actual figures are not given in the summary.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 04/17/2017 03:49 pm
https://link.springer.com/article/10.1007%2Fs00170-011-3423-2?LI=true

Quote
Tensile mechanical properties of selective laser-melted Hastelloy® X alloy in as-deposited condition and after hot isostatic pressing (HIP) have been studied at ambient and elevated temperatures. Room temperature four-point bending and tension–tension fatigue properties have also been investigated in as-deposited condition and after HIP. The yield strength of the as-deposited selective laser-melted Hastelloy® X specimen is higher than the heat-treated (hot forged) samples. The ultimate strength is also higher than that of the hot forged samples while the elongation property is lower. This can be attributed to its ultrafine microstructure caused by rapid solidification, which is characteristic of the selective laser melting process. It is also found that the mechanical properties (tensile and fatigue) do not vary with samples built in different bed locations.

I think this still does not address at all the comparison to a conventionally manufactured part that starts with a forged blank. Seems to me the summary calling the HIP treated part "hot forged" is confusing things.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 04/18/2017 12:00 pm
Surely comparing the different process strengths is mostly irrelevant - just use the one that is strong enough? And if its AM, then fill yer boots?

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 04/18/2017 01:31 pm
Surely comparing the different process strengths is mostly irrelevant - just use the one that is strong enough? And if its AM, then fill yer boots?
Anything can be strong enough if you make it beefier...

Saving mass is not the only consideration, but it's right up there at the top of the list.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rei on 04/18/2017 03:17 pm
Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

Again, additively manufactured metal parts are significantly inferior to forged metal parts.

https://link.springer.com/article/10.1007%2Fs00170-011-3423-2?LI=true

Quote
Tensile mechanical properties of selective laser-melted Hastelloy® X alloy in as-deposited condition and after hot isostatic pressing (HIP) have been studied at ambient and elevated temperatures. Room temperature four-point bending and tension–tension fatigue properties have also been investigated in as-deposited condition and after HIP. The yield strength of the as-deposited selective laser-melted Hastelloy® X specimen is higher than the heat-treated (hot forged) samples. The ultimate strength is also higher than that of the hot forged samples while the elongation property is lower. This can be attributed to its ultrafine microstructure caused by rapid solidification, which is characteristic of the selective laser melting process. It is also found that the mechanical properties (tensile and fatigue) do not vary with samples built in different bed locations.
What's the ultimate tensile strength in MPa of this printed sample?

I'm distrustful when actual figures are not given in the summary.

You could read more than the summary if you wanted, it's not an incredibly long paper  ;)  923-937 MPa, depending on where on the sample they tested.

Quote from: acsawdey
Seems to me the summary calling the HIP treated part "hot forged" is confusing things.

There's actually an additional category in there: SLM, SLM + HIP, and hot forged / no SLM.  SLM has the strongest UTS, followed by SLM + HIP (838-845 MPa), followed by hot forged (767 MPa). The images of the microstructure in figure 5 are telling; it makes very fine, very regular dendrites surrounded by precipitates, with the dendrites oriented in the building direction. After HIP the dendrites coarsen and become more irregular, while in the purely hot forged version, the microstructure is coarse grains.
 
HIP did however improve the fatigue life by removing cracks, decreasing porosity, eliminating embedded unmelted powder, etc. But it comes at a cost of tensile strength.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 04/18/2017 03:42 pm
You could read more than the summary if you wanted, it's not an incredibly long paper  ;)  923-937 MPa, depending on where on the sample they tested.

US$ 39.95 to read a 7-page paper? No thanks. But thank you for giving us a few numbers.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 04/18/2017 04:01 pm
Here's another paper you can grab a pdf of, it references the one Rei linked.

http://www.gruppofrattura.it/ocs/index.php/ICF/icf13/paper/view/11306/10685 (http://www.gruppofrattura.it/ocs/index.php/ICF/icf13/paper/view/11306/10685)

Shows the properties are strongly anisotropic with respect to the build direction.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DanielW on 04/18/2017 05:09 pm
Surely comparing the different process strengths is mostly irrelevant - just use the one that is strong enough? And if its AM, then fill yer boots?
Anything can be strong enough if you make it beefier...

Saving mass is not the only consideration, but it's right up there at the top of the list.

This is not true. There will always be important properties involved in "strength" that don't scale with "beefy" This is especially true for anythings that requires cooling.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 04/19/2017 01:49 am
This is a fascinating discussion. Can we draw any conclusions? How likely is it that AM state of the art will advance fast enough to rival forging by the time Raptor goes into serial production?  And even if not, SpaceX optimizes for cost. In this case, weight has a big leverage, presumably, but does that change the answer at all? 

Not sure if there's a better thread but maybe?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rei on 04/19/2017 12:30 pm
You could read more than the summary if you wanted, it's not an incredibly long paper  ;)  923-937 MPa, depending on where on the sample they tested.

US$ 39.95 to read a 7-page paper? No thanks. But thank you for giving us a few numbers.

If you don't have a subscription and can't get to a place that does, there's always sci-hub  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/19/2017 01:18 pm
I read the article. AM parts get higher strength than regular parts, but if you cold forge (cold draw) the metal, you get 1100MPa ultimate strength, which is a good 20% stronger than the figure they use in the paper (780MPa, I think?). Heat aging the metal also helps a lot.

So I feel vindicated. The right kind of forging definitely produces a much stronger part than a mere AM part, even if you HIP the AM part.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rei on 04/19/2017 07:58 pm
I read the article. AM parts get higher strength than regular parts, but if you cold forge (cold draw) the metal, you get 1100MPa ultimate strength, which is a good 20% stronger than the figure they use in the paper (780MPa, I think?). Heat aging the metal also helps a lot.

So I feel vindicated. The right kind of forging definitely produces a much stronger part than a mere AM part, even if you HIP the AM part.

Reference to that 1100 MPa figure just from cold rolling, if you would. I've been checking a variety of references for Hastelloy X and the only ones that show figures that high are from tempering.  And you can temper 3d prints, just like you can temper forged products.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Navier–Stokes on 04/21/2017 11:44 pm
New job posting for a Raptor Test Specialist (http://jobs.jobvite.com/spacex/job/oRr74fwd) at McGregor:
Quote
Responsibilities:
*    Work with design engineers to develop and document test procedures
*    50% hands on working with hardware, 50% control systems/operation work
*    Perform tests according to procedure
*    Design fixtures and adaptors needed to perform tests
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 04/22/2017 12:29 am
Any significance to specifying flight assemblies & hardware?

Quote
PREFERRED SKILLS AND EXPERIENCE:
>
Experience working on flight critical aerospace assemblies
>
ADDITIONAL REQUIREMENTS:

General physical fitness is required for some work areas, flight hardware is typically built in tight quarters and physical dexterity is required
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 04/22/2017 01:35 pm
I wouldn't get too excited about something likely massaged by HR
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: robert_d on 05/04/2017 02:04 am
My question is what conditions/factors must be accounted for if this new engine is to be restartable?
Will there be a separate restartable version? Does performance suffer overall? Is there extra weight involved for other equipment/fluids? What about power required before the engine can produce any of its own?

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/04/2017 11:18 pm
My question is what conditions/factors must be accounted for if this new engine is to be restartable?
Will there be a separate restartable version? Does performance suffer overall? Is there extra weight involved for other equipment/fluids? What about power required before the engine can produce any of its own?

They said it was spark ignited. The sparks probably ignite ignition torches which in turn ignites the pre-burners and the main chamber.  You can see the ignition leads on their CAD model.

This ignition approach would make all Raptors restartable assuming their propellants had enough head pressure.
Head pressure and an electrical power source is all that is required to start.

The start sequence is something like the following:
- crack valves and dribble in propellants to pre-chill the engine.
- open valves and propellants flow into their respective pre-burners.
- spark ignites stoichiometric mixture in torches.
- torches ignite pre-burners
- pre-burner exhaust spins turbines attached to propellant pumps. (one for methane, one for LOX)
- main chamber torch ignites gaseous propellants entering chamber.
- pumps start increasing pressure above head pressure and quickly climb to design pressure.

This requires detailed understanding of the combustion processes and the dynamics of the pumps, turbines and valves. It is a tightly choreographed dance.

John


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Apollo100 on 05/08/2017 05:52 pm
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

The Integrated Powerhead Demonstrator used liquid hydrogen propellant, so yes of course SpaceX must have changed the design for Raptor. Thanks for the pointer -- it was fascinating to read about IPD. I wonder how many Aerojet-Rocketdyne engineers are working at SpaceX now?

Hydrogen seems to behave pretty similarly to methane with regards to engine operation. Most of the methalox engines fired to date have been lightly modified hydrolox engines, not purpose-built designs.

Though I doubt there is much IPD heritage in Raptor

Thanks for the replies and apologies for the delayed response.... Given that SX acquired the IPD Final report, all of the drawings, and all of the hardware, I would imagine that there is quite a bit of IPD heritage in the Raptor engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 05/09/2017 12:42 pm
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

The Integrated Powerhead Demonstrator used liquid hydrogen propellant, so yes of course SpaceX must have changed the design for Raptor. Thanks for the pointer -- it was fascinating to read about IPD. I wonder how many Aerojet-Rocketdyne engineers are working at SpaceX now?

Hydrogen seems to behave pretty similarly to methane with regards to engine operation. Most of the methalox engines fired to date have been lightly modified hydrolox engines, not purpose-built designs.

Though I doubt there is much IPD heritage in Raptor

Thanks for the replies and apologies for the delayed response.... Given that SX acquired the IPD Final report, all of the drawings, and all of the hardware, I would imagine that there is quite a bit of IPD heritage in the Raptor engine.
I wouldn't. SpaceX learned the lessons and will implement the solutions in their own way.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 05/09/2017 12:54 pm
So, what is the proposed thrust SL and Vacuum?  I've seen it all over the map.  In pounds thrust, please.  I'm retired and grew up and used the English system all my life.  I compare it to old engines from the 1960's like the F-1 and H-1, etc. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 05/09/2017 02:13 pm
So, what is the proposed thrust SL and Vacuum?  I've seen it all over the map.  In pounds thrust, please.  I'm retired and grew up and used the English system all my life.  I compare it to old engines from the 1960's like the F-1 and H-1, etc.

R SL  685,000 LBS   3050 KN

Rvac  787,000 LBS   3500 KN

Source ITS presentation Sept 2016
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 05/10/2017 07:05 pm
That is more than I thought.  I though it was about 550,000 lbs. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 05/13/2017 05:33 pm
That is more than I thought.  I though it was about 550,000 lbs.

That was the # announced years before the September reveal.  Even before that it was up to F-1 levels.
In the BFR threads here I predicted the thrust upgrade in Elon's reveal and made the obvious (sun to rise in East tomorrow) prediction that all BFR #s would continue to evolve long after those on the September Tablets From The Mount.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 06/05/2017 01:21 pm
Is the Raptor being built with Merlin tooling?

I've read the Raptor is about the same size as Merlin.  I also read that Raptor will be 3 - 4,000 psi chamber pressure.  What is Merlin's chamber pressure?

If Raptor is going to have a much higher chamber pressure, the turbo pumps will be much stronger right?

I know my company has ran CNG in vehicles at 2-3000 psi to make a cylinder (welding size), handle the equivelant of 4-5 gallons of gasoline.  The compressor for a fleet of 25 vehicles is large.  I know the methane is liquid or LNG for the rocket.  Is SpaceX going to manufacture these turbo pumps or buy them off shelf?

If Raptor is going to use Merlin's tooling, how are they going to build the chamber for Raptor since it has to be stronger (thicker?) to handle the higher pressures?

I know the engine also seems to be much smaller than the BE-4.  Would it be lighter, thus higher thrust/weight ratio? 

Also, what is the throttle range going to be?

I know this stuff is in all the various threads somewhere.  I just wanted to get it all together in one place to have a more complete knowledge of the Raptor and it's capabilities.

Thanks
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Ictogan on 06/05/2017 02:12 pm
Is the Raptor being built with Merlin tooling?
Probably not, given that it's a very different engines(different fuels, different cycles, much higher chamber pressure).

I've read the Raptor is about the same size as Merlin.  I also read that Raptor will be 3 - 4,000 psi chamber pressure.  What is Merlin's chamber pressure?
Merlin 1D originally had 1,410psi chamber pressure, but it's been uprated two times since then. Now it's probably around 1,800psi.

If Raptor is going to have a much higher chamber pressure, the turbo pumps will be much stronger right?
More powerful engines usually means more powerful turbopumps.

I know my company has ran CNG in vehicles at 2-3000 psi to make a cylinder (welding size), handle the equivelant of 4-5 gallons of gasoline.  The compressor for a fleet of 25 vehicles is large.  I know the methane is liquid or LNG for the rocket.  Is SpaceX going to manufacture these turbo pumps or buy them off shelf?
AFAIK turbopumps for rocket engines are usually custom designs and manufactured by the engine manufacturers.

I know the engine also seems to be much smaller than the BE-4.  Would it be lighter, thus higher thrust/weight ratio?
We don't know, but it seems likely as Raptor will supposedly have a better TWR than M1D, which currently holds the record in that department. 

Also, what is the throttle range going to be?
Don't think there has been any information on this
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 06/05/2017 03:46 pm
Thanks, hopefully some of the others will be answered soon.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DOCinCT on 06/05/2017 08:46 pm
Thanks, hopefully some of the others will be answered soon.
According to the presentation from last year, Raptor will have a throttle range of 20 to 100% of thrust.  That is consistent with 3 engines landing a relatively empty upper stage (barely).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hamerad on 06/21/2017 02:50 am
Quote
Will be full regen cooled all the way out to the 3 meter (10 ft) nozzle diameter. Heat flux is nuts & radiative view factor is low.


https://twitter.com/elonmusk/status/877341165808361472 (https://twitter.com/elonmusk/status/877341165808361472)

Elon Musk on twitter. Probably already rumoured but now confirmed
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 06/21/2017 04:30 am
Quote
Will be full regen cooled all the way out to the 3 meter (10 ft) nozzle diameter. Heat flux is nuts & radiative view factor is low.


https://twitter.com/elonmusk/status/877341165808361472 (https://twitter.com/elonmusk/status/877341165808361472)

Elon Musk on twitter. Probably already rumoured but now confirmed

Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 06/21/2017 06:17 am
It also means it is a quite robust nozzle, as needed to survive reentry. Somewhat more heavy too. But that is the price for a reusable vac nozzle.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: nacnud on 06/21/2017 06:45 am
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)

Does it mean it has been shrunk? You could read the statement to mean that the last 2 feet of diameter increase is radiatively cooled.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 06/21/2017 07:08 am
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)

While this might be true, it's not something Elon wrote in this Tweet. The nozzle might well extend beyond the 3m diameter mark, but then with radiative cooling.

Edit: Just what nacnud wrote.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 06/21/2017 09:42 am
What expansion ratio does this 3m correspond to then?  40?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 06/21/2017 10:03 am
What expansion ratio does this 3m correspond to then?  40?
3m Raptor nozzle at ER 40 would produce F-1 class thrust so  3m nozzle dia. for the announced thrust of Raptor would indicate an ER of around 100-150.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 06/21/2017 10:46 am
What expansion ratio does this 3m correspond to then?  40?
3m Raptor nozzle at ER 40 would produce F-1 class thrust so  3m nozzle dia. for the announced thrust of Raptor would indicate an ER of around 100-150.

So, this would be the vac version?  Multiple engines on a second stage would preclude the radiative cooling used on single-engine Falcon S2s.

Note: Just realized this is all about Raptor Vac...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 06/21/2017 04:06 pm
What expansion ratio does this 3m correspond to then?  40?
3m Raptor nozzle at ER 40 would produce F-1 class thrust so  3m nozzle dia. for the announced thrust of Raptor would indicate an ER of around 100-150.

To narrow this down just a little, it should be an ER of 130 - 140 at 3m diameter to hit the expected thrust (3500 kN) and ISP (382 sec) with a full ER of 200.

They could lower the thrust target and get a ER of 200 with a 3 m nozzle.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 06/22/2017 03:04 pm
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)

Does it mean it has been shrunk? You could read the statement to mean that the last 2 feet of diameter increase is radiatively cooled.

That would a somewhat contrived reading, IMHO. 'Full regen' implies the whole thing, as opposed to simply 'regen' or 'partial regen'; 'nozzle diameter' implies the whole thing, not a measurement of a point on the nozzle.

It seems slightly more plausible that he simply forgot the .7
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 06/22/2017 04:06 pm
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)

Does it mean it has been shrunk? You could read the statement to mean that the last 2 feet of diameter increase is radiatively cooled.

That would a somewhat contrived reading, IMHO. 'Full regen' implies the whole thing, as opposed to simply 'regen' or 'partial regen'; 'nozzle diameter' implies the whole thing, not a measurement of a point on the nozzle.

It seems slightly more plausible that he simply forgot the .7

3.7 meters is 12 feet 2". Musk specifically gave an Imperial measurement of 10 feet in brackets, which is 3.048 meters.

Guess it comes down to if you think he gave 2 wrong/offhand numbers, or if it's changed and is now slightly smaller.

I'd take door #2 until proven otherwise.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 06/23/2017 12:48 am
This is getting absurd. People in the real world use rounded off numbers when speaking in vernacular context. In those circumstances, you cannot take a number with one significant digit and extrapolate a conversion to another number with four significant digits.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 06/23/2017 02:29 am
This is getting absurd. People in the real world use rounded off numbers when speaking in vernacular context. In those circumstances, you cannot take a number with one significant digit and extrapolate a conversion to another number with four significant digits.
And recall that the 12' number was based on measuring screenshots of slides in a presentation video.  Arguing over the last foot or two seems pointless.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 06/23/2017 08:47 am
Given what Gwynne Shotwell said yesterday about final thrust of Raptor, the 3m diameter seems to fit quite well.

Plus 3m means the nozzle fits into the Falcon interstage (duck and cover).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 06/23/2017 12:46 pm
Haven't picked engine size for Mars vehicle yet, will be 2-3 (probably less than 3) times the size of the sub-scale Raptor
EM had selected the Raptor size at 3.05MN SL when ITS was announced at IAC2016. Now SpaceX say they have not selected the Raptor size yet. SpaceX should select a larger not smaller Raptor size for ITS to stop the engine no. of the ITS system spiraling out of control. There are rumors that the final ITS design may end up larger than that announced at IAC2016.

I believe they work at the engine and home in on a size that is T/W efficient as well as efficient to build. Not on a preset thrust.

If they start with building a smaller ITS first they may later go for a bigger one. Bigger maybe as in 15m diameter, not necessarily more thrust. Bigger should mean they can land a larger payload on Mars with more refuelling runs and cargo delivery runs to LEO. That way the same hardware sent to Mars can do more. More work done with hardware doing many flights at the earth end. Addressing partly the concerns raised by Robert Zubrin.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 06/23/2017 03:34 pm
Given what Gwynne Shotwell said yesterday about final thrust of Raptor, the 3m diameter seems to fit quite well.

Plus 3m means the nozzle fits into the Falcon interstage (duck and cover).

I think it also makes sense because of the need to fit 6 Vac Raptors in the ITS spacecraft. With larger than 3 m nozzles, it makes it very tricky to fit everything inside. (in the schematic revealed, the nozzles were so large they practically touched each other and the outer heat shield) This is much more realistic, since they also need to fit the Vac Raptors inside the moving tail flaps of the heat shield. (shuttle-style)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 07/02/2017 10:49 am
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?

I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 07/02/2017 12:35 pm
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?

I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.

The BFS will also have to do a TMI burn after a long loiter in LEO. So they either need some sort of onboard propellant cooling capability, or the Raptors will have to be omnivorous and take whatever temperature propellants they are given.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 07/02/2017 12:42 pm
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?

I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.

The BFS will also have to do a TMI burn after a long loiter in LEO. So they either need some sort of onboard propellant cooling capability, or the Raptors will have to be omnivorous and take whatever temperature propellants they are given.

I think omivorous will be the way they have to go. From reading about supercooling on earth it seems very energy intensive and lots of equipment.
Nitrogen baths. etc.
On the other hand you have vacuum up there and that should provide any temperature you want with lower the pressure enough to get to the boiling point at supercooled temperatures.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 07/02/2017 03:03 pm
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?

I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.

The BFS will also have to do a TMI burn after a long loiter in LEO. So they either need some sort of onboard propellant cooling capability, or the Raptors will have to be omnivorous and take whatever temperature propellants they are given.

I think omivorous will be the way they have to go. From reading about supercooling on earth it seems very energy intensive and lots of equipment.
Nitrogen baths. etc.
On the other hand you have vacuum up there and that should provide any temperature you want with lower the pressure enough to get to the boiling point at supercooled temperatures.


I am especially worried about keeping it cold on the Martian surface. For the early missions it would be much simpler to accumulate the propellants in the spacecarft's tanks during production - and store there till the lauch. Above it is the inhabited crew quarters, which has to be kept warm... Pure evaporation cooling means loss.

On the other hand, the omnivorous option means significant loss of DeltaV. It is not clear, how the DeltaV values were calculated for the lecture.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 07/02/2017 04:13 pm
Maybe, the simplest solution is to tune the booster stage Raptors for chilled propellants and the spacecarft's ones to the normal propellant densitty. Obviously, this is even worse performance-wise.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/03/2017 01:58 am
Maybe, the simplest solution is to tune the booster stage Raptors for chilled propellants and the spacecarft's ones to the normal propellant density. Obviously, this is even worse performance-wise.
Propellant density and starting are both helped by cooling propellants to below their boiling points. Isp doesn't change.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/03/2017 02:02 am
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 07/03/2017 02:58 am
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   
According to some, turbines can move more propellant if it's denser. I guess that means turbine speed is more of a limiting factor than turbine power in this case.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 07/03/2017 03:25 am
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   

Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 07/03/2017 03:42 am
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   

Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?

Why wouldn't it be? Otherwise there would be no performance gain.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 07/03/2017 06:08 am
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   

Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?

lower vapour pressure of subcooled propellants means that pressure on suction sides of pump impellers can go lower before inducing damaging cavitation - so you can make more highly loaded pumps, or spin them faster for higher pressure ratios, or reduce the necessary tank pressure (pump inlet pressure) for tank mass savings.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/04/2017 10:07 pm
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   

Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?

Why wouldn't it be? Otherwise there would be no performance gain.

Because subcooling prop on-orbit or on the Mars surface is somewhat more challenging than on the launch pad, and the leg from staging to Earth orbit is more challenging than LEO to mars surface or Earth return empty.

I don't think subcooled props are strictly necessary for TMI or Earth return, but if they can solve long term boiling storage than long term subcooled isn't all that much more difficult, so they might do it. It does help with fast transits and next-synod reuse.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 07/04/2017 10:49 pm
So lox at 66k vapor pressure is .029 bar. So Mars surface at .005 bar and space at 0 bar will cool lox just fine. All you have to do is collect the gaseous oxygen boil off at .029 bar and compress it in a Linde liquefaction cycle and put it back in the tank. Takes energy but not the special stuff you do at 1 bar. Like vacuum pumps or LN2 baths.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 07/05/2017 12:29 pm
So lox at 66k vapor pressure is .029 bar. So Mars surface at .005 bar and space at 0 bar will cool lox just fine. All you have to do is collect the gaseous oxygen boil off at .029 bar and compress it in a Linde liquefaction cycle and put it back in the tank. Takes energy but not the special stuff you do at 1 bar. Like vacuum pumps or LN2 baths.

Good to know. What about LEO?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 07/05/2017 12:34 pm
So lox at 66k vapor pressure is .029 bar. So Mars surface at .005 bar and space at 0 bar will cool lox just fine. All you have to do is collect the gaseous oxygen boil off at .029 bar and compress it in a Linde liquefaction cycle and put it back in the tank. Takes energy but not the special stuff you do at 1 bar. Like vacuum pumps or LN2 baths.

Good to know. What about LEO?

My comment just has to do with active cooling using standard refrigeration equipment. The difference between LEO and deep space has to do with passive cooling and sun shades. Passive cooling would not work as well in LEO because of the radiation from earth. On the sunside there is not much sky to expose your passive radiators to. On the night side there is half the sky with no thermal radiation. Obviously a active cooling system still needs radiators to get rid of the heat so it will work better depending on the sky it is radiating to.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/05/2017 12:42 pm
My understanding was that they use subcooled propellant to eliminate cavitation in the turbopump. Full power would mainly be needed on earth ascent, both in the first and second stage. That can be provided with subcooled propellant on tanking. Can cavitation also be avoided with some throttling? For TMI full power would not be needed, also on Mars ascent it is not as essential.

If subcooled is needed in every phase, can you calculate, how much propellant would be wasted to cool propellant a few degrees below sea level pressure boiling temperature?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/05/2017 12:59 pm
My understanding was that they use subcooled propellant to eliminate cavitation in the turbopump. Full power would mainly be needed on earth ascent, both in the first and second stage. That can be provided with subcooled propellant on tanking. Can cavitation also be avoided with some throttling? For TMI full power would not be needed, also on Mars ascent it is not as essential.

If subcooled is needed in every phase, can you calculate, how much propellant would be wasted to cool propellant a few degrees below sea level pressure boiling temperature?

No propellant is required, just slightly better insulation, cooling and radiators than it takes to keep it from all boiling away in the first place. It depends on the location and cooling strategy. In deep space, passive cooling is likely more than sufficient for ZBO with methalox. On Mars, they will need some active cooling, but there are some heatsinks available so large radiators might not be necessary. The toughest place to do ZBO is in LEO.

Throttling results in a very slight hit to I_sp in vacuum, but the bigger hit by far is the loss of 17% of the propellant mass for the same tank volume at boiling densities. But ITS still has 5,900 m/s of delta-v with a 300 t payload and boiling methalox, which is enough for a fast-ish transit and landing - so they might decide to do boiling non-ZBO in LEO.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 07/05/2017 05:35 pm
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?   

Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?

Why wouldn't it be? Otherwise there would be no performance gain.

Because subcooling prop on-orbit or on the Mars surface is somewhat more challenging than on the launch pad, and the leg from staging to Earth orbit is more challenging than LEO to mars surface or Earth return empty.

I don't think subcooled props are strictly necessary for TMI or Earth return, but if they can solve long term boiling storage than long term subcooled isn't all that much more difficult, so they might do it. It does help with fast transits and next-synod reuse.

I still don't see the problem. If the propellant is no longer subcooled, it doesn't go away. (conservation of mass and all that) It is still there, just taking up more volume. The engines will certainly be able to handle a bit of temperature range.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/05/2017 09:28 pm
Engine pumps are very sensitive to vapor pressure and  tank pressure (head or otherwise). In space you don't have head pressure. On Mars, you obviously have less than on earth, but better than in space. In space, BFS will probably need to use either sub-cooled propellants, higher tank pressures or boost pumps or some combination. I was surprised that Raptor doesn't have them (yet). :^)

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/05/2017 10:38 pm
Engine pumps are very sensitive to vapor pressure and  tank pressure (head or otherwise). In space you don't have head pressure. On Mars, you obviously have less than on earth, but better than in space. In space, BFS will probably need to use either sub-cooled propellants, higher tank pressures or boost pumps or some combination. I was surprised that Raptor doesn't have them (yet). :^)

John

You better have head pressure whenever the engines are firing whether in space, on Mars, or on Earth - or else you have major problem: you aren't accelerating at all :D
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/06/2017 11:17 am
Of course, after they are started. Starting is the problem.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/06/2017 12:55 pm
Of course.

Didn't Elon say they were using multistage pumps on Raptor? The low pressure pump might be designed to handle lower vapor pressure without cavitation.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/06/2017 02:58 pm
I still don't see the problem. If the propellant is no longer subcooled, it doesn't go away. (conservation of mass and all that) It is still there, just taking up more volume. The engines will certainly be able to handle a bit of temperature range.

Assume you have filled the tank in orbit to capacity with subcooled propellant. Then the temperature drifts to the boiling point. Some of the propellant is going to go away unless you keep the vents closed. In that case it will stay until the tanks burst and it all goes away.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/06/2017 03:20 pm
I still don't see the problem. If the propellant is no longer subcooled, it doesn't go away. (conservation of mass and all that) It is still there, just taking up more volume. The engines will certainly be able to handle a bit of temperature range.

Assume you have filled the tank in orbit to capacity with subcooled propellant. Then the temperature drifts to the boiling point. Some of the propellant is going to go away unless you keep the vents closed. In that case it will stay until the tanks burst and it all goes away.

It will take 6 tanker launches to fill the tanks. If you can keep it subcooled for that long, what's stopping you from keeping it subcooled until you use it?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/06/2017 03:34 pm
It will take 6 tanker launches to fill the tanks. If you can keep it subcooled for that long, what's stopping you from keeping it subcooled until you use it?

Time maybe? They can fill it in a week with daily launches. But if it waits for months in LEO for the Mars window to open it will be hard to keep propellants subcooled without any active measures. Less hard while in interplanetary space away from IR emitting earth.

Edit: I was mostly repying to the "it does not go away".

There are ways to handle it. Fill up to boiling temperature, wait for departure time, with hopefully minimal boiloff. Subcool by opening to vacuum and have a last tanker fill up before departure. One tanker can probably do the topping off for several departing vehicles.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MP99 on 07/06/2017 08:01 pm
Prop tanks would be full (and so benefit from subcooling) only at Earth launch, TMI and Mars launch.

The two launches can rely on GSE to maintain temps.

Perhaps the refuelling tanker also carries subcooling equipment. Maybe that only the final fuelling delivery uses a special variant with the subcooling hardware, circulating prop until temps are low enough to fit in the full prop load.


There is actually a passive way for the tanker to re-cool ITS's prop load - by freezing its payload prop (in separate tanks from launch prop) before launch, then circulating ITS's prop load through it.

It would make tanker GSE a nightmare, though!

Cheers, Martin
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/06/2017 08:03 pm
Of course.

Didn't Elon say they were using multistage pumps on Raptor? The low pressure pump might be designed to handle lower vapor pressure without cavitation.

The methane pump is two stage and probably does not need a boost pump, but the LOX pump appears to be a single stage pump and may need one, or they might just hold higher pressure in the small landing tanks and could use these to start without a boost pump.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JasonAW3 on 07/06/2017 08:54 pm
It will take 6 tanker launches to fill the tanks. If you can keep it subcooled for that long, what's stopping you from keeping it subcooled until you use it?

Time maybe? They can fill it in a week with daily launches. But if it waits for months in LEO for the Mars window to open it will be hard to keep propellants subcooled without any active measures. Less hard while in interplanetary space away from IR emitting earth.

Edit: I was mostly repying to the "it does not go away".

There are ways to handle it. Fill up to boiling temperature, wait for departure time, with hopefully minimal boiloff. Subcool by opening to vacuum and have a last tanker fill up before departure. One tanker can probably do the topping off for several departing vehicles.

Actually, it might not be too much of a problem.  As I understand it, properly "doped", Carbon fiber makes a pretty good insulator.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cferreir on 07/19/2017 07:48 pm
What vehicle will use the Raptor? I know all about the ITS but the Raptor will be done well in advance of ITS and if it's only use is ITS then it seems like the economics of SpaceX won't work. Raptor has got to have more use than that. Is it only me or there is a BIG gap in the SpaceX launch family from F9 to ITS......
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rockets4life97 on 07/19/2017 07:54 pm
What vehicle will use the Raptor? I know all about the ITS but the Raptor will be done well in advance of ITS and if it's only use is ITS then it seems like the economics of SpaceX won't work. Raptor has got to have more use than that. Is it only me or there is a BIG gap in the SpaceX launch family from F9 to ITS......

It looks like SpaceX's plan is for F9, FH, and smaller-version of the proposed BFR/ITS. BFR looks to be the enabler of the satellite constellation and Mars. I'm not sure that Blue Origin's approach for New Glenn and then New Armstrong is so much different in capability when compared to FH and (the new) BFR.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: whitelancer64 on 07/19/2017 08:20 pm
Tune in for Elon Musk's presentation at this year's International Astronautical Congress (IAC) in September!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Formica on 07/24/2017 03:33 pm
The terms "methane" and "liquid natural gas" are sometimes being used interchangeably in discussions of Raptor engines. Is it safe to assume that SpaceX will be removing the other hydrocarbons in LNG to generate pure refined methane? Or is it possible they'll use LNG as is from commercial sources? It seems silly to even ask the question - I presume they will be purifying it - and that LNG is just being used as shorthand for "refined rocket grade liquid methane" in the same way that "kerosene" is sometimes used to refer to RP-1.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Jim on 07/24/2017 03:48 pm
The terms "methane" and "liquid natural gas" are sometimes being used interchangeably in discussions of Raptor engines. Is it safe to assume that SpaceX will be removing the other hydrocarbons in LNG to generate pure refined methane? Or is it possible they'll use LNG as is from commercial sources? It seems silly to even ask the question - I presume they will be purifying it - and that LNG is just being used as shorthand for "refined rocket grade liquid methane" in the same way that "kerosene" is sometimes used to refer to RP-1.

BO is using stock LNG for its engines and vehicles.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 07/24/2017 04:13 pm
The terms "methane" and "liquid natural gas" are sometimes being used interchangeably in discussions of Raptor engines. Is it safe to assume that SpaceX will be removing the other hydrocarbons in LNG to generate pure refined methane? Or is it possible they'll use LNG as is from commercial sources? It seems silly to even ask the question - I presume they will be purifying it - and that LNG is just being used as shorthand for "refined rocket grade liquid methane" in the same way that "kerosene" is sometimes used to refer to RP-1.

I have no further information on this, but I remember a nugget of information that was mentioned during the AMOS-6 investigation and that might be relevant in this context.

Somebody reported, that SpaceX did not use the same, refined and expensive LOX as NASA and used lower purity industrial one instead. Because of this I would assume that SpaceX will use the cheapest fuel acceptable. And if BO can use LNG for the BE-4 (as Jim mentioned), than I would assume, that SpaceX will do the same for Raptor (at least when launching from Earth).

BTW: Is there such a thing as "refined rocket grade liquid methane"? After all, according to astronautix.com there never was a production LOX/LCH4 engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 07/24/2017 04:17 pm
 LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. Commercial can also have up to 1% nitrogen in it.
 It shouldn't be too hard to pay for a little extra processing to get a consistent Methane/Ethane mix with most of the heavier stuff out.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: russianhalo117 on 07/24/2017 04:25 pm
LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. It can also have up to 1% nitrogen in it.
AFAIK, pure CH4 is planned for Raptor flight versions. This was stated previously in tweets during initial Raptor testing at McGregor Complex.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 07/24/2017 04:27 pm
LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. It can also have up to 1% nitrogen in it.
AFAIK, pure CH4 is planned for Raptor flight versions. This was stated previously in tweets during initial Raptor testing at McGregor Complex.
Yeah. I was thinking more of BO even though the question was specifically about Raptor. It's Jim's fault.
 Getting that last trace of Ethane out of raw gas is the hard part.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Jim on 07/24/2017 04:31 pm

Somebody reported, that SpaceX did not use the same, refined and expensive LOX as NASA and used lower purity industrial one instead. Because of this I would assume that SpaceX will use the cheapest fuel acceptable.


It all comes from the same plant in Mims, FL for NASA, ULA, Spacex etc.

The only place where refined LOX was used was for shuttle fuel cell reactant.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Port on 07/24/2017 04:49 pm
LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. It can also have up to 1% nitrogen in it.
AFAIK, pure CH4 is planned for Raptor flight versions. This was stated previously in tweets during initial Raptor testing at McGregor Complex.

elon stated that they'll be using lox and lch4 close to their freezing points, no problems with higher hydrocarbons since they'll freeze out before the rocket is filled (assuming propellant is already at temperature when loaded which only makes sense)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: strangequark on 07/24/2017 05:01 pm

BTW: Is there such a thing as "refined rocket grade liquid methane"? After all, according to astronautix.com there never was a production LOX/LCH4 engine.

Yes. Starting at Grade A, which is oddly enough the lowest quality grade.

For commercial LNG, liquefying gets rid of almost all the higher stuff, and undesirables like hydrogen sulfide. Ethane's the big remaining contaminant, and its properties are fairly similar. Even there, I've gotten commercial LNG, nothing special done to it, that was >99%. For these purposes, it is good enough, provided your source is reasonably consistent.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/24/2017 08:07 pm
If you have a storage tank with liquid natural gas, you can pull mostly pure methane from the middle.  Impurities will rise or fall in the storage tank. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/24/2017 08:08 pm
Will Raptor be the full 685,000 lb thrust engine or will the initial booster use less thrust? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 07/24/2017 09:22 pm
Will Raptor be the full 685,000 lb thrust engine or will the initial booster use less thrust?

Hopefully the next IAC slides may tell us the answers.  Opinions here only.  Ask on L2 and maybe one of the known insiders will tell although I'd wager Elon has a tight grip on leaks for the next 60 days.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/26/2017 01:00 pm
Some Raptor questions, can't find them anywhere else. 

1) How will the existing sub-scale engine be upgraded to a full thrust Raptor?  Is the combustion chamber and turbo pumps the same?  Just increasing speed of the pumps? 

2) Is the sub-scale engine smaller than the full thrust Raptor? 

3) How long will it take to go from sub-scale to full thrust?

4) With the possible revelation of a 9m BFR/ITS, will this use 42 sub-scale engines?  Or use full scale engines?

5) With the above revelation, can 42 sub-scale engines fit in a 9m BFR?

6) Will there possibly be an in between engine from the 225k lb thrust sub-scale to the 685k thrust full scale to power the 9m BFR/ITS? 

Thanks.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/27/2017 12:01 am
Some Raptor questions, can't find them anywhere else. 

1) How will the existing sub-scale engine be upgraded to a full thrust Raptor?  Is the combustion chamber and turbo pumps the same?  Just increasing speed of the pumps? 

2) Is the sub-scale engine smaller than the full thrust Raptor? 

3) How long will it take to go from sub-scale to full thrust?

4) With the possible revelation of a 9m BFR/ITS, will this use 42 sub-scale engines?  Or use full scale engines?

5) With the above revelation, can 42 sub-scale engines fit in a 9m BFR?

6) Will there possibly be an in between engine from the 225k lb thrust sub-scale to the 685k thrust full scale to power the 9m BFR/ITS? 

Thanks.

1&2) It probably has to be physically scaled up. Musk said it will handle a 150:1 nozzle on the test stand, and Shotwell said it's 2 to 3 times less thrust than they need for the Mars vehicle. Taking those in concert very strongly indicates a subscale (roughly 1/2 to 1/3 size) turbopump, thrust chamber, and nozzle throat (at least theoretically) capable of 30 MPa operating pressure.

3) Not quite as long as developing an engine from scratch, but still a while. Probably at least 4 years to first flight.

4) This isn't clear, but IMO full scale makes more sense on the booster. They might go subscale on the upper stage for engine-out ability and landing throttling.

5) Probably, yes. They can always lower the expansion ratio and make the nozzle slightly smaller.

6) What would they need that for?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 07/27/2017 12:11 am
...

3) Not quite as long as developing an engine from scratch, but still a while. Probably at least 4 years to first flight.

...

Since additive manufacturing is used for so much of this engine, making the transition to full scale on this 'scaleable' engine could be much faster than traditional builds.  I'd bet they can have full scale on the test stand within two years of first firing at subscale (NLT September 2018).  A year to test and validate their production processes (less than that from first firing to production run on subscale engine).  A flight qualified engine could be ready two years from now... 3-4 for first flight sounds about right.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/27/2017 12:17 am
Thanks a lot.  Interesting times we live in.  Hope we make it to Mars in my lifetime.  I'm 64.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 07/27/2017 02:13 am
...

3) Not quite as long as developing an engine from scratch, but still a while. Probably at least 4 years to first flight.

...

Since additive manufacturing is used for so much of this engine, making the transition to full scale on this 'scaleable' engine could be much faster than traditional builds. 
>

Putting a finer point on it, then Raptor Jeff Thornburg stated before a Congressional hearing that Raptor is a "highly-scalable" design. Perhaps this is related to it being a full-flow staged combustion engine as well as having some printed parts?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 07/27/2017 02:36 am
Since additive manufacturing is used for so much of this engine, making the transition to full scale on this 'scaleable' engine could be much faster than traditional builds.

We don't actually know how much and which parts of the engine are additively manufactured.  The only number I ever saw was about 40% by mass, and there were those comments from Elon's talk at NRO where he said using additive manufacturing on Raptor didn't work as well as it did on SuperDraco.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/27/2017 06:10 am
A more general question about Raptor. If I recall correctly the methane pump is a two stage pump. Would they use methane from the first stage to feed through the cooling channels? That way the pressure would not be too high and the second stage and the max pressure at the pump outlet could be somewhat smaller as it does not need to account for pressure loss in the cooling channel.

Or am I way off and the two stage is only for better geometry of the pump?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 07/27/2017 06:11 am
Wow.. Didn't expect that going from a subscale engine to full scale takes that long. What is the rational to do the subscale in the first place? It's practically useless if development of subscale plus the transition time to full scale is longer (and therefore more expensive) than going straight to full scale.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/27/2017 06:19 am
Wow.. Didn't expect that going from a subscale engine to full scale takes that long. What is the rational to do the subscale in the first place? It's practically useless if development of subscale plus the transition time to full scale is longer (and therefore more expensive) than going straight to full scale.

The Stennis test stand is not able to handle full scale components. I think the presently tested subscale engine is the scale that was tested at Stennis. The big question is how easily can they scale up from there? Their fluid dynamic simulations are top notch. But there is also the stress factor on materials. Not to forget the manufacturing of bigger components.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 07/27/2017 06:23 am
But they tested the raptor engine at McGregor if I recall correctly.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/27/2017 06:30 am
But they tested the raptor engine at McGregor if I recall correctly.

Yes, the engine is tested in McGregor. But it is based on components tested in Stennis.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 07/27/2017 07:51 am
Ok, but going back to my original question. When developing the fulls scale version, the components cant be tested in Stennis either, following your logic. So the lack of full scale components testing at Stennis doesnt help at all for subscale+fullscale instead of just fullscale. Unless you explicitly want to have the subscale Raptor engine as a separate engine family. But so far, I have not seen such an argument.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 07/27/2017 08:28 am
When developing the fulls scale version, the components cant be tested in Stennis either, following your logic.

Following the capabilities of the Stennis facility.

Building a subscale full engine still makes sense IMO, if they can utilize existing production facilities for Merlin engines to some extent. Like for the combusition chamber and the nozzle.

Establishing it works well on subscale at low cost is a good move before investing in completely new production facilities for the full size. Independently of wether they will ever fly a subscale engine or not.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 07/27/2017 10:43 am
When developing the fulls scale version, the components cant be tested in Stennis either, following your logic.

Following the capabilities of the Stennis facility.

Building a subscale full engine still makes sense IMO, if they can utilize existing production facilities for Merlin engines to some extent. Like for the combusition chamber and the nozzle.

Establishing it works well on subscale at low cost is a good move before investing in completely new production facilities for the full size. Independently of wether they will ever fly a subscale engine or not.

That doesnt add up either. Thats like loosing your key in the darkness and looking for it under the street lamp because thats where you can see stuff. My question is: why make a subscale Raptor in the first place if the full scale still takes so long to develop? It doesnt help full scale that your factory can produce the subscale at all. There must be some other rational behind it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/27/2017 11:26 am
The only reason I can see, is the Air Force money to develop an upper stage metholox engine.  This sub-scale engine can be made for a vacuum engine.  Capability is in the range of J2X or two BE-3's.  Someone said the way it is designed, that it can be scaled up easily.  Hopefully quickly to get the BFR/ITS on the road. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/27/2017 12:18 pm
The only reason I can see, is the Air Force money to develop an upper stage metholox engine.  This sub-scale engine can be made for a vacuum engine.  Capability is in the range of J2X or two BE-3's.  Someone said the way it is designed, that it can be scaled up easily.  Hopefully quickly to get the BFR/ITS on the road.

That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/27/2017 12:42 pm
SpaceX didn't (and apparently still doesn't) know how big they need the large Raptor to be. So they chose a size that allowed them to work with existing facilities at low cost and prove out the engine architecture.

If they hadn't done that, Raptor would still be 6 or 7 years from flying, instead of 3 or 4. If they see a need, they might be able to get the small one flying in closer to 2 years. I doubt they will do a 42 engine small-Raptor booster, but a 9 small-Raptor, 9 meter diameter BFS for suborbital EDL tests would make sense.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 07/27/2017 12:43 pm
The fact that SpaceX has been able to make the progress they have with Raptor, which has been tested much more than BE-4, is in large part due to starting out subscale.

But I suspect SpaceX will use subscale Raptor in at least some capacity for initial operations. Or at very least for ITS prototypes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/27/2017 01:23 pm
Where does SpaceX test this sub-scale engine?  Will they attempt to see what it's maximum thrust can be?

Also, can they test a full scale engine at McGregor?  Or do they need Stennis? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 07/27/2017 01:45 pm
Where does SpaceX test this sub-scale engine?  Will they attempt to see what it's maximum thrust can be?

Also, can they test a full scale engine at McGregor?  Or do they need Stennis? 
SpaceX tests their subscale Raptor at their McGregor test facility in Texas. I think that they will also test the full size Raptor at McGregor as well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 07/27/2017 02:22 pm
The only reason I can see, is the Air Force money to develop an upper stage metholox engine.  This sub-scale engine can be made for a vacuum engine.  Capability is in the range of J2X or two BE-3's.  Someone said the way it is designed, that it can be scaled up easily.  Hopefully quickly to get the BFR/ITS on the road.

That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.

John

Exactly, No one has done a FFSC Methane engine before.  Start small, learn how to run one, then scale up.

Smaller is smaller, faster to build, smaller test stand, less consumables, less damage if it goes boom.

The fact that they may end up with an US engine and that the USAF helped pay for it is a bonus.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RDMM2081 on 07/27/2017 05:51 pm
The only reason I can see, is the Air Force money to develop an upper stage metholox engine.  This sub-scale engine can be made for a vacuum engine.  Capability is in the range of J2X or two BE-3's.  Someone said the way it is designed, that it can be scaled up easily.  Hopefully quickly to get the BFR/ITS on the road.

That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.

John

Exactly, No one has done a FFSC Methane engine before.  Start small, learn how to run one, then scale up.

Smaller is smaller, faster to build, smaller test stand, less consumables, less damage if it goes boom.

The fact that they may end up with an US engine and that the USAF helped pay for it is a bonus.

Also to add because you didn't quite say it in so many words, but CHEAPER if/when it goes boom! 

My understanding is that early engine tests, especially when you are working with a new (to you/your design) engine cycle, there is every expectation that you'll blow at least a couple pieces up.  Blowing up smaller pieces is cheaper than blowing up larger pieces, and faster to rebuild those smaller pieces to resume testing and gathering data.  Plus all the other things you listed.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robert827 on 07/27/2017 06:36 pm
The only reason I can see, is the Air Force money to develop an upper stage metholox engine.  This sub-scale engine can be made for a vacuum engine.  Capability is in the range of J2X or two BE-3's.  Someone said the way it is designed, that it can be scaled up easily.  Hopefully quickly to get the BFR/ITS on the road.

That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.

John

Exactly, No one has done a FFSC Methane engine before.  Start small, learn how to run one, then scale up.

Smaller is smaller, faster to build, smaller test stand, less consumables, less damage if it goes boom.

The fact that they may end up with an US engine and that the USAF helped pay for it is a bonus.

I think I remember reading somewhere that sub-scale FFSC Raptor should scale up relatively easily.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 07/28/2017 10:28 am
If a full scale Raptor has a 100cm diameter.
And a scaled Raptor has a 75cm diameter.
How would the thrust scale, if they have the same nozzle pressure?

If it scales cubed the thrust would go from 685k Lbs to 289k Lbs.
If it scales squared it would go from 685k lbs to 385k lbs.

I assumed a 75% Nozzle diameter reduction to allow for the same BFR engine layout... yes I know it's very unlikely that the smaller BFR has 42 engines...  8)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/28/2017 11:14 am
Nozzle area is proportional to thrust, but thrust required is proportional to lift off mass.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/28/2017 01:06 pm
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/28/2017 01:51 pm
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

The subscale Raptor is a little smaller than Merlin but more powerful; the big one will be a little bigger than Merlin and much more powerful.

We don't have much to go on besides the info and pictures in this article:
https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/

Measuring the test stand photo gives a subscale Raptor nozzle exit diameter of ~0.9 m and a throat diameter of ~0.2 m:

https://www.nasaspaceflight.com/wp-content/uploads/2016/10/2016-10-03-000759-350x229.jpg

This is an expansion ratio of ~20:1, and the engine is about 80% Merlin's size but with a slightly bigger nozzle exit. The throat is actually substantially oversized for a 1,000 kN engine operating at 30 MPa, and I suspect that the 1,000 kN figure is for operation at somewhere around 20 MPa. Operating this engine at 30 MPa (if the turbopumps can achieve that) should produce closer to 1,500 kN.

Scaling the engine linearly by ~1.4x (~2x area ratio) and pushing the nozzle exit out to ~40:1 would yield a 3 MN engine at 30 MPa operating pressure. This engine would be about 120% Merlin size, but the nozzle diameter would be about 2x Merlin.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 07/28/2017 01:55 pm
Yes, I read the full scale sea level will be about 2m in diameter, while the vacuum version will be around 4m.  Just wondering the size of the sub-scale and if it could actually be put into production.  Maybe not optimal.  But a vacuum version could be used as a second stage for F9/FH to improve performance and begin reuse testing. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/28/2017 02:54 pm
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: GORDAP on 07/28/2017 02:59 pm
Yes, I read the full scale sea level will be about 2m in diameter, while the vacuum version will be around 4m.  Just wondering the size of the sub-scale and if it could actually be put into production.  Maybe not optimal.  But a vacuum version could be used as a second stage for F9/FH to improve performance and begin reuse testing. 

[bold above mine]  I don't think this can be right, at least for the SL version.  With the 42 engine arrangement shown in last year's reveal, you have a minimum of 7 engine bells across, plus two sizable gaps.  That means the full scale version was planned to be no more than 1.6 meters in diameter.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/28/2017 03:09 pm
Yes, I read the full scale sea level will be about 2m in diameter, while the vacuum version will be around 4m.  Just wondering the size of the sub-scale and if it could actually be put into production.  Maybe not optimal.  But a vacuum version could be used as a second stage for F9/FH to improve performance and begin reuse testing. 

[bold above mine]  I don't think this can be right, at least for the SL version.  With the 42 engine arrangement shown in last year's reveal, you have a minimum of 7 engine bells across, plus two sizable gaps.  That means the full scale version was planned to be no more than 1.6 meters in diameter.

I don't think we engine modelers ever solved that discrepancy. 1.7 vs 1.6 from the drawings. I believe the they took liberties to get the 42 engines to fit.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/28/2017 03:18 pm
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.

The observed size of the nozzle exit on the sub-scale engine is just over 0.9 meters, so that matches closely. But the nozzle throat appears too large for a 40:1 ER (or for that matter 1,000 kN thrust at 30 MPa).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 07/28/2017 04:42 pm
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.

According to a recent Musk tweet, the Vac Raptor nozzle size is now 3m.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 07/28/2017 07:23 pm
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.

According to a recent Musk tweet, the Vac Raptor nozzle size is now 3m.



I would expect it to become that size if the 75% scalled spaceship retained it's original engine layout.
6 x 3 meter vacuum raptors fit nicely in the 9 meter circle with a gimballing cluster of 3 x aprox. 1 meter sea level raptors.


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/29/2017 12:41 am
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.

The observed size of the nozzle exit on the sub-scale engine is just over 0.9 meters, so that matches closely. But the nozzle throat appears too large for a 40:1 ER (or for that matter 1,000 kN thrust at 30 MPa).

Yes, that fooled me. I think we must be looking at a plug to keep the critters out.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robert Willis on 08/02/2017 07:24 pm
Engines designed to burn liquid hydrogen, such as RD-0120 & RD-0146 have been extensively test fired running on liquid methane with little modification. Seeing as Raptor was originally planned to burn LH2, how difficult would it be to produce such an engine with a high degree of component commonality with the CH4 burning model currently under development? Please correct me if I'm wrong, but I would guess that an LH2 fueled Raptor would have lower thrust, but higher ISP than the CH4 powered baseline. Can anyone out there do some rough calculations/estimates?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Axiom on 08/07/2017 11:37 am
So, does anyone know what size the sub scale is?  Is it Merlin sized since they have the tooling?  Also, does anyone know what size the full scale Raptor will be.  I've seen some pictures but, they are not scaled, just guessing?

Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.

The observed size of the nozzle exit on the sub-scale engine is just over 0.9 meters, so that matches closely. But the nozzle throat appears too large for a 40:1 ER (or for that matter 1,000 kN thrust at 30 MPa).

Yes, that fooled me. I think we must be looking at a plug to keep the critters out.

John


Is it possible we are looking at the 'plug' of an ED-nozzle? Although I do not believe there has been any information to suggest SpaceX is pursuing an ED-nozzle, they are trying to create a state-of-the-art engine with Raptor, so I wouldn't rule it out.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 08/07/2017 01:59 pm
No.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 08/30/2017 10:17 pm
Anyone heard anything on the full scale Raptor development?  Got to have the engine before we get the 9m ITS going. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 08/30/2017 10:45 pm
Anyone heard anything on the full scale Raptor development?  Got to have the engine before we get the 9m ITS going.
Not necessarily. It's possible they'd start with some subscale Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: robert_d on 09/03/2017 12:59 pm
What is required to make the Raptor air-startable/restartable? Understand from the history of Constellation that it did not seem possible to modify the SSME to use as a second-stage. Isn't the Raptor more similar in concept to this engine than a normal gas generator engine? Might it be possible that SpaceX would need to develop a gas generator methane engine to generate the power to start the other (full/flow) engines. Somehow?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 09/03/2017 01:01 pm
No, because it will be designed to be air start able from the very beginning. Needs to air start for first and second stage both.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: robert_d on 09/03/2017 01:46 pm
No, because it will be designed to be air start able from the very beginning. Needs to air start for first and second stage both.
Thanks for that. Has it been discussed before? I was wondering especially whether a "tap off" might be a part of the solution, and if so, does that encroach on Blue Origin IP?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 09/03/2017 03:35 pm
Part of the initial reveal was that the raptor uses electrical spark ignition rather than TEA/TEB, which makes it infinitely restartable.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/05/2017 03:09 pm
Ok so I read entire 15pages but didn't find any discussion on this:

If I read it right, the methane/oxygen into main chamber are gasified already, where the gasification happened? preburner? As I find it hard to believe gasification happened during turbine stage.

Or I miss something?

Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Electric Paint on 09/05/2017 03:58 pm
Ok so I read entire 15pages but didn't find any discussion on this:

If I read it right, the methane/oxygen into main chamber are gasified already, where the gasification happened? preburner? As I find it hard to believe gasification happened during turbine stage.

Or I miss something?

Titus

If I understand staged combustion correctly, then yes there are preburners in the system.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rebel44 on 09/05/2017 05:17 pm
Ok so I read entire 15pages but didn't find any discussion on this:

If I read it right, the methane/oxygen into main chamber are gasified already, where the gasification happened? preburner? As I find it hard to believe gasification happened during turbine stage.

Or I miss something?

Titus

Here is pretty good video that explains (among other things) Full Flow Staged Combution cycle:
https://www.youtube.com/watch?v=4QXZ2RzN_Oo
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 09/05/2017 06:18 pm



Here is pretty good video that explains (among other things) Full Flow Staged Combution cycle:
https://www.youtube.com/watch?v=4QXZ2RzN_Oo

Another one I found very enlightening:
https://youtu.be/jheMusS0JwA
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/06/2017 11:10 am
Ok maybe I over-complex the thinking process  ;) ;) ;)

Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 09/06/2017 01:07 pm
In the pre burner. Liquid because supercritical gas.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/07/2017 03:01 pm
I'm trying to find pintle injector for gasified propellant, not much result...

An interesting find is patent by JAXA, I guess it's something to do with JAXA's LE-X engine development.
https://www.google.com/patents/US7703274


Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dave G on 09/10/2017 06:53 pm
Where does SpaceX test this sub-scale engine?  Will they attempt to see what it's maximum thrust can be?

Also, can they test a full scale engine at McGregor?  Or do they need Stennis? 
SpaceX tests their subscale Raptor at their McGregor test facility in Texas. I think that they will also test the full size Raptor at McGregor as well.

RAPTOR TEST EQUIPMENT ENGINEER: McGregor, TX
http://www.spacex.com/careers/position/211014

Quote
RESPONSIBILITIES:
• Engineering, design, analysis, material/component selection and procurement, construction, activation, and maintenance of test stands, tooling, and supporting infrastructure
• Provide support and direction to technicians during construction, activation, and maintenance of stands and equipment
• Support testing campaigns by operating ground propellant systems, reviewing data for system health, and modifying equipment or procedures as necessary
• Develop novel ways to streamline site-wide processes and increase the reliability and efficiency of testing operations
• Perform any additional tasks that ensure efficient and effective testing, as required
• It is sometimes necessary to perform hands-on work in all environments (heat, cold, rain), occasionally in tight quarters or at heights

BASIC QUALIFICATIONS:
• Bachelor’s degree in mechanical engineering, aerospace engineering or other engineering discipline

PREFERRED SKILLS AND EXPERIENCE:
• Master’s degree in mechanical or aerospace engineering
• 3+ years of relevant experience in an industrial setting
• Fundamental understanding, intuition, and aptitude of fluid and/or structural design and analysis
• Creative ability to imagine and design from scratch, while retaining low cost, reliability, efficiency, and maintainability
• Experience where quick-thinking and problem solving plays a critical role
• Good response to challenges posed by short deadlines
• Ability to work in a high-concentration, high-stress environment, under possible extended work hours
• Acute attention to detail, ability to see interactions with other systems to avoid problems
• Intermediate skill level using Windows Operating Systems
• Intermediate skill level using Microsoft Office
• Intermediate skill level using CAD (NX a plus)
• Experience with high pressure and cryogenic fluid systems and components
• Experience producing drawings for welders and machine shop fabrication
• Machining, welding, other fabrication techniques, and general hands-on experience
• Experience in FEA or CFD modeling and analysis, with ability to verify by simplified hand calculations
• Piping and pressure vessel design experience per ASME code, work with flanges, gaskets, fasteners
• Instrumentation, testing, data review and analysis, verification against a model

ADDITIONAL REQUIREMENTS:
• General physical fitness is required for some work areas, flight hardware typically is built in tight quarters and physical dexterity is required
• Physical effort including standing, lifting and carrying light weight such as materials or equipment. Must lift up to 30 pounds unassisted
• Occasionally exposed to work in extreme outdoor environments- heat, cold, rain
• Work performed in an environment requiring exposure to fumes, odors, and noise
• Must be available to work extended hours and weekends, which varies depending on site operational needs, flexibility required
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/29/2017 08:46 am
Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.

31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar

IAC 2016, it was 300bar, 3050kN
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Welsh Dragon on 09/29/2017 08:55 am
Still thinking they'll go up to 300 eventually.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 09/29/2017 09:07 am
Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.

31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar

IAC 2016, it was 300bar, 3050kN
So looks like what was termed subscale Raptor that is undergoing testing is now the size of Raptor they will use. Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/29/2017 09:11 am
Still thinking they'll go up to 300 eventually.
Surely they will, but 250bar->300bar only give thrust change as 1700kN to 2000kN
Thrust is mainly determined by prop flow rate.

Maybe SpaceX decide to go with test engine's "sub scale" size, instead of make it "full sized" ?

Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 09/29/2017 09:11 am
Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.

Still early days for Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/29/2017 09:17 am
Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.

31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar

IAC 2016, it was 300bar, 3050kN
So looks like what was termed subscale Raptor that is undergoing testing is now the size of Raptor they will use. Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.
That maybe true of the goal is reaching the original ITS(IAC2016) capability.

But with BFR(IAC2017) capability target, the original "IAC2016 full sized" Raptor will make the 2nd stage... either throttle-able down to 10% (not likly I think)... or cutdown engine number from 4+2 to 3+1... which will drop the redud-backup capability.

Or you use 2 different sized Raptor, not an idea solution I guess.

Interesting changes anyway.
Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 09/29/2017 09:30 am
Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.

31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar

IAC 2016, it was 300bar, 3050kN
So looks like what was termed subscale Raptor that is undergoing testing is now the size of Raptor they will use. Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.

Engines are now reliable (or can be made so), so no need to worry about the number of them being a source of failure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Dante80 on 09/29/2017 10:57 am
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: schaban on 09/29/2017 01:51 pm
I'm pretty sure this version of the Raptor will eventually produce 3 MN. Initially Merlin 1D had ~70% of block 5 levels; here's the difference is roughly the same.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 09/29/2017 01:59 pm
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!

Appears they need more or larger tanks if they can only get a 40 sec run.  They’ll need a lot more time running the Raptor.

The blue exhaust and shock diamonds were fantastic. 

Still have a hard time imagining 31 of them on a vehicle. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 09/29/2017 01:59 pm
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!

Appears they need more or larger tanks if they can only get a 40 sec run.  They’ll need a lot more time running the Raptor.

The blue exhaust and shock diamonds were fantastic. 

Still have a hard time imagining 31 of them on a vehicle.

Elon said that the tanks limited them to 100 seconds of run time.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 09/29/2017 02:04 pm
Looking at the size of the nozzle diameters, it looks like RaptorVac had the expansion ratio dropped from ~200 around 120-140. would explain the slightly larger drop in ISP than the SL version (375 from 381, vs 330 from 334)

SL is probably similar to old Raptor, maybe slightly less (37 vs 40)

(2.4^2 / 1.3^2 = 3.41)
(3.41 x 37 = 126)

Mixture ratio was also changed from 3.8:1 to 3.6:1. (860/240) Probably result of lower (250bar) chamber pressure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 09/29/2017 02:12 pm
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?

Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?

@ edit: Sorry, my mistake. Raptor actually operates fuel rich. My parsing of the ratio numbers were wrong.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 09/29/2017 02:17 pm
I'm pretty sure this version of the Raptor will eventually produce 3 MN. Initially Merlin 1D had ~70% of block 5 levels; here's the difference is roughly the same.

Probably not.

This engine, if it's maxed out to IAC2016 (chamber pressure increase from 250bar to 300bar, slight increase in ER, increase in O:F ratio) specs could probably hit a little over 2MN.

The original Raptor needed a ~1.7m bell diameter to hit 3MN.

The original Merlin engine only hit ~67bar. The current M1DFT is probably around 110bar, a ~65% increase.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/29/2017 02:28 pm
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?

Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
Graph is easier to explain :)
(using Rocket Propulsion Analysis :))

Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 09/29/2017 02:39 pm
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?

Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
Graph is easier to explain :)
(using Rocket Propulsion Analysis :))

Titus

Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..

@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/29/2017 02:49 pm
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?

Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
Graph is easier to explain :)
(using Rocket Propulsion Analysis :))

Titus

Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..

@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
Depending on what you define as "oxidizer-fuel" mid-point  ;D ;D ;D ;D

"Rich" is always "relatively compared to referencing point"  ;D ;D ;D ;D ;D ;D
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 09/29/2017 02:54 pm
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?

Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
Graph is easier to explain :)
(using Rocket Propulsion Analysis :))

Titus

Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..

@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
Depending on what you define as "oxidizer-fuel" mid-point  ;D ;D ;D ;D

"Rich" is always "relatively compared to referencing point"  ;D ;D ;D ;D ;D ;D

I think the reference point is the theoretical stoichiometric mixture ratio for complete combustion. Am I wrong?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 09/29/2017 03:03 pm
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!

Not only progressed, but progressed on the flight engine
This puts them squarely in the lead on next generation engines and vehicles.

Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: skel on 09/29/2017 03:52 pm
Not only progressed, but progressed on the flight engine
This puts them squarely in the lead on next generation engines and vehicles.

Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.

It is promising.

Do we have any public information about progress on the spark ignition. I couldn't see any initial green tint of TEA/TEB at startup in the most recent test video released with today's (IAC 2017) presentation.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 09/29/2017 03:57 pm
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?

Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
Graph is easier to explain :)
(using Rocket Propulsion Analysis :))

Titus

Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..

@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
Depending on what you define as "oxidizer-fuel" mid-point  ;D ;D ;D ;D

"Rich" is always "relatively compared to referencing point"  ;D ;D ;D ;D ;D ;D

No... It's relative to a balanced stochiometric burn.  Fuel rich means you have unburned fuel in the exhaust, Oxygen rich means you have unburned Oxygen.  Theoretically.  In practice you may have partial byproducts (think CO), but the principle stands.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 09/29/2017 04:05 pm
Not only progressed, but progressed on the flight engine
This puts them squarely in the lead on next generation engines and vehicles.

Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.

It is promising.

Do we have any public information about progress on the spark ignition. I couldn't see any initial green tint of TEA/TEB at startup in the most recent test video released with today's (IAC 2017) presentation.

It does have a beautiful blue/violet flame though.

Night launches for this rocket are going to be a treat.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 09/29/2017 04:07 pm
Not only progressed, but progressed on the flight engine
This puts them squarely in the lead on next generation engines and vehicles.

Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.

It is promising.

Do we have any public information about progress on the spark ignition. I couldn't see any initial green tint of TEA/TEB at startup in the most recent test video released with today's (IAC 2017) presentation.

It does have a beautiful blue/violet flame though.

Night launches for this rocket are going to be a treat.

As are day launches.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: GalacticIntruder on 09/29/2017 04:19 pm
Musk did say they ran mini-Raptor for 100 seconds, which is the max their test tanks can hold.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 09/29/2017 04:20 pm
So during the burn video there are 2 green flame episodes, one near the beginning, one at shutdown. Since these are not actually at startup, and supposedly Raptor uses spark ignition anyway, the only explanation I can think of is a bit of copper chamber or bell vaporizing.

The first frame of the early incident at 5:44 you just see a little streak by the bell, the next frame it's partway down the jet, then it's at the end of the jet. For some reason this frame is doubled. Then the next frame, no more green. However the vapor patterns along the ground don't suggest that any video is missing, they change from frame to frame in a consistent way.

Also, is it possible to tell the exhaust velocity from the spacing of the mach diamonds? They are extremely consistent right until shutdown starts around 6:20. You can see from one frame to the next that they are sliding to the right at that point.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 09/29/2017 04:22 pm
Endless Raptor gif, came out a little small:
(https://i.makeagif.com/media/9-29-2017/-fUtUN.gif)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 09/29/2017 04:34 pm
So during the burn video there are 2 green flame episodes, one near the beginning, one at shutdown. Since these are not actually at startup, and supposedly Raptor uses spark ignition anyway, the only explanation I can think of is a bit of copper chamber or bell vaporizing.

The first frame of the early incident at 5:44 you just see a little streak by the bell, the next frame it's partway down the jet, then it's at the end of the jet. For some reason this frame is doubled. Then the next frame, no more green. However the vapor patterns along the ground don't suggest that any video is missing, they change from frame to frame in a consistent way.

Also, is it possible to tell the exhaust velocity from the spacing of the mach diamonds? They are extremely consistent right until shutdown starts around 6:20. You can see from one frame to the next that they are sliding to the right at that point.

If I had to guess, if the test article is running around 200bar, then the ISP is probably around the low 320s.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 09/29/2017 04:48 pm
Endless Raptor gif, came out a little small:
(https://i.makeagif.com/media/9-29-2017/-fUtUN.gif)

Similarly, here's that green flame episode:

(https://i.makeagif.com/media/9-29-2017/s5Lpag.gif)

And shutdown:

(https://i.makeagif.com/media/9-29-2017/Z1rGm-.gif)

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 09/29/2017 05:07 pm
Endless Raptor gif, came out a little small:
(https://i.makeagif.com/media/9-29-2017/-fUtUN.gif)

Similarly, here's that green flame episode:

(https://i.makeagif.com/media/9-29-2017/s5Lpag.gif)

And shutdown:

(https://i.makeagif.com/media/9-29-2017/Z1rGm-.gif)
A quick google suggests that "incomplete combustion" can cool a methane flame from blue to yellow (http://www.elgas.com.au/blog/1585-why-does-a-gas-flame-burn-blue-lpg-gas-natural-propane-methane)
If they run fuel rich, the methane may take a few miliseconds longer to turn on and off, resulting in a slightly cooler (green flame) burst as the o2 turns off completely. (or before the O2 turns on)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 09/29/2017 05:47 pm
A quick google suggests that "incomplete combustion" can cool a methane flame from blue to yellow (http://www.elgas.com.au/blog/1585-why-does-a-gas-flame-burn-blue-lpg-gas-natural-propane-methane)
If they run fuel rich, the methane may take a few miliseconds longer to turn on and off, resulting in a slightly cooler (green flame) burst as the o2 turns off completely. (or before the O2 turns on)

I don't really think that's it, the yellow flame of incomplete methane combustion isn't going to be overlaid with enough blue to appear green.

This seems like either copper or boron ions in the flame. I suppose it is possible they are still using TEA/TEB ignition and it just spit out a couple more blobs of it during the burn as part of some kind of purge.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DOCinCT on 09/29/2017 06:15 pm
Still thinking they'll go up to 300 eventually.
At around 23 min into the presentation Elon comments on improvements in both ISP (add 10 units or so) and chamber pressure (to 300 bar).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 09/29/2017 06:19 pm




This seems like either copper or boron ions in the flame.

Are you implying engine rich combustion?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 09/29/2017 06:22 pm




This seems like either copper or boron ions in the flame.

Are you implying engine rich combustion?

If it's copper, yes. Merlin uses a plated copper combustion chamber and bell and I would assume Raptor is similar. I don't know that we could tell the difference between copper and boron with an uncalibrated camera.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 09/29/2017 06:27 pm
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 09/29/2017 06:34 pm


The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.

Blue might not have a choice. I read somewhere that they had to increase the size for Vulcan.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 09/29/2017 06:36 pm


The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.

Blue might not have a choice. I read somewhere that they had to increase the size for Vulcan.

ULA needs 550klbf for Vulcan. Blue won't build anything smaller than that.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mme on 09/29/2017 07:15 pm
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
My guess is that BO will stick with fewer, bigger engines.  At least I hope they do mostly because I like seeing a wide variety of approaches. :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 09/29/2017 09:21 pm
One aspect that hasn't been talked about much is that Raptor is gong to be able to throttle down to 20%... (Compared to Merlin 1D with ~40%)

This means that the lowest thrust on Raptor will be roughly the same as the lowest thrust on Merlin 1D, since Raptor is going to have ~2x the thrust capability.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 09/29/2017 10:08 pm
One aspect that hasn't been talked about much is that Raptor is gong to be able to throttle down to 20%... (Compared to Merlin 1D with ~40%)

This means that the lowest thrust on Raptor will be roughly the same as the lowest thrust on Merlin 1D, since Raptor is going to have ~2x the thrust capability.

Yes that’s quite important, if Merlin 1D could throttle down to 20%, Falcon 9 could probably hover, and thus also make a softer landing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Basto on 09/29/2017 10:35 pm
One aspect that hasn't been talked about much is that Raptor is gong to be able to throttle down to 20%... (Compared to Merlin 1D with ~40%)

This means that the lowest thrust on Raptor will be roughly the same as the lowest thrust on Merlin 1D, since Raptor is going to have ~2x the thrust capability.

Yes that’s quite important, if Merlin 1D could throttle down to 20%, Falcon 9 could probably hover, and thus also make a softer landing.

Hovering won’t necessarily make your landing softer. It just wastes fuel and diminishes payload capacity. If you do it right you use just enough fuel to zero velocity at the same moment you touch down.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ChaoticFlounder on 09/30/2017 01:01 am
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 09/30/2017 11:01 am
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.

BO will avoid N-1 type architectures like the plague unlike SpaceX.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 09/30/2017 11:08 am
Does anybody have a engine layout for 31 engines?
I have seen 1+8+16=25 but does the 31 engine layout just have 6 more engines around the outside?
So raptor SL is 2.8m...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 09/30/2017 11:27 am
Does anybody have a engine layout for 31 engines?
I have seen 1+8+16=25 but does the 31 engine layout just have 6 more engines around the outside?
So raptor SL is 2.8m...
Raptor SL nozzle is 1.3m dia. Most likely engine configuration for 31 engines is 1+6+12+12.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 09/30/2017 11:47 am
Similarly, here's that green flame episode:

And shutdown:

Nice animations. But I would be careful to draw any conclusions from the colour of the flame. It could be a white balance issue of the camera. I would say its even pretty likely given the dark background and the blue-red coloured flame which is imaged. The camera would adjust automatically by scaling the green up a bit due to the lack of green in the flame. When you then suddenly have a whitish component, it would look green.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 09/30/2017 01:59 pm
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?
Turbomachinery configuration info. is likely covered by ITAR so please don't expect SpX to release any info. on it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 09/30/2017 04:25 pm
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.

BO will avoid N-1 type architectures like the plague unlike SpaceX.

You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.

Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Norm38 on 09/30/2017 05:55 pm
Does anybody have a engine layout for 31 engines?
I have seen 1+8+16=25 but does the 31 engine layout just have 6 more engines around the outside?
So raptor SL is 2.8m...
Raptor SL nozzle is 1.3m dia. Most likely engine configuration for 31 engines is 1+6+12+12.

A drawing I saw on the discussion thread was 1+6+24. It wasn't two rings of 12, it was 24 staggered in a zig zag
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: butters on 09/30/2017 06:23 pm
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?

No, but one of the reasons why Raptor is the first production-intent FFSC (full flow staged combustion) engine is that it's difficult (perhaps prohibitively so) to test the fuel pump, oxidizer pump, and main combustion chamber in isolation from each other. Without resorting to extremely elaborate test stand hardware, the engine really has to be tested as a complete unit. This formidable upfront challenge has deterred engine manufacturers for decades.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mnelson on 09/30/2017 07:04 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nibb31 on 09/30/2017 07:11 pm
DJPledger has a weird fixation on the high number of engines on the BFR, comparing it to the ill-fated Soviet N-1 rocket. He's been reminded several times that modern technology has nothing to do with an overambitious and underfunded 1960's soviet design, assembled by underqualified and overworked personel with inexistant quality control, but he keeps on repeating his argument over and over again.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 09/30/2017 08:18 pm
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 09/30/2017 09:13 pm
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.

Every kg of mass removed from the engine is about 10 kg more to orbit.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 09/30/2017 10:07 pm
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.

Every kg of mass removed from the engine is about 10 kg more to orbit.

How does that work? Surely the most it could ever be is a 1:1 trade?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 09/30/2017 10:15 pm
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?

No, but one of the reasons why Raptor is the first production-intent FFSC (full flow staged combustion) engine is that it's difficult (perhaps prohibitively so) to test the fuel pump, oxidizer pump, and main combustion chamber in isolation from each other. Without resorting to extremely elaborate test stand hardware, the engine really has to be tested as a complete unit. This formidable upfront challenge has deterred engine manufacturers for decades.

If we assume that the layout has not changed since last year:
- 2 stage CH4 pump driven by a single stage turbine driven by a fuel rich pre-burner.
- 1 stage LOX pump driven by a single stage turbine driven by a oxygen rich pre-burner.
- LOX pump, pre-burner and turbine inline on top of main combustion chamber.
- CH4 pump, pre-burner and turbine off to the side.
- Apparently no boost pumps. Small internal tanks might be at slightly higher pressure to aid in starting.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Mongo62 on 09/30/2017 10:20 pm
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.

Every kg of mass removed from the engine is about 10 kg more to orbit.

Wouldn't it be the other way around for the first stage? 1 kg of first-stage mass reduction enables roughly 0.1 kg of extra payload to LEO.

For the second stage, it would be a 1:1 ratio.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 09/30/2017 10:32 pm
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.

Every kg of mass removed from the engine is about 10 kg more to orbit.

How does that work? Surely the most it could ever be is a 1:1 trade?

There are 6 engines on the upper stage, so one 1 kg off each of those is 6 more kg to orbit. There are 31 on the booster, but that trades at about 10:1 with payload mass. 31*1/10+6=9.1 kg of payload for each 1 kg saved on every engine. So about 10:1.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 09/30/2017 10:57 pm
What can we learn about the status of the Raptor development program from the details provided in the IAC talk (1200 seconds total, longest firing 100 seconds, 42 firings)?

A November 2007 SpaceX press release (http://www.businesswire.com/news/home/20071112005019/en/REPLACING-VIDEO-SpaceX-Completes-Development-Merlin-Regeneratively) indicates that the Merlin 1C development program included 3000 seconds total, longest firing 170 seconds, 125 firings.  Would the Raptor program be comparable, or would it be at a different scale?

Edit:  I note that a February 2008 SpaceX press release (http://www.spacex.com/press/2012/12/19/spacex-completes-qualification-testing-merlin-regeneratively-cooled-engine) indicates that the Merlin 1C qualification program comprised 1620+ seconds.

I can't readily find the Merlin 1D development program details.

Merlin 1D's qualification program included 1970 seconds among 28 firings (http://www.spacex.com/press/2013/04/13/spacexs-merlin-1d-engine-achieves-flight-qualification).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/01/2017 03:05 am
An early Merlin 1D development engine was installed on Grasshopper, right? I do wonder if we might see something similar with Raptor for testing the launch cradle.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/01/2017 03:45 am
What can we learn about the status of the Raptor development program from the details provided in the IAC talk (1200 seconds total, longest firing 100 seconds, 42 firings)?

A November 2007 SpaceX press release (http://www.businesswire.com/news/home/20071112005019/en/REPLACING-VIDEO-SpaceX-Completes-Development-Merlin-Regeneratively) indicates that the Merlin 1C development program included 3000 seconds total, longest firing 170 seconds, 125 firings.  Would the Raptor program be comparable, or would it be at a different scale?


Musk's comment that "tests could be much longer than [100s]" was of interest to me, given how confidently he stated that. I suspect the next Raptor testing we see might be with a larger scale article with higher capacity propellant tanks to enable live, full-length firings at 250 bar.

As for the length and number of firings, it's hard to know. Really not worthwhile to compare and contrast with Merlin regimes because of how different closed-cycle engines are from open-cyclers. Still, I'd generally expect at least as much testing (1200+s) with a full-size article, whenever that begins.

I'd recommend L2 if you can afford it (half off for students!), some amazing thread-relevant content and info on there.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: darkenfast on 10/01/2017 03:58 am
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers.  Does anyone have any information to contradict that?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: su27k on 10/01/2017 06:24 am
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers.  Does anyone have any information to contradict that?

Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/01/2017 06:40 am
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers.  Does anyone have any information to contradict that?

Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.

I think the stated fact of 1200 seconds total run time over 42 ignitions...
And the Raptor engine test stand has never been photographed damaged or looking like it suffered a RUD event all year...
Speaks volumes for one of either two things...
1) They can fix that test stand real damn quick on short notice...  ;D
2) Their design is sound and they are in good shape to meet timelines stated...  8)

Still... I have no info BO or AJ can't meet their stated deadlines either... 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/01/2017 06:52 am
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers.  Does anyone have any information to contradict that?

Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.

Yep. Despite the extremely successful-sounding testing stats, it was almost certainly only done with subscale Raptors.

We'll see if SpaceX provides more frequent updates over the next 6-12 months. I certainly hope/expect that they will if they are indeed planning on beginning the first BFR construction before H2 2018. My intuition tells me that there's no way that would happen unless tankage and engine designs were finalized and flight certified.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeAtkinson on 10/01/2017 07:16 am
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers.  Does anyone have any information to contradict that?

Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.

It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m  BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.

Because BFR is cheaper than the F9, FH and Dragon systems it is replacing, each year earlier that SpaceX can introduce BFR is worth a lot, perhaps $500-700 M, add the $1-2 B for development of a different scale engine and changing engines now to a bigger Raptor would be at considerable cost.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/01/2017 08:23 am
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.

BO will avoid N-1 type architectures like the plague unlike SpaceX.

You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.

Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
BO has the money to dev. an F-1 class engine that SpaceX does not have so SpaceX have downsized the Raptor to fit within their R&D budget.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/01/2017 08:28 am
SpaceX probably could build a bigger engine like BO, but they wouldn't be able to make it as awesome. Also, making it smaller means they think they can mass produce them like they've been doing Merlin. They won't need to make as many BFRs for a while, so having a bunch of engines still allows them to get that economy of scale going for them. It also helps give them the granularity to use the same engines on the spaceship part and have them small enough to be redundant.

Besides, they're going to try launching a rocket with 27 engines in a month or three. So it's not like it'll be unprecedented by the time they try it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: kevinof on 10/01/2017 08:34 am
Where's your evidence that they downsized because of budget?

The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.

BO will avoid N-1 type architectures like the plague unlike SpaceX.

You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.

Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
BO has the money to dev. an F-1 class engine that SpaceX does not have so SpaceX have downsized the Raptor to fit within their R&D budget.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/01/2017 09:21 am
Interesting quote from Elon Musk about why the Raptor engine is relatively small:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"

Apparently it has optimized to an even smaller thrust of 170 metric tons.



Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 10/01/2017 09:30 am


BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?


https://youtu.be/U9fkYIrRwbo
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: biosehnsucht on 10/01/2017 09:48 am
There were many problems with the N-1, but just having lots of engines wasn't one of them (other than controlling so many engines with the technology of the day was in no way trivial - but not an inherent problem with many engines).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/01/2017 09:58 am
N1 was a low funded hurry up job, to go to the Moon, no wonder it failed.

- Plumbing to feed the engines was the main problem, for the first failed attempts.
- Plumbing had to be reassembled on site.
- The engines where not all test fired beforehand.
- Inferior quality fuel was used.
- Bad commands came from the not so capable flight computer.

The fourth and final launch failed seconds before stage separation, extra depressing..


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 10/01/2017 10:06 am



- The engines where not all test fired beforehand.

As I understand it, none of the flight engines where test fired. They made them in batches and selected a random engine from the batch for testing. If it was good, the batch was declared good.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/01/2017 10:10 am



- The engines where not all test fired beforehand.

As I understand it, none of the flight engines where test fired. They made them in batches and selected a random engine from the batch for testing. If it was good, the batch was declared good.


Yes you’re right, once tested it’s probably unusable because of the bad fuel among other things.
Luckily Methalox burns very clean.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: OneSpeed on 10/01/2017 10:31 am
It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m  BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.

It would be great news if you are right about this. The subscale test Raptor (roughly 0.87mØ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mØ BFR SL engine is projected to be 1.7MN at 250 bar.  The test Raptor expansion appears ambient in the recent test footage. The ratio of nozzle areas is 1.3^2 / 0.87^2 = 1:2.23 (a 110% increase). Would the extra 50 bar (a 25% increase) prevent overexpansion with such a large nozzle? Would the increased chamber pressure and larger nozzle be sufficient to boost the test Raptor to 1.7MN, a 70% increase?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nibb31 on 10/01/2017 12:56 pm



- The engines where not all test fired beforehand.

As I understand it, none of the flight engines where test fired. They made them in batches and selected a random engine from the batch for testing. If it was good, the batch was declared good.


Yes you’re right, once tested it’s probably unusable because of the bad fuel among other things.
Luckily Methalox burns very clean.

They also never built a test stand for testing entire stages without committing to a full up launch. The vibration and control problems with the N-1 first stage would have been detected and fixed without blowing up the entire stack (and launch pad).

There were also quality control issues related with assembling the rocket in harsh conditions at Baikonur with underqualified and underequipped personnel. They didn't have the equipment and tooling that were available at the major aerospace facilities.

The whole project was a mess, and the number of engines was seriously the least of the issues.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mattstep on 10/01/2017 02:30 pm
In the 2017 IAC presentation, Musk places emphasis on the engine out capability for ITSy during landing. Do people think SpaceX considered that a required design criteria for the ship? If so, did the size of the engine drive the size of the ship, or visa versa?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/01/2017 03:02 pm
It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m  BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.

It would be great news if you are right about this. The subscale test Raptor (roughly 0.87mØ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mØ BFR SL engine is projected to be 1.7MN at 250 bar.  The test Raptor expansion appears ambient in the recent test footage. The ratio of nozzle areas is 1.3^2 / 0.87^2 = 1:2.23 (a 110% increase). Would the extra 50 bar (a 25% increase) prevent overexpansion with such a large nozzle? Would the increased chamber pressure and larger nozzle be sufficient to boost the test Raptor to 1.7MN, a 70% increase?

- I will be cutting numbers Monday for the sub-scale test Raptor and BFR Raptor, but it appears that relatively small (~1.15 linear) scaling up will be needed.

- I believe that engine out during BFS landing probably sized the engine and as a consequence resulted in 31 engines on the BFR. Also, the fact that they already had test fired an engine of about the right size.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 10/01/2017 03:41 pm
In the 2017 IAC presentation, Musk places emphasis on the engine out capability for ITSy during landing. Do people think SpaceX considered that a required design criteria for the ship? If so, did the size of the engine drive the size of the ship, or visa versa?
They wanted to be able to man-rate propulsive landing, which requires incredibly high reliability. Having a spare squares your reliability. (technically, squares your failure fraction, making it smaller)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 10/01/2017 04:11 pm
It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m  BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.

It would be great news if you are right about this. The subscale test Raptor (roughly 0.87mØ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mØ BFR SL engine is projected to be 1.7MN at 250 bar.  The test Raptor expansion appears ambient in the recent test footage. The ratio of nozzle areas is 1.3^2 / 0.87^2 = 1:2.23 (a 110% increase). Would the extra 50 bar (a 25% increase) prevent overexpansion with such a large nozzle? Would the increased chamber pressure and larger nozzle be sufficient to boost the test Raptor to 1.7MN, a 70% increase?

- I will be cutting numbers Monday for the sub-scale test Raptor and BFR Raptor, but it appears that relatively small (~1.15 linear) scaling up will be needed.

- I believe that engine out during BFS landing probably sized the engine and as a consequence resulted in 31 engines on the BFR. Also, the fact that they already had test fired an engine of about the right size.

John
As I play around RPA software, here is my "non mathematics" sense goes ...

Thrust is direct liner with mass flow rate. If you double the mass flow rate, you double the thrust.

Chamber pressure and chamber size goes reverse side based on the same thrust (mass flow rate).
So if you have the same chamber size, you move chamber pressure from 200bar to 300bar, you also move the mass flow rate by around 150%, which translate to around 153% thrust.

Nozzle doesn't affect your "design thrust" as when you do calculation, you already specific the design's exit pressure. Unless you want to change design's exit pressure while leave chamber size unchanged, otherwise you don't need to change the nozzle size.

So the real question about "upscale" is about "increase chamber size" or "increase chamber pressure".

---

Using, RPA to calculate, a 254mm dia throat sized chamber, running at 200bar chamber pressure, with flow exit at 1bar, give around 1698kN at SEA, and require 1133mm dia nozzle (flow exit at 1bar).

1133mm nozzle dia is within current discussion's 1-6-12-12 nozzle placement, which using 1.3m dia nozzle.
https://forum.nasaspaceflight.com/index.php?topic=43851.msg1729848#msg1729848


Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mgeagon on 10/01/2017 04:21 pm
Is the current test engine "sub-scale" or just under pressure fed? My take from Musk's talk was an increase from the currently tested 20 Mpa to 25 Mpa and eventually 30 Mpa. At no point does it appear any scaling up is called for or necessary. Why are we still calling the as tested Raptor "sub-scale"?

Edit: Corrected unit.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/01/2017 04:28 pm
In the 2017 IAC presentation, Musk places emphasis on the engine out capability for ITSy during landing. Do people think SpaceX considered that a required design criteria for the ship? If so, did the size of the engine drive the size of the ship, or visa versa?
Even if BFR or ship has a single benign engine failure and the mission is a total success then there will be a stand down period to find out and fix the issue. 31 engines on booster and 6 on ship will increase the risk of benign engine failures causing stand down periods which SpaceX may not be able to afford.

I think the size of Raptor was more driven by dev. costs and the desire to use one engine design throughout the ITS system which was driven by those costs and limited funding. If EM was given several tens of billions of dollars then he may have gone for the more traditional large engine for booster and small engine for ship/US.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/01/2017 04:30 pm
Is the current test engine "sub-scale" or just under pressure fed? My take from Musk's talk was an increase from the currently tested 20 bar to 25 bar and eventually 30 bar. At no point does it appear any scaling up is called for or necessary. Why are we still calling the as tested Raptor "sub-scale"?
Pressures you have quoted should be 200 bar, 250 bar, and 300 bar.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/01/2017 04:39 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.

If you bet one dollar that the 1st BFR mission is a complete success with no issues then you will win a fortune.

Lets all hope that Raptor works as advertised and that SpX don't lose any BFR's but I think that will be a long shot.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/01/2017 04:48 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.

Working principle and purpose is much more important than the number of something.

N-1 used differential thrusting for steering. And at least one of the failures were due lack of control.

BRF does not, it uses swiveling engines. It's immune to the "lose control authority due wrong engine failing".


And that "down time after engine failure" just means that after those downtimes, it will be much reliable than any competing rocket.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/01/2017 05:07 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.

Working principle and purpose is much more important than the number of something.

N-1 used differential thrusting for steering. And at least one of the failures were due lack of control.

BRF does not, it uses swiveling engines. It's immune to the "lose control authority due wrong engine failing".


And that "down time after engine failure" just means that after those downtimes, it will be much reliable than any competing rocket.

Left out test firing of engines.  Raptor will have it, N-1 didn't.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nibb31 on 10/01/2017 05:45 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture.

Says you. Why 30 and not 28 or 36 ?

Quote
Any booster with such high engine no. is likely to suffer engine failures.

An observation that is based on exactly one sample, where the actual number of engines was only a minor contributor to quality control nightmare that the N-1 program was.

By your reasoning, the Saturn V had 3 million parts, which was 3 million chances of a part failing and bringing down the Apollo program.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: butters on 10/01/2017 06:34 pm
Is the current test engine "sub-scale" or just under pressure fed? My take from Musk's talk was an increase from the currently tested 20 bar to 25 bar and eventually 30 bar. At no point does it appear any scaling up is called for or necessary. Why are we still calling the as tested Raptor "sub-scale"?

Raptor's main combustion chamber would be very difficult to test in isolation from its turbopumps, requiring elaborate test stand hardware to generate both fuel-rich hot gas and oxidizer-rich hot gas and force them into the injector at 200+ bar. The full-flow staged-combustion power cycle has a number of significant advantages, but Raptor is the first production-intent design, in large part because the upfront challenge of developing a startup sequence for the fully-integrated engine was a formidable deterrent.

Engineering is all about tradeoffs, and the downsides of FFSC are frontloaded in the development process. The designer has to get overcome early challenges in order to access the long-term benefits such as cooler/slower turbines and no bearing seals separating fuel and oxidizer. That Raptor already has 42 starts and stops under its belt should be seen as a very impressive achievement.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/01/2017 06:34 pm
Where's your evidence that they downsized because of budget?

The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.

BO will avoid N-1 type architectures like the plague unlike SpaceX.

You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.

Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
BO has the money to dev. an F-1 class engine that SpaceX does not have so SpaceX have downsized the Raptor to fit within their R&D budget.
Engine dev. costs increase faster than thrust so Raptor has been reduced in size to minimize dev. cost of BFR while keeping engine no. to what they think is acceptable which I don't agree with. BO has almost limitless funding so they can dev. a much bigger engine for their NA.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: butters on 10/01/2017 06:46 pm
Engine dev. costs increase faster than thrust so Raptor has been reduced in size to minimize dev. cost of BFR while keeping engine no. to what they think is acceptable which I don't agree with. BO has almost limitless funding so they can dev. a much bigger engine for their NA.

Raptor size was fixed by the desire to have two landing engines for reliability. A bigger engine would not allow for engine-out tolerance. Blue Origin doesn't have a reusable upper stage design at the moment, so they don't face this same consideration.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/01/2017 07:03 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?


https://www.youtube.com/watch?v=U9fkYIrRwbo

Lets all of us hope that SpaceX does not repeat this footage with their maiden BFR launch.

Please remember that Raptor has a much higher Pc than the NK-15 had so a Raptor failure has the potential to have an even bigger kaboom than an NK-15 failing.

BO has the funding to dev. an engine large enough for them to avoid the N-1 type architecture with their NA.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 10/01/2017 07:03 pm
How big of an engine Blue is going to make for New Armstrong has nothing to do with this thread.

How many engines you think is best for a launch vehicle has nothing to do with this thread.  That would be a discussion for the launch vehicle thread, not the engine thread.

We'll see if SpaceX provides more frequent updates over the next 6-12 months. I certainly hope/expect that they will if they are indeed planning on beginning the first BFR construction before H2 2018. My intuition tells me that there's no way that would happen unless tankage and engine designs were finalized and flight certified.

The haven't tested a flight model engine or tank yet, so I highly doubt they have been flight certified.

Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers.  Does anyone have any information to contradict that?

I don't think I'd call Blue's engine a "candidate" for Blue's launcher (which is really off-topic here anyway).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/01/2017 07:18 pm
BO will avoid N-1 type architectures like the plague unlike SpaceX.

Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.

If you bet one dollar that the 1st BFR mission is a complete success with no issues then you will win a fortune.

Lets all hope that Raptor works as advertised and that SpX don't lose any BFR's but I think that will be a long shot.

You're using the word "architecture" wrong.  The term you're looking for is "engine count".

To your method, F9 should also be losing engines and thereby crashing.

The key to a high engine-count booster is a) very high reliability engine, and b) a fault tolerant structure.

The second item is the interesting one.  How to make sure that engine failures are directed backwards and don't damage neighboring engines.  In this respect, smaller engines are easier, since the energy contained in the chamber is proportional to the pressure and the volume, so are linear in P, and cubic in l.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: drzerg on 10/01/2017 08:28 pm
so in aviation when some engine fails all fleet with this type of engine stop flying? no. they keep flying while investigation go on.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/01/2017 08:58 pm
Engine dev. costs increase faster than thrust so Raptor has been reduced in size to minimize dev. cost of BFR while keeping engine no. to what they think is acceptable which I don't agree with. BO has almost limitless funding so they can dev. a much bigger engine for their NA.

Raptor size was fixed by the desire to have two landing engines for reliability. A bigger engine would not allow for engine-out tolerance. Blue Origin doesn't have a reusable upper stage design at the moment, so they don't face this same consideration.

- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.

- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/02/2017 01:51 am
SpaceX is not afraid of "big".  They're building a 9 m spaceship, remember?

I'm sure it would have been cheaper to develop a 5 m ship and then make lots...  But the optimization said 9 m.  (Actually they wanted 12, but logistics and the desire to serve earth markets with the same ship said 9)

So unless there's some clear evidence otherwise, I'd go with what they said - that surprisingly, a smaller engine resulted in better overall T/W.

I personally also think that reliability goes up with smaller engines, since it's easier to prevent collateral damage.  We know 9 works under some conditions, but more is better.

With airplanes, we're very much at ease with the fact that engines fail sometimes, and usually (not always) the failure stays contained, and even when it doesn't, usually the plane survives.

We want to be at the same point - that engines failure is rare, and then rarely explosive, and ever rarer still is the case where the failure affects other engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ChaoticFlounder on 10/02/2017 02:18 am

- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.

- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.

John

Hi JW,

How did expected cost scale with:

1.  Number of parts?
2.  Chamber Pressure?
3.  Fuel / Oxidizer Choice?
4.  Engine Cycle?
5.  # of Turbopumps?
6.  Any other interesting tid bits I may have missed?

Also, how accurate do you believe these models were?  To use your example, how do you believe they verified cost scaled at ~(thrust)^(0.7)?

I've always been curious how well these models work and was just wondering what your opinion of them were?

Thanks,

C
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 10/02/2017 03:34 am
Also, SpaceX makes extensive use of 3D printers for their engines.  The size of Raptor may be limited to the largest 3D printer they could find.  As 3D printers get larger, the size of an engine could also increase.  3D printers lower production costs as it reduces labor costs.  It seems to also increase reliability. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/02/2017 04:02 am
Also, SpaceX makes extensive use of 3D printers for their engines.  The size of Raptor may be limited to the largest 3D printer they could find.  As 3D printers get larger, the size of an engine could also increase.  3D printers lower production costs as it reduces labor costs.  It seems to also increase reliability.
Youre overselling 3D printing here.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 10/02/2017 04:30 am
Also, SpaceX makes extensive use of 3D printers for their engines.  The size of Raptor may be limited to the largest 3D printer they could find.  As 3D printers get larger, the size of an engine could also increase.  3D printers lower production costs as it reduces labor costs.  It seems to also increase reliability.
Youre overselling 3D printing here.
My understanding is that additive manufacturing ("3d printing") is really good at the kind of complicated part that otherwise would have needed to be assembled out of multiple components. As I understand, turbopumps especially a re made of lots of that kind of component, and a FFSC engine has twice as many turbopumps as even a regular engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/02/2017 02:56 pm

- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.

- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.

John

Hi JW,

How did expected cost scale with:

1.  Number of parts?
2.  Chamber Pressure?
3.  Fuel / Oxidizer Choice?
4.  Engine Cycle?
5.  # of Turbopumps?
6.  Any other interesting tid bits I may have missed?

Also, how accurate do you believe these models were?  To use your example, how do you believe they verified cost scaled at ~(thrust)^(0.7)?

I've always been curious how well these models work and was just wondering what your opinion of them were?

Thanks,

C

I remember researching cost scaling with turbines and powerplant equipment

The number JW quoted (x^0.7) was one that came up very often. Most everything was between 0.6 and 0.8. Turbine equipment was definitely in that area, which would have the most relevance to rocket engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 10/02/2017 04:36 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: intrepidpursuit on 10/02/2017 05:59 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Tom Mueller said the M1C cost "some fraction of a million dollars" and that the M1D was the result of a huge amount of cost cutting R&D including the face shutoff architecture. This implies that the M1D is less than $500,000. Using the same math, that puts the booster at $42 million.

I know that type of number sounds preposterous, but I'm just quoting Mueller. I don't know how SpaceX is still in business, making money, and developing gigantic things unless they have been as successful at lowering costs as they imply.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/02/2017 06:47 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: tdperk on 10/02/2017 08:47 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.

"Now", only.  I await the progressive uprating.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 10/02/2017 08:54 pm
As they uprate to 300bar they may even be able to reduce engine count.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/02/2017 08:57 pm
As they uprate to 300bar they may even be able to reduce engine count.

Or increase the payload weight, like was done with Falcon 9

Rather have an increase from 150 ton to 200 ton cargo
Than 25 engines instead of 31 and redesign the plumbing
Load structure, guidance system etc.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 10/02/2017 10:33 pm



"Now", only.  I await the progressive uprating.

I can't wait for BFR Full Thrust followed by BFR Block 5.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/03/2017 01:13 am
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.

- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/03/2017 01:31 am

- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.

- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.

John

Hi JW,

How did expected cost scale with:

1.  Number of parts?
2.  Chamber Pressure?
3.  Fuel / Oxidizer Choice?
4.  Engine Cycle?
5.  # of Turbopumps?
6.  Any other interesting tid bits I may have missed?

Also, how accurate do you believe these models were?  To use your example, how do you believe they verified cost scaled at ~(thrust)^(0.7)?

I've always been curious how well these models work and was just wondering what your opinion of them were?

Thanks,

C

Thrust to the .7 power is a very rough ROM. It assumes similar design complexity, similar development processes, similar fabrication. If you dig into Transcost, you will see additional parameters that cover some of these. Its an historical data based trend, as are most costing estimates. It might get you +-30% range if used within a company with established processes. If used blindly +- 100%. Believe it or not, even that is better than nothing.

Most development cost is for man and machine time. Biggest unknown is how many design, build, test iterations each subsystem will have to go through during development. Each iteration of each subsystem takes time and money. The more complexity and uncertainty, the more time and money.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ChaoticFlounder on 10/03/2017 02:11 am
Wonderful info, thank you for sharing and I'm sure everyone else here appreciates it as well.

Another question I have for the forum, do we know of any other engines that have been used to date that inject both a gaseous oxidizer and gaseous fuel?

I say gaseous oxidizer and fuel because that's essentially what's coming in as a result of FFSC.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mnelson on 10/03/2017 02:53 am
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.

- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.

John

Really great, John. Thanks!

Is that a real picture of the demo engine or just something to use as an illustration? Do you have an inside source that has given you the current throat diameter or is that info generally available and I just missed it?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 10/03/2017 03:48 am
Wonderful info, thank you for sharing and I'm sure everyone else here appreciates it as well.

Another question I have for the forum, do we know of any other engines that have been used to date that inject both a gaseous oxidizer and gaseous fuel?

I say gaseous oxidizer and fuel because that's essentially what's coming in as a result of FFSC.

The 1960s Russian RD-270 (N2O4-UDMH, not flown),

https://en.m.wikipedia.org/wiki/RD-270

RD-270M (N2O4-pentaborane)

http://www.astronautix.com/r/rd-270m.html

the mid-2000's AJR Integrated Powerhead Demonstrator LOX-LH2, not a full engine),

https://en.m.wikipedia.org/wiki/Integrated_Powerhead_Demonstrator
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/03/2017 03:59 am
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.

- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.

John

Really great, John. Thanks!

Is that a real picture of the demo engine or just something to use as an illustration? Do you have an inside source that has given you the current throat diameter or is that info generally available and I just missed it?

No inside source. The picture was of the 2016 sized Raptor from SpaceX. The scale was based on NASA's CEA and methodology in their book SP-125. RPA software is similar. The assumption is that the Demo engine has the same layout. It may not, but the chamber and nozzle has to be very close to the size shown. These methods get you within a couple of percent given what information we've gotten from SpaceX.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: S.Paulissen on 10/03/2017 04:24 am
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.

- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.

John

Really great, John. Thanks!

Is that a real picture of the demo engine or just something to use as an illustration? Do you have an inside source that has given you the current throat diameter or is that info generally available and I just missed it?
Looks like a Merlin render.

If you have the thrust and chamber pressure and propellant you can reasonably estimate almost every dimension using basic rocket engine equations.  Though LW likely did something closer to a full simulation to match known performance values. I.e. Thrust, isp, chamber pressure, propellant type, expansion ratio, engine cycle etc.


Beat me to it, plus corrected me. Oh well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/03/2017 11:32 am
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.

"Now", only.  I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 10/03/2017 12:35 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.

"Now", only.  I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.
What is this fascination of yours with F-1 class rocket engines?

Raptor will not become F-1 class. F-1 class is way too big for BFR/ITS purposes. And then there is the affordability aspect of F-1 class engines: too d*rn expensive.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/03/2017 12:53 pm
John, at what chamber pressure would the 2017 engine produce 1000 kN?

Is it possible that the 1000 kN demo is the same turbopumps and chamber with a lower pressure rating and short nozzle?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 10/03/2017 12:55 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.

"Now", only.  I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.

Why bother? Lots of reliable engines does not increase the risk sufficiently to require big engines, which I think would make the whole system much harder to maintain, and lots more expensive.

I look on it like this. It's fairly easy to work on a small car engine. As they get larger, e.g. a truck engine,  they are more difficult to remove and handle, and require more and more specialist handling equipment, and more manpower. Harder to test as well, and also require more expensive production lines.

Not only that, but you start to lose economies of scale. If you are making lots of small engines, your production runs are larger, which usually means cheaper components, which also give more opportunity for manufacturing optimisation.

I can see the Raptor being not much more to build than the Merlin C, once they really get the production line going. Their sizes are fairly similar (from what I understand), there is of course a bit more material and complexity with SC, but I don't think it would be enough to really hike the price up hugely over the Merlin. Happy to be proven wrong on that one though.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Explorer on 10/03/2017 01:57 pm
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.

For F1 size that would have to be gigantic future systems. I seriously doubt SpaceX will use much bigger engines for the current BFR and the reason is simple. They want to use only one type of engine, so the upper stage will land using those. They also want at least two landing engines for redundancy. Combine that with the ability to throttle to 20% and you get a maximum thrust range. Exceed tolerable limits and landing the US will be a very exiting gambit.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/03/2017 02:41 pm
I would speculate that the ability to drive a special made "something" in under a booster sitting on the launch pad (yes into the trench) and then get personnel and a spare engine up close enough to change an engine out has been looked at as a future need...

The weight and size of a Raptor will have a profound impact on what that will look like in practice...
31 smaller engines makes this an easier proposition...

Remember... The long term goal is a booster will go on a launch pad and stay vertical until craned off and swapped at heavy maintenance time...
The pad will be where the booster stays rain or shine... as crazy as that seems...
(only leaving when in flight, making money)

They don't put all the big airliners in hangers when a storm rolls in...
They chock the wheels and set the brakes and leave it out in the storm..   ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/03/2017 02:53 pm
John, at what chamber pressure would the 2017 engine produce 1000 kN?

Is it possible that the 1000 kN demo is the same turbopumps and chamber with a lower pressure rating and short nozzle?

I assumed that the Demo engine has Pc = 20 MPa and delivered 1MN at SL. I also assumed an exit area of .97 m (I am re-accessing the Demo diameter).
Any of these could be in error.  I would appreciate input from anyone with better data or confirmation.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/03/2017 04:04 pm
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.

- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.

John

Nice!

Looks like my guess of ~37 and ~120 was pretty close for expansion ratios.

The only thing that may be wrong is I think they switched the mixture ratio from 3.8:1 to ~3.6:1. The spaceship per the presentation holds 860 tonnes of O2 and 240 tonnes of CH4. 860/240 = 3.583
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 10/03/2017 04:17 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.

"Now", only.  I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.

Why bother? Lots of reliable engines does not increase the risk sufficiently to require big engines, which I think would make the whole system much harder to maintain, and lots more expensive.

I look on it like this. It's fairly easy to work on a small car engine. As they get larger, e.g. a truck engine,  they are more difficult to remove and handle, and require more and more specialist handling equipment, and more manpower. Harder to test as well, and also require more expensive production lines.

Not only that, but you start to lose economies of scale. If you are making lots of small engines, your production runs are larger, which usually means cheaper components, which also give more opportunity for manufacturing optimisation.

I can see the Raptor being not much more to build than the Merlin C, once they really get the production line going. Their sizes are fairly similar (from what I understand), there is of course a bit more material and complexity with SC, but I don't think it would be enough to really hike the price up hugely over the Merlin. Happy to be proven wrong on that one though.

Oxygen rich systems could burn through anything quickly when failed. Engine redundency does not gurantee safty in this case.

Size of Merlin 1C is not big, it could be made to the scale of BE-4 without too much trouble on tooling. This reduce stage 1 engine count to ~15, being much conventional.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/03/2017 06:43 pm
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.

At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.

Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.

"Now", only.  I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.

Why bother? Lots of reliable engines does not increase the risk sufficiently to require big engines, which I think would make the whole system much harder to maintain, and lots more expensive.

I look on it like this. It's fairly easy to work on a small car engine. As they get larger, e.g. a truck engine,  they are more difficult to remove and handle, and require more and more specialist handling equipment, and more manpower. Harder to test as well, and also require more expensive production lines.

Not only that, but you start to lose economies of scale. If you are making lots of small engines, your production runs are larger, which usually means cheaper components, which also give more opportunity for manufacturing optimisation.

I can see the Raptor being not much more to build than the Merlin C, once they really get the production line going. Their sizes are fairly similar (from what I understand), there is of course a bit more material and complexity with SC, but I don't think it would be enough to really hike the price up hugely over the Merlin. Happy to be proven wrong on that one though.

Oxygen rich systems could burn through anything quickly when failed. Engine redundency does not gurantee safty in this case.

Size of Merlin 1C is not big, it could be made to the scale of BE-4 without too much trouble on tooling. This reduce stage 1 engine count to ~15, being much conventional.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: butters on 10/03/2017 07:01 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RotoSequence on 10/03/2017 07:06 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.

Quote from: SpaceX
Approximately one minute and 19 seconds into last night’s launch, the Falcon 9 rocket detected an anomaly on one first stage engine. Initial data suggests that one of the rocket’s nine Merlin engines, Engine 1, lost pressure suddenly and an engine shutdown command was issued immediately. We know the engine did not explode, because we continued to receive data from it. Our review indicates that the fairing that protects the engine from aerodynamic loads ruptured due to the engine pressure release, and that none of Falcon 9’s other eight engines were impacted by this event.

The armor has not been proven in an energetic explosion event.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 10/03/2017 07:10 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.

What's more, I believe this is what Musk was alluding to when he said that considering plumbing and everything the optimal engine size was coming in smaller. I'd be surprised if their optimization process didn't consider the weight of the armor cells around the engines.

I assume you're referring to the CRS-1 engine anomaly. Given the structural differences between Falcon 1.0 and the current incarnation, I'm not sure it's legitimate to say it's "demonstrated in flight" ...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mme on 10/03/2017 07:26 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.

Quote from: SpaceX
Approximately one minute and 19 seconds into last night’s launch, the Falcon 9 rocket detected an anomaly on one first stage engine. Initial data suggests that one of the rocket’s nine Merlin engines, Engine 1, lost pressure suddenly and an engine shutdown command was issued immediately. We know the engine did not explode, because we continued to receive data from it. Our review indicates that the fairing that protects the engine from aerodynamic loads ruptured due to the engine pressure release, and that none of Falcon 9’s other eight engines were impacted by this event.

The armor has not been proven in an energetic explosion event.
This seems like an argument for more smaller Raptors which would fail less energetically.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/03/2017 08:38 pm
It’s logical that future engines would optimize to even smaller sizes.
A rocket engine is basically a tool which converts chemical energy into kinetic energy.
How well it does this conversion is stated by specific impulse.
A rocket engine which has reached the theoretical maximum specific impulse can only optimize further by doing the same conversion job with less engine atoms.
If the “layer of highly efficient engines” is thinner and thus has less mass the rocket can carry more payload.
Future BFR’s with 200 or much more engines really wouldn’t surprise me at all.

To go a bit further in first principles thinking:
If the layer of highly efficient engines would only be 1 cm thick a T/W ratio of more than 10.000 is feasible.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/03/2017 09:21 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

There are faaaaar more variables in engine size vs engine count trade-offs... Not just the simplistic "more engines equal more danger" that you operate under.

I'm sure you would have advised SpaceX to not go ahead with with an single F-1 class engine instead of 9 Merlins for the Falcon 9. Thankfully they did not listen to such talk back then.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/03/2017 09:34 pm
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.

- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.

John

Nice!

Looks like my guess of ~37 and ~120 was pretty close for expansion ratios.

The only thing that may be wrong is I think they switched the mixture ratio from 3.8:1 to ~3.6:1. The spaceship per the presentation holds 860 tonnes of O2 and 240 tonnes of CH4. 860/240 = 3.583

Thanks. I have run 3.6 and 3.7 and it doesn't change things to much. Propellant density decreases a little. I might look at it again.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: alang on 10/03/2017 09:41 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

There are faaaaar more variables in engine size vs engine count trade-offs... Not just the simplistic "more engines equal more danger" that you operate under.

I'm sure you would have advised SpaceX to not go ahead with with an single F-1 class engine instead of 9 Merlins for the Falcon 9. Thankfully they did not listen to such talk back then.

Airlines have been keen to reduce the number of engines, but that's more to reduce maintenance costs. It will be some time before rocket engine reliability  a is so good that maintenance cost is the driving influence. Unlike aircraft, diversion for engine replacement is not yet an option.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 10/04/2017 12:50 am
Smaller engines have higher T/W due to less fluid pressurised volume/shorter flow path lengths - up to some limit set by minimum gauge constraints, and are also likely to have lower thermally induced stresses during start/stop caused by temperature gradients/heat flux through thinner walls - improving long term reusability.

Smaller engines might also have lighter thrust structures, and will lead to shorter interstages and landing legs, and possibly slightly less overall noise as well as greater redundancy, and more rapid throttling (to allow differential throttling steering, with benefits of eliminating flexible joints and actuators + less length), but might also have higher instrumentation and control mass overhead.

Being able to handle smaller engine components by hand will also reduce tooling, manufacturing, assembly and maintenance costs considerably.  And of course there are additional learning-curve advantages of making many smaller engines that make Merlin up to an order of magnitude cheaper per unit of thrust than eg RS68.

The ultimate limit to how small is optimal is likely down to the turbomachinery efficiency and turbine inlet temperatures needed to achieve the ~30MPa chamber pressure (limited by reusability issues caused by thermally induced stresses in thrust chamber wall created by through-thickness temperature gradient driven by heat flux).  Perhaps there is an additional limit created by Isp hit of making slightly cooler fuel-rich layer near thrust chamber wall.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/04/2017 11:29 am
Smaller engines have higher T/W due to less fluid pressurised volume/shorter flow path lengths - up to some limit set by minimum gauge constraints, and are also likely to have lower thermally induced stresses during start/stop caused by temperature gradients/heat flux through thinner walls - improving long term reusability.

Smaller engines might also have lighter thrust structures, and will lead to shorter interstages and landing legs, and possibly slightly less overall noise as well as greater redundancy, and more rapid throttling (to allow differential throttling steering, with benefits of eliminating flexible joints and actuators + less length), but might also have higher instrumentation and control mass overhead.

Being able to handle smaller engine components by hand will also reduce tooling, manufacturing, assembly and maintenance costs considerably.  And of course there are additional learning-curve advantages of making many smaller engines that make Merlin up to an order of magnitude cheaper per unit of thrust than eg RS68.

The ultimate limit to how small is optimal is likely down to the turbomachinery efficiency and turbine inlet temperatures needed to achieve the ~30MPa chamber pressure (limited by reusability issues caused by thermally induced stresses in thrust chamber wall created by through-thickness temperature gradient driven by heat flux).  Perhaps there is an additional limit created by Isp hit of making slightly cooler fuel-rich layer near thrust chamber wall.


Turbomachinery gains in efficiency with size. Cusp losses on the blades decrease along with the square/cube law.

Simple Brayton cycle gas turbines of a few hundred horsepower have thermodynamic efficiencies only in the teens, but large (400MW) gas turbines can reach 40% efficiency on the same cycle.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 10/04/2017 01:25 pm
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

Who says (apart from yourself) that 7-9 is the optimum number? How have these figures been determined?
Who says that the Raptor will have a problem with turbopumps burning through?
Who says ONE engine burning through will cause LOM?

You are inventing issues where none, so far, exist.

Given SpaceX have the most experience on this, and they are an actual rocket company, I'm inclined to think that 31 for the BFR is, whilst not necessarily optimum, a fairly good approximation to it. After all, if they thought it was a really bad idea, wouldn't it be a different number? They have already stated that T/W optimises for a smaller engine than they were expecting.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/04/2017 03:01 pm
Raptor video posted by SX
https://www.instagram.com/p/BZnLMNLloBa/?taken-by=spacex
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 10/04/2017 05:00 pm
I know some here feel like it's a rule of nature, but I think 79 engines is a few too many.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/04/2017 05:13 pm
I know some here feel like it's a rule of nature, but I think 79 engines is a few too many.

Matthew

79?  ???
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DreamyPickle on 10/04/2017 07:30 pm
SpaceX seems to have completely embraced a many-engine architecture and this is quite unique in the history of spaceflight. The falcon 9 was already an outlier when it came out with 9 Merlins. Proton has 6 engines, Saturn V had 5 but most rockets have 1 or 2. The soyuz has many nozzles but what you're actually looking at is a core stage and 4 boosters each with a single engine and 4 combustion chambers. The primary motivation for having multiple engines was usually just the difficulty of building a larger one, except for F9 and New Glenn where landing is also a factor.

A post higher up proposed that the sizing is based on the requirement to land on either of two upper-stage sea-level engines. This increases the complexity of first-stage plumbing but SpaceX doesn't seem to care, the penalty might be low or even non-existing. Historically rocket builders have frequently launched extremely expensive engine development programs in search of larger engines. Were they wrong to do so?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 10/05/2017 07:41 am
Turbomachinery gains in efficiency with size. Cusp losses on the blades decrease along with the square/cube law.

Simple Brayton cycle gas turbines of a few hundred horsepower have thermodynamic efficiencies only in the teens, but large (400MW) gas turbines can reach 40% efficiency on the same cycle.

Yes, but if you look at polytropic efficiencies of turbomachinery turbines and pumps then (leaving aside tip clearance issues) when scaling the efficiencies correlate strongly to the frictional losses you would get with flow through tubes of equivalent sizes - which is to say it is largely linked to the turbulent skin friction on the flow passage surfaces.  The 100's of kg/s mass flows of rocket turbopumps are typically at sufficiently high reynolds numbers that polytropic efficiencies will not be greatly impacted by a doubling or halving of the mass flow.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Patchouli on 10/05/2017 07:58 am

Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

Expect this as well they'll either reduce the size of BFS or move to a RD-180 thrust class engine which could be built by making a bigger turbo pump and using two Raptor combustion chambers.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 10/05/2017 08:30 am

Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

Expect this as well they'll either reduce the size of BFS or move to a RD-180 thrust class engine which could be built by making a bigger turbo pump and using two Raptor combustion chambers.

Which is ridiculous when you consider how Raptor is build.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/05/2017 01:03 pm

Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

Expect this as well they'll either reduce the size of BFS or move to a RD-180 thrust class engine which could be built by making a bigger turbo pump and using two Raptor combustion chambers.

Any evidence to cite, either moving to RD-180-class or multiple combustion chambers?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 10/05/2017 01:25 pm
Raptor cant transition to multiple combustion chambers. The turbo machinery for the LOX and its pre-burner is practically integrated into the combustion chamber. Ripping this apart means basically starting from scratch, trashing the thrust to weight ratio and introducing uncountable failure modes. Please stop with this.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/05/2017 02:18 pm

Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9.

7-9 is not optimum for the booster as it does not allow engine redundancy during landing.

And IAC2016-sized raptor on IAC2010-sized craft would have meant no redundancy for landing of the BFS.
That would have been very BAD for the reliability and safety of the craft.

Quote
More O2 rich turbopumps on booster = more risk of one burning through causing LOM.

Burning through what?

If one engine fails, it shuts down.

Or when some engine starts failing, it can be shut down before it "burns through".

More engines == smaller performance hit from engine shutdown, more engines can be shutdown without LOM, BETTER reliability.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Norm38 on 10/05/2017 02:28 pm
Given the 9m architecture that was presented, there is no reason to increase the size of the Raptor.  The landing engines don't need to be bigger, and the inner ring of 6 can't be reduced in number as that disrupts symmetry.
The only thing I can see is that if there were a Raptor version that was 2x to 3x larger, then the outer ring could go from 24 engines to 12 or 8.
SpaceX would only do that if they found that the cost savings was more than the cost to develop and carry two different engines.  It's probably not.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 10/05/2017 03:26 pm
As long as SpaceX want to...
1. use same engine core for both 1st/2nd stage
2. have 2 landing engine on 2nd stage as redundancy

We will continue see alot of engines in first stage.

UNLESS
1. You want a 10G rapid deceleration on 2nd stage landing...  :P :P :P :P
2. Some amazing tech allow you to throttle engine down to 5%...  8) 8) 8) 8)

Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/06/2017 12:34 am
BFS does not have enough sea-level thrust to cope with separation and landing in an abort  below some substantial fraction of its normal staging velocity, for good reasons.

In an abort, clearly the engines can be run a little harder, which will push flow separation out a little, and help somewhat, but even that won't help at some point.

Can this cycle of engine in principle be operated dramatically off-mixture, so as to dump either very cold (relatively) gas out of the nozzle (perhaps with damage), or even partially liquid, faster than normal due to the lower pressure drop.

I would assume the answer is no, but was wondering if I was wrong.

[edit] Or ... Well - the rather simpler option of large dump valves.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/06/2017 11:01 am
BFS does not have enough sea-level thrust to cope with separation and landing in an abort  below some substantial fraction of its normal staging velocity, for good reasons.

In an abort, clearly the engines can be run a little harder, which will push flow separation out a little, and help somewhat, but even that won't help at some point.

Can this cycle of engine in principle be operated dramatically off-mixture, so as to dump either very cold (relatively) gas out of the nozzle (perhaps with damage), or even partially liquid, faster than normal due to the lower pressure drop.

I would assume the answer is no, but was wondering if I was wrong.

[edit] Or ... Well - the rather simpler option of large dump valves.

In case BFS needs to do abort immediately after takeoff, it will probably use also the vacuum engines?

Flow separation would..  Damage the nozzles? Make the flow direction unstable, making the craft hard to steer precisely?

These are problems that are not immediately fatal when there are still 2 engines that can be used to steer the rocket and we are mostly just wanting to get higher and away from the failing first stage, not land precisely?

And probably the "170 tonne raptor" could be ran at something like 190 tonne emergency thrust that would decrease it's lifetime considerably.

So it would use the vacuum engines to help to get initial T/W higher and dump fuel faster, and when it has gained some altitude the vacuum engines would start to work better. And then when it has gotten much lighter and got due the consumed fuel it could use the atmospheric engines for the actual landing which needs to be precise

Though there is the problem that the initial T/W is less than one, which gives some "black zone" immediately after liftoff when the craft is lacking vertical velocity to stay in air long enough to get rid of extra fuel to become lighter, unless the "emergency thrust" is high enough to get T/W over one (something like >210 tonnes)




Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/06/2017 12:09 pm
7-9 is not optimum for the booster as it does not allow engine redundancy during landing.
Give Raptor the capability to throttle down to 10% or even less then a 7-9 engined booster can land on 3 engines, 2 engines, or the centre engine only. This would give redundancy on landing without going to some crazy engine no. on 1st stage.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 10/06/2017 01:22 pm
Raptor is smaller from what I see so as to drastically reduce its construction costs.  One by using 3D printers, and two being smaller makes it easier for crew to work on without excessive heavy equipment and specialty equipment.  So they decided on 31 engines, with a total thrust just under what the launch pad at the Cape can handle.  They are going to have 27 engines running for FH, so what is 4 extra for BFR?  It may also be cheaper for them to mass produce 200 Raptors vs 20 in the F-1 Saturn V class.  Especially if a Raptor cost $1 million apiece vs $50 million apiece for an F-1 class engine. 

They also may not want to go through the cost, time, and hassle of making the engine larger as it might add another 5 years before getting BFR.  Perfect what they have and 31 engines should be fine.  They have already made several 100 Merlins with only one failure, and it was shut down in flight, with no LOM.  Raptor is similar in size but twice the thrust, so it should be fine. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Patchouli on 10/06/2017 06:09 pm
Raptor cant transition to multiple combustion chambers. The turbo machinery for the LOX and its pre-burner is practically integrated into the combustion chamber. Ripping this apart means basically starting from scratch, trashing the thrust to weight ratio and introducing uncountable failure modes. Please stop with this.

Not necessarily as the RD170,180,and,181 share common heritage and come off the same line though the RD-170 was the first.

The specifics of whether this is practical to do with Raptor is unknown outside of Spacex but it would entitle a larger turbo pump and preburner but such an engine probably can come off the same line as the single chamber ones.

Keep in mind Raptor is still at a fairly early stage there is no full scale engine yet so the specifics can and probably will change.

7-9 is not optimum for the booster as it does not allow engine redundancy during landing.
Give Raptor the capability to throttle down to 10% or even less then a 7-9 engined booster can land on 3 engines, 2 engines, or the centre engine only. This would give redundancy on landing without going to some crazy engine no. on 1st stage.


I think how Falcon Heavy performs will influence the decision on what the final number of engines would be though the plumbing for 31 in one core is more complex then 9 in one core tough they could run multiple manifolds and intakes to reduce interaction.

One issue I still wonder about is can they protect the other engines from a catastrophic failure of one.
Don't say it can't happen today as this kind of failure has happened recently.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 10/06/2017 07:05 pm
Raptor cant transition to multiple combustion chambers. The turbo machinery for the LOX and its pre-burner is practically integrated into the combustion chamber. Ripping this apart means basically starting from scratch, trashing the thrust to weight ratio and introducing uncountable failure modes. Please stop with this.

Not necessarily as the RD170,180,and,181 share common heritage and come off the same line though the RD-170 was the first.

The specifics of whether this is practical to do with Raptor is unknown outside of Spacex but it would entitle a larger turbo pump and preburner but such an engine probably can come off the same line as the single chamber ones.

Keep in mind Raptor is still at a fairly early stage there is no full scale engine yet so the specifics can and probably will change.

(http://woosterphysicists.scotblogs.wooster.edu/files/2016/10/Raptor.png)

The lox turbopump is integrated into the combustion chamber. As focused as SpaceX are on good T/W, they are not going to go make that separate and require a heavy manifold, and they are not going to oversize it so the same turbopump can feed multiple chambers. That just doesn't make any sense.

Also from the 2017 IAC talk, the raptor they are testing is pretty much the size they will use, they are just going to increase the chamber pressure from 200 bar to 250 bar. Also since they've been testing this thing nearly once a week for a year now it's not really "early stage" any more.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 07:30 pm
Here is high-tech mspaint flow chart (as I understand it) of the raptor:

The key to Raptor's high T/W ratio is the minimization of high-pressure piping.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/06/2017 07:49 pm
Very nice drawing  :) I like it!

How high T/W? why is it not made public?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/06/2017 07:51 pm
Here is high-tech mspaint flow chart (as I understand it) of the raptor:

The key to Raptor's high T/W ratio is the minimization of high-pressure piping.

Doesn't the fuel flow go: first compressor stage -> regen nozzle cooling -> second compressor stage -> fuel preburner -> fuel turbine -> injectors?

You have it going through the regen channels after all compressor stage, meaning that the regen nozzle has to handle fuel at greater than chamber pressure. That's a LOT of high-pressure piping, which is what you say it's trying to minimize.

The 2nd stage fuel pump should be under the fuel preburner, and the fuel coming out of the regen should go right up into it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/06/2017 07:53 pm
Very nice drawing  :) I like it!

How high T/W? why is it not public?
Very likely Raptor's TWR has not yet been finalized so will not announce it until it is. May also be restricted by ITAR.

Also notice that SpX have only shown videos of Raptor firings in the dark to deliberately hide the engine's turbomachinery. Again could be down to ITAR.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/06/2017 07:54 pm
Very nice drawing  :) I like it!

How high T/W? why is it not public?
Very likely Raptor's TWR has not yet been finalized so will not announce it until it is. May also be restricted by ITAR.

Also notice that SpX have only shown videos of Raptor firings in the dark to deliberately hide the engine's turbomachinery. Again could be down to ITAR.

SpaceX published a CAD rendering of the engine. What something looks like isn't generally ITAR controlled.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/06/2017 07:58 pm
If Elon says at IAC 2016 on why such small engines? “similar engine size but 3 times the thrust” I asume T/W is around 600 because Merlin 1D is 200.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Patchouli on 10/06/2017 07:59 pm
Here is high-tech mspaint flow chart (as I understand it) of the raptor:

The key to Raptor's high T/W ratio is the minimization of high-pressure piping.

That clears it up I guess the most they could do without changing the flow characteristics is something along the lines of the LR87 which wouldn't be of much benefit except for packaging.

It seems the design would also help with manufacture and reduce the amount of ceramic coated parts needed on the O2 side by eliminating a lot of plumbing.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/06/2017 08:01 pm
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.
Raptor TWR of 600 is likely impossible but it may exceed 250.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 08:01 pm
Here is high-tech mspaint flow chart (as I understand it) of the raptor:

The key to Raptor's high T/W ratio is the minimization of high-pressure piping.

Doesn't the fuel flow go: first compressor stage -> regen nozzle cooling -> second compressor stage -> fuel preburner -> injectors?

You have it going through the regen channels after all compressor stage, meaning that the regen nozzle has to handle fuel at greater than chamber pressure. That's a LOT of high-pressure piping, which is what you say it's trying to minimize.

No, because you can get work from regenerative cooling.  Both the turbines run the brayton cycle (compression>Heat added>Turbine/work>exhaust) you use it in effect as a recuperator/mini expander cycle. If you didn't you'd have to work against it instead and it wouldn't be regenerative cooling.

Recuperation/regeneration is used on regular gas turbines as well.
(http://www.alentecinc.com/images/gtrecup.gif)

EDIT: derp, meant to show SSME:
http://pages.erau.edu/~ericksol/courses/sp210/images/ssme_schem.jpg


78% of the flow, by mass, is the oxidizer.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/06/2017 08:06 pm
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.
Raptor TWR of 600 is likely impossible but it may exceed 250.

Why imposible? It looks like the same weight or lighter even when the bell is bigger
Should be between 400 and 600
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 08:07 pm
If Elon says at IAC 2016 on why such small engines? “similar engine size but 3 times the thrust” I asume T/W is around 600 because Merlin 1D is 200.

Not likely, probably over 200 though.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/06/2017 08:09 pm
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.
Raptor TWR of 600 is likely impossible but it may exceed 250.

Why imposible?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 08:16 pm
Here is what I mean by minimizing high pressure piping;

The weight of pressure vessels scales with the the mass in the the pressure vessel and the pressure.

Here is the RD-180:
http://www.markelwood.com/images/spaceart/RD-180.jpg

Those two big pipes coming out of the top of the turbine are the O2 rich gas pipes. ~73% of the mass flow of the fuel is going through those two big pipes, it looks like it has to travel 2-3 meters before getting to the combustion chamber. In the Raptor, which is even more oxidizer-rich (oxidizer is ~78% of propellant), it only travels inches.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/06/2017 08:20 pm
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.
Raptor TWR of 600 is likely impossible but it may exceed 250.

Why imposible?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
It only travels inches as stated above.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 08:21 pm
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.
Raptor TWR of 600 is likely impossible but it may exceed 250.

Why imposible?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?

It's at much lower pressure
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/06/2017 08:23 pm
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?

How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/06/2017 08:24 pm
Here is what I mean by minimizing high pressure piping;

The weight of pressure vessels scales with the the mass in the the pressure vessel and the pressure.

Here is the RD-180:
http://www.markelwood.com/images/spaceart/RD-180.jpg

Those two big pipes coming out of the top of the turbine are the O2 rich gas pipes. ~73% of the mass flow of the fuel is going through those two big pipes, it looks like it has to travel 2-3 meters before getting to the combustion chamber. In the Raptor, which is even more oxidizer-rich (oxidizer is ~78% of propellant), it only travels inches.

Sure, I see what you mean (although pressure vessel mass scales with volume, not mass flow, and oxidizer is a lot denser at a given temperature and pressure). I still find it crazy that all those little pipes in the SSME regen nozzles were at 6000+ PSI.

I would think that the first stage fuel pump could get enough pressure to force the fuel through the regen channels and into the 2nd stage pump at a high enough rate. I guess not, because they would certainly save a lot more mass if it were possible.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/06/2017 08:26 pm
...
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

I wonder if any of the Raptor plumbing is carbon fiber overwrapped. It sure saves a lot of mass in COPVs, and not all the fluid flows are high temp.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/06/2017 08:31 pm
We will have to wait until or if SpX release the TWR figure for Raptor for us all to know what it is. They could decide to never release it and leave us all in the dark.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/06/2017 08:31 pm
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?

How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.

It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could easily say “similar sized engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust?
They are all bigger or have the same thrust. So why does he say that? Probably he meant wheight instead of size, then it makes sense, otherwise not.

But I dont know either, only that its “the highest TWR of any engine, of any kind”, so above 200.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/06/2017 08:38 pm
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?

How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.

It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could say “same size engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust
They are all bigger or have the same thrust. So why does he say that? Probably he meant by size, wheight, then it makes sense. Otherwise not
Raptor and M1D are completely different engines being FFSC LOx/LCH4 and GG LOx/RP-1 respectively. Raptor no longer has 3x M1D thrust more like 2x now. So your assumption of Raptor TWR of 600 based on Raptor having same size as M1D with 3x the thrust is not credible.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 08:38 pm
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?

How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.

It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could easily say “same size engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust
They are all bigger or have the same thrust. So why does he say that? Probably he meant by size, wheight, then it makes sense. Otherwise not

Size =/= weight. The Raptor is higher pressure, it will be denser, the metal parts much thicker, even if it's a similar physical size.

He was talking about the size of the combustion chamber, the Raptor has a higher expansion ratio though (~35 vs 16) so it's nozzle is larger (~0.9m vs ~1.3m).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/06/2017 08:53 pm
Really... we need to wait if or until SpaceX releases some specs at a later date TBD...

That said...
My guess of a mass of ~980kg (almost 1 metric ton) and a thrust stated at 170 to 190 metric tons
Puts Raptor in about the same thrust to weight ratio as Merlin 1D full thrust... 180 to 1
I will add my thought of "Good Enough" to this... No real need to try and beat that...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 09:03 pm
Really... we need to wait if or until SpaceX releases some specs at a later date TBD...

That said...
My guess of a mass of ~980kg (almost 1 metric ton) and a thrust stated at 170 to 190 metric tons
Puts Raptor in about the same thrust to weight ratio as Merlin 1D full thrust... 180 to 1
I will add my thought of "Good Enough" to this... No real need to try and beat that...  ;)

I'm basing my prediction of >200:1 on this:
https://www.youtube.com/watch?v=tdUX3ypDVwI
FF to 5:45

"So the, the Raptor Engine will be the highest thrust to weight engine, we believe, on any engine of any kind ever made" - Elon Musk.

Merlin 1D is already at ~200:1
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/06/2017 09:07 pm
200:1 is the planned 300 bar later version... in my opinion...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/06/2017 09:10 pm
200:1 is the planned 300 bar later version... in my opinion...

I guess we'll see....  ;D ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DreamyPickle on 10/06/2017 09:11 pm
If Elon says at IAC 2016 on why such small engines? “similar engine size but 3 times the thrust” I asume T/W is around 600 because Merlin 1D is 200.
You're taking things way too literally. In this context "similar size" does not mean "equal mass". WTF?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/07/2017 02:48 am
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.

And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?

How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.

It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could easily say “similar sized engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust?
They are all bigger or have the same thrust. So why does he say that? Probably he meant wheight instead of size, then it makes sense, otherwise not.

But I dont know either, only that its “the highest TWR of any engine, of any kind”, so above 200.

It is similar in size to the Merlin 1D and 3 times the thrust because it's about 3 times the pressure! It is made from similar materials as the Merlin (Copper alloy and high temperature Nickel alloys) so it will weigh approximately 3 times as much because both are basically very complex pressure vessels! Hence; thrust to weight stays about the same. 

Also, I am pretty sure that ZachF's rough sketch of the propellant flow is correct.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/07/2017 03:02 am
I posted this earlier, but here is a picture showing sizes of the Demonstrator engine, the 2016 engine and the new smaller 2017 engine. The 2017 Raptor appears to be about a 15% scale up of the Demonstrator Raptor. Today I re-estimated the demonstrator engine exit diameter from the best picture we have. I think it is closer to .94 m which would make its expansion ratio closer to 25:1 instead of 26:1. I am also working up a Pc = 3000 psi engine.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 10/07/2017 03:08 am
I wonder if any of the Raptor plumbing is carbon fiber overwrapped. It sure saves a lot of mass in COPVs, and not all the fluid flows are high temp.

No chance, too much mismatch in thermal expansion coefficients, near impossible to do wrapping due to poor accessibility, potential fire danger around oxygen, and carbon fibre doesn't have the ability to handle more than about 200-250°C during reentry.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/07/2017 03:12 am
I wonder if any of the Raptor plumbing is carbon fiber overwrapped. It sure saves a lot of mass in COPVs, and not all the fluid flows are high temp.

No chance, too much mismatch in thermal expansion coefficients, near impossible to do wrapping due to poor accessibility, potential fire danger around oxygen, and carbon fibre doesn't have the ability to handle more than about 200-250°C during reentry.
Carbon-carbon, on the other hand...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 10/07/2017 04:14 am
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/07/2017 11:44 am
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 10/07/2017 01:16 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.

Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 10/07/2017 02:44 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.

Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 10/07/2017 02:53 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.

Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.

But it wouldn't contribute to thrust, it would be ejected from the severed regen channels and burn somewhere behind the vehicle.
It would also cut off cooling to the remaining part of the nozzle and the chamber, leading to very rapid engine failure.
So any sort of jettisonable nozzle is going to have to address this anyway by redirecting the coolant pathway.

Edited to add: simply chopping off the nozzle would actually lower T:W because all that fuel is lost rather than going to the combustion chamber. So I would assert that even if it is only for use in dire emergencies, it is essential that any sort of nozzle jettison capability must be accompanied by a redirect of the regen pathway. Not impossible, I'm sure, just an added complication.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 10/07/2017 10:01 pm
Raptor is on the path of needing to exceed LM ASC engine reliability. That's a tall order to fill. (If those point to point transport on Earth graphics are "real", likely engine reliability would have to approach commercial transport turbofan reliability, which is three orders plus of magnitude higher.) To prove this would require extreme testing/use/reuse.

One could "concern troll" that if AJR/BO can't test to such, then SX couldn't ever do such, omitting the fact that they seem to be able to meet reliability margins above industry norms.

Copying this over from another thread, I always assumed that SpaceX tested Merlin 1D extensively development versus industry norms.  But as I posted above, Merlin 1C 's development program was about 3,000 seconds of firing (http://www.spacex.com/press/2012/12/19/spacex-completes-development-merlin-regeneratively-cooled-rocket-engine).  I'm taking a look at the SSME Block III upgrade proposal (https://forum.nasaspaceflight.com/index.php?topic=35269.msg1312483#msg1312483), which quoted 38,000 seconds of firing.

What are the industry norms on development testing?  Was the SSME Block III proposal especially gold-plated?  Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/07/2017 10:54 pm
What are the industry norms on development testing?  Was the SSME Block III proposal especially gold-plated?  Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?

Some things to educate yourself with:

Test and Evaluation Guideline for Liquid Rocket Engines (http://www.dtic.mil/dtic/tr/fulltext/u2/a554916.pdf)

Liquid Rocket Engine Flight Certification (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910018936.pdf)

In general, look at the acceptance criteria of contracts for vehicles engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: groundbound on 10/08/2017 05:46 am
What are the industry norms on development testing?  Was the SSME Block III proposal especially gold-plated?  Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?

Some things to educate yourself with:

Test and Evaluation Guideline for Liquid Rocket Engines (http://www.dtic.mil/dtic/tr/fulltext/u2/a554916.pdf)

Liquid Rocket Engine Flight Certification (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910018936.pdf)

In general, look at the acceptance criteria of contracts for vehicles engines.

Very interesting, thanks. One thing I picked up is "Testing should demonstrate margin on maximum specified operating life." If you read that literally and simplistically, then all the claims for BFR booster design life imply an extremely long test program.

I'm guessing that there are other ways to verify that particular margin than 400,000 sec (or whatever) of test time.  :o 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/08/2017 07:36 am
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?

What’s the Block 5 Merlin ?
Is it Merlin 1E ?

Is it 145% more heavy?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/08/2017 01:50 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?

What’s the Block 5 Merlin ?
Is it Merlin 1E ?

Is it 145% more heavy?

The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9.  It is basically the same mass engine AFAIK.  The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 10/08/2017 03:07 pm
What are the industry norms on development testing?  Was the SSME Block III proposal especially gold-plated?  Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?

Some things to educate yourself with:

Test and Evaluation Guideline for Liquid Rocket Engines (http://www.dtic.mil/dtic/tr/fulltext/u2/a554916.pdf)

Liquid Rocket Engine Flight Certification (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910018936.pdf)

In general, look at the acceptance criteria of contracts for vehicles engines.

Very interesting, thanks. One thing I picked up is "Testing should demonstrate margin on maximum specified operating life." If you read that literally and simplistically, then all the claims for BFR booster design life imply an extremely long test program.

For sure, it is interesting to take those two documents together, because it shows that none of the engines detailed in the NASA/Richards document (SSME, F-1, J-2, RL-10, LR87, and LR91) had qualification requirements that demonstrated margin.  The SSME had a design life of 27,000 seconds, but a qualification requirement of only 5,000 seconds, at least for the first iteration in the late 70s/early 80s.

That said, a NASA 2011 powerpoint (https://www.nasa.gov/pdf/553045main_Space_Shuttle_Main_Engine_Van_Hooser.pdf) at page 15 appears to show a development/qualification/testing program taken in its entirety to be robust.  Before its first flight, SSME had on the order of 145,000 seconds over 700 test firings.  The design program seems to have been a bit rocky.  Perhaps that necessitated starting over the design testing a lot.

The NASA/Richards document is very good and succinct.  It states clearly that well into the modern rocket age (it is undated but perhaps in the late 80s), there were no industry/government-wide rules and requirements for design and certification and that processes were historically based and heuristic.  Basically, you do design testing until you are satisfied that you are done.  And so no two design testing campaigns are identical.

Edit:  The NASA/Richards document also shows that none of the rocket engines listed were tested for FOD ingestion.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/08/2017 03:13 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?

What’s the Block 5 Merlin ?
Is it Merlin 1E ?

Is it 145% more heavy?

The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9.  It is basically the same mass engine AFAIK.  The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'

A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 10/08/2017 05:43 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?

What’s the Block 5 Merlin ?
Is it Merlin 1E ?

Is it 145% more heavy?

The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9.  It is basically the same mass engine AFAIK.  The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'

A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.

John

Understood, but the original question was asking for the BFS Abort system thread. While running engines at higher than rated values greatly reduces the life of the engine, I am asking whether it may, in a dramatic situation, contribute to an increase in the life of the payload, given what we know about expansion ratios and TWR. :p
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/08/2017 06:13 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?

What’s the Block 5 Merlin ?
Is it Merlin 1E ?

Is it 145% more heavy?

The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9.  It is basically the same mass engine AFAIK.  The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'

A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.

John

Understood, but the original question was asking for the BFS Abort system thread. While running engines at higher than rated values greatly reduces the life of the engine, I am asking whether it may, in a dramatic situation, contribute to an increase in the life of the payload, given what we know about expansion ratios and TWR. :p

Falcon 9 Block 5 will have higher thrust than Falcon 9 Full Thrust. We don’t know if the 145%, is higher than planned operation.  Since we don’t know how much more thrust the Fuller than Full thrust will be...


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MP99 on 10/09/2017 06:54 am
Very nice drawing  :) I like it!

How high T/W? why is it not made public?
Musk said it was the best ever, so better than M1D.

Cheers, Martin

Sent from my Nexus 6 using Tapatalk

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/09/2017 04:33 pm
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?

What’s the Block 5 Merlin ?
Is it Merlin 1E ?

Is it 145% more heavy?

The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9.  It is basically the same mass engine AFAIK.  The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'

A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.

John

Understood, but the original question was asking for the BFS Abort system thread. While running engines at higher than rated values greatly reduces the life of the engine, I am asking whether it may, in a dramatic situation, contribute to an increase in the life of the payload, given what we know about expansion ratios and TWR. :p

Falcon 9 Block 5 will have higher thrust than Falcon 9 Full Thrust. We don’t know if the 145%, is higher than planned operation.  Since we don’t know how much more thrust the Fuller than Full thrust will be...

But we do know -- the Block 5 engines are to provide 190,000 lbf of thrust.  The 145% is just a test to destruction (or margin verification) as previously stated.  Please stop trying to make the Block 5 M-1D a 145% rating.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nilof on 10/11/2017 04:02 am
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.

Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.

But it wouldn't contribute to thrust, it would be ejected from the severed regen channels and burn somewhere behind the vehicle.
It would also cut off cooling to the remaining part of the nozzle and the chamber, leading to very rapid engine failure.
So any sort of jettisonable nozzle is going to have to address this anyway by redirecting the coolant pathway.

Edited to add: simply chopping off the nozzle would actually lower T:W because all that fuel is lost rather than going to the combustion chamber. So I would assert that even if it is only for use in dire emergencies, it is essential that any sort of nozzle jettison capability must be accompanied by a redirect of the regen pathway. Not impossible, I'm sure, just an added complication.

Does the extension skirt actually need active cooling, or is it radiatively cooled? I'd expect the exhaust to be rather cold when it has expanded 30 times or so...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 10/11/2017 08:03 am
Question for the thread:  I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.

Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.

But it wouldn't contribute to thrust, it would be ejected from the severed regen channels and burn somewhere behind the vehicle.
It would also cut off cooling to the remaining part of the nozzle and the chamber, leading to very rapid engine failure.
So any sort of jettisonable nozzle is going to have to address this anyway by redirecting the coolant pathway.

Edited to add: simply chopping off the nozzle would actually lower T:W because all that fuel is lost rather than going to the combustion chamber. So I would assert that even if it is only for use in dire emergencies, it is essential that any sort of nozzle jettison capability must be accompanied by a redirect of the regen pathway. Not impossible, I'm sure, just an added complication.

Does the extension skirt actually need active cooling, or is it radiatively cooled? I'd expect the exhaust to be rather cold when it has expanded 30 times or so...

Quote
Will be full regen cooled all the way out to the 3 meter (10 ft) nozzle diameter. Heat flux is nuts & radiative view factor is low.

https://mobile.twitter.com/elonmusk/status/877341165808361472?lang=en-gb
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Oersted on 10/12/2017 07:52 pm
Gwynne Shotwell Q&A. The quote below is not verbatim but from notes by Reddit-user "Sticklefront": https://www.reddit.com/r/spacex/comments/75ufq9/interesting_items_from_gwynne_shotwells_talk_at/

"What is the biggest obstacle to the BFR's success?

The composite tanks will be challenge, but we are doing it already. We are currently building a larger raptor right now, and currently have a scaled version of raptor on the test stands. Harder than the rocket, though, will be where poeple are going to live, what will life be like, what will they do there? Also, while the choice of fuel for the BFR was constrained by resource availability on Mars, it is no accident that the final choice of methane is the cheapest energy source here on earth. This will greatly facilitate the economics side of things."
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Manabu on 10/13/2017 03:25 pm
Historically rocket builders have frequently launched extremely expensive engine development programs in search of larger engines. Were they wrong to do so?
Probably not.

Historically, rockets used analog computers to control the engines. More engines increase complexity more than linearly, which means both heavier and more difficult to design avionics. Think N1 KORD as an extreme and failed example.

Miniaturized digital computers you can program, optical cables, etc, diminish the mass and complexity of such systems significantly, allowing more engines to be used economically. So, probably since the 90s, multiple engine rockets became more viable.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/13/2017 04:25 pm
Historically rocket builders have frequently launched extremely expensive engine development programs in search of larger engines. Were they wrong to do so?
Probably not.

Historically, rockets used analog computers to control the engines. More engines increase complexity more than linearly, which means both heavier and more difficult to design avionics. Think N1 KORD as an extreme and failed example.

Miniaturized digital computers you can program, optical cables, etc, diminish the mass and complexity of such systems significantly, allowing more engines to be used economically. So, probably since the 90s, multiple engine rockets became more viable.

I disagree... Many engines have been a viable option since the beginning of the space age.

Just look at the Saturn I(B), it flew fine with 8 engines in the 60's. R-7 has flown with 5 engines for decades. The problem with the N-1 was primarily a lack of testing. More engines have been a viable option for many decades, and it took SpaceX to break that industry trend of "fewer is better, one is best" that was reaching absurd levels in the last few decades.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: intrepidpursuit on 10/13/2017 10:40 pm
The F1 was built by hand with people welding thousands of parts together to make mechanical works of art. The RL-10 was made basically the same way. When the manufacturing method and complexity doesn't scale, you want to make as big an engine as you can so there are fewer places to go wrong.

Now we have the capability to mass produce something the size of a Raptor, which brings the cost per unit down dramatically. The effort goes into building the tooling rather than the engine. Now, if you build a bigger engine you are building bigger tooling that will get less use, so you get more cost and less reliability inherently. If you can build a smaller engine that is an order of magnitude more reliable than the old hand built engines then you want to design for manufacturability.

I think the difference is in the way engines are built now and the way they can be modeled. Economy of scale shifts to favoring quantity over size.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 10/13/2017 11:38 pm
As has been noted about the Merlin engines in the F-9, lots of engines means a much more rapid acquisition of reliability data. A successful flight of the booster will pile up somewhere around 4500 seconds of engine time. Engine data will be amassed three plus times faster than on the F-9, 31 times faster than on Atlas or Vulcan.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 10/15/2017 08:08 pm
Future BFR’s with 200 or much more engines really wouldn’t surprise me at all.
Raptor no. on BFR system has been determined by the need for the ship to have engine out capability for landing and using single engine design throughout the system while keeping complexity to the minimum required level. For future larger BFR's just scale up Raptor thrust with BFR system mass to keep engine no. same as current BFR.

EM said in his recent Reddit AMA that Raptor can easily be scaled from 1MN to 1.7MN at SL so further scale ups of Raptor for future larger BFR systems should not be too difficult.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 10/15/2017 08:37 pm
Excerpts from Musk's AMA yesterday related to Raptor.

------

Q:  Why was Raptor thrust reduced from ~300 tons-force to ~170 tons-force?
A:  We chickened out
A:  The engine thrust dropped roughly in proportion to the vehicle mass reduction from the first IAC talk. In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you've lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.

Q:  Could you update us on the status of scaling up the Raptor prototype to the final size?
A:  Thrust scaling is the easy part. Very simple to scale the dev Raptor to 170 tons.

The flight engine design is much lighter and tighter, and is extremely focused on reliability. The objective is to meet or exceed passenger airline levels of safety. If our engine is even close to a jet engine in reliability, has a flak shield to protect against a rapid unscheduled disassembly and we have more engines than the typical two of most airliners, then exceeding airline safety should be possible.

That will be especially important for point to point journeys on Earth. The advantage of getting somewhere in 30 mins by rocket instead of 15 hours by plane will be negatively affected if "but also, you might die" is on the ticket.

Q:  Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?
A:  The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).

Q:  Will the BFR autogenous pressurization system be heat exchanger based?  You told us previously that the BFR will eliminate the use of Helium and use hot oxygen and hot CH4 to auto-pressurize the propellant tanks.  Can you tell us more about this new system, will it involve heating the propellants at the engines via heat exchangers and routing the hot gas back to the tanks via pipes, or will they use some other method?  If it's heat exchanger based, will all Raptor engines have heat exchangers?
A:  We plan to use the Incendio spell from Harry Potter (http://harrypotter.wikia.com/wiki/Fire-Making_Spell)
A:  But, yes and probably

Q:  Will Raptor engines be (metal-) 3D printed?
A:  Some parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.

Q:  Can BFS vacuum-Raptors be fired at sea level pressure?
A:  The "vacuum" or high area ratio Raptors can operate at full thrust at sea level. Not recommended.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 10/15/2017 09:05 pm
Quote from: EM AMA
The flight engine design is much lighter and tighter, and is extremely focused on reliability. The objective is to meet or exceed passenger airline levels of safety. If our engine is even close to a jet engine in reliability, has a flak shield to protect against a rapid unscheduled disassembly and we have more engines than the typical two of most airliners, then exceeding airline safety should be possible.

The NASA/Richards document continues to be golden.  Thanks again, Space Ghost.  It shows a jet fighter engine qualification requirement to be 150 hours (540,000 seconds), or roughly two orders of magnitude more than the original SSME qualification requirement.

The 150 hour requirement also appears to be replicated in the FAA type certification requirements for endurance testing (https://www.ecfr.gov/cgi-bin/text-idx?SID=eed43786296c5051130faf9170d05790&mc=true&node=pt14.1.33&rgn=div5#se14.1.33_149).  Perhaps because Raptor only fires for a short time compared to jet engines, the qualification requirements arrived at for Raptor may be less, at least in duration.

Edit:  Reliability for modern jet engines seems pretty extreme.  GE's G90 powerplant (used on the Boeing 777) is said to have an in-flight shutdown rate of one per million engine flight hours (https://www.geaviation.com/press-release/ge90-engine-family/record-year-worlds-largest-most-powerful-jet-engine).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: yokem55 on 10/15/2017 10:40 pm
Quote
Q:  Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?
A:  The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).
I'm curious what kind of ignition these will use. I would think even sparker ignition might be too slow. Could they have a hot ignition coil in the combustion chamber that stays heated when it is anticipated that they might need to be fired?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/15/2017 10:43 pm
Quote from: EM AMA
The flight engine design is much lighter and tighter, and is extremely focused on reliability. The objective is to meet or exceed passenger airline levels of safety. If our engine is even close to a jet engine in reliability, has a flak shield to protect against a rapid unscheduled disassembly and we have more engines than the typical two of most airliners, then exceeding airline safety should be possible.

The NASA/Richards document continues to be golden.  Thanks again, Space Ghost.  It shows a jet fighter engine qualification requirement to be 150 hours (540,000 seconds), or roughly two orders of magnitude more than the original SSME qualification requirement.

The 150 hour requirement also appears to be replicated in the FAA type certification requirements for endurance testing (https://www.ecfr.gov/cgi-bin/text-idx?SID=eed43786296c5051130faf9170d05790&mc=true&node=pt14.1.33&rgn=div5#se14.1.33_149).  Perhaps because Raptor only fires for a short time compared to jet engines, the qualification requirements arrived at for Raptor may be less, at least in duration.

Edit:  Reliability for modern jet engines seems pretty extreme.  GE's G90 powerplant (used on the Boeing 777) is said to have an in-flight shutdown rate of one per million engine flight hours (https://www.geaviation.com/press-release/ge90-engine-family/record-year-worlds-largest-most-powerful-jet-engine).
Long haul intercontinental jet flights run the engines 100 to 1000 times longer per flight then a P2P rocket would.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Gliderflyer on 10/16/2017 08:49 am
Quote
Q:  Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?
A:  The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).
I'm curious what kind of ignition these will use. I would think even sparker ignition might be too slow. Could they have a hot ignition coil in the combustion chamber that stays heated when it is anticipated that they might need to be fired?
I don't have any inside information, but I would bet they will use normal spark torch igniters. I have worked with them before, and they have a pretty fast response time that should be more than adequate for an RCS system.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 10/16/2017 09:51 am
Quote
Q:  Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?
A:  The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).
I'm curious what kind of ignition these will use. I would think even sparker ignition might be too slow. Could they have a hot ignition coil in the combustion chamber that stays heated when it is anticipated that they might need to be fired?
I don't have any inside information, but I would bet they will use normal spark torch igniters. I have worked with them before, and they have a pretty fast response time that should be more than adequate for an RCS system.

The Morpheus moon lander testbed uses spark ignition and it does high frequency firing bursts. I guess a 5 or 10t thruster will not be quite as fast but it does not need to be.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kaputnik on 10/16/2017 11:32 am
How will they get the required tank pressure if they only need a short burst? Seeing as they need to build up some energy in the heat exchanger.
(Unless he was serious about the Harry Potter thing)...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/16/2017 11:58 am
How will they get the required tank pressure if they only need a short burst? Seeing as they need to build up some energy in the heat exchanger.
(Unless he was serious about the Harry Potter thing)...

You don’t need high tank presure just high chamber pressure.
For this they could use an air driven hydraulic pump, like this one.
Instant respons high presure, up to 7000 bar.
It’s similar to a turbo pump but much lower flow, and better presure control.

http://www.haskel.com/products/pneumatic-pumps/liquid-pumps/

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/16/2017 12:23 pm
How will they get the required tank pressure if they only need a short burst? Seeing as they need to build up some energy in the heat exchanger.
(Unless he was serious about the Harry Potter thing)...

There will be separate high pressure gaseous LOX and CH4 tanks for RCS. Liquid propellants can easily be electrically pumped in and vaporized by heating. Heating could be electrical, chemical or, if engines are running, tapped off  the engines.

The heat exchangers on the Raptor are primarily for autogenesis main tank pressurization.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/16/2017 04:00 pm
Are RCS motors fed with gaseous propellants or liquid? I seem to remember someone stating that they would be gaseous, but now I am not so sure.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/16/2017 04:07 pm
Are RCS motors fed with gaseous propellants or liquid? I seem to remember someone stating that they would be gaseous, but now I am not so sure.

John

Musk said at 2016 IAC that they would be gaseous. That hasn't changed, as far as I know.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Oersted on 10/16/2017 04:11 pm
From the Reddit AMA:

QUESTION 1

Why was Raptor thrust reduced from ~300 tons-force to ~170 tons-force?

One would think that for (full-flow staged combustion...) rocket engines bigger is usually better: better surface-to-volume ratio, less friction, less heat flow to handle at boundaries, etc., which, combined with the target wet mass of the rocket defines a distinct 'optimum size' sweet spot where the sum of engines reaches the best thrust-to-weight ratio.

Yet Raptor's s/l thrust was reduced from last year's ~300 tons-force to ~170 tons-force, which change appears to be too large of a reduction to be solely dictated by optimum single engine TWR considerations.
What were the main factors that led to this change?

Elon Musk initial reply:

We chickened out

Elon Musk follow-up reply:

The engine thrust dropped roughly in proportion to the vehicle mass reduction from the first IAC talk. In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you've lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.

Redditor comment:

You can't land on moon using 3MN engine

Elon Musk reply:

Yes, you can. - Bob, the Builder
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/16/2017 08:33 pm
From the Reddit AMA:
Redditor comment:

You can't land on moon using 3MN engine

Elon Musk reply:

Yes, you can. - Bob, the Builder

It's funny because current Raptor is < 2MN. Elon could have just been being silly (he was self-admittedly drunk :D) but also lends credence to the idea that SpaceX will likely eventually move to a > 3MN Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: CuddlyRocket on 10/17/2017 09:01 am
In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you've lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.

I'm interested in the phrase "a third medium area ratio Raptor engine". Everyone's taken this to mean a third SL engine; but, if so, why didn't he just say so? (Are SL engines referred to as medium area ratio engines?) But, I wonder if he actually meant a third type of engine, with an area ratio between that of the SL and Vac engines. One that will work at sea-level without the usual adverse consequences of attempting to run a Vac engine at those atmospheric pressure, but one that is more efficient than the SL engines when used in regimes of lower atmospheric pressure (though not as efficient as a Vac engine).

Are they proposing to land on the Moon and Mars using the SL engines? If so, such an engine would be more efficient. How high will the BSF be when it separates from the booster and could such an engine be useful at such an altitude? I'm not a rocket engineer (does it show! :) ), but could there be benefits from having such an intermediate engine?

The obvious argument against is having a third type of engine, with the design and manufacturing complexity etc. I suppose this depends on the level of commonality with the SL and/or Vac engines, and whether any benefits are worth the cost.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DreamyPickle on 10/17/2017 10:24 am
I'm interested in the phrase "a third medium area ratio Raptor engine". Everyone's taken this to mean a third SL engine; but, if so, why didn't he just say so? (Are SL engines referred to as medium area ratio engines?) But, I wonder if he actually meant a third type of engine, with an area ratio between that of the SL and Vac engines. One that will work at sea-level without the usual adverse consequences of attempting to run a Vac engine at those atmospheric pressure, but one that is more efficient than the SL engines when used in regimes of lower atmospheric pressure (though not as efficient as a Vac engine).

There is some speculation in this thread (https://forum.nasaspaceflight.com/index.php?topic=42003.0) that the 2016 ITS already had 3 types of engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/17/2017 12:09 pm
In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you've lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.

I'm interested in the phrase "a third medium area ratio Raptor engine". Everyone's taken this to mean a third SL engine; but, if so, why didn't he just say so? (Are SL engines referred to as medium area ratio engines?) But, I wonder if he actually meant a third type of engine, with an area ratio between that of the SL and Vac engines. One that will work at sea-level without the usual adverse consequences of attempting to run a Vac engine at those atmospheric pressure, but one that is more efficient than the SL engines when used in regimes of lower atmospheric pressure (though not as efficient as a Vac engine).

Are they proposing to land on the Moon and Mars using the SL engines? If so, such an engine would be more efficient. How high will the BSF be when it separates from the booster and could such an engine be useful at such an altitude? I'm not a rocket engineer (does it show! :) ), but could there be benefits from having such an intermediate engine?

The obvious argument against is having a third type of engine, with the design and manufacturing complexity etc. I suppose this depends on the level of commonality with the SL and/or Vac engines, and whether any benefits are worth the cost.

I thought the phrasing was odd as well. If they put a 70:1 ER engine in the middle and two landing engines on either side, this would increase their allowable landing weight and increase BFS thrust to weight during ascent. This would be an improvement if BFS's T/W was a little low to begin with.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Bynaus on 10/17/2017 12:43 pm
I was confused at first, but I think what this means is:

"large area ratio engine" = vacuum-engine
"medium area ratio engine" = sea-level-engine
"small area ratio engine" = Raptor-based RCS thruster (?)

I think overall it just means that he wants have these engine descriptions capture an unchanging property of the engine (the area ratio) as opposed to the conditions they are used (which isn't as clear cut).

I don't know if the "small" version is really an RCS thruster, but that seems to make most sense.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: nacnud on 10/17/2017 12:51 pm
Would there be any difference in the best area ratio for an air start raptor that then needs to land versus a ground start raptor that then needs to land?

Small, ground start and landing
Medium, air start (well space really) and landing
Large, vacuum optimized
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomic on 10/17/2017 02:55 pm
I'm interested in the phrase "a third medium area ratio Raptor engine".

Booster might have a lower ER Raptor then the landing Raptors on BFS? ISP doesn't matter as much for the booster, can squeeze more engines on to it with a smaller nozzle and save a bit of weight.

So BFR with low ER raptor, BFS has mid size and vac?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/17/2017 07:45 pm
I'm interested in the phrase "a third medium area ratio Raptor engine".

Booster might have a lower ER Raptor then the landing Raptors on BFS? ISP doesn't matter as much for the booster, can squeeze more engines on to it with a smaller nozzle and save a bit of weight.

So BFR with low ER raptor, BFS has mid size and vac?

That is my interpretation. The booster needs a smaller ER nozzle to pack more engines into a small space.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: intrepidpursuit on 10/19/2017 08:47 pm
I'm interested in the phrase "a third medium area ratio Raptor engine".

Booster might have a lower ER Raptor then the landing Raptors on BFS? ISP doesn't matter as much for the booster, can squeeze more engines on to it with a smaller nozzle and save a bit of weight.

So BFR with low ER raptor, BFS has mid size and vac?

As far as I can tell, the situation where you need the most efficiency out of the center engines on the BFS is on takeoff. That will never happen at earth sea level. It may be the biggest ratio they can get away with at SL so it is better optimized for mars and moon operations.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rockets4life97 on 10/20/2017 03:15 am
Blue Origin recently reported a successful test of the B-4. Does this put the B-4 ahead or behind raptor in terms of development? or is it hard to compare given SpaceX's decision to test a subscale engine first?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/20/2017 04:43 am
Blue Origin recently reported a successful test of the B-4. Does this put the B-4 ahead or behind raptor in terms of development? or is it hard to compare given SpaceX's decision to test a subscale engine first?

Both Raptor and BE-4 engines are in the early stages of test. To advance a test program for large engines like these is a long term program, not the kind of thing you can handicap like a horse race.

They also have different goals, chamber pressures, and scale of size. The interesting commonality is the combustion of largely similar propellants at similar mass flows.

SX has about a year lead on operating a more complex and scaleable engine against BO getting one to operate (an enormous achievement nonetheless). Both of these are "firsts" in different ways - SX in hydrocarbon FFSC globally, and BO in the first non-Russian, non-Ukrainian ORSC engine.

BE-4 is intended for use in multiple vehicles, so from this POV the tests progress building confidence in a single design instance at high fidelity. Raptor as currently implemented is a compact, extremely high chamber pressure engine apparently not intended for use, but to allow many derivatives that will be used,  to be rapidly developed for an exotic two stage vehicle intended to land on other moons/planets. So the first follows an immediate path of critical development/review, while the other's path requires even more reliability/application over a more elaborate development path.

So they are necessarily hard to compare. Even if ULA selects BE-4 by end of year, the nature of the Vulcan engine role isn't the same as with NG nor that of the Raptors that will actually be flown.

And even with the first flown BE-4's ... that will occur far enough into the future, that understanding where Raptor in an actual vehicle will be in comparison, isn't possible.

So the best one can do is compare test programs now.

In short - Raptor has a lead on time/reliability. They're both at about the same thrust currently, though this will change as duration and power level increases with next steps. Both have gotten by a major achievement in start-up/shutdown.

Next big steps for each - Raptor needs to move on to a flight scale/quality engine (it was wise to do a 1MN one first, makes this next step easier),  BE-4 needs to expand its operating range to its design limits while stably functioning and thus proving that its model of operation matches its actual function. Both are tall orders.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/20/2017 05:20 am

They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.

Are you seriously claiming that BE-4 is only running at about 40% thrust currently?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mnelson on 10/20/2017 05:49 am

They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.

Are you seriously claiming that BE-4 is only running at about 40% thrust currently?

Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 10/20/2017 06:06 am

They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.

Are you seriously claiming that BE-4 is only running at about 40% thrust currently?

Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/

Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: butters on 10/20/2017 06:17 am

They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.

Are you seriously claiming that BE-4 is only running at about 40% thrust currently?

Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/

Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.

Was there evidence that the Raptor engine tested is physically subscale relative to the 9m BFR proposal? It's certainly being tested below design thrust at 200 bar chamber pressure vs. 250-300 bar design target. But is it any smaller in physical dimensions?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/20/2017 11:36 am
Raptor Demo engine exit diameter was a little under a meter (~.94) as measured from photograph and was supposedly putting out 1MN at 20 MPa. If we take that as a given, then the Raptor Engine will be about a 15% larger throat diameter.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/20/2017 02:20 pm

They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.

Are you seriously claiming that BE-4 is only running at about 40% thrust currently?

Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/

Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.

Yes, there is a big difference. Running subscale at 3000 psi is way harder than full scale at 1000 psi.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 10/20/2017 02:27 pm
Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.

Yes, there is a big difference. Running subscale at 3000 psi is way harder than full scale at 1000 psi.

This. They have not even reached the chamber pressures that M1D runs at, much less Raptor or RD-180. Also livingjw's calculations show that it's not a very big scale-up that SpaceX needs, only 15%.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 10/20/2017 03:38 pm
Does any one have more on this tweet earlier today:

Quote
SpaceX gets another $40.8 million in Pentagon funding for Raptor engine

https://twitter.com/R_Wall/status/921257396797870080 (https://twitter.com/R_Wall/status/921257396797870080)

The tweet doesn't appear to be a reply to anything else, sounds like new money?

Edit: forgot to say that Robert Wall is aerospace reporter for WSJ
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 10/20/2017 03:45 pm
Does any one have more on this tweet earlier today:

Quote
SpaceX gets another $40.8 million in Pentagon funding for Raptor engine

https://twitter.com/R_Wall/status/921257396797870080 (https://twitter.com/R_Wall/status/921257396797870080)

The tweet doesn't appear to be a reply to anything else, sounds like new money?

Edit: forgot to say that Robert Wall is aerospace reporter for WSJ

https://www.defense.gov/News/Contracts/Contract-View/Article/1348379/ (https://www.defense.gov/News/Contracts/Contract-View/Article/1348379/)
Space Exploration Technologies Corp., Hawthorne, California, has been awarded a $40,766,512 modification (P00007) for the development of the Raptor rocket propulsion system prototype for the Evolved Expendable Launch Vehicle program.  Work will be performed at NASA Stennis Space Center, Mississippi; Hawthorne, California; McGregor, Texas; and Los Angeles Air Force Base, California; and is expected to be complete by April 30, 2018.  Fiscal 2017 research, development, test and evaluation funds in the amount of $40,766,512 are being obligated at the time of award.  The Launch Systems Enterprise Directorate, Space and Missile Systems Center, Los Angeles AFB, California, is the contracting activity (FA8811-16-9-0001).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/20/2017 03:48 pm
Does any one have more on this tweet earlier today:

Quote
SpaceX gets another $40.8 million in Pentagon funding for Raptor engine

https://twitter.com/R_Wall/status/921257396797870080 (https://twitter.com/R_Wall/status/921257396797870080)

The tweet doesn't appear to be a reply to anything else, sounds like new money?

Edit: forgot to say that Robert Wall is aerospace reporter for WSJ

https://spaceflightnow.com/2016/03/07/ulas-candidates-to-replace-rd-180-engine-win-air-force-funding/

Quote
ULA has agreed to initially add $40.8 million under the terms of the government award.
I wonder if some wires have been crossed.

There has of course been the recent air force proposal for a call for bids to do various rocket development stuff, but that was not due for a while yet.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 10/20/2017 04:34 pm
I really hope this 40 million doesn't reignite Raptor upper stage fever.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/20/2017 06:20 pm
I really hope this 40 million doesn't reignite Raptor upper stage fever.

But they do plan to fly a Raptor based upper stage in a few years - BFS.  8)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Jakusb on 10/20/2017 09:33 pm
Related to this?
 https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/ (https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/20/2017 09:43 pm
Related to this?
 https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/ (https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/)

Which was just announced, and has some long way to go until the contracts are awarded.
Any decision to award anything at this time would be extraordinarily vulnerable to challenge, if not flat-out illegal. (unsure on the latter).

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 10/20/2017 09:44 pm
Related to this?
 https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/ (https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/)

No, that is still in RFP stage.  The thread for that is here: USAF RFP for new EELV Launch Service Agreements (2017-10-05) (http://forum.nasaspaceflight.com/index.php?topic=43924.0)

This money is a continuation of their previous contract from January 2016:
Quote
Space Exploration Technologies, Corp. (SpaceX), Hawthorne, California, has been awarded a $33,660,254 other transaction agreement for the development of the Raptor rocket propulsion system prototype for the Evolved Expendable Launch Vehicle (EELV) program. This agreement implements Section 1604 of the Fiscal Year 2015 National Defense Authorization Act, which requires the development of a next-generation rocket propulsion system that will transition away from the use of the Russian-supplied RD-180 engine to a domestic alternative for National Security Space launches. An other transaction agreement was used in lieu of a standard procurement contract in order to leverage on-going investment by industry in rocket propulsion systems. This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles. The locations of performance are NASA Stennis Space Center, Mississippi; Hawthorne, California; and Los Angeles Air Force Base, California. The work is expected to be completed no later than Dec. 31, 2018. Air Force fiscal 2015 research, development, test and evaluation funds in the amount of $33,660,254 are being obligated at the time of award.  SpaceX is contributing $67,320,506 at the time of award. The total potential government investment, including all options, is $61,392,710. The total potential investment by SpaceX, including all options, is $122,785,419. This award is the result of a competitive acquisition with multiple offers received. The Launch Systems Enterprise Directorate, Space and Missile Systems Center, Los Angeles Air Force Base, California is the contracting activity (FA8811-16-9-0001).
.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/20/2017 11:03 pm
...
Quote
Air Force fiscal 2015 research, development, test and evaluation funds in the amount of $33,660,254 are being obligated at the time of award.  SpaceX is contributing $67,320,506 at the time of award The total potential government investment, including all options, is $61,392,710. .
.

$61M-$33M does not equal $40M. I have not looked if more money has been allocated.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/21/2017 12:01 am
Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.

Yes, there is a big difference. Running subscale at 3000 psi is way harder than full scale at 1000 psi.

This. They have not even reached the chamber pressures that M1D runs at, much less Raptor or RD-180. Also livingjw's calculations show that it's not a very big scale-up that SpaceX needs, only 15%.

Precisely. Not to mention the fact that FFSC is definitively more difficult than ORSC, which has a long and successful heritage. BE-4's thrust to weight ratio is going to be less than impressive, even if it is an impressive technical accomplishment as a whole.

Raptor is all about efficiency  and reliability. As Musk put it in the AMA last weekend, "thrust scaling is the easy part.... very simple to scale the dev Raptor to 170 tons", the focus now has moved on to optimizing for reliability and TWR.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 10/21/2017 01:33 am
$61M-$33M does not equal $40M. I have not looked if more money has been allocated.

It looks like they did increase the amount at some point, but we may need to wait for this latest one to flow through the systems into the publicly available databases (govtribe or fpds.gov) before we can even have a chance of figuring out where it stands now.

Here is the original contract and the latest mod in June 2017.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: watermod on 10/21/2017 03:20 am
looking at the dates... did they buy the test engine?
Its only a few months start to end.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 10/21/2017 07:53 am
When the tweet about the $40.8B was posted on the FB fan group self described SpaceXer Phillip Aubin replied,

https://www.facebook.com/groups/spacexgroup/permalink/10155943228621318/

"All I can say is: The people who complain the most are the ones NOT putting payloads into orbit on a monthly basis, if not even shorter."
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 10/21/2017 08:45 pm
Here's a write-up of the additional Raptor funding and current status of Raptor development:

Quote
Air Force adds more than $40 million to SpaceX engine contract
by Jeff Foust — October 21, 2017

http://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/ (http://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rik ISS-fan on 10/21/2017 08:54 pm
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.

What I find typical is that: 33.6mln + 67.3 mln = 100.9mln development cost for 1MN raptor.
?what was the prometheus engine development going to cost?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/21/2017 09:06 pm
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.

What I find typical is that: 33.6mln + 67.3 mln = 100.9mln development cost for 1MN raptor.
?what was the prometheus engine going to cost?

🤔🤔🤔🤔🤔🤔🤔🤔🤔🤔🤔🤔🤔🤔🤔

So many issues with this comment, don't even know where to begin. Contract you cited actually shows that the completion date has been accelerated to April 2018, six months from now. Raptor did tests of the preburner in 2015, AR-1 literally only completed its first preburner test this year. Raptor has 1200+ seconds of firing, AR-1 has zero seconds.


Taken directly from Musk's mouth and educated estimates in this very thread, Raptor scaling is of little concern and the physical scaling needed is less than 20%.

Also, AR-1 is being developed by Aerojet-Rocketdyne, not ULA. Hence AR.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Navier–Stokes on 10/21/2017 09:08 pm
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.

What I find typical is that: 33.6mln + 67.3 mln = 100.9mln development cost for 1MN raptor.
?what was the prometheus engine development going to cost?
I fail to see any evidence of which to draw such an extreme conclusion from.

Stennis was included in original contract as well as the modification. In fact, under the modification, McGregor has actually been added to the locations of performance.
Quote from: CR-203-17
The locations of performance are NASA Stennis Space Center, Mississippi; Hawthorne, California; and Los Angeles Air Force Base, California.
Quote from: CR-008-16
Work will be performed at NASA Stennis Space Center, Mississippi; Hawthorne, California; McGregor, Texas; and Los Angeles Air Force Base, California[.]
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/21/2017 09:56 pm
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.

What?... Proof it's behind?... No!...

The test unit fired 46+ times in the last year retired a LOT of unknowns for SpaceX and Tom Mueller on this engine architecture...
The design and layout of pumps and turbines is sound...
The method to light it off, control thrust, and shut down the engine is sound...
The stated run times over the last year indicates the cooling of key parts is not in question...
All at the 200 bar chamber pressures stated... already...

SO... it's thought they only need to physically scale it about 12% (chamber and throat)
And then work up to 250bar with 300bar as an end goal (and design for such)

Start over... Stennis... Full redesign... 2 years... HA!!!

This thing will be flying on a BFS "Grasshopper" test bed... BEFORE 2 years from now...  ;)
And that IS an upper stage of a rocket...  :)

On edit...
The USAF (and we the taxpayers providing the funds) are getting one heck of a return on investment in Raptor tech...
This is Tom Mueller (and his team) doing what they do best...
Designing the best value, lowest cost thing able to turn Methane and Oxygen into delta/v...
Low cost comes from using it over and over and over... reusability...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 10/22/2017 03:10 am
Something I was trying to read the tea leaves about and that is total spending that will have been done on Raptor thru April 2018 looks to be at >$300M. The $300M is ~$105M from the AF and $195M from SpaceX just during this contract duration from FY2016 thru April 2018. So my speculation is that Raptor development up to start of production will likely end up being ~$500M. This would include quite a bit of spending that had occurred prior to the AF contract. Probably easily $150M total spread over multiple years.

In all I would expect that a 380Klbf production prototype test article is in testing by April 2018. Once such a test article has successfully completed testing then production is not far off (as in months away not years). If production starts in 3Q2018 I would expect about 1 year later the first flight article production engines being delivered for qualification and acceptance testing. With assembly into a flight thrust structure occurring after 3Q2019. This then suggests a flight vehicle could be ready for its testing phase to begin in 2Q2020.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Mike Jones on 10/22/2017 09:42 am
How did you get to 105 M$ investment from US Air Force ? They only communicated on 33+40 M$ contracts to SpaceX. And only 66 M$ investment from SpaceX has been confirmed so far in the frame of this OTA with USAF. Do you have complementary information ? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 10/22/2017 03:30 pm
How did you get to 105 M$ investment from US Air Force ? They only communicated on 33+40 M$ contracts to SpaceX. And only 66 M$ investment from SpaceX has been confirmed so far in the frame of this OTA with USAF. Do you have complementary information ?
The original AF commitment of the contract was $95M over a execution period covering 3 years. This option execution just made some modifications by adding McGregor and increasing the total by ~$8M by upping the amount on this option from $33M to $40.7M. Making the new total of AF funding increasing from the $95M to ~$105M I think the new value may really be closer to $103M.

So it is not really that much of an increase over what the AF already was planning to spend.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/22/2017 06:54 pm
Wow, I poked a hornet's nest of dickishness on Twitter. Really disappointed at how consistently arrogant Blue's own purported engineers are, especially publicly so on Twitter.


In response to Jeff Foust's simple Raptor funding article (http://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/), a BO propulsion engineer commented (utterly unsolicited, I might add)
"Or they can pay $0 for a more reliable engine that produces more thrust and already tested at full scale 🙃". (https://twitter.com/KeganB_18/status/921893011092582401) An extraordinary statement that is really hard to rationally parse for an engine that has fired for no more than 3 seconds at half thrust and experienced at least one serious failure during testing.

Another BO employee chimed in, "How exactly is a subscale version of an engine "ages closer" to flight readiness than a full scale version of an engine?" (https://twitter.com/Chenzo_13/status/922145178885881857)

Me: "That 1MN Raptor has been tested for 100s nonstop and > 1200s total should be self-explanatory" (https://twitter.com/13ericralph31/status/922152819532034048)
Me: "I would also be a fool to totally discount a company's CTO saying that it is "simple to scale the dev Raptor to 170 tons""

BO guy: "You'd also be a fool to just blindly believe everything that person says when they've proven to not do things when they say they will." (https://twitter.com/Chenzo_13/status/922157750792024064)
BO guy: "But since you admittedly have no tech expertise, sure just believe what others say. I have technical expertise and know it's not that simple" (https://twitter.com/Chenzo_13/status/922158002223792128)


It doesn't exactly take a genius to understand that Musk has a habit of understating the difficulty of doing relatively hard things, but both of these BO engineers were dead-set on a single 3s 50% thrust firing of a full-scale engine indicating that BE-4 was somehow closer to flight-readiness than subscale Raptor, with (probably multiple) successful ~100s hot-fires and more than 1200s total. It boggles the mind.

I really want to cheer on Blue Origin but s*** like this makes it rather difficult to support a company with such a seemingly arrogant culture. These are anecdotes, of course, I can only hope that they are representative of a tiny minority. But it's starting to feel like Jeff "Welcome to the club" Bezos managed to only hire clones of himself...



Edit: Someone requested links, and links you shall have! Some additional entertaining public quotes from BO employees below, too.

BO guy: "These are same people who thought it'd be simple to just strap on side boosters to falcon 9 and poof now we have falcon heavy. Not the case."

Me: "Ah yes, the ole FH strawman 😉 If we're that off topic, let's just wait until Blue has reached orbit NET '20 and take stock of the industry."

BO guy: "It's not a straw man. You used appeal to authority fallacy by saying "oh well CTO said this so it must be true." I used FH to disprove that."  (https://twitter.com/Chenzo_13/status/922162647071563776)


Me (coulda had a little more tact but c'est la vie): "1200 seconds > 3 seconds." (https://twitter.com/KeganB_18/status/921893011092582401)

BO guy: Lmk when they test full scale for 3 seconds lol

Me: Don't get me wrong, I'm thrilled for BO and wish you guys the best of luck, but Raptor is ages closer to flight readiness.

BO guy: That's cool and all but my argument is BE-4 costs 0 taxpayer dollars. So I'm all for that.

Me: [The] USAF costs taxpayers $160b a year whether or not it includes rounding-error funding for RPS. Rockets are cool regardless of funding sources.


Me: "Blue's consistent arrogance is truly disappointing. The willingness to discount actual launch providers doesn't befit reasonable people."

Me: "It doesn't take a technical expert to understand that orbital rocketry provides more experience than sub-Mach 4 flight regimes

BO guy: "It's hilarious that you think Blue is the arrogant company" (https://twitter.com/Chenzo_13/status/922162761248915456)

"So I'll go on actually changing the future of spaceflight by working in the industry and you can go on "covering" stuff. Have a nice day." (https://twitter.com/Chenzo_13/status/922163395524042752)

Right in the journalism :'(


F i n .
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/22/2017 07:36 pm
Wow, I poked a hornet's nest of dickishness on Twitter. Really disappointed at how consistently arrogant Blue's own purported engineers are, especially publicly so on Twitter.


In response to Jeff Foust's simple Raptor funding article (http://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/), a BO propulsion engineer commented, "Or they can pay $0 for a more reliable engine that produces more thrust and already tested at full scale 🙃". An extraordinary statement that is really hard to rationally parse for an engine that has fired for no more than 3 seconds at half thrust and experienced at least one serious failure during testing.

Another BO employee chimed in, "How exactly is a subscale version of an engine "ages closer" to flight readiness than a full scale version of an engine?"

Me: "That 1MN Raptor has been tested for 100s nonstop and > 1200s total should be self-explanatory"
Me: "I would also be a fool to totally discount a company's CTO saying that it is "simple to scale the dev Raptor to 170 tons""

BO guy: "You'd also be a fool to just blindly believe everything that person says when they've proven to not do things when they say they will."
BO guy: "But since you admittedly have no tech expertise, sure just believe what others say. I have technical expertise and know it's not that simple"


It doesn't exactly take a genius to understand that Musk has a habit of understating the difficulty of doing relatively hard things, but both of these BO engineers were dead-set on a single 3s 50% thrust firing of a full-scale engine indicating that BE-4 was somehow closer to flight-readiness than subscale Raptor, with (probably multiple) successful ~100s hot-fires and more than 1200s total. It boggles the mind.

I really want to cheer on Blue Origin but s*** like this makes it rather difficult to support a company with such a seemingly arrogant culture. These are anecdotes, of course, I can only hope that they are representative of a tiny minority. It's almost as if Jeff "Welcome to the club" Bezos managed to only hire clones of himself...

They have a bigger engine with more thrust to match their ego, it’s understandable.
What they seem to forget is that their 7 big engines have less thrust than 31 small engines.


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 10/22/2017 07:36 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/22/2017 07:44 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/22/2017 09:11 pm
Who do you think will post a 30+ second video to YouTube as proof of actual goal reached first?
(full design production thrust stated, for long enough to show it will not melt parts (engine rich exhaust) and/or RUD)

A) Blue Origin with a ULA Vulcan SL spec BE-4 firing at 2450 kN thrust for 30+ seconds...

B) SpaceX with a BFS/BFR SL spec Raptor firing at 1700 kN thrust for 30+ seconds...

Based on what I have seen to date and the leadership within and the culture of the two companies employees...
B is my guess... 

BUT... enough of that BO bashing... when will the first full scale chamber and throat Raptor be fired up?...
I'm thinking springtime... at the latest...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ChaoticFlounder on 10/22/2017 09:20 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

I don't know if you really understand what he's saying:

What he's saying is that the Blue Origin engine is full size ... Combustion Chamber, injectors, preburner injectors, pumps, everything is full size, that requires no drawing changes (in theory) when they go to flight, full thrust SHOULD be as easy as opening up the fuel valve to the preburner to let the turbopump spin faster... i don't know if this is what they're using to alter pump speed (inevitably engine power level) or not... whether this happens or not is yet to be seen

SpaceX's Raptor is: what, help me out here ... 80% geometrically the size of the flight engine size, this means new part numbers for the combustion chamber, injector, preburner(s) injectors, pumps, if i'm understanding what Elon has said correctly, everything has to be geometrically scaled up to reach flight engine size, that is not a small task, also, dynamic similitude in fluid mechanics doesn't mean you multiply or divide everything by 0.80 ...

^ my $0.02

C


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/22/2017 09:29 pm
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 10/22/2017 09:38 pm
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
https://twitter.com/SciGuySpace/status/921106486272675840
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: whatever11235 on 10/22/2017 09:45 pm
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.

Space or orbit? ;D
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/22/2017 09:46 pm
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
https://twitter.com/SciGuySpace/status/921106486272675840
That's okay. He doesn't actually follow SpaceX very closely.

Also, I worded what I said very particularly.

But there's a pretty decent probability that one of those 4 will beat it. But I doubt more than 1.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Torbjorn Larsson, OM on 10/22/2017 10:19 pm
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
https://twitter.com/SciGuySpace/status/921106486272675840

Berger is trying to compare apples and pears by pointing at citrus (stress issues in multi-body launchers).

Between SX and BO, the former has launched multi-engine, multi-stage rockets to LEO. And while SX has developed almost all the key elements for BFR/BFS into LEO [excluding the refuel maneuver for other uses] in some form or other, BO has done little. Maybe BO can compete with SLS ME2, maybe they will all be close when the combusted fuel hit the launch pad, maybe they will spread over many years, maybe some will fail. But Berger is out on a fishing expediting for bad analysis.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/22/2017 10:20 pm
I think Blue will be able to compete eventually. I just think SpaceX is ahead with BFR.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 10/22/2017 11:07 pm
I think Blue will be able to compete eventually. I just think SpaceX is ahead with BFR.
Wandering somewhat OT. But anyway, if BFR was just a much much larger version of F9 then I would agree that BFR would be ahead and would get top flight first. But BFR is much more complex with more testing gates to successfully pass than what NG has to.

As far as engines go they are both from the standpoint of going into "production" in the near future about even. With both Raptor an BE-4 both likely starting production of flight units around mid 2018.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/22/2017 11:18 pm
I'd say SpaceX is a year ahead with Raptor and is generally faster at executing anyway.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 10/22/2017 11:41 pm
The BFR would be similar the F9 booster.  It is the BFS that is going to take time.  They may have the booster ready 2 years before the BFS.  It could launch 4 F9 upper stages in a cluster for second stage or stages going to different orbits.  Probably wouldn't be worth it, but, they could build a big expendable upper stage to get some things launched before BFS is ready. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/22/2017 11:46 pm
They're going to develop and test BFS before the booster.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/22/2017 11:57 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

I don't know if you really understand what he's saying:

What he's saying is that the Blue Origin engine is full size ... Combustion Chamber, injectors, preburner injectors, pumps, everything is full size, that requires no drawing changes (in theory) when they go to flight, full thrust SHOULD be as easy as opening up the fuel valve to the preburner to let the turbopump spin faster... i don't know if this is what they're using to alter pump speed (inevitably engine power level) or not... whether this happens or not is yet to be seen

SpaceX's Raptor is: what, help me out here ... 80% geometrically the size of the flight engine size, this means new part numbers for the combustion chamber, injector, preburner(s) injectors, pumps, if i'm understanding what Elon has said correctly, everything has to be geometrically scaled up to reach flight engine size, that is not a small task, also, dynamic similitude in fluid mechanics doesn't mean you multiply or divide everything by 0.80 ...

^ my $0.02

C

Thanks, your thoughts are appreciated.

I completely agree, and that's largely how I understood the situation. Scaling up both physical dimensions and chamber pressure by 15-25% is not said and done by any means, and the complexity of RPS and plumbing necessitate that it will be more difficult than "enlarging the CAD model by 15%", as one of the BO employees condescendingly suggested.

Howeverrrrr, I also have little doubt that SpaceX has been iterating and exploring full scale Raptor hardware during the 12+ months they've been testing its scaled prerequisites, thus learning many lessons about running an integrated 1MN methalox FFSC engine. Dozens of times and at considerable duration, as well. (Also some L2 info that strengthens this feeling, but can't say more)

Given how little Raptor will have to grow to reach its current operational performance specs, as well as SpaceX's vast (compared to BO) experience producing rocket propulsion systems, it seems implausible to say that BE-4 is closer to flight readiness because they successfully fired a full sized engine for 3 seconds, after suffering at least one major hardware failure.

Another main difference I perceive simply lies in SpaceX's decision to begin with subscale testing. They've developed some level of expertise with Raptor, even if it may not all remain applicable after scaling thrust by an additional 70%. BO has a sum total of 3 seconds of experience testing an integrated engine, even if it's full scale. Their test program could proceed utterly flawlessly, but that seems improbable. I'm sure SpaceX has had to deal with many issues with scale Raptors over 40+ tests, and I would bet money that a lot of the lessons learned with scale Raptor will transfer to full scale testing.

Again, I am self-admittedly not a technical expert. I don't currently have time to do so, due to school, but my hope is to build a decent foundation of the basics of rocketry and RPS when I have the free time. What minimal reading I've done has informed the above opinions, and I welcome any and all criticisms and corrections, as well as complete refutation. Just trying to better understand things and tweak my intuition along the way.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/23/2017 12:19 am
Between SX and BO, the former has launched multi-engine, multi-stage rockets to LEO. And while SX has developed almost all the key elements for BFR/BFS into LEO [excluding the refuel maneuver for other uses] in some form or other, BO has done little. Maybe BO can compete with SLS ME2, maybe they will all be close when the combusted fuel hit the launch pad, maybe they will spread over many years, maybe some will fail. But Berger is out on a fishing expediting for bad analysis.
I think Blue will be able to compete eventually. I just think SpaceX is ahead with BFR.

I have to agree. Blue Origin has been around for literally two decades, have manufactured a handful of suborbital rockets, flown those a handful of times, and have failed to travel beyond Mach 4. SpaceX has had plenty of missteps with Falcon 1 and Falcon 9, but orbital rocketry is f***** hard, and at this point they are already normalizing routine recovery and reuse.

ULA and Arianespace may wave their launch records around with the humility of pop musicians, but the reality is that they've been flying orbital rockets that suffered plenty of failures for the better part of half a century, and Atlas 5, Delta IV, and Ariane 5 are the results. SpaceX has been in the business for 8-9 years total and have only experienced two complete failures. BO has a longggg road ahead of themselves, and excessive arrogance and a lack of humility will only serve to make it even rockier.

As Robert Heinlein definitely 100% said, "The Karman line didn't exist in 1950 but once you're there, you're maybe 20% of the way to orbit."
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: QuantumG on 10/23/2017 12:48 am
I tell ya, if ULA start flying a vehicle with a Raptor engine I'm going to have to go buy some new ice skates.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: groundbound on 10/23/2017 02:32 am
I tell ya, if ULA start flying a vehicle with a Raptor engine I'm going to have to go buy some new ice skates.

Slightly OT but I would guess it is (slightly) more probably that SpaceX sells Merlins to someone else if they find themselves in dire need of another revenue stream. This assumes they could find a customer that would put it to a use that was not in direct competition.

The wildcard in all of this is the US government. They can potentially offer enough $$ to convince SpaceX to do things they would not otherwise.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kansan52 on 10/23/2017 02:49 am
Does the new money the Air Force possibly invest in Raptor mean SX can progress faster?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 10/23/2017 05:06 am
Does the new money the Air Force possibly invest in Raptor mean SX can progress faster?
Only if they were ressource constrained before. Otherwise I would go with the old adage: Throwing manpower at a late project makes it later.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 10/23/2017 05:08 am
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.

Unmanned SLS will get there first. Then you'll see manned BFS on Luna and SLS will never fly again.

Most likely none of those others will ever fly again. Why sail galleons when container ships suddenly show up?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/23/2017 05:49 am
Does the new money the Air Force possibly invest in Raptor mean SX can progress faster?
Only if they were ressource constrained before. Otherwise I would go with the old adage: Throwing manpower at a late project makes it later.

Yeah, I wouldn't point to the additional $10m of AF money as anything remarkable. SpaceX is contractually required to invest twice as much as the AF, so it's up to around $300m total if the AF contract is taken at face value. I'm sure the money is helpful, but I doubt SpaceX is dependent upon it.

FWIW, I expect we'll see full scale testing begin before the end of 2017.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 10/23/2017 06:07 am

FWIW, I expect we'll see full scale testing begin before the end of 2017.

If they had a full-scale engine almost ready for testing I expect Musk would have shown off some pictures of full-scale hardware at IAC. The fact that he didn't might suggest that they are still a ways off.

FWIW, I wouldn't be surprised if we didn't see a full scale Raptor test fire until late 2018 or even 2019. These things take time.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/23/2017 06:57 am

FWIW, I expect we'll see full scale testing begin before the end of 2017.

If they had a full-scale engine almost ready for testing I expect Musk would have shown off some pictures of full-scale hardware at IAC. The fact that he didn't might suggest that they are still a ways off.

FWIW, I wouldn't be surprised if we didn't see a full scale Raptor test fire until late 2018 or even 2019. These things take time.

Given the fact that it would barely be appreciably larger, I doubt it. Musk's comment during the AMA also suggests that full scale testing is imminent, like months away.

I'd bet money that full scale hardware ready for testing already exists and full scale preburner testing has already begun.

But just pure speculation at this point. That's it from me!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 10/23/2017 07:28 am
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.

Unmanned SLS will get there first. Then you'll see manned BFS on Luna and SLS will never fly again.

Most likely none of those others will ever fly again. Why sail galleons when container ships suddenly show up?
Although Raptor will help revolutionize access to space it will not cause the death of all other (new) launch systems.
For example: Ariane 6 will fly, and after it a next generation Ariane vehicle will as well. The reason is simple: Europe wants it's own independently assured access to space.
When the original Ariane vehicle was being developed there was a lot of pressure from the United States to stop that development. The thinking was that Europe could get all the launch services they ever needed by buying them from the United States.
Europe developed Ariane regardless, despite that seeming to be the more expensive option.
The same applies to China.
So, once BFR/BFS is flying, there will still be vehicles such as Ariane 6 and Long March (insert a number here).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/23/2017 07:36 am
The real question here, is which engine will create the best rocket?
A few heavy 2400 kN BE4’s or many light 1700 KN Raptors?
We miss some data to answer that question, for instance the mass of both engines.

Thats probably question number one, the airforce wanted to know...

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: KelvinZero on 10/23/2017 07:53 am
Although Raptor will help revolutionize access to space it will not cause the death of all other (new) launch systems.
For example: Ariane 6 will fly, and after it a next generation Ariane vehicle will as well. The reason is simple: Europe wants it's own independently assured access to space.
Off topic, but I think very true, other nations will not cede space to SpaceX, they will invest what is needed to catch up. That is the really exciting time. Landing on mars is not as fundamental a milestone to me as the moment we see China test their first grasshopper, and the age of reusable rockets is here no matter how badly SpaceX may stuff up in any future endeavour.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 10/23/2017 07:57 am
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.

Unmanned SLS will get there first. Then you'll see manned BFS on Luna and SLS will never fly again.

Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 10/23/2017 09:14 am
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.

I think that's optimistic. I am confident BFR/RFS are going to succeed, but let's remember, SX hasn't even put a manned Dragon into orbit yet.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 10/23/2017 09:22 am
Although Raptor will help revolutionize access to space it will not cause the death of all other (new) launch systems.
For example: Ariane 6 will fly, and after it a next generation Ariane vehicle will as well. The reason is simple: Europe wants it's own independently assured access to space.
When the original Ariane vehicle was being developed there was a lot of pressure from the United States to stop that development. The thinking was that Europe could get all the launch services they ever needed by buying them from the United States.
Europe developed Ariane regardless, despite that seeming to be the more expensive option.
The same applies to China.
So, once BFR/BFS is flying, there will still be vehicles such as Ariane 6 and Long March (insert a number here).

...I think very true, other nations will not cede space to SpaceX, they will invest what is needed to catch up. That is the really exciting time. Landing on mars is not as fundamental a milestone to me as the moment we see China test their first grasshopper, and the age of reusable rockets is here no matter how badly SpaceX may stuff up in any future endeavour.

Agreed. I should have said other U.S. LVs. It will be hard to compete. Old space is too calcified and most make more money on aircraft. BO could eventually compete, but now Bezos has to think about reusable upper stages in order to do so. I expect to see serious espionage and hacking attempts by the Russians and Chinese to gain SpaceX's technology. Once it becomes obvious that BFR/BFS/Raptor are going to be successful, I expect to see Congress and the Pentagon lay restrictions and protections against that technology making its way to other nations.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AbuSimbel on 10/23/2017 09:34 am
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.

I think that's optimistic. I am confident BFR/RFS are going to succeed, but let's remember, SX hasn't even put a manned Dragon into orbit yet.
You say it like it's indicative of incompetence... no private company has ever put a manned spacecraft into orbit ever, and in the case of BFS/BFR SpaceX can set their own requirements and distribute their milestones in developing a manned ship as they like.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RotoSequence on 10/23/2017 10:08 am
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.

I suspect Elon means sub-orbital flight testing, with the first orbital flights not happening until they've ironed out the kinks.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 10/23/2017 10:27 am
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.

I suspect Elon means sub-orbital flight testing, with the first orbital flights not happening until they've ironed out the kinks.

He said "orbital speed"
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 10/23/2017 11:37 am
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

Falling in to the same Dunning-Kruger hole. You are calling them arrogant and irrational, and yet, by your own admission, THEY are the experts. Which probably means they are not irrational or arrogant, but in fact know more about it than you/other commentators do. I KNOW they know more about it than I do, which is why I let them design rocket engines without me putting my oar in.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 10/23/2017 12:08 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

Falling in to the same Dunning-Kruger hole. You are calling them arrogant and irrational, and yet, by your own admission, THEY are the experts. Which probably means they are not irrational or arrogant, but in fact know more about it than you/other commentators do. I KNOW they know more about it than I do, which is why I let them design rocket engines without me putting my oar in.
Wrong take away.
The BO guys know all about the BO engine (BE-4)
The SpaceX guys know all about the SpaceX engine (Raptor)
The BO guys do NOT know all about the SpaceX engine (Raptor)
The SpaceX guys do NOT know all about the BO engine (BE-4)

Any time that a BO guy says that their engine is better (or further ahead in development) than the SpaceX engine he is stating an ASSUMPTION.
Any time a SpaceX guy says their engine is better (or further ahead in development) than the BO engine he is stating an ASSUMPTION.

Anyone else, armchair engineers included, best not comment on aspect like "better" or "further along in development" at all.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MP99 on 10/23/2017 02:49 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

I don't know if you really understand what he's saying:

What he's saying is that the Blue Origin engine is full size ... Combustion Chamber, injectors, preburner injectors, pumps, everything is full size, that requires no drawing changes (in theory) when they go to flight, full thrust SHOULD be as easy as opening up the fuel valve to the preburner to let the turbopump spin faster... i don't know if this is what they're using to alter pump speed (inevitably engine power level) or not... whether this happens or not is yet to be seen

SpaceX's Raptor is: what, help me out here ... 80% geometrically the size of the flight engine size, this means new part numbers for the combustion chamber, injector, preburner(s) injectors, pumps, if i'm understanding what Elon has said correctly, everything has to be geometrically scaled up to reach flight engine size, that is not a small task, also, dynamic similitude in fluid mechanics doesn't mean you multiply or divide everything by 0.80 ...

^ my $0.02

C

Thanks, your thoughts are appreciated.

I completely agree, and that's largely how I understood the situation. Scaling up both physical dimensions and chamber pressure by 15-25% is not said and done by any means, and the complexity of RPS and plumbing necessitate that it will be more difficult than "enlarging the CAD model by 15%", as one of the BO employees condescendingly suggested.

Howeverrrrr, I also have little doubt that SpaceX has been iterating and exploring full scale Raptor hardware during the 12+ months they've been testing its scaled prerequisites, thus learning many lessons about running an integrated 1MN methalox FFSC engine. Dozens of times and at considerable duration, as well. (Also some L2 info that strengthens this feeling, but can't say more)

Given how little Raptor will have to grow to reach its current operational performance specs, as well as SpaceX's vast (compared to BO) experience producing rocket propulsion systems, it seems implausible to say that BE-4 is closer to flight readiness because they successfully fired a full sized engine for 3 seconds, after suffering at least one major hardware failure.

Another main difference I perceive simply lies in SpaceX's decision to begin with subscale testing. They've developed some level of expertise with Raptor, even if it may not all remain applicable after scaling thrust by an additional 70%. BO has a sum total of 3 seconds of experience testing an integrated engine, even if it's full scale. Their test program could proceed utterly flawlessly, but that seems improbable. I'm sure SpaceX has had to deal with many issues with scale Raptors over 40+ tests, and I would bet money that a lot of the lessons learned with scale Raptor will transfer to full scale testing.

Again, I am self-admittedly not a technical expert. I don't currently have time to do so, due to school, but my hope is to build a decent foundation of the basics of rocketry and RPS when I have the free time. What minimal reading I've done has informed the above opinions, and I welcome any and all criticisms and corrections, as well as complete refutation. Just trying to better understand things and tweak my intuition along the way.
ISTM the test Raptor falls squarely within the definition of a prototype. This from wiki:

"A prototype is an early sample, model, or release of a product built to test a concept or process or to act as a thing to be replicated or learned from. It is a term used in a variety of contexts, including semantics, design, electronics, and software programming. A prototype is generally used to evaluate a new design to enhance precision by system analysts and users.Prototyping serves to provide specifications for a real, working system rather than a theoretical one."

In my experience (software) the luxury of building a pathfinder and then following it up with a production system using lessons learned should result in a better end product - better performing, more reliable, more usable. It's also a great way to do an Agile development.

Cheers, Martin

Sent from my Nexus 6 using Tapatalk

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 10/23/2017 03:07 pm
This thread is starting to wander a bit.  It's not the SpaceX vs. Blue (or anyone else) thread.  There's another thread for that somewhere.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 10/23/2017 03:45 pm
The Raptor is being developed in similar manor that the M1D was done.

First there was a prototype that was used to test assumptions and update the design models. Then the updated design models were used to create a production design with a resulting fairly accurate physical set of specifications in thrust ISP and weight. What happened was the prototype was a design with a probability of a  higher level of successful operation. But then the data gathered is used to "tighten" the engine design models such that when entered a set of constraints like thrust, bell size, ISP, TC pressure, and a few others the design software developed will generate parts drawings that has a high probability to create an engine with almost those exact specifications that will work!

Thus at the end of the prototype testing a specification for a production engine which was delivered 4 months later actually met the design specifications. It also had a near trouble free testing.

So to is the Raptor strategy. The Raptor is at a similar point that the M1D was at in April 2012. And that less than 6 months later production engines were proceeding successfully thru qualification testing and flight units were being delivered and successfully acceptance tested.

Then just a little more than 2 years from the April 2012 point a F9v1.1 flew.

So it is possible that a tanker/cargo version of the SC could start into tests late 2019 or early 2020.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 10/23/2017 05:38 pm
I don't know, the jump from a subscale, lower chamber pressure, development Raptor to a full scale Raptor with ~2x the thrust seems a lot more like the jump from M1C to M1D, rather than the jump from a prototype M1D to a production M1D.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mattstep on 10/24/2017 03:53 am

(trimmed)

I'd bet money that full scale hardware ready for testing already exists and full scale preburner testing has already begun.

But just pure speculation at this point. That's it from me!

Do we expect that the larger pre-burner unit will require a new round of testing? Would that likely occur at Stennis again? Is that something that the public could gain insight into if it is underway or has already completed?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 10/24/2017 04:33 am
I don't know, the jump from a subscale, lower chamber pressure, development Raptor to a full scale Raptor with ~2x the thrust seems a lot more like the jump from M1C to M1D, rather than the jump from a prototype M1D to a production M1D.

I agree for the most part.   However, they have a lot more company experience, likely have a larger team with a bigger budget.   It is a fuel and combustion cycle combination no one has ever used before.  So that’s big.  If they were doing a new RP1 engine that was 2-3 Times larger it would be fun to see how fast they could go. 

The first Raptor would have taught them a lot about how to start, run and shut down this engine.  So they’ll have a great start with the production engine.  I’m sure they’ll continue to test and learn on the prototype until there is nothing left to learn.   

As someone else pointed out, I wouldn’t be surprised if it isn’t already in some level of construction or even testing.   

I would not be surprised if there was a qualified engine in 2 years time.   But they’ll need more tanks at the test stand first 😜
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/24/2017 05:35 am

(trimmed)

I'd bet money that full scale hardware ready for testing already exists and full scale preburner testing has already begun.

But just pure speculation at this point. That's it from me!

Do we expect that the larger pre-burner unit will require a new round of testing? Would that likely occur at Stennis again? Is that something that the public could gain insight into if it is underway or has already completed?

I sincerely doubt Stennis will be used again during Raptor development. Correct me if I'm wrong, anyone, but AFAIK Stennis was only used for preburner tests because it came with approximately $1m in incentives and allowed for hot gas testing capabilities unique to it.

Now that SpaceX has a mature Raptor test facility at McGregor (and magnitudes better funding from the AF), I see no reason they would use Stennis again. The E-2 test stand they used was only rated for 100 klbf thrust, anyways :D

Edit: Thanks, guckyfan.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 10/24/2017 06:50 am
Stennis has unique capabilities. The test stand delivers hot gases and allows for tests of preburners or fuel injection by itself. A normal teststand like in McGregor can only test full engines or at least full powerheads.

My understanding was that the components tested in Stennis used the full capacity and larger preburners or fuel injectors could not be tested there. I may be wrong on this one.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/24/2017 07:15 am
Stennis has unique capabilities. The test stand delivers hot gases and allows for tests of preburners or fuel injection by itself. A normal teststand like in McGregor can only test full engines or at least full powerheads.

My understanding was that the components tested in Stennis used the full capacity and larger preburners or fuel injectors could not be tested there. I may be wrong on this one.

Never a bad excuse to reread this gem (https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/), which basically states that even the size of the subscale test article was decided based on the limits of individual component testing at Stennis. Good catch.

Quote
Since the final thrust level of the Raptor had not been settled, it was decided that the first integrated test engine would be a 1MN sub-scale engine.

It enabled the full testing at Stennis E2 and allowed for the development of robust startup and shutdown sequences, characterize hardware durability and anchor analytical models that would be used for future designs.

Perhaps most interesting, however, are these two paragraphs, derived from comments Mueller made in 2014 and the author's general knowledge of propulsion:
Quote
Once the final engine thrust was defined, the engine could be scaled up with relative ease. The full flow cycle is very helpful in that sense and the 1MN thrust level would already be considered a big engine.

With the production engines – as currently envisioned – it would need to triple its thrust. Not trivial, but still within what could be considered highly representative as a demonstrator.

If 1-3MN is "not trivial" but able to be done with "relative ease", then I can only imagine the same is true for 1-1.7MN.

Forgot how awesome that article was. I would love an updated version for 2017-18 :D
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 10/24/2017 12:03 pm
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.

I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.

and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.

Falling in to the same Dunning-Kruger hole. You are calling them arrogant and irrational, and yet, by your own admission, THEY are the experts. Which probably means they are not irrational or arrogant, but in fact know more about it than you/other commentators do. I KNOW they know more about it than I do, which is why I let them design rocket engines without me putting my oar in.
Wrong take away.
The BO guys know all about the BO engine (BE-4)
The SpaceX guys know all about the SpaceX engine (Raptor)
The BO guys do NOT know all about the SpaceX engine (Raptor)
The SpaceX guys do NOT know all about the BO engine (BE-4)

Any time that a BO guy says that their engine is better (or further ahead in development) than the SpaceX engine he is stating an ASSUMPTION.
Any time a SpaceX guy says their engine is better (or further ahead in development) than the BO engine he is stating an ASSUMPTION.

Anyone else, armchair engineers included, best not comment on aspect like "better" or "further along in development" at all.

And yet they are all rocket engine engineers, which gives them a greater insight in to the issues involved, whether at BO or SpaceX. Of course there are assumptions, but they are very educated assumptions from people who are experts in the field, which should give them greater validity than most of the people who comment on it, including myself.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 10/24/2017 01:56 pm
And yet they are all rocket engine engineers, which gives them a greater insight in to the issues involved, whether at BO or SpaceX. Of course there are assumptions, but they are very educated assumptions from people who are experts in the field, which should give them greater validity than most of the people who comment on it, including myself.
They are also extremely biased (as they should be!) and posting off-the-cuff (as you do) on twitter.  Frankly, it's juvenile and embarrassing, and not a good look.  Fortunately the engine looks great and I look forward to SpaceX and BO trying to one-up each other, hopefully for some time to come.

Now, can we all get back to Raptor, pretty please?  (Not intended to call out anyone specifically, and I include myself in that request).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 10/24/2017 04:46 pm
Now, can we all get back to Raptor, pretty please?  (Not intended to call out anyone specifically, and I include myself in that request).

Yes please.

Does anyone tellif the Raptor test bay in McGregor that is being outfitted is maybe being out fitted for pre-burner testing of the full scale raptor?

That would seem like a logical place to conduct that testing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/24/2017 05:44 pm
Now, can we all get back to Raptor, pretty please?  (Not intended to call out anyone specifically, and I include myself in that request).

Yes please.

Does anyone tellif the Raptor test bay in McGregor that is being outfitted is maybe being out fitted for pre-burner testing of the full scale raptor?

That would seem like a logical place to conduct that testing.

There's some tangential L2 info related to your question.

Unrelated to the L2 info, given how successful the subscale test program has been, I wouldn't be surprised if SpaceX jumped headfirst into full scale integrated testing. Could start with a burp test this time around, if they want to be extra cautious.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/24/2017 06:53 pm
The real question here, is which engine will create the best rocket?
A few heavy 2400 kN BE4’s or many light 1700 KN Raptors?
We miss some data to answer that question, for instance the mass of both engines.

Thats probably question number one, the airforce wanted to know...

There is only 30% difference between the sizes of your two engines -- one is not 'light' requiring 'many' and the other 'heavy' requiring 'few.'  IF New Glenn used Raptors for same thrust, there would be ten Raptors instead of seven BE-4s... quite similar engine counts.  You're probably mentally comparing the 'light' vehicle New Glenn with the 'heavy' vehicle BFR... the latter could lift 5.56 NG payloads per flight, hence the larger number of engines.  (Compare the 38.9 BE-4 flights needed to lift the same payload as the 31 Raptors.)

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gin455res on 10/24/2017 08:10 pm
Does anyone know if every system on the test Raptor is subscale?


(eg. w/could one run full size pumps at reduced power on a subscale chamber and nozzle?)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 10/24/2017 08:30 pm
Now, can we all get back to Raptor, pretty please?  (Not intended to call out anyone specifically, and I include myself in that request).

Yes please.

Does anyone tellif the Raptor test bay in McGregor that is being outfitted is maybe being out fitted for pre-burner testing of the full scale raptor?

That would seem like a logical place to conduct that testing.

There's some tangential L2 info related to your question.

Unrelated to the L2 info, given how successful the subscale test program has been, I wouldn't be surprised if SpaceX jumped headfirst into full scale integrated testing. Could start with a burp test this time around, if they want to be extra cautious.

Good point, Build, test, iterate is how EM likes to work. 

If the testing results have closely matched the design and expectation then jumping to a more complete next step may work well.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/24/2017 09:07 pm
Good point, Build, test, iterate is how EM likes to work. 

If the testing results have closely matched the design and expectation then jumping to a more complete next step may work well.

Yep. It's also clear that a major goal of subscale testing was to drastically improve SpaceX's modeling of FFSC and their specific implementation. It also allows them to develop a mature understanding of how to operate a FFSC methalox engine and tells them what to look out for in terms of wear, combustion instability, etc.

The fact that they do not appear to have unintentionally destroyed any integrated subscale articles is truly a testament to how accurate their preliminary models must have been. I'm excited to see how the next several months of testing play out, the fact that the most recent USAF payout accelerated the completion date from August to April 2018 bodes well for full scale testing in the very near term. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/25/2017 01:17 am
Hypothetically ...

Why one might want to do a subscale FFSC first?
 * Easier to wrestle with its complexity when compact
 * Easier to get basic performance figures
 * Easier to find materials issues/limitations
 * Easier to reach maximum design chamber pressure
 * Ability to adapt vehicle economics/scale as needed

Why one might want to do a full scale ORSC first?
 * Get to actual operating environment soonest
 * Prove actual design goals and correct functioning at scale
 * Prove CFD and actual engine to be flown agree
 * Determine operating margins to support specific missions
 * Establish a baseline & engineering change correlation as one approaches flight
 * Prove to "stakeholders" that the engine under test does exactly what you say it does
 * Gradually increase thrust/duration while proving stable combustion

Why might AF "follow on" with FFSC after "proof of concept"?
 * To see if results are a fluke or the "scaling" works
 * To understand how predictable the engine is at the extreme limits, as it scales

Note - different agendas/processes/risks, and very proud/experienced people who are very competitive.

This is true worldwide too. Not just these two.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/25/2017 01:30 am
It's /barely/ even "subscale." It's at 80% thrust (better than Blue Origin's 50% thrust...), i.e. 200 bar vs 250 bar. Currently tested to 1MN. At "full throttle" it'd be 1.25MN. That's effectively ~75% scale. Correct me if I'm wrong, but that's about the same difference between the original Merlin 1D and the full thrust Merlin 1D on block 5.

They could probably even just use these 75% scale ones initially with a performance hit.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Patchouli on 10/25/2017 01:37 am
The Raptor is being developed in similar manor that the M1D was done.

First there was a prototype that was used to test assumptions and update the design models. Then the updated design models were used to create a production design with a resulting fairly accurate physical set of specifications in thrust ISP and weight. What happened was the prototype was a design with a probability of a  higher level of successful operation. But then the data gathered is used to "tighten" the engine design models such that when entered a set of constraints like thrust, bell size, ISP, TC pressure, and a few others the design software developed will generate parts drawings that has a high probability to create an engine with almost those exact specifications that will work!

Thus at the end of the prototype testing a specification for a production engine which was delivered 4 months later actually met the design specifications. It also had a near trouble free testing.

So to is the Raptor strategy. The Raptor is at a similar point that the M1D was at in April 2012. And that less than 6 months later production engines were proceeding successfully thru qualification testing and flight units were being delivered and successfully acceptance tested.

Then just a little more than 2 years from the April 2012 point a F9v1.1 flew.

So it is possible that a tanker/cargo version of the SC could start into tests late 2019 or early 2020.

Keep in mind Merlin 1D was an evolutionary improvement from Merlin 1C.
Raptor is an all new engine using a different cycle and propellants.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/25/2017 01:56 am
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (135 bar and 2,450 kN) if both were pushing the same booster and upper stage.

The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 10/25/2017 03:44 am
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.

The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.

Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.

Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: HIP2BSQRE on 10/25/2017 04:05 am
Is there a poll as to when we expect to see a full scale production Raptor engine?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/25/2017 05:11 am
My sincere apologies for unintentionally starting the BE-4 v. Raptor debate ;D Not to say it can't be fruitful, but we should probably move it into a separate dedicated discussion thread to avoid this one straying too far from Raptor.

Is there a poll as to when we expect to see a full scale production Raptor engine?

I added a poll! Feel free to let me know if you think I missed important options.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jded on 10/25/2017 08:57 am
Sorry if this is a wrong thread or this has been discussed already...

I was wondering about using BFS as it's own escape system and I understand one of the problems is that complex engines with turbopumps need more time to start running. But what order of magnitude are we talking about? What could be the possible minimum time they might achieve for Raptor startup sequence? Miliseconds, seconds, a minute?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 10/25/2017 11:26 am
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.

The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.

Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.

Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.
Yes, I meant 135 bar, 67 is the test pressure. Fixed my post.

With chamber pressure known and some some conservative assumptions on mixture,   efficiency, and expansion, ISP is straightforward to calculate.

I assumed that 2x BE-4 and 4x demo Raptor weigh the same, which is likely being very generous to the BE-4.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 10/25/2017 01:54 pm
Sorry if this is a wrong thread or this has been discussed already...

I was wondering about using BFS as it's own escape system and I understand one of the problems is that complex engines with turbopumps need more time to start running. But what order of magnitude are we talking about? What could be the possible minimum time they might achieve for Raptor startup sequence? Miliseconds, seconds, a minute?

Since nobody here has Raptor specs I'd estimate based on other large liquid fueled engines a couple/few seconds to reach full thrust.  BFS low T/W is an issue.  Gotta shut down that BFR first.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 10/25/2017 02:03 pm
Raptor measurements and specs are earlier in this thread as well as on L2.  It is not much bigger than Merlin, so it would seem to be much less mass than BE-4.  A BFR with 31 of these subscale engines would get it in Saturn V range.  Probably 75 to 100 tons to LEO with a full up BFR/BFS with subscale engines. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 10/25/2017 04:03 pm
Raptor measurements and specs are earlier in this thread as well as on L2.  It is not much bigger than Merlin, so it would seem to be much less mass than BE-4.  A BFR with 31 of these subscale engines would get it in Saturn V range.  Probably 75 to 100 tons to LEO with a full up BFR/BFS with subscale engines.

NOT the specs on startup response times which was the question.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 10/25/2017 04:51 pm
That is a good question.  IF the BFS can startup fast enough to escape a RUD on the first stage. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 10/25/2017 05:20 pm
I added a poll! Feel free to let me know if you think I missed important options.

I don't see a poll in the polls section yet.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/25/2017 05:41 pm
I added a poll! Feel free to let me know if you think I missed important options.

I don't see a poll in the polls section yet.

Very odd! It's at the top of the page for me, FWIW.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/25/2017 06:01 pm
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.

The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.

Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.

Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.

Can anyone with engine design background estimate the weight of both engines?

My layman estimation is:
400kg for Raptor 1000kN  (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)

Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mme on 10/25/2017 06:18 pm
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.

The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.

Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.

Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.

Can anyone with engine design background estimate the weight of both engines?

My layman estimation is:
400kg for Raptor 1000kN  (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)

Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
If you are going for a "wisdom of the crowds" estimate it works better if it's done blind. Otherwise you get herding.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/25/2017 06:48 pm
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.

The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.

Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.

Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.

Can anyone with engine design background estimate the weight of both engines?

My layman estimation is:
400kg for Raptor 1000kN  (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)

Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
If you are going for a "wisdom of the crowds" estimate it works better if it's done blind. Otherwise you get herding.

I could have answered I’m part of the crowd, but I don’t consider the members of this particular discussion forum having a herd mentality. Quite the (sometimes annoyingly) opposite actually  :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: intrepidpursuit on 10/25/2017 07:00 pm
I could have answered I’m part of the crowd, but I don’t consider the members of this particular discussion forum having a herd mentality. Quite the (sometimes annoyingly) opposite actually  :)

This crowd is more hive than heard. Somebody just has to do a dance and everyone gets to work on the problem.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 10/25/2017 07:02 pm
Suggest you look up the published dry weight of the Merlin 1D SL and maybe rethink your numbers...

That said... I have already posted (somewhere here) a dry weight guess of ~980kg for a single SL Raptor (1700kN) as fitted to BFR...

On edit...
Source of said guess I made...
https://forum.nasaspaceflight.com/index.php?topic=41363.msg1732978#msg1732978 (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1732978#msg1732978)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 10/25/2017 07:21 pm
Suggest you look up the published dry weight of the Merlin 1D SL and maybe rethink your numbers...

That said... I have already posted (somewhere here) a dry weight guess of ~980kg for a single SL Raptor (1700kN) as fitted to BFR...

On edit...
Source of said guess I made...
https://forum.nasaspaceflight.com/index.php?topic=41363.msg1732978#msg1732978 (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1732978#msg1732978)

Far more likely to be near correct than guesses of half the mass.  T/W of 350 or over 300 are unrealistic.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: intrepidpursuit on 10/25/2017 08:04 pm
Elon said he believes the Raptor will have the best T/W ratio of any engine ever made, but said it hesitantly as if it is only just true. Their M1D holds the title currently at 180:1 (well maybe as high as 199:1 if the weight hasn't grown with the block 5 thrust upgrade). That implies to me that Raptor will be close to that but slightly better. Considering Raptor is as complex an engine as it is, it should have a lower thrust to weight (see the SSME at ~54:1), so to be better than the comparatively dead simple M1D is incredibly impressive already.

If you assume a minimum of 200:1 T/W and the a thrust of 1,900kN, that puts Raptor's max weight around 975kg.
Title: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 10/25/2017 08:12 pm
Elon said he believes the Raptor will have the best T/W ratio of any engine ever made, but said it hesitantly as if it is only just true. Their M1D holds the title currently at 180:1 (well maybe as high as 199:1 if the weight hasn't grown with the block 5 thrust upgrade). That implies to me that Raptor will be close to that but slightly better. Considering Raptor is as complex an engine as it is, it should have a lower thrust to weight (see the SSME at ~54:1), so to be better than the comparatively dead simple M1D is incredibly impressive already.

If you assume a minimum of 200:1 T/W and the a thrust of 1,900kN, that puts Raptor's max weight around 975kg.

Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.

I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.

140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/25/2017 08:45 pm
Elon said he believes the Raptor will have the best T/W ratio of any engine ever made, but said it hesitantly as if it is only just true. Their M1D holds the title currently at 180:1 (well maybe as high as 199:1 if the weight hasn't grown with the block 5 thrust upgrade). That implies to me that Raptor will be close to that but slightly better. Considering Raptor is as complex an engine as it is, it should have a lower thrust to weight (see the SSME at ~54:1), so to be better than the comparatively dead simple M1D is incredibly impressive already.

If you assume a minimum of 200:1 T/W and the a thrust of 1,900kN, that puts Raptor's max weight around 975kg.

Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.

I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.

140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.

What do you think is the weight of the 1000kN sub scale engine?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 10/25/2017 08:57 pm

What do you think is the weight of the 1000kN sub scale engine?


If I had to guess a range, somewhere between 700-1300kg.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 10/25/2017 08:59 pm
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.

I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.

140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.

https://en.wikipedia.org/wiki/RD-270
N2O4/UDMH 190:1 thrust to weight ratio full flow stages combustion at 26MPa chamber pressure in 1969 Russian design.  2018 SpaceX can probably do better than that given advances in analytical tools with their thrust-chamber top integrated LOX turbopump even if LOX/CH4 bulk fuel density is only around 0.82 vs 1.12 for UDMH/N204
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: schaban on 10/25/2017 11:30 pm
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.

I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.

140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.

https://en.wikipedia.org/wiki/RD-270
N2O4/UDMH 190:1 thrust to weight ratio full flow stages combustion at 26MPa chamber pressure in 1969 Russian design.  2018 SpaceX can probably do better than that given advances in analytical tools with their thrust-chamber top integrated LOX turbopump even if LOX/CH4 bulk fuel density is only around 0.82 vs 1.12 for UDMH/N204

I'm not sure if they actually achieved that, though.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/25/2017 11:41 pm
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.

I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.

140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.

https://en.wikipedia.org/wiki/RD-270
N2O4/UDMH 190:1 thrust to weight ratio full flow stages combustion at 26MPa chamber pressure in 1969 Russian design.  2018 SpaceX can probably do better than that given advances in analytical tools with their thrust-chamber top integrated LOX turbopump even if LOX/CH4 bulk fuel density is only around 0.82 vs 1.12 for UDMH/N204

I'm not sure if they actually achieved that, though.

There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.

http://www.russianspaceweb.com/ur700.html

Edit: Hmmm. Found some scholarly articles that contradict that claim: "Moreover, the testing of the RD-270, a key element of the revamped rocket, was not producing satisfactory results. All the 27 test firings conducted between October 1967 and July 1969 ended in some kind of failure before development of the engine was suspended in August 1969."

Hendrickx, Bart. “Heavy Launch Vehicles of the Yangel Design Bureau—Part 1.” Journal of the British Interplanetary Society 63, no. 2 (2010): 50.

Regardless, SpaceX is already at 200 MPa, and their stated goal is to get to 250 and then eventually surpass 300 as operational Raptor matures like Merlin. A bonkers chamber pressure is within reach.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: schaban on 10/26/2017 12:20 am

There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.

http://www.russianspaceweb.com/ur700.html

there's a thread dedicated to this engine on Novosti Kosmonavtiki forum:
http://novosti-kosmonavtiki.ru/forum/messages/forum13/topic4559/message176620/#message176620

you may try google translate it; it states that all test fires ended in failures, sometimes serious. All they achieved is a startup at reduced pressure and smaller bell.

Author of the message on forum claims to quote it from official Energomash history book

Exact russian quote:
Quote
Всего за период с октября 1967 г. по июль 1969 г. было проведено 27 огневых испытаний двигателя РД-270. Все они имели аварийный исход.

Google translate:
Quote
In total, during the period from October 1967 to July 1969, 27 fire tests were performed on the RD-270 engine. All of them had an emergency outcome.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/26/2017 12:33 am

There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.

http://www.russianspaceweb.com/ur700.html

there's a thread dedicated to this engine on Novosti Kosmonavtiki forum:
http://novosti-kosmonavtiki.ru/forum/messages/forum13/topic4559/message176620/#message176620

you may try google translate it; it states that all test fires ended in failures, sometimes serious. All they achieved is a startup at reduced pressure and smaller bell.

Author of the message on forum claims to quote it from official Energomash history book

Exact russian quote:
Quote
Всего за период с октября 1967 г. по июль 1969 г. было проведено 27 огневых испытаний двигателя РД-270. Все они имели аварийный исход.

Google translate:
Quote
In total, during the period from October 1967 to July 1969, 27 fire tests were performed on the RD-270 engine. All of them had an emergency outcome.

Ha, I must have found an author that translated the same source :D Edited my post before I saw this.

Sorry, mods, OT :-X
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Patchouli on 10/26/2017 12:50 am

There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.

http://www.russianspaceweb.com/ur700.html

there's a thread dedicated to this engine on Novosti Kosmonavtiki forum:
http://novosti-kosmonavtiki.ru/forum/messages/forum13/topic4559/message176620/#message176620

you may try google translate it; it states that all test fires ended in failures, sometimes serious. All they achieved is a startup at reduced pressure and smaller bell.

Author of the message on forum claims to quote it from official Energomash history book

Exact russian quote:
Quote
Всего за период с октября 1967 г. по июль 1969 г. было проведено 27 огневых испытаний двигателя РД-270. Все они имели аварийный исход.

Google translate:
Quote
In total, during the period from October 1967 to July 1969, 27 fire tests were performed on the RD-270 engine. All of them had an emergency outcome.

If I remember the RD-270 had some combustion instability issues that were never fully solved before the program was canceled.

On my vote on when we'll see a full scale Raptor on the test stand I think mid 2018 to early 2019.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/26/2017 03:33 am
Can anyone with engine design background estimate the weight of both engines?

My layman estimation is:
400kg for Raptor 1000kN  (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)

Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values

No, you don't get closer to the truth based on just opinions and data pulled from thin air. (Like your numbers) More people doing the same doesn't make the conclusion more valid.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 10/26/2017 04:16 am
Can anyone with engine design background estimate the weight of both engines?

My layman estimation is:
400kg for Raptor 1000kN  (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)

Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values

No, you don't get closer to the truth based on just opinions and data pulled from thin air. (Like your numbers) More people doing the same doesn't make the conclusion more valid.

Rather unintuitive , but there is actually some veracity to the "wisdom of the crowd" concept (https://www.ncbi.nlm.nih.gov/pmc/articles/PMC3107299/). The effect is far more pronounced when in highly selective groups like these forums, too :)

Nevertheless, this thread 100% should lean more towards discussion based on available data (like mass estimates from known thrust and TWR figures above) than base speculation.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 10/26/2017 04:51 am
Rather unintuitive , but there is actually some veracity to the "wisdom of the crowd" concept (https://www.ncbi.nlm.nih.gov/pmc/articles/PMC3107299/). The effect is far more pronounced when in highly selective groups like these forums, too :)

True, and it's pretty amazing to review the credentials of some of the participants here. Not just one or two rocket designers, but a bunch. So yeah. Our polls are often pretty predictive too.

Quote
Nevertheless, this thread 100% should lean more towards discussion based on available data (like mass estimates from known thrust and TWR figures above) than base speculation.

Absolutely, as much fun as the alternative might be. There are other threads
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 10/26/2017 05:45 am
In this case I think it’s questionable that reverse TWR calculation will lead to the correct mass.
It’s more objective to use the Cad renderings to estimate mass, how unlikely the TWR outcome might be.
Maybe an anonymous poll will be more predictive, for wisdom of the crowd. There the non-con-formative estimation won’t be called out or ridiculed upon.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aero on 10/26/2017 02:21 pm
I'll bet that some here have used the Delphi method to predict the solution to types of problems.

https://en.wikipedia.org/wiki/Delphi_method (https://en.wikipedia.org/wiki/Delphi_method)

We have on this forum the correct assortment of experts to answer this current question, all we need is a moderator and a clear definition of the question to be answered. Then a rule for the cut-off of course.

aero
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/26/2017 07:02 pm
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
You can't say this yet.

What you can say is that they are likely to exceed RD-180 (possibly even RD-270 BTW) in some ways given what they currently show. That's fair. But in truth all of this is missing the point.

You have to judge engines "apples to apples". That's really hard here. You can judge test programs, ambitions, project close rates, accomplishments. Any where you can keep the units straight.

Raptor's ambition is not BE-4. (Although by quirk of AF contract, technically Raptor as an propulsion asset could benefit other AF related work thus Vulcan thus the perception of overlap, which I'm certain does not go down well with the Blue engine team.) Likewise BE-4's ambitions, both near term and longer term, are not Raptors.

I would caution all, especially the engine teams, to bear that in mind.

BE-4 will likely eclipse RD-180 (a very hard act to follow) in all ways, including in vacuum thrust.

Raptor will have to exceed BE-4's T/W ratio while having a hundred times better reliability under more extreme conditions, w/o rework, to accomplish its objective. Also, it will likely always beat BE-4 on chamber pressure.

BE-4 will always have less wear on its fewer working parts. It may be cheaper to manufacture.

To do this will take years/decades.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 10/26/2017 09:26 pm
Does anyone have breakdown on masses for an existing high pressure high performance design - like the SSME or RD191 for the major subcomponents like nozzle, thrust chamber, turbopumps?

Having some idea of nozzle and combustion chamber mass would probably bring more clarity to what is possible for overall engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/26/2017 09:31 pm
...
BE-4 will likely eclipse RD-180 (a very hard act to follow) in all ways, including in vacuum thrust.
...

Not sure where this comes from... 2x BE-4 about equals RD-180, and BE-4 is shooting for less extreme operating parameters across the board with respect to RD-180.

What do you mean by eclipse -- replace, or exceed RD-180 performance cross the board?  If the latter, I believe Blue has stated otherwise.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/26/2017 09:57 pm
...
BE-4 will likely eclipse RD-180 (a very hard act to follow) in all ways, including in vacuum thrust.
...

Not sure where this comes from... 2x BE-4 about equals RD-180, and BE-4 is shooting for less extreme operating parameters across the board with respect to RD-180.
That's the current goal.

All they need to do to satisfy ULA is a reliable, proven version of just that. Certain that is in the crosshairs.

But I was referring to the continuing agenda.

Which likely means "gradatim ferociter" as applied to taking the current design as far as it is possible to go, given the base technology. ULA doesn't need much in the way of engine reuse (some for testing clearly), but NG definitely does.

And the lower chamber pressure was meant to secure the ability to reach ULA's goals in a timely manner, butonce that is done, you either/both reduce the design or increase the chamber pressure, improve combustion efficiencies/velocity, improve power pack mass flows, add throttle range, ...

You also characterize the engine for broader applications. Like say vacuum/US.

Quote
What do you mean by eclipse -- replace, or exceed RD-180 performance cross the board?  If the latter, I believe Blue has stated otherwise.

They'll still use 2 on Vulcan, as opposed to the single engined, dual chambered RD-180 on Atlas V. (Please note than the RD-180 is a variant of the four chamber RD-170 (now RD-171M). (If you want to compare "apples to apples", use the single chamber RD-191 to BE-4. Or, conceivably one could do a dual chamber variant of the BE-4 to directly compare to the RD-180, which would not make sense for other reasons.)

But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Johnnyhinbos on 10/27/2017 07:11 am
Wait, what thread is this?

And while I’m grousing, could we get a thread title change? “ITS” bothering me... (<- see what I did there?)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 10/27/2017 01:18 pm
...
But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.

Understand where you're going with this, but don't think Blue has the push-the-tech-to-the-limit DNA that is driving Raptor. 
Started with much lower goals and will end with much lower performance, IMO. 
Whatever you mean by 'effectiveness' (T/W, ISP, ?) will be lower, too -- lower than RD-180 and Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/28/2017 01:53 am
...
But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.

Understand where you're going with this, but don't think Blue has the push-the-tech-to-the-limit DNA that is driving Raptor. 
Started with much lower goals and will end with much lower performance, IMO. 
Whatever you mean by 'effectiveness' (T/W, ISP, ?) will be lower, too -- lower than RD-180 and Raptor.

These teams all think they're doing the best engine, second to none.

But the first priority for BE-4 after reaching basic performance targets is ... reliable, reproducible, deterministic. Otherwise ULA can't use it. When they can use it, it doesn't have to be anything more than that to allow Vulcan to displace Atlas/Delta (with Centaur V ...). Doesn't have to eclipse Raptor/RD-180.

Yes, they could sit on their hands then. "Mission accomplished"  ::) But that's just the start for Vulcan/NG.

Given the processes they use, they retain the above, at no added risk, ... but increase chamber pressure/iSP/duration/margin. Vulcan vehicle/avionics/operations adapts to encompass this in missions.

RD-180 also has been gradually improving, as a mature design. But mature designs already have lesser bounds, as to go further, you have to add risk with significant changes, which likely exceed the scope of the business. (There is work on other RD-170 variants pressing.) What vehicles would those newer scoped engines fly in? Not Atlas/Vulcan.

Now circle back to Raptor - this thread.(BTW, all engine teams are acutely aware of the others work, practically in real time.) These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe. And the engines will have to on a single mission critically fire dozens of times without incident. Unlike BE-4/RD-180, who will have 2-3 critical burns in flight at max (Vulcan just one!). And since those burns will occur after extremely high delta-v targets (again unlike the other two), propulsion efficiency/iSP has to exceed the other two, to retain the advantage of the rocket equation with remaining propellant - to allow the vehicle architecture to realize its design goals.

Very different engines. BE-4, unlike RD-180 but like Merlin 1D, will have a vacuum variant. However, RaptorVac isn't in the same league. We're talking about optimization for in space propulsion as the majority of its role (the NG/NA architecture pursues hydrolox for this purpose), with a scaled engine to match that need without additional stages but with refueling.

(Note that ULA is backing off ACES and instead going for a expanded Centaur V. They can't get the "buy in" to fund the rest, which likely will be factored in incrementally as capability is desired.)

So they point to be made with this is that the different approaches by SX, BO, ULA in vehicles/engines is not in them being  more/less talented/aggressive/creative/experienced/... its instead the nature of what they are attempting to bring to bear.

For ULA its a next generation Atlas without past baggage, leveraging as much of the future as the parents will let them.

For BO its in entering the partially reusable LV business at Ariane/SX level of capability/flight frequency.

For SX its in a fully reusable vehicle with interplanetary HSF capability in excess of SLS.

Back to RD-170 variants for comparison - likely enhanced single chamber for Angara, methalox four chamber derivative for Russian SHLV like SLS. They'll not need much more than per chamber what RD-180 already does. And reuse isn't yet on the map.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 10/28/2017 11:37 am
I think I know the answer.
Quote
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/28/2017 12:44 pm
I think I know the answer.
Quote
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
Full flow means you are extracting more energy for pumping.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 10/28/2017 11:17 pm
I think I know the answer.
Quote
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
Full flow means you are extracting more energy for pumping.
Advantages of a full-flow staged combustion cycle engine system (https://arc.aiaa.org/doi/abs/10.2514/6.1997-3318)

The chief point here is that the separate OR/FR paths are at less pressure than the combustion chamber, and there is no interpropellant seal to fail at extreme pressure or transient flow. The maximum chamber pressure is thus set by the design limits of the combustion chamber and injector(s).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ChaoticFlounder on 10/28/2017 11:38 pm
I think I know the answer.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).

Hopefully this will help you understand a little bit better the design choices they're making

so, pump power requirement is directly proportional to mass flowrtate and output pressure (P1)

P1 is going to be your highest pressure in the system and will directly dictate your chamber pressure (Pc)

so, turbine work per unit mass is equal to

Tw/m = Cp*(T4 - T5)

where:

Cp = Specific Heat
T4 = Turbine Inlet Temperature
T5 = Turbine Outlet Temperature

This is pulled directly from this site:  https://www.grc.nasa.gov/www/k-12/airplane/powtrbth.html

When you multiply this specific turbine work by your preburner mass flow rate, you get shaft power (Ps)

This shaft power is used to drive the pump;  therefore, pump output pressure is directly related to (preburner mass flow)*(deltaT) across your turbine

deltaT is limited by your materials, so your variable to change is mass flow and that is what the FFSC does in spades

Essentially, it's saying that your theoretical pressure limit is going to be around twice what your theoretical pressure limit will be with FRSC / ORSC <- assuming T4 is the same and your propellant densities are similar

All, please correct me if I said anything wrong and ask if you have questions.

C


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: IainMcClatchie on 11/11/2017 08:03 am
How hot is the partially combusted oxygen in the turbopump?

In the following quick analysis, I'm assuming a chamber pressure of 25 MPa and that the turbine and compressor are 100% efficient, with no heat leakage into the incoming LOX at 66 K.  I tried a range of preburner pressures from 35 to 60 MPa.  I got very, very low temperatures.

The output of the compressor is a subcritical liquid around 72 K.  The preburner appears to heat it up to at most room temperature.  At 50 MPa turbine inlet temperature, the turbine inlet is at -17 C.  Flameholding will be a challenge.  The preburner is going to need to burn with a small fraction of the LOX before mixing the result with the bulk of the LOX, a more radical version of the burner cans in turbofan engines.

There is a substantial benefit to running the preburner at high pressure: the turbine extracts more energy, and so less propellant is burned in the preburner.  That means everything from the preburner output to the injector face gets more dense and therefore smaller.  In particular, the volume gets smaller faster than the pressure goes up, so the figure of merit for a pressure vessel, which is pressure*volume, goes DOWN at higher preburner pressure, while at the same time you get a small bump in exhaust velocity from burning more of the propellant in the main chamber..

That means an engine with a higher preburner pressure can be lighter weight, which seems counterintuitive to me.

Turbine energy extracted varies from 7% at 35 MPa to 20% at 60 MPa.

I wonder how useful it is to fully mix the preburner.  More energy can be extracted from a hotter stream, so that for a given chamber pressure, less propellant can be used by the preburners and a higher chamber temperature and better exhaust velocity can be achieved.  With these ridiculously low temperatures, there is gobs of temperature headroom.  It seems totally feasible to triple the energy extracted by the turbopump.

I don't think they are going to need ceramic coatings for this thing.  The turbine bits do have to deal with high pressure oxygen, but surely that's more benign than hot oxygen.

I've attached my spreadsheet if anyone wants to check my math.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 11/11/2017 05:19 pm
Looking at the above... I'm kind of surprised and then again not...  ???

I mean, my take on this FFSC cycle is... Use the "energy" stored in the liquefied prop as it "boils" back to a vapor to drive a turbine powering the pump...

I always figured the highest pressure (2x+ chamber) would be found between the pump outlet and the "heater" in line...
I don't like to call it a preburner... as we are not sure it actually will use much if any of the LOX flow rushing by...
I like to simplify it, by picturing a burner can fed gaseous oxygen and gaseous methane from other onboard sources...
The vaporizers for these flows MAY be the burner can itself separately fed high psi liquid prop from upstream (so it can be controlled)...
The exhaust from the preburner mixes with the cold LOX and heats it enough to flash it all to oxygen "steam"...
It's still at nearly the same high pressure as it enters the turbines and the pressure is then dropped across the turbines converting the energy of the much expanded flow into mechanical energy to drive the pumps... 

LOL... Yes... it's kind of like taking a fire hose and hooking it up to a steam turbine...   ;D
Just got to put a big enough heater in line to convert all the water to steam before it reaches the turbine...  :o

Anyway... that's how I wrapped my head around FFSC Raptor inner workings...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 11/11/2017 06:10 pm
That doesn't make sense.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: yokem55 on 11/11/2017 07:21 pm
SpaceX Veterans Day Commemoration pic on Twitter has the McGregor vets pictured in the Raptor Test cell.


https://twitter.com/SpaceX/status/929441494494208000
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: IainMcClatchie on 11/11/2017 07:52 pm
It seems the preburner has to mix the flame exhaust into the LOX pretty well to avoid two-phase hitting the turbine.  So much for making the turbine more efficient.

That said, the oxygen preburner is using 2-3% of the total energy flow of the propellant, most of that just for the phase change rather than temperature change.

Note that the oxygen is hardly *flashing* to "steam".  This diagram shows that the transition from liquid to supercritical is very smooth at 25 MPa.  50 MPa will be very smooth as well.

Neither oxygen nor methane is squishy in their liquid forms, so the turbopump shaft power is 15-20 megawatts instead of some even more ludicrous number necessary for a hydrogen compressor.

The turbopumps in the Raptor will be amazingly small.  The fluid density coming from the preburner will be well over half that of water, so the volume is under 800 liters/sec.  I'm not sure what the right velocities are, but at 100 m/s, that's a cross section of 80 cm^2.  I doubt it ever goes in a pipe but if it did it would be 10 cm diameter.

Does anyone have some rules of thumb for determining the turbopump compressor and impeller sizes?  The tip speed needed for a single-stage 60 MPa centrifugal pump is 304 m/s, which is just below the speed of sound in cold LOX.  Liquid methane has a much higher speed of sound, well over 1 km/s.  If a single-stage impeller can do the job, I'm wondering if the turbopumps can eliminate the shaft and stator vanes entirely and just consist of an impeller and turbine back-to-back, with the burner cans around the periphery.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 11/11/2017 08:13 pm
Agreed... "Steam" was a term used in gross error...
BUT, I was about to use a water analog, so I used it to help connect the two...  :-[

I understand gas turbines and steam turbines (air and water working fluids) and was just trying to relate it to a LOX working fluid...

My bad...  :P
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 11/11/2017 08:31 pm
The aforementioned test stand pic.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 11/12/2017 04:16 am
The aforementioned test stand pic.

Full-scale engine?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 11/12/2017 05:15 am
The aforementioned test stand pic.

Full-scale engine?

It's hard to judge from this picture, and the difference between subscale and full scale is going to be extremely small, ~10-20% as modelers on this thread have estimated. I expect Musk will tweet about it whenever SpaceX conducts the first "full scale" tests or at least the first 250 bar test.

But he suggested that the main hurdles between the test article and flight engine would be aggressive mass reduction and reliability improvements, so no guarantee that anything more than the test firing videos we've been given will be deemed tweetworthy.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 11/12/2017 01:03 pm
Re: cold turbine exhaust temperatures.

It appears to me you want the turbine exhaust at all thrust regimes to be above the condensation temperature of steam. At 20..30 bar steam should condense below 230..240 celsius.

After the preburner/heater the propellant is a mix ture including combustion products.

No?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 11/12/2017 02:32 pm
The aforementioned test stand pic.

Full-scale engine?

L2 discussion on that very issue
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 11/12/2017 05:35 pm
Re: cold turbine exhaust temperatures.

It appears to me you want the turbine exhaust at all thrust regimes to be above the condensation temperature of steam. At 20..30 bar steam should condense below 230..240 celsius.

After the preburner/heater the propellant is a mix ture including combustion products.

No?

Yes and No... (my opinion)
Yes there will be some small amount of water vapor in the combustion products downstream of the burner can...
BUT...
Note the heat added to the stream is only enough to phase change the easily pumped, not compressible, sub cooled LOX into squishy. expandable, but supercritical at high pressure gaseous Oxygen  in good enough condition to be expanded over a turbine section from say 600 bar down to 350 bar on it's way across say a final 50 bar drop across the chamber injector into a running, firing 300 bar rocket combustion chamber...
Literally... the temps at the oxygen turbine may be room temperature... and the turbine design will have to allow for some liquid droplet (lox or otherwise) to pass thru harmlessly I believe...

The startup sequence on this must be very interesting... I must say...  ???

The fact the above picture of a pristine Raptor development test stand with a year old development engine  (assuming same basic assy) still in place is a HUGE achievement...
Tom Mueller and his group figured out how to start and stop this thing without it blowing up...  8)

Later edit... fixed my bar numbers above off by factor 10 (woops)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 11/12/2017 05:46 pm
From my understanding water droplets in a gas turbine is a bad thing, abrasive blasting or some such effect.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 11/12/2017 05:50 pm
From my understanding water droplets in a gas turbine is a bad thing, abrasive blasting or some such effect.

Yes it is... very much so in both gas and steam turbines with their sharp and delicate edges that are easily damaged by water or ice...
Somehow they must have a design that can handle it...  ;)

On edit...
I am also quite sure this design and control information will never see the light of day as long as SpaceX has say over their designs and intellectual property... BUT, we can speculate...  :)

Much later edit...
I'm wondering if maybe the turbine has the reverse of gas turbine industry typical and has passageways in the blades to allow HEATING the turbine somewhat using hot gas tapped from the upstream burner assy so as to heat the blades enough to vaporize on contact any stray droplets that find their way into contact with the turbine...  ???
If so... that's some interesting stuff right there...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: IainMcClatchie on 11/12/2017 08:00 pm
Hominans,

You raise an excellent point.

Here's the graph of water density vs temperature at 25 MPa (nominal combustion chamber pressure).  As you say, the stuff is going supercritical around 650 K.

I'm really not sure if this is a terrible problem or not.

At 25 MPa, the density difference between supercritical and liquid is smooth.  If you've got a two-phase mixture, the liquid isn't going to separate out quickly at all.  I'm guessing water droplet separation isn't a big problem in high pressure steam turbines.

At 1 MPa, the density difference between vapor and liquid is not smooth.  The liquid is 170x denser than the vapor, and 10x more viscous.  A two-phase mixture is going to separate.  This means that drops of water can get ballistic velocity relative to the overall flow, and I'd guess that's what causes the erosion.  I know low pressure steam turbines have systems for removing water from the steam flow in the middle of the multi-stage turbine.  That said, they do tolerate some condensation, it's just important to get it out expeditiously.

At 25 MPa and 221 K, the density difference between ice and oxygen isn't large, about 2:1 to 3:1 (don't have a handy reference for high pressure ice density).  I wouldn't think the ice would form large crystals, but I'd get pretty nervous about the stuff accumulating on surfaces in the preburner and then flaking off, or worse still plugging injector nozzles.

If the preburner pressure is turned down below 40 MPa, the turbine outlet temp can be above 273.16 K, where NIST's data starts for liquid water.  With a preburner pressure of 35 MPa, turbine inlet temp is 335 K, turbine outlet temp is 306 K, and at the outlet the density of oxygen is 325 g/liter and the density of water is 1007 g/liter, about 3x larger.  Maybe that wouldn't separate fast enough to be troublesome.

All this trouble makes me wonder if you might have the compressor generate two streams.  One much smaller stream goes through the preburner and generates supercritical fluid hot enough that the turbine exhaust is above the autoignition temperature for LOX-methane (550 C?).  This exhaust is mixed with similarly hot methane at the top of the combustion chamber.  The main LOX and methane paths stay cool and get injected into the combustion chamber slightly downstream (maybe just centimeters) of the torch face.  The idea would be to eliminate the need for recirculation-based ignition inside the combustion chamber, which might make for smoother combustion.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: IainMcClatchie on 11/12/2017 09:30 pm
Autoignition temperature for methalox is lower at high pressure.  This article http://onlinelibrary.wiley.com/doi/10.1002/bbpc.19950990110/abstract (http://onlinelibrary.wiley.com/doi/10.1002/bbpc.19950990110/abstract) reports a stoichiometric mix spontaneously ignites at 900 K at 0.1 MPa and 660 K at 110 MPa.

So at 25 MPa autoignition might happen at 800 K.

This is a lot hotter than I was hoping for.  The corresponding turbine inlet temperature is quite high and so all the relaxed metal requirements of FFSC are gone.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 11/12/2017 10:20 pm
On the Methane side of Raptor...
I've always thought of it as a modified expander cycle with the preburner there to kick start it from cold and add some heat to vaporize the LNG full flow pre turbine once running...

In short... BOTH turbines will run at near room temps once going... (my opinion)  ;)

That said... the hard part of Raptor is starting it... (I think)
I'm thinking a supply of very high pressure gaseous oxygen and gaseous methane is needed to bring Raptor to life from a cold start...
700 bar room temp COPV's anyone?...  :o

All preburners (and the RCS system) share this common supply (maybe with some redundancies)

Once a Raptor is running... It can be tapped to refill such a bottle supply and keep it topped up...
(tap high pressure liquid into a small "boiler" to batch flash it into the higher pressure of the storage system)
Batch boilers may be electric heated... I'm not sure on that... 

It's all a system... thinking system and not just a rocket engine here...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 11/13/2017 06:23 pm
Near the end of this video, 2 new clips of raptor including startup (18 video frames between the start of what is shown and main ignition)

https://youtu.be/TXYh4re0j8M?t=2m45s (https://youtu.be/TXYh4re0j8M?t=2m45s)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 11/13/2017 06:34 pm
2:44 or so... 18 frames or so of the start up and light off...
very interesting to pause and then use < and > to step thru it back and forth (from the YouTube webpage)

Again I must say... the secret sauce has got to be the light off sequence...  8)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 11/13/2017 06:56 pm
Again I must say... the secret sauce has got to be the light off sequence...  8)

Indeed. With FFSC, aside from the metallurgy needed for Raptor's operating temp and pressure, a reliable and simple startup procedure is arguably the most difficult problem that has to be solved.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 11/13/2017 07:22 pm
Near the end of this video, 2 new clips of raptor including startup (18 video frames between the start of what is shown and main ignition)

https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s (https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s)

Check the comment tree:
Elon Musk: Can confirm, pretty cool place.
Jeff Bezos: Want me to show you a real rocket?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 11/13/2017 07:27 pm
Near the end of this video, 2 new clips of raptor including startup (18 video frames between the start of what is shown and main ignition)

https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s (https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s)

Check the comment tree:
Elon Musk: Can confirm, pretty cool place.
Jeff Bezos: Want me to show you a real rocket?

There are youtube users "Elon Musk," "Jeff Bezos," and "Tory Bruno" who all joined youtube September 16, 2017 ... parody accounts?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 11/13/2017 07:27 pm
Check the comment tree:
Elon Musk: Can confirm, pretty cool place.
Jeff Bezos: Want me to show you a real rocket?

I'm guessing that's not actually Elon Musk, Jeff Bezos, and Tory Bruno.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 11/13/2017 08:57 pm
GIF of the startup sequence:

(https://j.gifs.com/[email protected])
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 11/13/2017 09:02 pm
Steady state #1:

(https://j.gifs.com/[email protected])
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 11/13/2017 09:09 pm
Steady state #2:

(https://j.gifs.com/[email protected])
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 11/14/2017 01:58 am
And full res screenshots!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: IainMcClatchie on 11/14/2017 08:19 am
Is that a green flash when the Raptor starts up?  TEA/TEB shot?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AbuSimbel on 11/14/2017 09:37 am
Is that a green flash when the Raptor starts up?  TEA/TEB shot?
Isn't it spark ignited?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 11/14/2017 10:16 am
Green could be a camera artifact? Or maybe they are using TEA/TEB in development (I would think not)? Maybe it's used for ground starts? That doesn't make a lot of sense to me, you would think you'd want the spark igniter tested a lot under varying conditions (it has to work for supersonic retropropulsion, as well as in a vacuum, in near vacuum but earth composition atmosphere,, in a thin but mostly CO2 atmosphere, etc....)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 11/14/2017 11:44 am
Maybe the igniter or some other part of the engine contains copper?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 11/14/2017 12:08 pm
Drew a little doodle for size comparison of Raptor compared to some others. Not finished yet, will add more engines and get a scan later.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 11/14/2017 12:17 pm
There was green flash in the middle of of the 40 second test fire that was revealed before. 
https://www.reddit.com/r/SpaceXLounge/comments/73v4yh/comment/dntc2ku

Hopefully just a camera effect but perhaps engine rich combustion? Fuel impurities?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 11/14/2017 12:51 pm
Green is usually impurities.  It could be running LNG, not pure methane. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 11/14/2017 01:20 pm
I think it's TEA-TEB.  Spark igniters are external (?right?) and there's no evidence of one present.  There's also no illumination that would suggest sparks prior to ignition.  SpaceX is very comfortable with and experienced at using TEA-TEB.  The idea they would use it for the prototype engine seems very unsurprising to me as a result.

As far as the green flash in the middle of the previous fire, maybe some residue somehow?  That seems unlikely.  Or a leak.  That would be troubling, not really for the engine itself, but more for the test program.  So that's definitely a data point going the other direction.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 11/14/2017 02:59 pm
As the green at the Raptor test stand or the merlin test stand? because we KNOW merlin uses TEA/TEB
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 11/14/2017 03:01 pm
Raptor.  As you say we know TEA-TEB is used for Merlin ignition.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 11/14/2017 03:31 pm
I think it's TEA-TEB.  Spark igniters are external 

I believe you are thinking of the pyros visible around the main engines at the base of STS. Those were not there to ignite the engines, their purpose was to burn any stray H2 before ignition. Raptor ignition device will be internal.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 11/14/2017 03:36 pm
I think it's TEA-TEB.  Spark igniters are external (?right?) and there's no evidence of one present.  There's also no illumination that would suggest sparks prior to ignition.  SpaceX is very comfortable with and experienced at using TEA-TEB.  The idea they would use it for the prototype engine seems very unsurprising to me as a result.

As far as the green flash in the middle of the previous fire, maybe some residue somehow?  That seems unlikely.  Or a leak.  That would be troubling, not really for the engine itself, but more for the test program.  So that's definitely a data point going the other direction.

I rather imagine that the spark ignitors they'll use will be a torch type one like this:

https://twitter.com/fineri/status/930355129454178306 (https://twitter.com/fineri/status/930355129454178306)
(https://pbs.twimg.com/media/DOlIjihXcAE1LKe.jpg)

Also I think the green in the raptor startup sequence is a camera artifact (chromatic abberation, or sensor artifacts) due to the fact that when the main chamber lights the video is massively overexposed as the camera frantically dials down the exposure in the next few frames.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: notsorandom on 11/15/2017 05:10 am
Could the green be a very slight engine rich combustion?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 11/15/2017 08:45 am
Could the green be a very slight engine rich combustion?

Most likely it's just a camera artifact from the sudden increase in brightness.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/15/2017 08:37 pm
Could the green be a very slight engine rich combustion?

Most likely it's just a camera artifact from the sudden increase in brightness.

I think so too. I remember distinctly that Elon said in the 2016 presentation that Raptor has a little torch inside that is spark ignited which in turn ignites the main combustion cycle. Creating a prototype with hypergolics makes no sense since this is one of the core problems in engine development.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 11/15/2017 10:22 pm
Could the green be a very slight engine rich combustion?

Most likely it's just a camera artifact from the sudden increase in brightness.

I think so too. I remember distinctly that Elon said in the 2016 presentation that Raptor has a little torch inside that is spark ignited which in turn ignites the main combustion cycle. Creating a prototype with hypergolics makes no sense since this is one of the core problems in engine development.

Also, with his Mars ambitions, Elon needs Raptor to be able to be restarted many times without servicing or refilling TEA/TEB supplies. And, spark ignition of an oxygen/methane mixture is a pretty well understood problem at this point -- millions of stoves, furnaces, and water heaters in service doing it every day.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 11/16/2017 11:56 am
FWIW there are plenty of kerosene/oxygen refrigerators out there in the world as well, although they tend to use pilot lights, not spark (or TEA/TEB!) ignition.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: abaddon on 11/16/2017 04:18 pm
To be clear, nobody thinks that TEA/TEB might be used for the production engine.  We know it will use spark igniters.

I've mostly come around to it not being TEA/TEB in the prototype either.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/23/2017 07:25 am
Spinning the wheel a bit further (pun intended), how does the Raptor actually start? I mean, spark ignition or not, it needs to spin up its turbines. Following the ongoing discussion on the Merlin:

[...]
The LOX and RP-1 tanks are pre pressurized with helium. 
High pressure helium spins up the turbo pump.  LOX and RP-1 are ignited by TEA-TEB in the gas generator and  takes over from the helium.  The propellants meet in the combustion chamber and are also ignited by TEA-TEB.
[...]

But the Raptor doesnt have high pressure helium available. Its tanks are autogenous pressurization. So how do the turbine wheels of Raptor start? I do have ideas how it could be done but I dont want to wildly speculate. Does anyone has info on that?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 11/23/2017 09:16 am
They said it uses autogenous  pressurization, so use some of those gases.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/23/2017 10:43 am
They said it uses autogenous  pressurization, so use some of those gases.

I am pretty sure the tank pressure provided by autogenous pressurization system is not enough to start the spin of the turbines. If that was the case, F9 would be able to do the same with LOX and RP1 but they use high pressure helium instead. Probably a lot of it. But I am not an expert and happy to be proven wrong.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hopalong on 11/23/2017 11:23 am
IMHO the CH4 and Oxygen gas would be fed into the turbine combustion chamber and lit. The resulting combustion gases will spin the turbine. There may be a few ticks involved in getting enough initial pressure in the CH4 and GO2 supply lines to the pump turbines, but I understand that there is a lot of dark arts in starting a full flow engine.  :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/23/2017 01:06 pm
Spinning the wheel a bit further (pun intended), how does the Raptor actually start? I mean, spark ignition or not, it needs to spin up its turbines. Following the ongoing discussion on the Merlin:

[...]
The LOX and RP-1 tanks are pre pressurized with helium. 
High pressure helium spins up the turbo pump.  LOX and RP-1 are ignited by TEA-TEB in the gas generator and  takes over from the helium.  The propellants meet in the combustion chamber and are also ignited by TEA-TEB.
[...]

But the Raptor doesnt have high pressure helium available. Its tanks are autogenous pressurization. So how do the turbine wheels of Raptor start? I do have ideas how it could be done but I dont want to wildly speculate. Does anyone has info on that?

They said it was spark ignited. The sparks probably ignite ignition torches which in turn ignites the pre-burners and the main chamber.  You can see the ignition leads on their CAD model.

This ignition approach would make all Raptors restartable assuming their propellants had enough head pressure.
Head pressure and an electrical power source is all that is required to start.

The start sequence is something like the following:
- crack valves and dribble in propellants to pre-chill the engine.
- open valves and propellants flow into their respective pre-burners.
- spark ignites stoichiometric mixture in torches.
- torches ignite pre-burners
- pre-burner exhaust spins turbines attached to propellant pumps. (one for methane, one for LOX)
- main chamber torch ignites gaseous propellants entering chamber.
- pumps start increasing pressure above head pressure and quickly climb to design pressure.

This requires detailed understanding of the combustion processes and the dynamics of the pumps, turbines and valves. It is a tightly choreographed dance.

John

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/23/2017 08:48 pm
This ignition approach would make all Raptors restartable assuming their propellants had enough head pressure.
Head pressure and an electrical power source is all that is required to start.

The start sequence is something like the following:
- crack valves and dribble in propellants to pre-chill the engine.
- open valves and propellants flow into their respective pre-burners.
- spark ignites stoichiometric mixture in torches.
- torches ignite pre-burners
- pre-burner exhaust spins turbines attached to propellant pumps. (one for methane, one for LOX)
- main chamber torch ignites gaseous propellants entering chamber.
- pumps start increasing pressure above head pressure and quickly climb to design pressure.

This requires detailed understanding of the combustion processes and the dynamics of the pumps, turbines and valves. It is a tightly choreographed dance.

John

Thats exactly what I am interested in. So initially, the propellant flows through the not-jet-rotating pumps until it reaches the preburner, is than ignited. It therefore puts pressure onto the turbine which starts to turn. But at the same time, the preburner also puts pressure back up the pumps and into the tanks. Because the pumps are not yet rotating. They are about to start rotating but they dont do it yet. It looks to me like a hen and a egg problem. How can you start the turbines/pumps under these conditions? Are there valves in front of the preburner that quickly close once some propellant is in the preburners and push it out the turbine only to open a fraction of a second later to allow new fuel to reach the preburner and further turn the turbine? And now my thought process looks like a moebius strip...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 11/24/2017 12:04 am
This ignition approach would make all Raptors restartable assuming their propellants had enough head pressure.
Head pressure and an electrical power source is all that is required to start.

The start sequence is something like the following:
- crack valves and dribble in propellants to pre-chill the engine.
- open valves and propellants flow into their respective pre-burners.
- spark ignites stoichiometric mixture in torches.
- torches ignite pre-burners
- pre-burner exhaust spins turbines attached to propellant pumps. (one for methane, one for LOX)
- main chamber torch ignites gaseous propellants entering chamber.
- pumps start increasing pressure above head pressure and quickly climb to design pressure.

This requires detailed understanding of the combustion processes and the dynamics of the pumps, turbines and valves. It is a tightly choreographed dance.

John

Thats exactly what I am interested in. So initially, the propellant flows through the not-jet-rotating pumps until it reaches the preburner, is than ignited. It therefore puts pressure onto the turbine which starts to turn. But at the same time, the preburner also puts pressure back up the pumps and into the tanks. Because the pumps are not yet rotating. They are about to start rotating but they dont do it yet. It looks to me like a hen and a egg problem. How can you start the turbines/pumps under these conditions? Are there valves in front of the preburner that quickly close once some propellant is in the preburners and push it out the turbine only to open a fraction of a second later to allow new fuel to reach the preburner and further turn the turbine? And now my thought process looks like a moebius strip...
I would expect the combustion chamber and turbine would be wider than the pipe into  the combustion chamber and pump connected to the turbine, so the shock wave only applies a few square CM of the pressure wave back toward the tanks, but many more times the pressure foreward toward the combustion chamber.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/24/2017 04:10 am
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/24/2017 06:38 am
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.

It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: STS-200 on 11/24/2017 01:47 pm
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.

It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure?
It doesn't increase the flow rate (mass-per-second). It does increase the velocity of the propellants.
This happens because the propellant is heated in the preburner - it may simply be a gas getting hotter, or a liquid vaporising to gas. Either way, the volume increases, requiring the propellant to accelerate so mass flow rate remains the same.
Dynamic pressure - the pressure caused by he motion of the gas - rises, as the gas is moving faster. Static pressure falls to compensate, as total pressure stays constant (in an ideal device, in the real world it will always drop a bit).


Quote
Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.
Tank head start is possible, but it's a slow process that could easily give rough starts/problems with startup sequences. Venting Helium (or other start gases) through the turbine is something that can be precisely controlled, is highly predictable and spins up the turbine very quickly.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/24/2017 02:13 pm
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.

It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.

- My mistake. I should have said: Everything downstream of the PRE-BURNER has a lower total pressure. The static pressure will rise in the pre-burner as combustion products back pressure the turbine.  But, it cannot increase above the pressure upstream of the pre-burner injectors.

- The start mode I outlined is what NASA SP-125 (pg 68) calls as "main tank head start". If this type of start takes too long (> 3 seconds or so) a "turbine spin start" may be added to the system to decrease the start time. I do not know which method the Merlin uses. I would guess a "main tank head start". Does anyone know?

Also see SP-125 pg 181 for different types of starts.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/24/2017 03:02 pm
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.

It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.

- My mistake. I should have said: Everything downstream of the PRE-BURNER has a lower total pressure. The static pressure will rise in the pre-burner as combustion products back pressure the turbine.  But, it cannot increase above the pressure upstream of the pre-burner injectors.

- The start mode I outlined is what NASA SP-125 (pg 68) calls as "main tank head start". If this type of start takes too long (> 3 seconds or so) a "turbine spin start" may be added to the system to decrease the start time. I do not know which method the Merlin uses. I would guess a "main tank head start". Does anyone know?

Also see SP-125 pg 181 for different types of starts.

I need to digest all this but I think the Merlin has a face start sequence (I have no idea what that means). I remember Mueller in an interview reporting that they blew up 100 engines before they go it right.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 11/24/2017 03:53 pm
I need to digest all this ...

Semmel - I think the key question here is, ignoring the actual combustion chamber, does the powerpack run a thermodynamic cycle.

If yes, then startup would be like a jet engine's, which cannot be done by simply "lighting it up".

But I think the power pack is different. The power extracted from the exhaust is not used to pump fuel into the pack, but to pump it into the combustion chamber. I think that's why it is possible.

However, with all the phase changes that are going on, it is far from trivial, and the explanation upthread is too simplistic - I wouldn't take it literally.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cppetrie on 11/24/2017 05:11 pm
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.

It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.

- My mistake. I should have said: Everything downstream of the PRE-BURNER has a lower total pressure. The static pressure will rise in the pre-burner as combustion products back pressure the turbine.  But, it cannot increase above the pressure upstream of the pre-burner injectors.

- The start mode I outlined is what NASA SP-125 (pg 68) calls as "main tank head start". If this type of start takes too long (> 3 seconds or so) a "turbine spin start" may be added to the system to decrease the start time. I do not know which method the Merlin uses. I would guess a "main tank head start". Does anyone know?

Also see SP-125 pg 181 for different types of starts.

I need to digest all this but I think the Merlin has a face start sequence (I have no idea what that means). I remember Mueller in an interview reporting that they blew up 100 engines before they go it right.

Face shut-off.....not face start. AIUI the engine is shut down by closing the pintle completely on the pintle injector. The center of the injector carries the oxidizer and the outside ring carries RP-1.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/24/2017 07:22 pm
I need to digest all this ...

Semmel - I think the key question here is, ignoring the actual combustion chamber, does the powerpack run a thermodynamic cycle.

If yes, then startup would be like a jet engine's, which cannot be done by simply "lighting it up".

But I think the power pack is different. The power extracted from the exhaust is not used to pump fuel into the pack, but to pump it into the combustion chamber. I think that's why it is possible.

However, with all the phase changes that are going on, it is far from trivial, and the explanation upthread is too simplistic - I wouldn't take it literally.

- By powerpack I assume you mean a turbo-pump assembly with its associated gas generator. Yes, it runs a thermodynamic cycle.

- This "powerpack" can be started with only main tank head pressure and igniters, but may be slow to spool up. If this is the case a "spin turbine" may be added. The tank pressure is the initial motive force.

- The power extracted from the exhaust (along with the tank pressure ~3 atms) is used to pump fuel into the powerpack as well as the main chamber.

- The gas generator, or pre-burner, gasifies the propellants either fuel rich or oxidizer rich. This is well understood. I fail to see the problem in my start sequence?

- According to Sutton, the F1, MA-3 and SSME are all started using "tank head" starting.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 11/24/2017 08:38 pm
I need to digest all this ...

Semmel - I think the key question here is, ignoring the actual combustion chamber, does the powerpack run a thermodynamic cycle.

If yes, then startup would be like a jet engine's, which cannot be done by simply "lighting it up".

But I think the power pack is different. The power extracted from the exhaust is not used to pump fuel into the pack, but to pump it into the combustion chamber. I think that's why it is possible.

However, with all the phase changes that are going on, it is far from trivial, and the explanation upthread is too simplistic - I wouldn't take it literally.

- By powerpack I assume you mean a turbo-pump assembly with its associated gas generator. Yes, it runs a thermodynamic cycle.

- This "powerpack" can be started with only main tank head pressure and an igniters, but may be slow to spool up. If this is the case a "spin turbine" may be added. The tank pressure is the initial motive force.

- The power extracted from the exhaust (along with the tank pressure ~3 atms) is used to pump fuel into the powerpack as well as the main chamber.

- The gas generator, or pre-burner, gasifies the propellants either fuel rich or oxidizer rich. This is well understood. I fail to see the problem in my start sequence?

- According to Sutton, the F1, MA-3 and SSME are all started using "tank head" starting.

John

I'm not sure about the cycle.

In a jet engine, you have a clear "cycle", since the far field inlet and outlet conditions are sinked into the same atmosphere, and mechanical power extracted from the exhaust goes into compressing the inflow.

Here, the conditions in the far field inlet are simple the tanks (with head pressure), and the outlet goes into the combustion chamber.   Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.

If you include the combustion chamber, then far field outlet conditions are the cold hard cynical vacuum of space.

If there's no cycle, then in theory you could just "light it up", but as you say, practicalities may dictate that the spin up will be impractically slow.

Whichever way, I don't think it's an intractable problem. For all that I know, they might put an electrical motor on the shaft...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 11/24/2017 09:33 pm
For all that I know, they might put an electrical motor on the shaft...

That was the idea I had and didn't voice because of the danger of baseless speculation. Once running, the motor would generate electricity to heat up and gasify some of the propellant in the tanks to create the autogenous pressure. Safes the running of hot fuel pipes in favor of electrical cables. No idea what is lighter but it probably would safe a lot of headaches with the hot pure oxygen.

Again, total speculation on my part and probably wrong.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MegabytePhreak on 11/24/2017 09:56 pm
Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.
No, the the powerpack fuel is also pumped by the powerpack. This is certainly a requirement for raptor, since in FFSC the pressure in the preburner must be greater than chamber, and since all fuel goes through the preburners, there would be nowhere else to pump to.

For a GG like Merlin, it may not be 100% necessary theoretically , but it would suck for ISP to run the pumps on just 3 atm of pressure drop, which is all you get if you rely on tank head to push fuel into the preburner. The turbopump would also need to be physically  much larger.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/25/2017 12:26 am

I'm not sure about the cycle.

In a jet engine, you have a clear "cycle", since the far field inlet and outlet conditions are sinked into the same atmosphere, and mechanical power extracted from the exhaust goes into compressing the inflow.

Here, the conditions in the far field inlet are simple the tanks (with head pressure), and the outlet goes into the combustion chamber.   Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.

If you include the combustion chamber, then far field outlet conditions are the cold hard cynical vacuum of space.

If there's no cycle, then in theory you could just "light it up", but as you say, practicalities may dictate that the spin up will be impractically slow.

Whichever way, I don't think it's an intractable problem. For all that I know, they might put an electrical motor on the shaft...

- I'm sure of the cycles both gas generator and pre-burners.

- Before starting, the main chamber is at what ever the outside pressure is (which could be vacuum).

- Mechanical power extracted from a gas generator's or pre-burner's exhaust all goes into pumping the propellants!

- The propellants then either go to the main chamber or gas generator / pre-burner for combustion.
 In a gas generator cycle only a small portion of the propellants is burnt and it is exhausted separately from the main chamber. In a pre-burner a larger portion of the propellant is burnt and it is exhausted into the main chamber. The pre-burner obviously needs to be at pressure higher than the main chamber.

- No electric motors. Pumps for large rocket engines require 10s of thousands of horse power.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 11/25/2017 01:21 am
For all that I know, they might put an electrical motor on the shaft...

That was the idea I had and didn't voice because of the danger of baseless speculation. Once running, the motor would generate electricity to heat up and gasify some of the propellant in the tanks to create the autogenous pressure. Safes the running of hot fuel pipes in favor of electrical cables. No idea what is lighter but it probably would safe a lot of headaches with the hot pure oxygen.

Again, total speculation on my part and probably wrong.
Saying something is possible or even a good idea is not speculation...

Saying something "might be in place", or "may have happened" is.

So wrt electric drive, it'll be heavier than a gas starter, but much more reliable and controllable.

I have no doubt it was on the trade table, but who knows what they ended up with.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 11/25/2017 01:26 am

I'm not sure about the cycle.

In a jet engine, you have a clear "cycle", since the far field inlet and outlet conditions are sinked into the same atmosphere, and mechanical power extracted from the exhaust goes into compressing the inflow.

Here, the conditions in the far field inlet are simple the tanks (with head pressure), and the outlet goes into the combustion chamber.   Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.

If you include the combustion chamber, then far field outlet conditions are the cold hard cynical vacuum of space.

If there's no cycle, then in theory you could just "light it up", but as you say, practicalities may dictate that the spin up will be impractically slow.

Whichever way, I don't think it's an intractable problem. For all that I know, they might put an electrical motor on the shaft...

- I'm sure of the cycles both gas generator and pre-burners.

- Before starting, the main chamber is at what ever the outside pressure is (which could be vacuum).

- Mechanical power extracted from a gas generator's or pre-burner's exhaust all goes into pumping the propellants!

- The propellants then either go to the main chamber or gas generator / pre-burner for combustion.
 In a gas generator cycle only a small portion of the propellants is burnt and it is exhausted separately from the main chamber. In a pre-burner a larger portion of the propellant is burnt and it is exhausted into the main chamber. The pre-burner obviously needs to be at pressure higher than the main chamber.

- No electric motors. Pumps for large rocket engines require 10s of thousands of horse power.

John
Depends where you draw the boundary of the control space.

If around the powerpack, then no, power doesn't go to pump propellant into it.  It is fed by head pressure.  Power goes into pumping into the main chamber.

If around the whole motor, then yes, but then a rocket engine as a whole - does it run a thermo cycle?  I'm not sure.  It's very different from a jet engine...

Anyway, yes, electric motors would have to be huge or only act as primers of some sort...  and they would have been visible in the CAD models.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RoboGoofers on 11/25/2017 01:36 am
An electric starter/generator could be used to power electric gimbaling or electric grid fin actuation (only when the turbine is spinning, of course). I have a hard time believing they'll stick with open hydraulics for bfr/bfs
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/25/2017 12:19 pm
- I have not seen or heard of anything that indicates the use of an electric motor on either Merlin or Raptor.

- Again, the gas generator or pre-burner is fed from the output of the pumps!

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/25/2017 01:00 pm
Depends where you draw the boundary of the control space.

If around the powerpack, then no, power doesn't go to pump propellant into it.  It is fed by head pressure.  Power goes into pumping into the main chamber.

If around the whole motor, then yes, but then a rocket engine as a whole - does it run a thermo cycle?  I'm not sure.  It's very different from a jet engine...

Not that much different.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 11/25/2017 03:06 pm
Depends where you draw the boundary of the control space.

If around the powerpack, then no, power doesn't go to pump propellant into it.  It is fed by head pressure.  Power goes into pumping into the main chamber.

If around the whole motor, then yes, but then a rocket engine as a whole - does it run a thermo cycle?  I'm not sure.  It's very different from a jet engine...

Not that much different.

John

They are similar, and turbojets use electric motors to spin up the turbines to get the compressors feeding air pressure. Why couldn't a FFSC engine spin the turboshaft with a motor, just to get greater than tank head pressure in the preburner?

A key difference is that a turbojet has zero pressure differential between the inlet and the burner before spinning the compressor up, while the rocket has several atmospheres (~50 psi) of pressure pushing oxidizer into the burner. So a turbojet can't do a head start, while a SC rocket engine can.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/25/2017 04:40 pm

They are similar, and turbojets use electric motors to spin up the turbines to get the compressors feeding air pressure. Why couldn't a FFSC engine spin the turboshaft with a motor, just to get greater than tank head pressure in the preburner?

A key difference is that a turbojet has zero pressure differential between the inlet and the burner before spinning the compressor up, while the rocket has several atmospheres (~50 psi) of pressure pushing oxidizer into the burner. So a turbojet can't do a head start, while a SC rocket engine can.
- Yes, you could use an electric motor to spin up the turbine for faster starting, or you could use another high pressure gas source, or you could just use tank head pressure like the F1, MA-3 and SSME. I'm just saying that I have seen no evidence that the Raptor or Merlin use any type of spin up system. Does anyone know different?

- Yes, without a pressure difference the turbojet needs something to spin it up, but if the turbojet had high pressure at its compressor face as happens in forward flight, it can start without a starter motor.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ChaoticFlounder on 11/26/2017 12:48 am
To corroborate what JW was saying

https://blogs.nasa.gov/J2X/2013/12/

And as JW mentioned, take a look at SP-125 and SP-8107 and Rocket Propulsion Elements by George Sutton

Jet engines and LPRE's are very similar as well, see Kuznetzov Design Bureau and NK-33 / NK-15

C

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: biosehnsucht on 11/26/2017 12:58 am
They said it uses autogenous  pressurization, so use some of those gases.

I am pretty sure the tank pressure provided by autogenous pressurization system is not enough to start the spin of the turbines. If that was the case, F9 would be able to do the same with LOX and RP1 but they use high pressure helium instead. Probably a lot of it. But I am not an expert and happy to be proven wrong.

Pretty sure the issue is that RP-1 can't be autogenously pressurized the way that Methane and Oxygen can be. So you could in theory use autogenous pressurization of the LOx tank on Falcon 9 but you'd still need Helium for the RP-1 side of things...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: lrk on 11/26/2017 03:41 am
I'm just saying that I have seen no evidence that the Raptor or Merlin use any type of spin up system. Does anyone know different?

Merlin uses a spin up system driven by high-pressure helium.  The actual valves used to control the flow of fuel and LOX are spring-actuated (built into the pintle in the case of the combustion chamber, not sure about the preburner but presumably it has something similar?), so in order for fuel to even be injected the pressure must be high enough, which requires first spinning up the turbopump somehow. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 11/26/2017 12:50 pm
I'm just saying that I have seen no evidence that the Raptor or Merlin use any type of spin up system. Does anyone know different?

Merlin uses a spin up system driven by high-pressure helium.  The actual valves used to control the flow of fuel and LOX are spring-actuated (built into the pintle in the case of the combustion chamber, not sure about the preburner but presumably it has something similar?), so in order for fuel to even be injected the pressure must be high enough, which requires first spinning up the turbopump somehow.
Thank you.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 12/08/2017 03:15 am
Any new information from SpaceX on Raptor development? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 12/14/2017 09:54 am
Any new information from SpaceX on Raptor development? 
Not heard anything. Perhaps you should not expect any more new info. on Raptor dev. until IAC2018 knowing SpaceX.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mgeagon on 01/30/2018 07:35 am
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology? Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future? It seems Blue is going for a slightly more proven ORSC methalox design, and is slowly making some progress, but even that seems years ahead of any new motor on the horizon.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 01/30/2018 08:35 am
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology? Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future? It seems Blue is going for a slightly more proven ORSC methalox design, and is slowly making some progress, but even that seems years ahead of any new motor on the horizon.



ESA is in fact pursuing a Methane-fueled engine as part of the Future Launcher Preparatory Program (FLPP):

http://www.esa.int/Our_Activities/Space_Transportation/Prometheus_to_power_future_launchers (http://www.esa.int/Our_Activities/Space_Transportation/Prometheus_to_power_future_launchers)

http://spacenews.com/frances-prometheus-reusable-engine-becomes-esa-project-gets-funding-boost/ (http://spacenews.com/frances-prometheus-reusable-engine-becomes-esa-project-gets-funding-boost/)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 01/30/2018 09:51 am
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology?

Engines are not developed just to use new technology. Engines are developed because they are needed and because someone is willing to pay the development costs.

And methalox is not an optimal booster propellant. Kerolox allows lighter tanks, and for similar engine, allows better T/W. (however, FFSC may be easier with methalox than kerolox)

Methalox is a very good compromize between booster propellant and upper stage propellant, when only single propellant type is desired for both.

Quote
Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future?

Russia IS planning a methane-fueled rocket to replace most of their rockets

https://en.wikipedia.org/wiki/Soyuz-7_(rocket)

China has series of branch new efficient ORSC engines. They don't need a better engine right now. Developing a new engine would delay their new rockets by many years, and would not make them much better.

SpaceX needed a heavier engine than Merlin, and Blue Origin needed a heavier and more booster-optimized (better T/W, better T/$) engine than BE-3.
When they were anyway developing new booster engines, they decided to go to metlalox.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Cheapchips on 01/30/2018 12:41 pm

And methalox is not an optimal booster propellant. Kerolox allows lighter tanks, and for similar engine, allows better T/W.

Methalox is better than kerolox for heavy/rapid reuse as it avoids the coking issues that kerolox creates.  That must have been part of the decision by both SX & Blue to chose methane. 

As you point out, why pay the development cost if Methalox is of no benefit.  Only SX and Blue are really targeting reuse at this point.  Doesn't make a whole lot of sense for anyone else yet.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 01/30/2018 01:02 pm
It does make sense for everyone else, they're just too hide bound to realize that reuse is the new reality. No point spending billions on a new expendable launcher that effectively assumes SpaceX and Blue will fail at reuse. That's a stupid gamble at this point (and I think some are realizing it now).

Better to fly out existing expendable launchers.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Cheapchips on 01/30/2018 01:14 pm

I should have phrased that "doesn't make a whole lot of sense to anyone else".   
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TrevorMonty on 01/30/2018 02:30 pm
Methalox also allows for elimination of He for tank pressurisation. Makes for cheaper tank, plumbing and associated ground systems. Especially in SpaceX case another point of failure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 01/30/2018 03:18 pm
Methlox is also prime propellant for depots, on-orbit refueling, and long duration spaceflight.
I believe it was a propellant option for the Lunar Lander (Altair) envisioned during Constellation for many of these reasons, but the Methlox engine technology was too immature and NASA went back to RL-10s and Hydrolox.

Edit: Updated choice of Methlox option details for Lunar lander.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 01/30/2018 05:39 pm
The "problem" in the metaphorical mind of the bureaucracy, is that FFSC, Methane, Reuse, and orbital refueling are all a package- they all solve each others problems.

But changing that many things at once is anathema to Oldspace. Test ONE thing until you know it wont fail, then move to the next. Right now they're working on the ACES orbital refueling with their tried and true hydrogen expansion cycle booster, and wont touch the others till the solve, for instance, hydrogen leaking through solid metal over long duration missions.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 01/30/2018 06:12 pm
Methalox is not the end-all be-all of propellants. For SpaceX's BFR architecture it makes sense for a variety of reasons, including ISRU on Mars. For other in-space applications hydrolox may be a better option overall.

FFSC makes higher chamber pressures easier to achieve than ORSC, but at the expense of increased complexity (with all the cost and reliability impacts that entails).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 01/30/2018 06:27 pm
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage...  :o

BUT... can't make propane on mars...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 01/30/2018 06:41 pm
>
BUT... can't make propane on mars...  ;)

Sez here you can, using Fischer-Tropsch Synthesis.

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140002709.pdf
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 01/30/2018 07:02 pm
Methane is basically natural gas. I think Blue actually wants to use commodity natural gas as fuel for the BE-4. With reuse the price of the fuel becomes a significant part of the overall launch price.

I think both, SpaceX and Blue Origin chose methane among other reasons because it is the cheapest practical rocket fuel.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 01/30/2018 07:29 pm
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage...  :o

BUT... can't make propane on mars...  ;)

Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577

But methane is perfectly acceptable, especially for higher delta-v stages.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 01/30/2018 09:55 pm
>
BUT... can't make propane on mars...  ;)

Sez here you can, using Fischer-Tropsch Synthesis.

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140002709.pdf

I stand corrected... you can make it on Mars with a bit of fine tuning of the processes...  ;)

This same paper also pointed out what the EU folks had said...
(which my Google-Fu has failed to locate and link to sadly today)
Propane and Lox (both deep sub cooled) seems to make the best overall rocket system...  :)
(pages 6 and 7 in that pdf)

But yes... Methane is close enough and cheaper here on earth... likely easiest to make on Mars... 
-------------------------------------------------------------
On edit (aside note)
This other pdf I stumbled across and read some time ago... and now can't find... In a nutshell...
Was a report by some study group in the EU rocket program recommending what the future Ariane 7 or 8 system should be based on...
In short, it suggested copy the Falcon 9 systems ideas, but develop parts based on Propane and Lox...
Engines... tank sizes... all optimized to sub cooled PropLox...
They also said 9 engines on stage 1 and copy the stage one recovery like SpaceX...
Single same type engine stage 2 (like F9) but also added a 3rd stage using hypergols
Overall it was a weird dry read, and now it seems to be gone from the web...

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 01/30/2018 11:16 pm
Subcooled propane with Falcon 9 recovery IS a good architecture. They should cancel Ariane 6 and go straight to it, over-sized first stage that would allow upper stage reuse down the road.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TorenAltair on 01/30/2018 11:41 pm
@John Alan
Do you mean those studies of the German Aerospace Center?   http://elib.dlr.de/114430/2/PresentationIAC-17%20-%20D2.4.3_f.pdf
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 01/30/2018 11:46 pm
Let's maybe get closer to talking about Raptor and less Blue and ULA ???
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Space Ghost 1962 on 01/30/2018 11:46 pm
Subcooled propane with Falcon 9 recovery IS a good architecture. They should cancel Ariane 6 and go straight to it, over-sized first stage that would allow upper stage reuse down the road.
Won't work.

They need a LV in the interim. (They already painted themselves into a corner 8 years ago.) How they get it, they're committed to unfortunately.

Perhaps if they survive, they might come around to something like you suggest, although they are not at all bought into it.

(IMHO, a bit of siege mentality. Perhaps some signs of this show in various areas.) What made them slow to the game before, is the same thing that slows them to successive "game ups". Thus there's this "moving target" effect.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 01/30/2018 11:52 pm
Maybe they should start with a 1 ton RLV? That would avoid the Blue and SpaceX niche and perhaps allow them the beat RocketLab and others on price.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 01/30/2018 11:56 pm
@John Alan
Do you mean those studies of the German Aerospace Center?   http://elib.dlr.de/114430/2/PresentationIAC-17%20-%20D2.4.3_f.pdf

That's not it directly...
HOWEVER... it reads like that was written after the prior study (I can't find) was published...
And it seems to be a follow on presentation based on that prior dry wordy study...

Nice Find...  :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TorenAltair on 01/31/2018 12:05 am
Quote from: John Alan

That's not it directly...
HOWEVER... it reads like that was written after the prior study (I can't find) was published...
And it seems to be a follow on presentation based on that prior dry wordy study...

Nice Find...  :)

You'll find the whole study at http://elib.dlr.de/114430/
Actually you can find a lot of similar studies at the DLR servers. Their ftp servers never forget anything (Experiencing a Vulcain full run engine test from 200 meters away is something you won't forget as well ;) )
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 01/31/2018 12:47 am
This whole subtopic has me wondering something...  ???

Once SpaceX gets deep into Raptor[1] production and flying BFR/BFS... say 10+ years from now...  8)
And the Falcon System is winding down production and usefulness to SpaceX and the USA...

Would SpaceX offer to license the design rights and prints for the F9/FH and M1D engine to the Europeans?...  :o
On the condition they convert it to a scPropLox setup and call it something else... Ariane 9 maybe...  :D

The EU gets a better/cheaper Rocket system to serve themselves and their members...
SpaceX get a big wad of cash to invest in the BFS/BFR system build out...

Obviously... they would be the only area of the world I would let get ahold of this tech...
(On edit)... I guess I would include Japan in the list of ok to sell to also...

[1] getting back on topic Lar... I swear...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 01/31/2018 04:16 am
[1] getting back on topic Lar... I swear...  ;)
long way round the barn but ok :)

That said I can't see them licensing F9 tech and tooling.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: mgeagon on 01/31/2018 01:44 pm
That said I can't see them licensing F9 tech and tooling.
OTOH, the Russians did license their 50 year old Soyuz to the Europeans, so you never know. In fact, why not? If you have already reduced launch cost by an order of magnitude with your methalox FFSC powered RLV?

A better question is why is the ESA bothering with Soyuz or potentially Falcon 9 in the future? What the heck is a "Europeanized" Soyuz anyways? The answer lies somewhere within providing launch services without the need for any R & D fixed costs and timely availability.

In Alaska, there are still cargo airlines using DC-6s, because they continue to fill a very specific niche that newer airplanes never met. It is still amazing seeing those old birds taking off, but one is clearly struck by the inefficiency of those thirsty, smokey, noisy old radial engines.

http://www.evertsair.com/pages/aircraft/aircraft.php

Back to the Raptor, can it be said that methane is the cleanest, most easy to use at lox temperatures, easiest to find and refine propellant in the solar system (including Earth)?

Mark Eagon
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 01/31/2018 04:29 pm
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage...  :o

BUT... can't make propane on mars...  ;)

Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577

But methane is perfectly acceptable, especially for higher delta-v stages.

Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 01/31/2018 06:02 pm
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage...  :o

BUT... can't make propane on mars...  ;)

Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577

But methane is perfectly acceptable, especially for higher delta-v stages.

Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.

Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...

Purified Propane could be brought in at outside temp in pressurized tanks... Infinite shelf life...
(I'm picturing 40ft ISO container/tanks floated into port on cargo ships and trucked to the launch site)
Then run the propane thru a sub cooler system and load on the rocket at atmospheric pressure before launch...

Point was, both NASA and the Germans published papers (links up above) that indicate a scPropLox Rocket has the best mass fraction and ISP trade off of any system they have studied...  IF your starting over clean sheet, you may want to look at this, was my take away from both of them...

My 2nd point was... there is no reason I can see that the F9 basic design couldn't be reconfigured to fly with scPropLox and likely would gain some additional performance for the trouble of doing so...
SpaceX has already said BFS/BFR is our future... F9/FH will die someday as obsolete...

3rd and last point was... SpaceX could stand to make some big bucks by selling either the designs OR just sell prebuilt rocket assemblies to the EU or Japan as an additional income stream down the road...

Have to remember the big picture in the world...
Many countries maintain a space and rocket launching ability for national pride and me too standings...
They will not give that up... even when SpaceX gains a monopoly with BFS/BFR...

Meanwhile... No new news on Raptor [1]

[1] on topic...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hobbes-22 on 01/31/2018 06:41 pm
That said I can't see them licensing F9 tech and tooling.
OTOH, the Russians did license their 50 year old Soyuz to the Europeans, so you never know. In fact, why not? If you have already reduced launch cost by an order of magnitude with your methalox FFSC powered RLV?

A better question is why is the ESA bothering with Soyuz or potentially Falcon 9 in the future? What the heck is a "Europeanized" Soyuz anyways? The answer lies somewhere within providing launch services without the need for any R & D fixed costs and timely availability.


Russians didn't license Soyuz (i.e. allow ESA/Arianespace to build them). They just sell finished Soyuz rockets.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aero on 01/31/2018 07:01 pm
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"

Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?

It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Zed_Noir on 01/31/2018 07:26 pm
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"

Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?

It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?

Variable exhaust nozzle for a FFSC engine? I think that's too much cutting edge tech to implement with the Raptor. It is impressive enough that SX got a working large FFSC engine. Lets not tempt fate.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 01/31/2018 08:08 pm
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"

Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?

It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?
Reminds me of Burt Rutan not wanting a liquid cooled engine for Voyager because it weighed an extra 15 pounds on a plane where they were scrounging for ounces of weight savings. Then someone figured the extra efficiency of the engine would mean 1,000 pounds less fuel used on the trip.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Patchouli on 01/31/2018 08:30 pm

Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...

Purified Propane could be brought in at outside temp in pressurized tanks... Infinite shelf life...
(I'm picturing 40ft ISO container/tanks floated into port on cargo ships and trucked to the launch site)
Then run the propane thru a sub cooler system and load on the rocket at atmospheric pressure before launch...

Point was, both NASA and the Germans published papers (links up above) that indicate a scPropLox Rocket has the best mass fraction and ISP trade off of any system they have studied...  IF your starting over clean sheet, you may want to look at this, was my take away from both of them...

My 2nd point was... there is no reason I can see that the F9 basic design couldn't be reconfigured to fly with scPropLox and likely would gain some additional performance for the trouble of doing so...
SpaceX has already said BFS/BFR is our future... F9/FH will die someday as obsolete...

3rd and last point was... SpaceX could stand to make some big bucks by selling either the designs OR just sell prebuilt rocket assemblies to the EU or Japan as an additional income stream down the road...

Have to remember the big picture in the world...
Many countries maintain a space and rocket launching ability for national pride and me too standings...
They will not give that up... even when SpaceX gains a monopoly with BFS/BFR...

Meanwhile... No new news on Raptor [1]

[1] on topic...  ;)


Propane might make the ground systems a little easier and definitely would save on the power bill for refrigeration since the bulk propellant doesn't require any kind of active cooling unless the booster is using sub cooled propellants and then only before loading.
Plus it could be used to get a little more total delta V out of a booster since it's denser than LNG.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 01/31/2018 10:48 pm
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage...  :o

BUT... can't make propane on mars...  ;)

Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577

But methane is perfectly acceptable, especially for higher delta-v stages.

Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.

Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...

Purified Propane could be brought in at outside temp in pressurized tanks... Infinite shelf life...
(I'm picturing 40ft ISO container/tanks floated into port on cargo ships and trucked to the launch site)
Then run the propane thru a sub cooler system and load on the rocket at atmospheric pressure before launch...

Point was, both NASA and the Germans published papers (links up above) that indicate a scPropLox Rocket has the best mass fraction and ISP trade off of any system they have studied...  IF your starting over clean sheet, you may want to look at this, was my take away from both of them...

My 2nd point was... there is no reason I can see that the F9 basic design couldn't be reconfigured to fly with scPropLox and likely would gain some additional performance for the trouble of doing so...
SpaceX has already said BFS/BFR is our future... F9/FH will die someday as obsolete...

3rd and last point was... SpaceX could stand to make some big bucks by selling either the designs OR just sell prebuilt rocket assemblies to the EU or Japan as an additional income stream down the road...

Have to remember the big picture in the world...
Many countries maintain a space and rocket launching ability for national pride and me too standings...
They will not give that up... even when SpaceX gains a monopoly with BFS/BFR...

Meanwhile... No new news on Raptor [1]

[1] on topic...  ;)
Ugh, don't call it scPropLox. It's ambiguous whether you mean propane or propylene.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 01/31/2018 11:01 pm
There has been precious little discussion of Raptor in this thread lately.  Mentioning Raptor in the last sentence of your off-topic post doesn't suddenly make it relevant.  Where did I leave that trimmer...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aero on 01/31/2018 11:30 pm
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"

Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?

It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?

Variable exhaust nozzle for a FFSC engine? I think that's too much cutting edge tech to implement with the Raptor. It is impressive enough that SX got a working large FFSC engine. Lets not tempt fate.

This might not be the right thread to use, but I don't find a more appropriate one.

But my continuing thought process leads me to wonder how much different the variable exhaust nozzle engine tech would between the Raptor and the Merlin. Especially since they would know going in that the ultimate target is the Raptor. If the tech carries over from a merlin to a raptor, then SpaceX has several, even many flight-proven F9 Block 3 and 4 first stages that they could use as test platforms. Importantly, the nozzle would be tested all the way from sea level pressure to vacuum. After recovering the first stage again, they would have the variable exhaust nozzle hardware to treat just as though they had used a permanent test stand. And if the DUT (engine) blew, well, there are enough engines to (probably) bring it back. With real-world flight data. And the blown DUT didn't destroy a test stand. That would be a much more productive use of the superseded first stage rocket models than simply ditching them in the ocean or scrapping them. And if they didn't manage to recover the blown engine, it gets ditched in the ocean. No great loss except for some valuable data.

And this is tech that cries to be developed and has for 50 years.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 02/01/2018 01:07 am
They would launch these from where?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aero on 02/01/2018 02:03 am
They would launch these from where?
Flordia of course. If it is thought that untested hardware is high risk, then recall that three of the engines are restartable and they are not needed for lift-off without the mass of the fully fueled second stage atop the booster. Launch with the DUT quiescent, then start engines with modified nozzles after the booster clears the beach. After they are somewhat proven then launch normally just as the FH is doing.

The variable nozzle test program would be a sequence, starting in Texas, one or more test launches from Florida, then a demo launch, then transition to the Raptor nozzle development/test. The transition to Raptor nozzle development would be made at the optimal point in testing based on risk/reward trade-offs. Of course, Raptor testing would start again on the test stand but by the time this happens, the reward will be much better understood.

The simplest thing that occurs to me would be a computer/computer program under the cap over the interstage simulating the second stage as it interacts/controls the first stage.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 02/01/2018 02:15 am
Ok, does anyone have any updates on Raptor?  Have they tested it to a higher pressure yet? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 02/01/2018 08:33 am
Ok, does anyone have any updates on Raptor?  Have they tested it to a higher pressure yet? 
No more news on Raptor as far as I know. We will likely have to wait until IAC2018 for the next Raptor update.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 02/01/2018 11:58 am
Ok, does anyone have any updates on Raptor?  Have they tested it to a higher pressure yet? 
No more news on Raptor as far as I know. We will likely have to wait until IAC2018 for the next Raptor update.
Or a BFS update...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JoeyOak on 02/07/2018 04:29 pm
On the Methane side of Raptor...
I've always thought of it as a modified expander cycle with the preburner there to kick start it from cold and add some heat to vaporize the LNG full flow pre turbine once running...

In short... BOTH turbines will run at near room temps once going... (my opinion)  ;)

That said... the hard part of Raptor is starting it... (I think)
I'm thinking a supply of very high pressure gaseous oxygen and gaseous methane is needed to bring Raptor to life from a cold start...
700 bar room temp COPV's anyone?...  :o

All preburners (and the RCS system) share this common supply (maybe with some redundancies)

Once a Raptor is running... It can be tapped to refill such a bottle supply and keep it topped up...
(tap high pressure liquid into a small "boiler" to batch flash it into the higher pressure of the storage system)
Batch boilers may be electric heated... I'm not sure on that... 

It's all a system... thinking system and not just a rocket engine here...

Is there a commonly recognized name for this startup sequence? "Expander pressure vessel-boosted bootstrap startup"? :)

The RS-25 uses "pure" bootstrap startup, but the start sequence is slow, three to six seconds according to blogs.nasa.gov/J2X/2014/01/24/inside-the-leo-doghouse-light-my-fire/ (http://blogs.nasa.gov/J2X/2014/01/24/inside-the-leo-doghouse-light-my-fire/), which seems a bit slow for inflight and landings. Boosting the spin-up with pressurized gaseous propellant seems quite ingenious to me.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 02/07/2018 04:42 pm
On the Methane side of Raptor...
I've always thought of it as a modified expander cycle with the preburner there to kick start it from cold and add some heat to vaporize the LNG full flow pre turbine once running...

In short... BOTH turbines will run at near room temps once going... (my opinion)  ;)

That said... the hard part of Raptor is starting it... (I think)
I'm thinking a supply of very high pressure gaseous oxygen and gaseous methane is needed to bring Raptor to life from a cold start...
700 bar room temp COPV's anyone?...  :o

All preburners (and the RCS system) share this common supply (maybe with some redundancies)

Once a Raptor is running... It can be tapped to refill such a bottle supply and keep it topped up...
(tap high pressure liquid into a small "boiler" to batch flash it into the higher pressure of the storage system)
Batch boilers may be electric heated... I'm not sure on that... 

It's all a system... thinking system and not just a rocket engine here...

Is there a commonly recognized name for this startup sequence? "Expander pressure vessel-boosted bootstrap startup"? :)

The RS-25 uses "pure" bootstrap startup, but the start sequence is slow, three to six seconds according to blogs.nasa.gov/J2X/2014/01/24/inside-the-leo-doghouse-light-my-fire/ (http://blogs.nasa.gov/J2X/2014/01/24/inside-the-leo-doghouse-light-my-fire/), which seems a bit slow for inflight and landings. Boosting the spin-up with pressurized gaseous propellant seems quite ingenious to me.

It's likely SpaceX (and Tom Mueller in particular) have a name for the Raptor startup sequence... never made public to my knowledge...
Please be aware that what I posted back in November was my speculation on how Raptor worked (and started) based on what we know so far about it's design...
In short... my opinion...
It's an engine with four sources of prop to run...
Super high psi gaseous Methane and Oxygen to power the pre-burners...
Super cold ~50psi liquid LNG and LOX... to feed the main chamber via pumps...
Just my opinion... nothing more...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 02/09/2018 05:00 pm
Quote from: John Alan

That's not it directly...
HOWEVER... it reads like that was written after the prior study (I can't find) was published...
And it seems to be a follow on presentation based on that prior dry wordy study...

Nice Find...  :)

You'll find the whole study at http://elib.dlr.de/114430/
Actually you can find a lot of similar studies at the DLR servers. Their ftp servers never forget anything (Experiencing a Vulcain full run engine test from 200 meters away is something you won't forget as well ;) )

Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage...  :o

BUT... can't make propane on mars...  ;)

Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577

But methane is perfectly acceptable, especially for higher delta-v stages.

Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.

Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...

Purified Propane could

A large part that appears to be missing from the discussion, even from that german paper, is the comparative performance of the propellants as pressurants. The german paper cites a modelling tool not available for review: PMP 1.0 http://elib.dlr.de/113634/

Their conclusions on the benefits of propane versus methane could be flipped if that modeling is incorrect. They both have about the same launch vehicle mass, higher density might make upper stage reusability more difficult?

Besides the viscosity advantages of methane over propane (and cost), there is also performance as pressurization gas to consider. Methane has the handicap of having larger tank volumes to pressurize, but it has the two advantages of higher pressure and lower molecular weight.

https://i.imgur.com/fmDdyAe.png
Compiled from publicly available data, propane's curve would be two vertical dividers to the right, crossing 1 bar at around 230 kelvin. I hope to improve the spreadsheet soon. Edit: assume pressurization performance improves the further to the left (closer to helium) your substance is on this metric. Other metrics will apply as well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 02/09/2018 05:45 pm
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology? Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future? It seems Blue is going for a slightly more proven ORSC methalox design, and is slowly making some progress, but even that seems years ahead of any new motor on the horizon.

No clue about the Indians or the Japanese, but I'm pretty sure for the rest it could be due to a shortage of young engineering talent without the experience of a previous successful  engine design under their belt form the recent past. Smaller talent, experience and legacy pool to work with. That and lack of private space flight industry, they're all variants of public sector contractors and subsidised munitions factories. Taking risks isn't public sector style usually.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 02/09/2018 05:49 pm
BTW, "subscale" is in the eye of the beholder. Current Raptor would work fine in a prototype BFS doing Grasshopper-like hoops.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 02/09/2018 08:37 pm
BTW, "subscale" is in the eye of the beholder. Current Raptor would work fine in a prototype BFS doing Grasshopper-like hoops.

Agreed, EM has shown with rockets and cars that he prefers to get something done and iterate.  He said they'd have a BFS flying.  He didn't say it would have the final Raptor or be the final design (edit: of the BFS).

I could see them building something that flys and has the equivalent of a Merlin 1A engine.  Get it up, learn and iterate.

Compared to how NASA has spent $10 of billions in the last 30 years on vehicles that have never left the ground I prefer the SpaceX development method.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 02/09/2018 09:37 pm
BTW, "subscale" is in the eye of the beholder. Current Raptor would work fine in a prototype BFS doing Grasshopper-like hoops.

Agreed, EM has shown with rockets and cars that he prefers to get something done and iterate.  He said they'd have a BFS flying.  He didn't say it would have the final Raptor or be the final design (edit: of the BFS).

I could see them building something that flys and has the equivalent of a Merlin 1A engine.  Get it up, learn and iterate.

Compared to how NASA has spent $10 of billions in the last 30 years on vehicles that have never left the ground I prefer the SpaceX development method.

Let's be fair, Ares IX flew a sum total of one time ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 02/10/2018 09:01 pm
Let's be fair, Ares IX flew a sum total of one time ;)
And that was a subscale booster and lacked a second stage...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 02/10/2018 09:13 pm
Let's be fair, Ares IX flew a sum total of one time ;)
And that was a subscale booster and lacked a second stage...

HEY, it still flew successfully :D consider it a ~$40 billion test flight, of maybe 2-4, assuming SLS flies even a few times.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 02/11/2018 01:22 am
A smaller raptor would mean an even smaller initial vehicle.

Not going to happen.

They could drop in 1000 kN Raptors on the 2017 BFS and still launch with up to half a fuel load.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 02/11/2018 12:44 pm
Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 02/11/2018 03:14 pm
A smaller raptor would mean an even smaller initial vehicle.

Not going to happen.

They could drop in 1000 kN Raptors on the 2017 BFS and still launch with up to half a fuel load.

They would still need to flight qualify them. I think they will flight qualify the intended version, even if they initially fly them below full pressure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 02/11/2018 08:13 pm
Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.
There were interim versions of Merlin used on actual Falcon 9 missions, and I don't think I need to tell you that falcon 9 changed substantially during its service life.

They could use the subscale Raptor for BFR initially, but the vehicle (like Falcon 9) would have to be smaller.

no, it would not have to be smaller. It could launch with only partially fueled tanks.

When they were developing Merlin 1A, falcon 1 and falcon 9 1.0, they did not know much more thrust they will eventually get from the updated later merlin engine variants, and they really, really had to get SOMETHING flying.

So they made the craft for those engines they had at that point.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 02/13/2018 07:14 pm
Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.
There were interim versions of Merlin used on actual Falcon 9 missions, and I don't think I need to tell you that falcon 9 changed substantially during its service life.

They could use the subscale Raptor for BFR initially, but the vehicle (like Falcon 9) would have to be smaller.

no, it would not have to be smaller. It could launch with only partially fueled tanks.

When they were developing Merlin 1A, falcon 1 and falcon 9 1.0, they did not know much more thrust they will eventually get from the updated later merlin engine variants, and they really, really had to get SOMETHING flying.

So they made the craft for those engines they had at that point.

Or even just fueling the landing tanks, if the point is to test the propulsion and structure and TPS. Those look like they would hold a few hundred tonnes of methalox. Even 100 tonnes would be enough to get to the Karman line and back.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 02/18/2018 09:42 pm
Falcon 9 never flies with partially-fueled tanks.

It sort-of-did-if-you-look-at-it-right.

In that early ones had lower amounts of LOX and RP1 on liftoff.
(densification and tank stretch).

It's not quite insane to imagine that at some point there might be a reasonable trade where it was easier to leave the tanks a bit long, and later upgrade the engines.

Of course, this did not happen with F9.

http://www.parabolicarc.com/2016/12/01/year-russian-launch-failure/
Quote
Dec. 5, 2010   Proton-M/ Blok-DM-3   Uragan-M #739 Uragan-M #740
Uragan-M #741   Failure   Rocket failed to reach orbital velocity after upper stage overfilled with propellant.

Seems to imply that all tanks do not always fly full.
This post from 2013 suggests it was overfilled by 1.5 tons (https://forum.nasaspaceflight.com/index.php?topic=11127.msg1044806#msg1044806)
I could not immediately find a primary report.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 02/18/2018 10:08 pm
Moving the goalposts.
Launching with half-size tanks is not the same as launching with full-size tanks that are half-empty.

You also seem to think that SpaceX will build a BFR core with undersized engines, and then swap out those engines later on with the full-size version instead of just building a larger core for the larger engines.

Quote
It's not quite insane to imagine that at some point there might be a reasonable trade where it was easier to leave the tanks a bit long, and later upgrade the engines.
is very far from a ringing endorsement of the concept.

Vehicles have flown which have partially filled the tanks, rather than - for example - shrunk the tank for the mission.

I think they'll design the system to fly with the engines they think they are very likely to have at construction time, even if those engines are not quite ready yet.

May this in some unlikely contingencies result in an early test vehicle that flies best with underfilled tanks, sure.

It would however also not surprise me at all if the Raptor has now hit the IAC benchmarks, and is working on more.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: su27k on 02/19/2018 02:00 am
Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.
There were interim versions of Merlin used on actual Falcon 9 missions, and I don't think I need to tell you that falcon 9 changed substantially during its service life.

They could use the subscale Raptor for BFR initially, but the vehicle (like Falcon 9) would have to be smaller.

no, it would not have to be smaller. It could launch with only partially fueled tanks.

When they were developing Merlin 1A, falcon 1 and falcon 9 1.0, they did not know much more thrust they will eventually get from the updated later merlin engine variants, and they really, really had to get SOMETHING flying.

So they made the craft for those engines they had at that point.
Falcon 9 never flies with partially-fueled tanks.

The first stage flew 15 times with partial fuel load under Grasshopper and F9R-Dev1 program, this is what the original comments are about: SpaceX can test full scale BFS by flying sub-orbital trajectory using underpowered Raptors and partial fuel load, which fits well with what Elon Musk suggested (testing BFS like Grasshopper).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 02/19/2018 02:17 am
Grasshopper and F9R-dev1 flew with partially filled tanks. For Grasshopper-like tests of BFS, BFS will probably be part filled.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Chris Bergin on 02/19/2018 02:57 am
A reminder to you new people (and long time people). If you aren't civil, you lose your post. Simple as that. Don't like that, go to Facebook. No one is going to be uncivil here. Small trim.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 02/19/2018 02:57 am
The Grasshopper test vehicle used a significantly lower thrust (development) variant of Merlin 1D than is in use today, and even the current Raptor is only somewhat lower thrust than the planned operational Raptor. This isn’t “weaseling,” it’s factual.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: FinalFrontier on 02/19/2018 03:21 am
The Grasshopper test vehicle used a significantly lower thrust (development) variant of Merlin 1D than is in use today, and even the current Raptor is only somewhat lower thrust than the planned operational Raptor. This isn’t “weaseling,” it’s factual.
SpaceX has supposedly done a fair amount of engine firings on the current raptor already, so there is probably a decent data set showing what the current engine is capable of. I do not believe this is publicly available at this time, however it would be interesting to know where we are in thrust and specific impulse right now vs the current plan for 'operational' raptor.

We do know the current chamber pressure is lower, Elon mentioned what they achieved so far but I forget where he talked about this.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 02/19/2018 03:46 am
SpaceX has supposedly done a fair amount of engine firings on the current raptor already, so there is probably a decent data set showing what the current engine is capable of. I do not believe this is publicly available at this time, however it would be interesting to know where we are in thrust and specific impulse right now vs the current plan for 'operational' raptor.

We do know the current chamber pressure is lower, Elon mentioned what they achieved so far but I forget where he talked about this.

IAC.
Quote
The test engine currently operates at 200 atmospheres, 200 bar, the flight engine will be at 250 bar, and then we believe over time we could probably get that to a little over 300 bar.

It would not surprise me to learn they've hit 250 already, but who knows.
There is a big gap between 'it works on the test stand' and 'we are leaping in with both feet at the highest pressure we can justify'.

The margin, if other things don't go wrong is very large (in space terms) for BFS/R two stage.

Post FH conference -
Quote
I think we might, if we get lucky, be able to do short hop flights with the spaceship part of BFR maybe next year.

Maybe we'll get more on Raptor next year.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: FinalFrontier on 02/19/2018 04:14 am
SpaceX has supposedly done a fair amount of engine firings on the current raptor already, so there is probably a decent data set showing what the current engine is capable of. I do not believe this is publicly available at this time, however it would be interesting to know where we are in thrust and specific impulse right now vs the current plan for 'operational' raptor.

We do know the current chamber pressure is lower, Elon mentioned what they achieved so far but I forget where he talked about this.

IAC.
Quote
The test engine currently operates at 200 atmospheres, 200 bar, the flight engine will be at 250 bar, and then we believe over time we could probably get that to a little over 300 bar.

It would not surprise me to learn they've hit 250 already, but who knows.
There is a big gap between 'it works on the test stand' and 'we are leaping in with both feet at the highest pressure we can justify'.

The margin, if other things don't go wrong is very large (in space terms) for BFS/R two stage.

Post FH conference -
Quote
I think we might, if we get lucky, be able to do short hop flights with the spaceship part of BFR maybe next year.

Maybe we'll get more on Raptor next year.

I just went back and found all the IAC stuff/other details beat me to the edit button.
In any case, supposedly the maximum total accumulated testing time is 1200 seconds across 42 test firings with the longest being around 100 seconds limited by the tankage at the stand at the time. This of course is all assumed to be at the 200 bar mark for the test engine or engines.

At the very least 250 bar should be achievable, if this is correct so far, by sometime this year but it's hard to say if they already got there or not from the outside looking in. 250 is what is required for the flight engine so once they have tested reliably at this pressure the flight engine is more or less ready. I do not see any reason why they would bother putting the 200 bar test stand variant on the BFShopper test vehicle since they seem to be pretty close to the first iteration flight engine.

Also, it's probably fair to assume raptor may have various derivations through its life as the design is improved similar to the merlin engine, this was pointed out earlier and is a good assumption imho.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DaveH62 on 02/19/2018 11:09 pm
If they used the sub scale raptor in an upper stage now, does anyone have a good idea of the additional mass enabled for geostationary orbit? I assume low earth orbit would have less gain, but could be substantial if used as US for the falcon heavy.
Would a Raptor US also be easier to start testing for reuse?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 02/19/2018 11:13 pm
If they used the sub scale raptor in an upper stage now, does anyone have a good idea of the additional mass enabled for geostationary orbit? I assume low earth orbit would have less gain, but could be substantial if used as US for the falcon heavy.
Would a Raptor US also be easier to start testing for reuse?

 Raptor Upper Stage consolidated thread   (https://forum.nasaspaceflight.com/index.php?topic=42861.msg1674190#msg1674190)

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 03/01/2018 04:36 pm
Any news on Raptor testing above sub-scale?  Seems like Raptor has to be finished first before BFR/BFS can really get started. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 03/01/2018 08:01 pm
Any news on Raptor testing above sub-scale?  Seems like Raptor has to be finished first before BFR/BFS can really get started. 
No news as far as I know. SpaceX are being very secretive regarding Raptor dev. with updates only during the annual IAC events.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/02/2018 01:40 am
Any news on Raptor testing above sub-scale?  Seems like Raptor has to be finished first before BFR/BFS can really get started.
I think they can do BFS tests with subscale Raptor just fine (while full scale is being finished up). Don't know if they would bother to do so, though.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 03/02/2018 11:31 am
We are only talking about roughly a 15% linear scale up. I think they will be full size. Why invest in creating multiple engines of a size you don't plan on developing fully? Ultimately, a waste of time and manpower.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 03/02/2018 02:29 pm
We are only talking about roughly a 15% linear scale up. I think they will be full size. Why invest in creating multiple engines of a size you don't plan on developing fully? Ultimately, a waste of time and manpower.

John
My theory is that the BFR/BFS specs were in flux, so to avoid blocking engine dev on BFR dev they told the propulsion guys to just pick a reasonable number and go with it.  While they were doing engine dev they settled on a 15% larger number in the BFR dev process, but to avoid the distraction of a continually moving target the raptor guys are going to keep going with the subscale engine to the end of their dev cycle (whatever that involves), instead of pivoting instantly to match whatever # the BFR guys prefer this week.

They'll plan the work in a reasonable manner to balance the disruption of regular pivots with the potential for effort wasted developing the "wrong size" engine.

Tom Mueller proved with the Merlin he can handle a long-timeframe engine dev process that continually squeezes out more performance from a base engine.  I'm sure they have a similar long-range plan worked out for Raptor as well.  I don't know exactly when that plan calls for a physical hardware scale-up, but I'm sure it will happen at an appropriate time.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 03/02/2018 11:10 pm
That makes no sense.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RotoSequence on 03/02/2018 11:55 pm
That makes no sense.

It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 03/03/2018 04:53 am
That makes no sense.
CScott's?  Makes perfect sense, exactly as he described it.

-----
ABCD: Always Be Counting Down
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 03/03/2018 05:03 am
That makes no sense.

It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
They are exploring the unknown.  During engine development there's some uncertainty in exactly how much thrust a given physical scale will be able to provide, and during BFR development there's uncertainty in how much the rocket will need to weigh.  You have to untie the knot somehow, and you can't throw away your work and start from scratch every time any component diverges from its initial estimates. We only get snapshots of progress once a year, and at that point the Raptor on the test stand was underperforming the BFR design specs by 15%. (Or the BFR was 15% overweight.) Who knows, Tom might squeeze out more performance from the subscale Raptor (or the BFR might slim down) and the final version will only be (say) a 10% scale up.

It does assume that they got the initial estimate close enough that the final scale up won't be too different from the initial version.  They're not talking about a half-scale or quarter-scale raptor here.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 03/03/2018 01:33 pm
That makes no sense.

It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.

Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 03/03/2018 02:01 pm
Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?

I don't think 'scale down' is the right word.
BFS/R was scaled down to fit market needs.
This meant that for engine out and packing reasons, you pretty much need that size of engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Darkseraph on 03/03/2018 02:02 pm
SpaceX recieved quite a bit of money of the air force to develop raptor as an upperstage engine for F9/FH. The current scale of the subscale version could have been set by this requirement. They may have well have scaled down the planned raptor for BFR to be close to the development version to save money. Which is really smart if so. 2016 ITS was essenially being pitched to NASA for funding and they didn't bite. The scale had to come down with less funding. The lionshare of government funds for BFR will likely come from the next iteration of the U.S Air Force's EELV program. At least that program has no equivalent of the SLS. When are the proposals for that due again?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 03/03/2018 02:05 pm
There is some misunderstanding here. The sea level thrust scale up is about 70%. The throat, combustion chamber and pumps is about 15% linear scale increase. Throat area scale is 1.15^2 = 1.32. Chamber pressure goes from 2000 psi to 2500 psi for an increased pressure ratio of 1.25. This would require about a 12% increase in pump speed. The thrust scale up due to the increased pressure and throat area is 1.25 x 1.32 = 1.65. An increased expansion ratio from about 26 to 35 gets you to 1.7 SL thrust scale ratio.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 03/03/2018 02:07 pm
Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?

I don't think 'scale down' is the right word.
BFS/R was scaled down to fit market needs.
This meant that for engine out and packing reasons, you pretty much need that size of engine.

I read it as "scaled down", but not because of engine size, or else they'd have ended up at the development engine size.

They first chose a development engine size long before detail design of the ship was done.

They later chose a ship size which makes sense as a first generation. 12 m was larger than necessary and introduced difficulties that they chose to avoid.

Larger ships will follow, but all in good time.

The chosen ship size dictated the flight engines will be larger than development engines by said 15%.

----
ABCD: Always Be Counting Down
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 03/03/2018 02:09 pm
That makes no sense.

It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.

Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?

I think they did.  See the size comparison on my chart.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: su27k on 03/03/2018 02:12 pm
The sub-scale engine size is determined by the Stennis test stand size, it's in the article started this thread: https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/

Quote
While incapable of handling the full size of the expected Raptor engine unit, the Stennis test stand enabled the individual testing of each subcomponent of the 1MN scaled prototype that SpaceX currently has at its test facility in McGregor, Texas.

Quote
Since the final thrust level of the Raptor had not been settled, it was decided that the first integrated test engine would be a 1MN sub-scale engine.

It enabled the full testing at Stennis E2 and allowed for the development of robust startup and shutdown sequences, characterize hardware durability and anchor analytical models that would be used for future designs.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 03/09/2018 09:55 pm
That makes no sense.

It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.

Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?

I think they did.  See the size comparison on my chart.

John

Correlation is not causation. The vehicle architecture is not being designed to match the engine, but the expected economic of the entire operation. Looking it from the opposite direction is the path to being mislead. Size of vehicle is not determined by one parameter (difficulty of engine scaling) but by a hundred.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: OneSpeed on 03/10/2018 11:07 am
Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?

I think they did.  See the size comparison on my chart.

John

Correlation is not causation. The vehicle architecture is not being designed to match the engine, but the expected economic of the entire operation. Looking it from the opposite direction is the path to being mislead. Size of vehicle is not determined by one parameter (difficulty of engine scaling) but by a hundred.

The 2017 BFR is fundamentally a 3/4 scale model of the 2016 ITS.

DimensionRatio
Linear0.75
Area0.5625
Mass0.421875

Thrust is proportional to area, so if you scale the ITS engines by 0.75, the thrust is scaled by 0.5625.
The ITS SL engines were 3000 kN, and the BFR SL engines will be 1700 kN.
3000 * 0.5625 = 1687 kN, pretty close to 1700.

The ITS liftoff mass was 10500 mT, so multiply by 0.421875 for a 0.75 scale model.
10500 * 0.421875 = 4429 mT, also pretty close to 4400 mT.

42 * 3000 kN engines would have given ITS a T/W ratio of 42 * 3000 / 10500 * 9.8 = 1.22.
For BFR, the number of engines required for the same T/W ratio would be 4400 * 1.22 * 9.8 / 1700 = 30.95, very close to 31.

I suspect the correlation is no accident ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 03/10/2018 01:21 pm
That makes no sense.

It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.

Didn't they scale down the size of the production engines to be very close to the engine they already had under test?  Or did I just dream that 2017 IAC presentation?

I think they did.  See the size comparison on my chart.

John

Correlation is not causation. The vehicle architecture is not being designed to match the engine, but the expected economic of the entire operation. Looking it from the opposite direction is the path to being mislead. Size of vehicle is not determined by one parameter (difficulty of engine scaling) but by a hundred.

Didn't mean to imply it was the only reason, just one of many.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 03/17/2018 07:34 pm
Is it possible that some of the pre-burner( or main injectors) structure is copper like many other engines. and some local impingement occurs during start and shutdown?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 03/23/2018 12:44 pm
Is it possible that some of the pre-burner( or main injectors) structure is copper like many other engines. and some local impingement occurs during start and shutdown?

Don't know about the pre-burners, but the main chamber is a a copper alloy. Probably GRCo-84 or similar.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Navier–Stokes on 03/26/2018 03:59 pm
Quote from: https://twitter.com/jeff_foust/status/978295808679464968
Jeff Foust‏ @jeff_foust 

If you squint at this chart, you can see ongoing and planned test activity at Stennis by Aerojet Rocketdyne, Relativity, Stratolaunch and SpaceX, among others.

8:41 AM - 26 Mar 2018
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 03/26/2018 06:58 pm
Quote from: https://twitter.com/jeff_foust/status/978295808679464968
Jeff Foust‏ @jeff_foust 

If you squint at this chart, you can see ongoing and planned test activity at Stennis by Aerojet Rocketdyne, Relativity, Stratolaunch and SpaceX, among others.

8:41 AM - 26 Mar 2018
What is the "SpaceX combustion device"?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 03/26/2018 07:06 pm
Quote from: https://twitter.com/jeff_foust/status/978295808679464968
Jeff Foust‏ @jeff_foust 

If you squint at this chart, you can see ongoing and planned test activity at Stennis by Aerojet Rocketdyne, Relativity, Stratolaunch and SpaceX, among others.

8:41 AM - 26 Mar 2018
What is the "SpaceX combustion device"?

Most likely full scale preburners, or subscale injector testing for Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 12:23 am
I thought this was more interesting. What's "Mars Lander"?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 03/27/2018 12:38 am
I thought this was more interesting. What's "Mars Lander"?

Whatever it is, it's not likely to have anything to do with Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 12:50 am
I thought this was more interesting. What's "Mars Lander"?

Whatever it is, it's not likely to have anything to do with Raptor.
Why the heck not?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 01:05 am
The first flight of Falcon 9 had a failed restart of the Vacuum Merlin because roll control froze up in vacuum, a problem they probably could've caught in a test like this. At the time (correct me if I'm wrong), this facility had been basically mothballed, thus they would've had to spend like hundreds of millions to restart it or something vs the ~$40 million price of another Falcon 9 launch which they may have had to do anyway. So at the time, it was a good trade, even especially in retrospect.

Apparently now the facility is being used by upper stages again, which means SpaceX won't have to pay the full cost of un-mothballing. Also, a BFR and BFS cost about an order of magnitude more than a Falcon 9 v1.0, and SpaceX would have to wait another 26 months to try again, thus pushing back their crewed flight another synod.

Pretty sure it makes sense to test for integrated-Raptor landing failure modes in Mars-like conditions.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 03/27/2018 01:17 am
The first flight of Falcon 9 had a failed restart of the Vacuum Merlin because roll control froze up in vacuum, a problem they probably could've caught in a test like this. At the time (correct me if I'm wrong), this facility had been basically mothballed, thus they would've had to spend like hundreds of millions to restart it or something vs the ~$40 million price of another Falcon 9 launch which they may have had to do anyway. So at the time, it was a good trade, even especially in retrospect.

Apparently now the facility is being used by upper stages again, which means SpaceX won't have to pay the full cost of un-mothballing. Also, a BFR and BFS cost about an order of magnitude more than a Falcon 9 v1.0, and SpaceX would have to wait another 26 months to try again, thus pushing back their crewed flight another synod.

Pretty sure it makes sense to test for integrated-Raptor landing failure modes in Mars-like conditions.

That facility won't support a single Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 01:26 am
The first flight of Falcon 9 had a failed restart of the Vacuum Merlin because roll control froze up in vacuum, a problem they probably could've caught in a test like this. At the time (correct me if I'm wrong), this facility had been basically mothballed, thus they would've had to spend like hundreds of millions to restart it or something vs the ~$40 million price of another Falcon 9 launch which they may have had to do anyway. So at the time, it was a good trade, even especially in retrospect.

Apparently now the facility is being used by upper stages again, which means SpaceX won't have to pay the full cost of un-mothballing. Also, a BFR and BFS cost about an order of magnitude more than a Falcon 9 v1.0, and SpaceX would have to wait another 26 months to try again, thus pushing back their crewed flight another synod.

Pretty sure it makes sense to test for integrated-Raptor landing failure modes in Mars-like conditions.

That facility won't support a single Raptor.
Based on what, exactly?

And keep in mind: The facility CAN be upgraded, and the slide directly implies that it will.

The facility can handle up to 400,000lbf, which is roughly the same as a vacuum Raptor: https://www1.grc.nasa.gov/facilities/isp/

...and, like Merlin, Raptor will almost certainly start life out at lower thrust. (and the initial Mars landing will involve throttling down the engines)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 03/27/2018 02:39 am
Adjusting the fiscal year quarters to actual calendar dates, doesn't that test window close at the end of March 2020? Additionally, doesn't the Mars 2020 rover have a NET of July 2020?

3-4 months to button up and ship it?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 03:28 am
Adjusting the fiscal year quarters to actual calendar dates, doesn't that test window close at the end of March 2020? Additionally, doesn't the Mars 2020 rover have a NET of July 2020?

3-4 months to button up and ship it?
That would be the Mars rover, not the "Mars Lander." The 2020 rover uses the Skycrane stage, just like MSL, therefore there is no "lander" part.

The only US "Mars Lander" I can think of, then, is BFS.

...there is the European/Russian lander, but that is, as you say, a 2020 mission (July) so there wouldn't be enough time to test it there in 2020.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 03/27/2018 04:56 am
https://cosmosmagazine.com/space/nasa-prepares-mars-missions
Quote
To reach Mars, a spacecraft must travel through the cold vacuum of space for nine months. These extreme conditions are recreated at NASA’s Glenn Research Centre in Cleveland. In Vacuum Chamber 5, powerful pumps suck air out of the 4.5 metre tall space. Even a spacecraft’s thrusters can be tested here – it’s designed to handle the heat and gas they generate. Panels cooled to -262°C line the walls, chilling the thruster’s exhaust. Pumps lining the bottom of the tank concentrate and recycle the precious xenon gas that the thrusters blast out as propellant.


The timeline is odd for Mars 2020 though.

I find references to wheels and nuclear work being done for rovers at Glenn.
Perhaps final installation or preparation for the RTG would make more sense than thrusters or BFS/raptor tests.

With planned schedule, there will be very little need for Glenns vacuum test capability, as actual (possibly suborbital) space works just fine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 12:15 pm
That’s a different vacuum chamber. ISPF is a lot bigger.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/27/2018 11:06 pm
...probably not BFS.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Torbjorn Larsson, OM on 03/29/2018 11:44 am
https://cosmosmagazine.com/space/nasa-prepares-mars-missions
Quote
To reach Mars, a spacecraft must travel through the cold vacuum of space for nine months. These extreme conditions are recreated at NASA’s Glenn Research Centre in Cleveland. In Vacuum Chamber 5, powerful pumps suck air out of the 4.5 metre tall space. Even a spacecraft’s thrusters can be tested here – it’s designed to handle the heat and gas they generate. Panels cooled to -262°C line the walls, chilling the thruster’s exhaust. Pumps lining the bottom of the tank concentrate and recycle the precious xenon gas that the thrusters blast out as propellant.


The timeline is odd for Mars 2020 though.

I find references to wheels and nuclear work being done for rovers at Glenn.
Perhaps final installation or preparation for the RTG would make more sense than thrusters or BFS/raptor tests.

With planned schedule, there will be very little need for Glenns vacuum test capability, as actual (possibly suborbital) space works just fine.

The "recycle the precious xenon gas" part indicates the chamber is dedicated for ion engine test. RTGs would be one way of providing the electric power needed.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 03/29/2018 09:07 pm
https://cosmosmagazine.com/space/nasa-prepares-mars-missions
Quote
To reach Mars, a spacecraft must travel through the cold vacuum of space for nine months. These extreme conditions are recreated at NASA’s Glenn Research Centre in Cleveland. In Vacuum Chamber 5, powerful pumps suck air out of the 4.5 metre tall space. Even a spacecraft’s thrusters can be tested here – it’s designed to handle the heat and gas they generate. Panels cooled to -262°C line the walls, chilling the thruster’s exhaust. Pumps lining the bottom of the tank concentrate and recycle the precious xenon gas that the thrusters blast out as propellant.


The timeline is odd for Mars 2020 though.

I find references to wheels and nuclear work being done for rovers at Glenn.
Perhaps final installation or preparation for the RTG would make more sense than thrusters or BFS/raptor tests.

With planned schedule, there will be very little need for Glenns vacuum test capability, as actual (possibly suborbital) space works just fine.

The "recycle the precious xenon gas" part indicates the chamber is dedicated for ion engine test. RTGs would be one way of providing the electric power needed.

So would Kilopower, potentially at  higher power levels.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 03/29/2018 11:39 pm
So would Kilopower, potentially at  higher power levels.
The RTG I mentioned was in the context of wondering if the Mars 2020 lander RTG was going to be installed there.
Neither Kilopower or RTGs help meaningfully for a craft the size of BFS around earth-mars orbits, plus it is not unlikely that either would cost a substantial fraction of a whole BFS.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 03/30/2018 03:29 am
I'm pretty sure it's not SpaceX for the "Mars Lander." My speculation was incorrect.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 04/07/2018 02:48 pm
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?

I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 04/07/2018 02:54 pm
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?

I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.

Are the turbopumps and combustion chamber different between the sea level and vac versions of the Merlin?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 04/07/2018 03:28 pm
Are the turbopumps and combustion chamber different between the sea level and vac versions of the Merlin?

I don't know. It is just whenever the theme comes up it is maintained that they are very different. Usually in the context of someone asking if SL Merlins could be converted to vac Merlin, the answer is emphatically "No, they are too different". The obvious difference is that the outlet of the gas generator is channeled into the nozzle extension to provide a cooling film. That may cause differences in pressure balances that force changes all the way.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 04/07/2018 03:42 pm
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?

I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.

Are the turbopumps and combustion chamber different between the sea level and vac versions of the Merlin?

The combustion chamber is integral to part of the nozzle (immediately downstream from the throat), which is definitely different on MVac. The turbopumps are probably slightly different as well, though I don't know specifics.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: groundbound on 04/10/2018 01:30 am
I'm not sure this belongs here or in the BFR thread:

Can we infer by the arrival of parts of BFS tooling that Raptor passed a SpaceX internal milestone a few months ago?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/10/2018 01:48 am
I'm not sure this belongs here or in the BFR thread:

Can we infer by the arrival of parts of BFS tooling that Raptor passed a SpaceX internal milestone a few months ago?
No. I think the timing is due to finally getting the permission to use that piece of port land. Raptor has been pretty far along in testing for a while, now. There's not a ton of uncertainty there.

I'd say it has a lot more to do with maturity of the BFR/BFS design than Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 04/10/2018 11:44 am
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?

I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.

- The combustion chamber / nozzle initial nozzle angles are different for different expansion ratios. A higher expansion ratio require a higher initial turning angle.

- I don't know of reasons why the pumps or throat area would be different, and many reasons for them being the same.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 04/24/2018 11:34 pm
Was it previously announced that a larger Raptor was under development?
I could easily have missed it...

Quote
Speaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.
https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/

Since this is six-month old news, we should be seeing something on the stand yet this year...  The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 04/25/2018 12:15 am
Was it previously announced that a larger Raptor was under development?
I could easily have missed it...

Quote
Speaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.
https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/

Since this is six-month old news, we should be seeing something on the stand yet this year...  The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.

It is well know that the engine that have been testing is subscale development version of Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vaporcobra on 04/25/2018 12:26 am
Was it previously announced that a larger Raptor was under development?
I could easily have missed it...

Quote
Speaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.
https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/

Since this is six-month old news, we should be seeing something on the stand yet this year...  The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.

It would also be pretty damn hard to tell that the "larger" Raptor had been swapped in, too. If livingjw is more or less correct, the final-ish 1700kN would require a physical scale-up of maybe 10-20%. That's deeeeeep in the weeds for the photos we have at hand (both L2 and now public).

Attached my favorite of the Raptor bays.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 04/25/2018 12:29 am
Was it previously announced that a larger Raptor was under development?
I could easily have missed it...

Quote
Speaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.
https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/

Since this is six-month old news, we should be seeing something on the stand yet this year...  The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.

It would also be pretty damn hard to tell that the "larger" Raptor had been swapped in, too. If livingjw is more or less correct, the final-ish 1700kN would require a physical scale-up of maybe 10-20%. That's deeeeeep in the weeds for the photos we have at hand (both L2 and now public).

Attached my favorite of the Raptor bays.
Yup. Makes sense to scale up before you have production tooling made, but honestly, they probably could get away with the dev engine just fine if they had to. It's barely any smaller.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 04/25/2018 06:56 pm
Agree that the 'smaller' engine would be fine for early BFS testing. 
Maybe we'll hear about full scale engine NLT IAC 2018.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/09/2018 12:48 pm
Making the Raptor engine 10-20% larger is not to get it to produce the 1700kN of thrust, so it meets the specs of IAC 2017.
The “non-scalled” engine which will be revealed at IAC2018, probably has more thrust than 1700 kN since TED talk animations show the BFR has become higher since IAC2017.

The exact specs don’t dictate the design. The specs follow out of the optimal design.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 05/09/2018 01:03 pm
Making the Raptor engine 10-20% larger is not to get it to produce the 1700kN of thrust, so it meets the specs of IAC 2017.
The “non-scalled” engine which will be revealed at IAC2018, probably has more thrust than 1700 kN since TED talk animations show the BFR has become higher since IAC2017.

The exact specs don’t dictate the design. The specs follow out of the optimal design.

Falcon development history shows that 'optimal design' is achieved after many iterations (i.e., the end point), not some starting point.  BFR/BFS will most likely be exactly this -- many iterations -- since there is so much innovation required.  The last rocket with this amount of innovation was Saturn V; there was much iteration leading up to that 'optimal design.'
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/09/2018 02:40 pm
Exactly! If the Merlin engine increased its thrust by design optimization over the years, so will the Raptor.
A full scale Raptor with a spec of 1700kN is therefor not etched in stone.

The Falcon 9 diameter was determined by road transport limitations and the Merlin engine upper size was determined by engine lay-out so that 9 could fit in nicely.
When Merlin got more thrust they made the rocket higher, which increased payload capability also.

If we see BFR 2018 being higher and Raptor 2018 having more thrust, history repeats itself, just faster
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 05/09/2018 08:48 pm
Today, Josh Brost stated that Raptor had been test fired for "several thousand seconds."  Not really new news or by any means unexpected, but it is good to hear confirmation that progress is being made.

Video at 3:06.

https://livestream.com/viewnow/HumanstoMars2018/videos/174579723
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 05/19/2018 02:18 pm
Attached are two pictures of the Raptor test stand taken 1 year apart. The first one taken in early 2017  shows some evidence of operation can be seen in the gravel apron and close in in the "grass". The other picture taken a year later shows evidence of much heaver erosion on the Gravel apron and "grass damage out about 1000 feet.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: testguy on 05/19/2018 07:09 pm
Attached are two pictures of the Raptor test stand taken 1 year apart. The first one taken in early 2017  shows some evidence of operation can be seen in the gravel apron and close in in the "grass". The other picture taken a year later shows evidence of much heaver erosion on the Gravel apron and "grass damage out about 1000 feet.

First photo may just show that the only testing up to that point was only cold flow tests.  BTW great photos.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 05/19/2018 08:21 pm
Attached are two pictures of the Raptor test stand taken 1 year apart. The first one taken in early 2017  shows some evidence of operation can be seen in the gravel apron and close in in the "grass". The other picture taken a year later shows evidence of much heaver erosion on the Gravel apron and "grass damage out about 1000 feet.

First photo may just show that the only testing up to that point was only cold flow tests.  BTW great photos.

The IAC 2016 (September 2016) featured a Raptor hot firing from that stand, so there was at least one power test, but likely many more by early 2017 (first image).  Suspect the difference is early low power tests vs later full power for that engine.  42 tests between Sep 2016 and Sep 2017, so second image had at least that many, plus reports that tests were continuing past IAC 2017.

By the way, some engine, probably derived from that 'sub-scale' design, will be flying next year on BFS.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 05/20/2018 07:42 am
The last rocket with this amount of innovation was Saturn V; there was much iteration leading up to that 'optimal design.'

Keep in mind that engineers had already optimized far beyond that point when the program was cancelled. F-1A was already done and certified @ 1.8M lb thrust. Designs for new blocks of Saturn V with extended tanks had been completed. Solid and liquid boosters for the S1 had been designed. A perusal of Encyclopedia Astronautica (http://www.astronautix.com) in this area is fascinating.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 05/20/2018 01:42 pm
The last rocket with this amount of innovation was Saturn V; there was much iteration leading up to that 'optimal design.'

Keep in mind that engineers had already optimized far beyond that point when the program was cancelled. F-1A was already done and certified @ 1.8M lb thrust. Designs for new blocks of Saturn V with extended tanks had been completed. Solid and liquid boosters for the S1 had been designed. A perusal of Encyclopedia Astronautica (http://www.astronautix.com) in this area is fascinating.

One could say that FH and BFR are 'optimizations' beyond Falcon 9 Block 5.  The counter-example to this iterative development is Shuttle -- it began with a great burst of innovation, and then stagnated/died before optimization.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 05/24/2018 07:59 pm
https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 05/25/2018 06:55 am
I guess this is new...

Quote
"I don’t want to say too much. We’re building up the test stand right now. We’ve got the first flight version of that engine in work. We’ve been running the development engine quite a bit. It’s running great," Mueller told the audience.

From GeekWire (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DigitalMan on 05/25/2018 12:19 pm
Given their aggressiveness, I expect this version will be extremely lean and mean.

edit: I hope this is an indication they are on track to potentially do BFS tests on or near schedule.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: sleepy-martian on 05/25/2018 03:57 pm
I guess this is new...

Quote
"I don’t want to say too much. We’re building up the test stand right now. We’ve got the first flight version of that engine in work. We’ve been running the development engine quite a bit. It’s running great," Mueller told the audience.

From GeekWire (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/)
Would the current test stand be destroyed by a full thrust raptor?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/25/2018 04:59 pm
https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018

Ive been waiting for this number, for a while now, since I know it will be insanely high.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 05/25/2018 05:10 pm
I guess this is new...

Quote
"I don’t want to say too much. We’re building up the test stand right now. We’ve got the first flight version of that engine in work. We’ve been running the development engine quite a bit. It’s running great," Mueller told the audience.

From GeekWire (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/)

It’s a little hard to tell what “we’ve got the first flight version of that engine in work” means. Is it being built? Is it still being designed?

If they’re still building the test stand, it looks like full scale raptor testing is a ways off. Will the BFS hopper use the sub scale development engines?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 05/25/2018 05:37 pm
https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018

Ive been waiting for this number, for a while now, since I know it will be insanely high.
It almost doesn't matter.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)

1800 tons of thrust for BFS would mean eleven tons of engines at this figure. Halving it to 300:1 only drops five tons.
5 seconds more ISP would make up that limit.

Not that more margin isn't better of course.
But if doubling weight lets you make a simple trade for reliability - it would make sense.

(it is regrettably unlikely to be this simple)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 05/25/2018 06:36 pm
It almost doesn't matter.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)
SL Merlins are now at 185:1 T/W at SL and close to 200:1 in a vacuum.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/27/2018 10:28 pm
https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018

Ive been waiting for this number, for a while now, since I know it will be insanely high.
It almost doesn't matter.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)

1800 tons of thrust for BFS would mean eleven tons of engines at this figure. Halving it to 300:1 only drops five tons.
5 seconds more ISP would make up that limit.

Not that more margin isn't better of course.
But if doubling weight lets you make a simple trade for reliability - it would make sense.

(it is regrettably unlikely to be this simple)


High T/W of the Merlin engine was important for making the Falcon 9 rocket reusable

https://youtu.be/PK0kTcJFnVk

@27.20


Its the Final Mass that determines Delta V in the Tsiolkovsky Rocket equation

https://en.m.wikipedia.org/wiki/Tsiolkovsky_rocket_equation


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 05/28/2018 12:13 am

Its the Final Mass that determines Delta V in the Tsiolkovsky Rocket equation

https://en.m.wikipedia.org/wiki/Tsiolkovsky_rocket_equation
Yes, it is.
And five tons is basically irrelevant in a vehicle with 85 tons total weight, with 150 tons payload, and 20 tons or so landing fuel reserve.

As I said - five seconds more ISP would make it a worthwhile trade. (I did the rocket equation calculation)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/28/2018 06:28 am
I see your logic, but It’s not only the 5 tons of the BFS engines than needs to be substracted from the 150 tons payload to LEO. Its also the extra weight of the 31 booster engines, and the extra weight of
extra landing fuel for the heavier booster that (partly) needs to be subtracted.
It’s quite a difficult calculation actually.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 05/28/2018 10:10 am
https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018

Ive been waiting for this number, for a while now, since I know it will be insanely high.
It almost doesn't matter.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)

1800 tons of thrust for BFS would mean eleven tons of engines at this figure. Halving it to 300:1 only drops five tons.
5 seconds more ISP would make up that limit.

Not that more margin isn't better of course.
But if doubling weight lets you make a simple trade for reliability - it would make sense.

(it is regrettably unlikely to be this simple)

While reliability is important, the company shows continued preference for high thrust and T/W ratio due to overall vehicle mass ratio benefits over ISP gains. They'll take ISP gains where they can get it, and it can't be neglected, but they obviously prefer thrust over specific impulse. Otherwise they would not be pushing to match and exceed the T/W ratio of the already staggering almost 200:1 rato of Merlin.

11 tonnes for 6 or 7 Raptors is also way too much, it's nearly quadruple the weight of Merlins. Thrust increase is only double, and chamber pressure increase is only triple. While weight may well scale worse than linear, Raptor also makes use of plumbing geometry optimizations, the mass growth from Merlin to Raptor should be nowhere near fourfold.

And we have official confirmation that the T/W ratio is not going to be worse than Merlin. Debating overweight Raptors is no longer meaningful speculation outside of alternate history.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 05/28/2018 10:14 am
I see your logic, but It’s not only the 5 tons of the BFS engines than needs to be substracted from the 150 tons payload to LEO. Its also the extra weight of the 31 booster engines, and the extra weight of
extra landing fuel for the heavier booster that (partly) needs to be subtracted.
It’s quite a difficult calculation actually.
It is actually not.
Yes, there are more engines on the booster, but they have limited impact due to the low staging velocity and high mass of required fuel to do retroburns and landing, as well as the fact they are on a first stage meaning approximately only 10% mass penalty is typical, rather than around 100% for second stages.

Landing fuel is in the range of 20 tons for BFS, landing empty, landing fuel with 5 tons more mass adds somewhat less than one ton.

If you add five tons of engine mass on S2, and 20 tons on S1, you reduce payload by around 6 tons, if you are returning empty.

Having said that, I don't believe this is a trade they're looking at - it's just that while it is an advance, and five tons (or whatever) extra payload is nice, it does not make a meaningful difference to if the vehicle can work or not for Mars if the engines weigh as much as, or half as much as Merlin per unit thrust.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 05/29/2018 01:23 am
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 05/29/2018 01:39 am
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.

To be sure. However, that doesn't necessarily make a meaningful difference in whether the vehicle can perform the Mars mission.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/29/2018 05:39 am
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.

To be sure. However, that doesn't necessarily make a meaningful difference in whether the vehicle can perform the Mars mission.


Elon talks in the video about the reusability penalty on payload in, “percentage of mass to orbit” going from 4% to 2%.
And that until 2012 he didnt think it was possible to have a reusable mars mision.

Just see how that litlle landing fuel in Falcon 9 propagates in payload capability penalty vs Falcon 9 in expendable mode. suppose Falcon 9 had engines which where half the mass with the same thrust, the booster would need much less landing fuel.

Whats the percentage of mass to orbit of BFR vs Falcon 9 and what improved this? My guess is T/W of Raptor and Carbon fiber.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 05/29/2018 10:23 am
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.
To reiterate - I did a more-or-less reasonable calculation assuming mars-entry like delta-v for earth return of BFS (700m/s, which ties in neatly with the capacity of the shown tanks), return velocity of BFR being akin to F9S1, ... counting the structure, published ISP, ...

And came up with only a little over 5 tons, for an increase of 5 tons in S2 engines, and a proportionate one in S1 engines.

As a very obvious thought experiment, if you have 5 tons of extra engines on S2, and 20 tons of extra engines on S1, that absolutely can't be worse than removing 20 tons from your payload capability of 150 tons up and return.

It is measurable, but not significant.

it is only significant if payloads already in the 130-150(*) ton range are already designed and can't be changed, or if the change in price of 15% makes or breaks the business cost.

* (not actually 130 ton or 5% more like 144 and 5%).

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 05/29/2018 10:41 pm
Anner J. Bonilla
@annerajb
How's raptor testing going?

Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.

https://twitter.com/elonmusk/status/1001565360783474688
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 05/30/2018 12:09 am
Anner J. Bonilla
@annerajb
How's raptor testing going?

Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.

https://twitter.com/elonmusk/status/1001565360783474688

Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor?  I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 05/30/2018 12:44 am
Mid 200's I could maybe believe. once fully developed out (8 years+)...
350:1... Good Grief, what alien stuff is it made out of...  :o
I have a real hard time believing they can reach over 300:1 with 2 axis steering system included...  :P
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 05/30/2018 04:53 am
Anner J. Bonilla
@annerajb
How's raptor testing going?

Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.

https://twitter.com/elonmusk/status/1001565360783474688

Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor?  I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?

Where do you get the "easily take the thrust to weight crown" from?? We know Mueller implies that it will be better than Merlin, but the easily part seems like you are getting overly excited and reading something into it which is not based in fact.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: biosehnsucht on 05/30/2018 05:16 am


Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor?  I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?

Are you sure you're not remembering numbers thrown around for chamber pressure (300 bar sounds familiar) ?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/30/2018 05:26 am
Anner J. Bonilla
@annerajb
How's raptor testing going?

Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.

https://twitter.com/elonmusk/status/1001565360783474688

Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor?  I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?

I was the one who predicted at least 350 based on the cad drawings of the design and chamber pressure.
Made a bet for $50 dollars on it, in this forum.

Another way to approximate T/W of the Raptor engines is to calculate the % mass to orbit of the whole rocket.

Theoretical T/W maximum of a rocket engine can become very high, the smaller you make your engines while retaining high Isp. Practical limit of T/W of methane burning engine would be around 10.000.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 05/30/2018 05:43 am
Anner J. Bonilla
@annerajb
How's raptor testing going?

Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.

https://twitter.com/elonmusk/status/1001565360783474688

Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor?  I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?

I was the one who predicted at least 350 based on the cad drawings of the design and chamber pressure.
Made a bet for $50 dollars on it, in this forum.

Another way to approximate T/W of the Raptor engines is to calculate the % mass to orbit of the whole rocket.

Theoretical T/W maximum of a rocket engine can become very high, the smaller you make your engines while retaining high Isp. Practical limit of T/W of methane burning engine would be around 10.000.

Yes, but that is just your speculation. Not even in the same galaxy as a verified fact. I'm trying to clear out the fog of confusion about what we know and do not know. Be careful about how you phrase things, because someone already took that 350 as a "known fact".

Remember you already lost a bet about the number of engines on the BFR.  :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/30/2018 11:34 am
Last year I speculated based on historic data trends and a little physics that the Raptor would have a T/W of somewhere between 160 - 190 and that 220 might even be possible. This includes pumps, pre-burners, plumbing and actuators. I see no evidence that would indicate a thrust to weight of 350 or higher.  10,000??? Where did that come from?

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Doesitfloat on 05/30/2018 02:36 pm
I remember reading of micro thrusters with T/W in that range. They were tiny.  Made by removing material from thin sheet of steel.  They made a 2-D engine with two propellant inputs combined combustion chamber and vac nozzle.
If I remember it was all theoretical as the  thruster had no ignition, cooling, and had to be supplied with pressurized gas.
So not a main engine.
I do however think the idea of smaller high pressure engines is correct.  With smaller engines they are able to make fault tolerant systems; instead of the expensive fault proof current practice. I also agree that a smaller engine can run at higher pressure.  I did some quick FEA guesses at the Merlin, Raptor, and Rutherford ( Electron) engines.
The Guesses boiled down to thin material was better at heat flux making smaller engines stronger and able to handle higher pressure/mass flow.
I realize this goes against historical trends, but so does 9 engines on a boost stage.  They have had plenty of time to change that...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 05/30/2018 05:14 pm
I saw these also, including valves turbo pumps and sensors
With t/w of 1000. (See pic)


Interesting Quote from Elon Musk:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"


http://spacenews.com/elon-musks-ask-me-anything-qa-just-the-space-parts/

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 05/30/2018 05:35 pm
I saw these also, including valves turbo pumps and sensors
With t/w of 1000. (See pic)


Interesting Quote from Elon Musk:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"


http://spacenews.com/elon-musks-ask-me-anything-qa-just-the-space-parts/

Sigh. You are linking two things that have no connection. Stop conflating "this is theoretically possible based on paper X" with "this is what SpaceX will do".
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 05/30/2018 05:41 pm
The theoretical performance of micro-engines really doesn't have anything to do with this thread.  If you want to speculate about alternative architectures for a propulsion system there is probably a better thread elsewhere on the forum.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 05/30/2018 06:14 pm
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x 4-5x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.


Edit: There's no need for kick pumps if the full fuel flow goes through the regen. There are if they don't. There probably is more stages to the "main pumps" than just one due to limited head rise per stage/pump.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 05/30/2018 06:24 pm
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.

Both Musk and Mueller have made it pretty clear that they expect Raptor to have higher TWR than Merlin 1D.

There are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 05/30/2018 06:52 pm
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.

Both Musk and Mueller have made it pretty clear that they expect Raptor to have higher TWR than Merlin 1D.

There are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.

Relevant Quote from https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 05/30/2018 07:35 pm

Interesting Quote from Elon Musk:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"


 I was talking to Baldusi about that.
 There are lots of factors to how heavy you make various engine parts, but the simplest rule to start with is the square/cube thing. If you make an engine twice as big in all dimensions, it weighs 8 times as much but only has 4 times the thrust.
 Of course, it's way more complicated since the thickness of various components won't always exactly double, but it's a good basis for understanding why lots of smaller engines can be better than a few big ones.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 05/30/2018 07:54 pm
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.

Both Musk and Mueller have made it pretty clear that they expect Raptor to have higher TWR than Merlin 1D.

Yes, at some point, but not necessary initially. The thrust of the first version may be lower than the thrust of the final version which will beat merlin 1D in T/W.

Quote
There are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.

... but instead of liquid medium-pressure fuel pipes there are fuel-rich higher-pressure gas pipes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 05/30/2018 08:55 pm
Quote
There are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.

... but instead of liquid medium-pressure fuel pipes there are fuel-rich higher-pressure gas pipes.

The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DistantTemple on 05/30/2018 10:47 pm
Unfortunately, though relevant in general EM's comment you quote is from and AMA in 2015. Since that is before a lot of development, including the clip of raptor in the 2016 IAC, and the change to BFS in 2017, etc and more recent comments from EM and Tom Mueller, its effectively ancient history. Other posters seem not to have noticed, and seem to be reacting to it as if its current news! And thus confusing the picture.

edit: 5th Jan 2015... 3 1/2 years ago.

snip...
Interesting Quote from Elon Musk:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"

http://spacenews.com/elon-musks-ask-me-anything-qa-just-the-space-parts/
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 01:23 am

Interesting Quote from Elon Musk:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"


 I was talking to Baldusi about that.
 There are lots of factors to how heavy you make various engine parts, but the simplest rule to start with is the square/cube thing. If you make an engine twice as big in all dimensions, it weighs 8 times as much but only has 4 times the thrust.
 Of course, it's way more complicated since the thickness of various components won't always exactly double, but it's a good basis for understanding why lots of smaller engines can be better than a few big ones.

The physics of rocket weight is primarily the physics of pressure vessels. Pressure vessel mass is proportional to volume and pressure: Mass = c*vol*press where c is a constant which depends on the shape and material properties. The mass of main combustion chamber, pre-burners, plumbing including turbo-pump bodies all are primarily stressed by pressure. Larger engines require less cooling due to the square-cube scaling difference. Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines.  All the physics points to less performance and less thrust to weight for smaller engines. Having said that, performance and thrust to weight are relatively flat from 100,000 - 2,000,000 lbf thrust range.
 
   The performance and thrust to weight of extremely small rocket engines as shown in an earlier post will suffer significant Reynolds number induced losses. Turbo-pumps of that size will have very poor efficiency. The mass of flat slab sides compared to cylindrical pressure vessels of the same volume and pressure are much higher as is the wetted area that has to be cooled. Do we have a link to the paper? Their numbers are hard to fathom and would like to read the original paper.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 02:09 am
OK, I skimmed the paper:

http://www.las.inpe.br/~jrsenna/AerospaceMEMS/Propulsao/S&Aav2997p1-7.pdf

They are weighing only the 1.2 gram main combustion chamber. No manifolds or plumbing. No pumps. No ignition. No nothing accept the chamber. The manifold it was mounted in for testing appears to be many times that weight. The above paper only tested it to 10% of its design pressure, so most of their conclusions are 10x extrapolations, though they may have done testing at higher pressures not reported in this paper. No attempt was made to predict turbo-pump efficiency, but I suspect they would be better of with a piston type pump for this size rocket because they are less effected by low Reynolds numbers. I see nothing in this paper relevant to large reusable rockets.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 05/31/2018 03:20 am
...Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines.  All the physics points to less performance and less thrust to weight for smaller engines...
...this isn't quite true. Particularly if you look at the nozzle. The nozzle mass scales as dimension cubed, but thrust only scales as dimension squared. There are other, more subtle examples. But if you look at nozzle mass, clusters of small engines win handily over one big nozzle (past a certain size). They're much shorter, too, for the same expansion ratio.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 05/31/2018 08:26 am
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg


I have seen this image many times but only now looked at it in detail. Some things make me wonder..
* Where is the fuel turbine? I can sort of see a connection of the preburner to the injector face but there should be a turbine somewhere in between. Cant see it.
* The fuel regen out splits to the oxygen preburner and into the fuel preburner. But at the bottom. Looks strange to me, how can they inject the fuel into the preburner at a different position than the main fuel flow?
* I assume the fuel regen in has to be connected to the exit of the fuel pumps. I cant find that connection.
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.

I love the Raptor design. The fact that they mounted the oxygen powerpack directly ontop of the injector seems revolutionary in oxygen rich staged combustion. I havent seen that trick before. Might be the primary reason why that engine is so compact and might achieve a high TWR. The RD-191 (http://www.russianspaceweb.com/images/rockets/angara/rd191/infograph_1.jpg) might be the closest (active) relative (even though its not FFSC) and is a monster in comparison. Probably this sort of trick is only possible in FFSC engines. But even the RD-270 (http://www.b14643.de/Spacerockets/Specials/Russian_Rocket_engines/RD-270.jpg) has lots of thick pipes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 11:54 am
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg


.... Some things make me wonder..
* Where is the fuel turbine? I can sort of see a connection of the preburner to the injector face but there should be a turbine somewhere in between. Cant see it.
* The fuel regen out splits to the oxygen preburner and into the fuel preburner. But at the bottom. Looks strange to me, how can they inject the fuel into the preburner at a different position than the main fuel flow?
* I assume the fuel regen in has to be connected to the exit of the fuel pumps. I cant find that connection.
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.

- The fuel turbine is in the upper part of the fuel preburner. The flow is upward towards the fuel pump. The exhaust does a 180 degree turn downward and exits the horizontal pipe to the MCC.
- The regen. fuel exits the MCC near the injector face and part way down the expansion nozzle. They both appear to be routed to the bottom of the fuel pre-burner. A smaller pipe branches off just before one of the regen. fuel pipe enters the pre-burner and heads to a valve on its way to the Lox pre-burner.
- Fuel out of the pump must be hidden behind the pump. I am assuming it goes straight to the manifold at the throat. From the throat, fuel is directed upward towards the main injectors and downward to the expansion nozzle.
- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.

The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 12:13 pm
...Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines.  All the physics points to less performance and less thrust to weight for smaller engines...
...this isn't quite true. Particularly if you look at the nozzle. The nozzle mass scales as dimension cubed, but thrust only scales as dimension squared. There are other, more subtle examples. But if you look at nozzle mass, clusters of small engines win handily over one big nozzle (past a certain size). They're much shorter, too, for the same expansion ratio.

- The mass of most of the expansion portion of the nozzle is relatively low pressure. Its mass, to a large degree is determined by regenerative cooling, which is proportional to dimension squared.
- Clusters of small engines win handily? What is the physics of this? Do you know of a paper or tests comparing one large regenerativly cooled LRE MCC versus many smaller one?

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 05/31/2018 12:56 pm
...Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines.  All the physics points to less performance and less thrust to weight for smaller engines...
...this isn't quite true. Particularly if you look at the nozzle. The nozzle mass scales as dimension cubed, but thrust only scales as dimension squared. There are other, more subtle examples. But if you look at nozzle mass, clusters of small engines win handily over one big nozzle (past a certain size). They're much shorter, too, for the same expansion ratio.

- The mass of most of the expansion portion of the nozzle is relatively low pressure. Its mass, to a large degree is determined by regenerative cooling, which is proportional to dimension squared.
- Clusters of small engines win handily? What is the physics of this? Do you know of a paper or tests comparing one large regenerativly cooled LRE MCC versus many smaller one?

John
Aware of no “published” study, although surely some exists. This comes from scaling laws (I’ve analyzed this in detail) and is borne out in historical example. Compare the bottom of N1 with the bottom of Saturn V.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 05/31/2018 01:15 pm
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg


.... Some things make me wonder..
* Where is the fuel turbine? I can sort of see a connection of the preburner to the injector face but there should be a turbine somewhere in between. Cant see it.
* The fuel regen out splits to the oxygen preburner and into the fuel preburner. But at the bottom. Looks strange to me, how can they inject the fuel into the preburner at a different position than the main fuel flow?
* I assume the fuel regen in has to be connected to the exit of the fuel pumps. I cant find that connection.
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.

- The fuel turbine is in the upper part of the fuel preburner. The flow is upward towards the fuel pump. The exhaust does a 180 degree turn downward and exits the horizontal pipe to the MCC.
- The regen. fuel exits the MCC near the injector face and part way down the expansion nozzle. They both appear to be routed to the bottom of the fuel pre-burner. A smaller pipe branches off just before one of the regen. fuel pipe enters the pre-burner and heads to a valve on its way to the Lox pre-burner.
- Fuel out of the pump must be hidden behind the pump. I am assuming it goes straight to the manifold at the throat. From the throat, fuel is directed upward towards the main injectors and downward to the expansion nozzle.
- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.

The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.

John

Interesting, does this mean the entire fuel flow goes through the regen system? Because I cant imagine how else they would come up with a downward pointing pump system and an upwards pointing pre-burner system. Or is there an other reason that the flow does a 180 turn?

The 1400 F (760 C) you mention is the fuel pre-burner exhaust temperature? What do you think is the LOX pre-burner exhaust temperature?

And I undertsand that the device that is connected to both, the "LOX to Fuel Pre-burner" and the "Fuel to LOX Pre-burner" is where the fuel flow is controlled. It controls how much combustion happens in the pre-burners, hence how strong the tourbine goes, hence how strong the pumps pump and thus throttles the engine (Its probably much more complicated than that). But what does the pipe to the right of that do? On second looks, it doesnt tab into the fuel regen out pipe. Is it a LOX overpressure relieve, that can dump access LOX to the side of the nozzle? It probably goes beyond the point it is drawn there if that was the case.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 04:41 pm
This drawing was posted a couple of years ago by one of the guys. I think it is pretty close to the internal layout.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 04:51 pm
Raptor's fuel pre-burner and turbo pump is probably similar to the SSME fuel pump. Turn it upside down and remove a pump stage and at turbine stage.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: kraisee on 05/31/2018 06:27 pm
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg

At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.

My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.

Ross.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: kraisee on 05/31/2018 06:34 pm
I love the Raptor design. The fact that they mounted the oxygen powerpack directly ontop of the injector seems revolutionary in oxygen rich staged combustion. I havent seen that trick before.

It somewhat reminds me a bit of the German P111 engine, which stacked the turbine(s) directly above, and exhausted straight into the MCC.

Whatever the inspiration, SpX clearly looked around at all the previous design alternatives, considered ideas that most other companies would consider 'out of the box' and picked a fairly aggressive, and very efficient, solution for this particular engine. They're clearly still innovating very strongly.

I wonder who actually first proposed with this particular approach? Tom? Elon? Will we ever find out? :)

Ross.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 05/31/2018 08:44 pm
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg

At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.

My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.


Ross.

Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F. Also, it doesn't look like the MCC is using a pintle injector. Wrong shape. Looks more like a planar array of coaxial injectors, of course its hard to tell.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 05/31/2018 11:11 pm
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg



At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.

My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.


Ross.

Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F. Also, it doesn't look like the MCC is using a pintle injector. Wrong shape. Looks more like a planar array of coaxial injectors, of course its hard to tell.

John

I think the RD-170 injector design gives a good idea on what the injectors will look like. The differences of course will be due to FF and liquid methane vs RP1 and full "gas" phase injection in the MCC. and the use of partitions  in the MCC may be unnecessary.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 06/01/2018 02:14 pm


I think the RD-170 injector design gives a good idea on what the injectors will look like. The differences of course will be due to FF and liquid methane vs RP1 and full "gas" phase injection in the MCC. and the use of partitions  in the MCC may be unnecessary.

Could you care to explain your thought process on the injector geometry choice and why you think the "partitions" wouldn't need it?

edit: I ask because your statements are contrary to what I believe. For the preburners, the RD 170 preburner injector style geometry might well be used, but I don't agree with the main injector.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Doesitfloat on 06/01/2018 03:04 pm
When Spacex was testing their pintle injector at Stennis,  that  made sense since both the Kestrel and all versions of the Merlin used the face shutoff pintle for the MCC. 
While we have no info from the Raptor team, on their choice of injector. I would be surprised if they went in a different direction than what they had built with the Merlin engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 06/01/2018 03:15 pm
When Spacex was testing their pintle injector at Stennis,  that  made sense since both the Kestrel and all versions of the Merlin used the face shutoff pintle for the MCC. 
While we have no info from the Raptor team, on their choice of injector. I would be surprised if they went in a different direction than what they had built with the Merlin engines.

That caption in the picture is wrong. It's a oxygen rich preburner hot fire in Stennis. It's not a cold flow test, and they wouldn't even go to Stennis to do a cold flow test. The statement about the pintle injector being used is unsubstantiated.


edit: also physics drives what kind of injectors get used. One does not simply go with one style of injector just because they used it before for a totally different purpose.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/01/2018 03:38 pm


I think the RD-170 injector design gives a good idea on what the injectors will look like. The differences of course will be due to FF and liquid methane vs RP1 and full "gas" phase injection in the MCC. and the use of partitions  in the MCC may be unnecessary.

With gas phase injection most of the cyclic excitation caused by the liquid interaction in the burn area will be absent.

Could you care to explain your thought process on the injector geometry choice and why you think the "partitions" wouldn't need it?

edit: I ask because your statements are contrary to what I believe. For the preburners, the RD 170 preburner injector style geometry might well be used, but I don't agree with the main injector.

I would agree as detail shown is RP1 liquid injecting outside the Hot oxidizer
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 06/01/2018 06:32 pm

Interesting Quote from Elon Musk:

"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"


 I was talking to Baldusi about that.
 There are lots of factors to how heavy you make various engine parts, but the simplest rule to start with is the square/cube thing. If you make an engine twice as big in all dimensions, it weighs 8 times as much but only has 4 times the thrust.
 Of course, it's way more complicated since the thickness of various components won't always exactly double, but it's a good basis for understanding why lots of smaller engines can be better than a few big ones.

The physics of rocket weight is primarily the physics of pressure vessels. Pressure vessel mass is proportional to volume and pressure: Mass = c*vol*press where c is a constant which depends on the shape and material properties. The mass of main combustion chamber, pre-burners, plumbing including turbo-pump bodies all are primarily stressed by pressure. Larger engines require less cooling due to the square-cube scaling difference. Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines.  All the physics points to less performance and less thrust to weight for smaller engines. Having said that, performance and thrust to weight are relatively flat from 100,000 - 2,000,000 lbf thrust range.
 
   The performance and thrust to weight of extremely small rocket engines as shown in an earlier post will suffer significant Reynolds number induced losses. Turbo-pumps of that size will have very poor efficiency. The mass of flat slab sides compared to cylindrical pressure vessels of the same volume and pressure are much higher as is the wetted area that has to be cooled. Do we have a link to the paper? Their numbers are hard to fathom and would like to read the original paper.

John
Goes to show that SpaceX sized Raptor to minimize dev. costs of BFR/BFS system by having just one engine design throughout the system due to budget limitations. If EM had the money of JB then I bet SpaceX would have dev. a larger Raptor for BFR with a smaller version for BFS. Dev. costs do not equate to op. costs.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 06/01/2018 06:41 pm
Goes to show that SpaceX sized Raptor to minimize dev. costs of BFR/BFS system by having just one engine design throughout the system due to budget limitations. If EM had the money of JB then I bet SpaceX would have dev. a larger Raptor for BFR with a smaller version for BFS. Dev. costs do not equate to op. costs.

Just goes to show that having no money constraints is often counterproductive.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 06/01/2018 07:13 pm
Multiple things here.

1. Raptor has a gas/gas injector combustion, merlin a liquid/liquid. That difference alone tells you that no history on development can be used as viable guidance.

2. Raptor is optimized for thrust to weight and engine out capability on the second stage. Larger engines are not automatically preferable. No matter the development cost.

3. I love this thread. Thanks for this lively and informative discussion! Especially thanks to living!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DistantTemple on 06/01/2018 08:43 pm
This comment
Quote from: Interesting Quote from Elon Musk
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"

Is from 5th Jan 2015... 3 1/2 years ago and there has been progress since then.

http://spacenews.com/elon-musks-ask-me-anything-qa-just-the-space-parts/
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 06/01/2018 08:56 pm
Could the progress be, that thrust to weight optimized further to an even lower thrust of 170 metric tons? 8)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DistantTemple on 06/01/2018 10:28 pm
Could the progress be, that thrust to weight optimized further to an even lower thrust of 170 metric tons? 8)

Ah yes I see. Lesson to self: check data before posting! 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 06/02/2018 02:49 am
I remembered reading (in this thread?) a few years back, that Raptor would not use a pintle injector, but a shower head injector. I can not remember whether that statement was based on actual information from SpaceX or speculation. Time goes by and my brain can only retain so much. The shower head portion stuck, the rest did not :(
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/02/2018 05:42 pm
I don't think SpaceX has ever said what the Raptor injector types are. The CAD rendering make me think the main injectors are coaxial type in a shower head arrangement. No clues, that I know of, concerning the pre-burners.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: john smith 19 on 06/02/2018 08:07 pm
OK, I skimmed the paper:

http://www.las.inpe.br/~jrsenna/AerospaceMEMS/Propulsao/S&Aav2997p1-7.pdf

They are weighing only the 1.2 gram main combustion chamber. No manifolds or plumbing. No pumps. No ignition. No nothing accept the chamber. The manifold it was mounted in for testing appears to be many times that weight. The above paper only tested it to 10% of its design pressure, so most of their conclusions are 10x extrapolations, though they may have done testing at higher pressures not reported in this paper. No attempt was made to predict turbo-pump efficiency, but I suspect they would be better of with a piston type pump for this size rocket because they are less effected by low Reynolds numbers. I see nothing in this paper relevant to large reusable rockets.

John
AFAIK they never actually got this plan to work. The numbers looked impressive (100g thrust from a 1.2g structure that's 83.3:1. Not bad for V 0.1 tech) but as you noted they didn't include the pumps. AFAIK they never did. A thesis on why they failed indicate pumping losses were a big part of it and the structure could not take the pressure.

I'd agree, piston based pumps at this scale would be a good idea.  Silicon is attractive if you need semiconductors but it's not that good an idea at the mesoscale.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 06/03/2018 03:12 am
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg

At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.

My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.


Ross.

Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F. Also, it doesn't look like the MCC is using a pintle injector. Wrong shape. Looks more like a planar array of coaxial injectors, of course its hard to tell.

John

Exactly. Also the CH4 won't go directly from regen to the MCC, it first will be partially burned in the preburner and sent through the turbine with much higher temperature and volume. It might flash from supercritical to gas in the preburner, but definitely won't in the MCC.

Injection to the MCC will definitely be gas-gas which I think makes a pintle highly unlikely.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 06/04/2018 08:01 am
https://i.imgur.com/ld7z2Fn.jpg


.... Some things make me wonder..
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.


- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.

The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.

John

It does look substantial but I don't think the thick 'pipe' connecting "Lox to fuel preburner" to the lower "Fuel Regen out" is part of the flow system. 
It looks to be part of the sensing and control system used to control reagent supply to the three reaction chambers (preburners and main combustion chamber).

Seems the reagent supply valve controls take delta pressure(?) measurements using these (right to left):

> Heated methane pump output pressure before reaction chamber against cold oxygen pump output pressure before same reaction chamber. 
> Cold oxygen pump output pressure [same pipe] against main combustion chamber hot oxygen* injection manifold pressure. 
> Heated methane pump output pressure against either ??cold oxygen?? or ??cold methane?? pump inlet pressure

*that's actually oxygen rich gas mix between oxidizer preburner-turbine before main combustion chamber, not pure unreacted propellant like the other sensor inputs appear to be.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 06/05/2018 07:05 am
https://i.imgur.com/ld7z2Fn.jpg


.... Some things make me wonder..
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.


- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.

The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.

John

It does look substantial but I don't think the thick 'pipe' connecting "Lox to fuel preburner" to the lower "Fuel Regen out" is part of the flow system. 
It looks to be part of the sensing and control system used to control reagent supply to the three reaction chambers (preburners and main combustion chamber).

Seems the reagent supply valve controls take delta pressure(?) measurements using these (right to left):

> Heated methane pump output pressure before reaction chamber against cold oxygen pump output pressure before same reaction chamber. 
> Cold oxygen pump output pressure [same pipe] against main combustion chamber hot oxygen* injection manifold pressure. 
> Heated methane pump output pressure against either ??cold oxygen?? or ??cold methane?? pump inlet pressure

*that's actually oxygen rich gas mix between oxidizer preburner-turbine before main combustion chamber, not pure unreacted propellant like the other sensor inputs appear to be.

Ok, that makes sense. So the setting on this differential pressure measurement would control the LOX/LCH4 mixture ratio.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 06/05/2018 07:17 am
Ok, that makes sense. So the setting on this differential pressure measurement would control the LOX/LCH4 mixture ratio.

To be accurate, these measurements could help control Three LOX/LCH4 mixture ratios in three reaction chambers: both of the turbine-preburners as well as the main combustion chamber.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 06/05/2018 07:43 am
The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F.

Please see the two quotes above. These are estimates of livingjw about the two pre-burner exit temperatures. Maybe before going on, I should ask.. how do you get to these? To an untrained person like me, these temperatures seem quite high. I have no experience with these and my thermodynamics is long ago (and not particularly great to begin with), so please be patient with me for the following.

The pre-burner burns LOX/LCH4 in just the right amount to run the turbine that drives the pumps and to provide the pressure AFTER the turbine that gets injected into the main combustion chamber. The pressure has to be higher than the pressure in the combustion chamber, so thats some considerable energy.
Here is the part that I am not comfortable with and could be completely wrong: My understanding is that the efficiency of the engine depends on the expansion ratio of the propellents, which depends on the temperature differential of the propellant going in the combustion chamber and the leaving temperature. So basically you want the incomming propellant as cold as possible and the outgoing propellant as hot as possible. Can you set this right please? I have the feeling there is some error in my understanding.

I one could compute the temperatures backwards. Starting with the chamber pressure and thrust of Raptor. One could compute the pressure after the turbines and the mass flow rate. This would inform about how strong the pumps would have to work and how much chemical energy has to be released inside the pre-burners, which would inform about the exit temperature and the mixture ratio inside the pre-burners. Maybe there is a better way of doing that?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/05/2018 01:31 pm
A major source  of energy is the just the expansion of the Lox to gas. The O2 side Turbine inlet temps  (in this thread) estimated anywhere from  "room temp" to 1000 F  (including calculations ).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 06/05/2018 01:33 pm
The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F.

Please see the two quotes above. These are estimates of livingjw about the two pre-burner exit temperatures. Maybe before going on, I should ask.. how do you get to these? To an untrained person like me, these temperatures seem quite high. I have no experience with these and my thermodynamics is long ago (and not particularly great to begin with), so please be patient with me for the following.

The pre-burner burns LOX/LCH4 in just the right amount to run the turbine that drives the pumps and to provide the pressure AFTER the turbine that gets injected into the main combustion chamber. The pressure has to be higher than the pressure in the combustion chamber, so thats some considerable energy.
Here is the part that I am not comfortable with and could be completely wrong: My understanding is that the efficiency of the engine depends on the expansion ratio of the propellents, which depends on the temperature differential of the propellant going in the combustion chamber and the leaving temperature. So basically you want the incomming propellant as cold as possible and the outgoing propellant as hot as possible. Can you set this right please? I have the feeling there is some error in my understanding.

I one could compute the temperatures backwards. Starting with the chamber pressure and thrust of Raptor. One could compute the pressure after the turbines and the mass flow rate. This would inform about how strong the pumps would have to work and how much chemical energy has to be released inside the pre-burners, which would inform about the exit temperature and the mixture ratio inside the pre-burners. Maybe there is a better way of doing that?

There aren't many metals that can survive above 1400F. Personally I think the fuel rich outlet will be more around 1100F-1200F.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/05/2018 02:53 pm
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 06/05/2018 03:26 pm
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.

How common are single crystal alloy casting shops? Rene N5 and N6 are single crystal alloys. I presumed they were primarily only done in house at big OEMs like P&W and Rolls.

SpaceX is probably going to be forging or casting big portions of the turbine and nozzle with more conventional alloys like Inconel 718, Rene 41/Haynes R41, Haynes 188, 230, 282, ect.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 06/05/2018 03:32 pm
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.

How common are single crystal alloy casting shops? Rene N5 and N6 are single crystal alloys. I presumed they were primarily only done in house at big OEMs like P&W and Rolls.

SpaceX is probably going to be forging or casting big portions of the turbine and nozzle with more conventional alloys like Inconel 718, Rene 41/Haynes R41, Haynes 188, 230, 282, ect.

Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)

Quote
Some parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 06/05/2018 04:27 pm
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.

How common are single crystal alloy casting shops? Rene N5 and N6 are single crystal alloys. I presumed they were primarily only done in house at big OEMs like P&W and Rolls.

SpaceX is probably going to be forging or casting big portions of the turbine and nozzle with more conventional alloys like Inconel 718, Rene 41/Haynes R41, Haynes 188, 230, 282, ect.

Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)

Quote
Some parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.

Should be some metal that can't burn in pure oxygen, similar to Monel 400, but with thermal strength near or better than 718.

Monel 400 absolutely can't burn in pure oxygen, but too soft at high temp.

718 is mechanically and thermally enough to FFSC, which is cooler than ORSC and FRSC.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 06/05/2018 04:37 pm
The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F.

Please see the two quotes above. These are estimates of livingjw about the two pre-burner exit temperatures. Maybe before going on, I should ask.. how do you get to these? To an untrained person like me, these temperatures seem quite high. I have no experience with these and my thermodynamics is long ago (and not particularly great to begin with), so please be patient with me for the following.

The pre-burner burns LOX/LCH4 in just the right amount to run the turbine that drives the pumps and to provide the pressure AFTER the turbine that gets injected into the main combustion chamber. The pressure has to be higher than the pressure in the combustion chamber, so thats some considerable energy.
Here is the part that I am not comfortable with and could be completely wrong: My understanding is that the efficiency of the engine depends on the expansion ratio of the propellents, which depends on the temperature differential of the propellant going in the combustion chamber and the leaving temperature. So basically you want the incomming propellant as cold as possible and the outgoing propellant as hot as possible. Can you set this right please? I have the feeling there is some error in my understanding.

I one could compute the temperatures backwards. Starting with the chamber pressure and thrust of Raptor. One could compute the pressure after the turbines and the mass flow rate. This would inform about how strong the pumps would have to work and how much chemical energy has to be released inside the pre-burners, which would inform about the exit temperature and the mixture ratio inside the pre-burners. Maybe there is a better way of doing that?

There aren't many metals that can survive above 1400F. Personally I think the fuel rich outlet will be more around 1100F-1200F.
Fuel rich operation temp of CH4 are limited by soot production, above certain temp CH4 decompose to C, yet below temp limit of 718 blades.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 06/05/2018 04:50 pm
Fuel rich operation temp of CH4 are limited by soot production, above certain temp CH4 decompose to C, yet below temp limit of 718 blades.

Isn't that temperature around 1300F?


Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)

Quote
Some parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.

I was referring to the fuel turbomachinery when talking about the forging and cast materials.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/05/2018 11:55 pm
I am currently using 860K ( 1548R or ~1100F ) for both pre-burner exit temperatures. This is for the 250 bar chamber pressure. These temperatures will have to be raised when they evolve towards 300 bar, so I am giving them some temperature margin.

Looks like the pre-burner pressures are around 345 bar (~5000 psi).

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 06/06/2018 12:18 am
I hope at some point they reveal the temps and pressures at various points in the cycle on Raptor...  ???
My guess is they will not discuss... because of ITAR...  :P

My guess is the temps posted above (in many postings) are higher then reality pre turbine...

The key in my mind... is they are pumping a liquid and expanding a gas to power it...

What videos exist of either the powerhead tests or the prelight sequence hint to me the gas is way colder then you all think coming out...

I still think that Raptor uses a separate very high pressure gaseous methane and O2 tankage and supply (shared with RCS) to light up and provide the prop at each pre-burner to make it work...
The gross flow being pumped on both sides is not being burned at all, but just heated to a gas form of modest temp by the pre-burner exhaust products and radiated heat ...
This gas only has to be hot enough to not condense much going across the power turbines pressure drop...
I believe the real kicker will be not the 300 Bar chamber pressure, but the much higher pre-turbine pressures that exist between the pump outlets and the turbine inlets on both sides...  :o
Likely eye watering Bar numbers... but at temps that make exotic materials unnecessary...  8)
Hence the need for the separate high pressure burner supply, and a method and apparatus to replenish same at a useful rate to maintain power during a long burn...

I guess time will tell... my 2 cents...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/06/2018 01:15 am
- Pump power in must equal turbine power out. Lower pre-burner temperatures require higher pressure drops across the turbine and corresponding higher pump pressure. To minimize peak cycle pressure you need to maximize temperature consistent with material properties, engine life and growth.

- I also think they might use RCS high pressure tank propellants for spinning up the turbines during starting. It would be much faster than a boot strap start.

John

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Johnnyhinbos on 06/06/2018 01:37 am
Just for reference, modern common rail diesel engines pump fuel at up to 2,500 bar (yes bar - 36,000 psi) at ambient temp. They do so reliably and continuously. In very rough, variable conditions.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/06/2018 03:23 am
Pulled a couple of graphs showing power and pressure trends with turbine pressure ratio and different turbine entry temperature. As you drop the turbine temperature, power goes up a little, but pump pressure goes up a bit faster. Higher power and higher pressure will increase weight.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 06/06/2018 06:18 am
Fuel rich operation temp of CH4 are limited by soot production, above certain temp CH4 decompose to C, yet below temp limit of 718 blades.

Isn't that temperature around 1300F?


Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)

Quote
Some parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.

I was referring to the fuel turbomachinery when talking about the forging and cast materials.
Yes, both CH4 and 718 works up to 1300F.
So 1100-1200F have equally good safety margin for both.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 06/06/2018 07:24 pm
Question, for those smarter then me on this...  ???

How many grams of gaseous Methane and grams of gaseous O2 burned at a stoichiometric ratio in a "preburner/heater" would generate enough heat to then raise ONE whole Kilogram of either Liquid Methane OR Liquid Oxygen (likely are close but slightly different numbers) from a sub cooled at liquid nitrogen temperatures to a warm gas at say 204C (400F)...

My guess is... I don't know...  ???

On edit...
My goal with this question is to back figure how many grams of prop is needed burned to then vaporize the entire full flow staged combustion prop flow to reach the declared thrust values they have claimed...
As a percentage... I am thinking that many may be surprised how small a % number it is...  ;)

Point is... this is likely not like a typical rocket engine where a gas generator creates hot exhaust gases to drive the turbine (Merlin, F1, etc)
This is more like a power plant where heat is used to phase change something from a liquid to a gas and that then drives the turbine (much more mass at a much cooler temp)...
The nice part is the turbine exhausts can then be mixed and combusted to then make a real rocket out of it...

I may be crazy... but this is how I see Raptor... you got to forget SSME and every other rocket engine to date.
It's a different animal in many ways...  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 06/06/2018 07:39 pm
I'm only a mechanical engineer, but the numbers for pressures, temperatures, mass flow rates, total power and material properties are all outstanding and hard to wrap ones head around. 

That this is even possible is amazing.  The fact a private company is doing it with their own funding, wow, just wow.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 06/06/2018 09:08 pm
Question, for those smarter then me on this...  ???

How many grams of gaseous Methane and grams of gaseous O2 burned at a stoichiometric ratio in a "preburner/heater" would generate enough heat to then raise ONE whole Kilogram of either Liquid Methane OR Liquid Oxygen (likely are close but slightly different numbers) from a sub cooled at liquid nitrogen temperatures to a warm gas at say 204C (400F)...

My guess is... I don't know...  ???

On edit...
My goal with this question is to back figure how many grams of prop is needed burned to then vaporize the entire full flow staged combustion prop flow to reach the declared thrust values they have claimed...
As a percentage... I am thinking that many may be surprised how small a % number it is...  ;)

Point is... this is likely not like a typical rocket engine where a gas generator creates hot exhaust gases to drive the turbine (Merlin, F1, etc)
This is more like a power plant where heat is used to phase change something from a liquid to a gas and that then drives the turbine (much more mass at a much cooler temp)...
The nice part is the turbine exhausts can then be mixed and combusted to then make a real rocket out of it...

I may be crazy... but this is how I see Raptor... you got to forget SSME and every other rocket engine to date.
It's a different animal in many ways...  ;)

Methane rich or oxygen rich GG is already "combustion and boiling", no need of decomposing kerosene, etc.
Typical mix ratio of fuel rich methane gg: 0.3~0.4
Typical mix ratio of typical kerosene ORSC: 30~40
Both means stoichiometric combustion gas : cooling gas around 1:10.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/07/2018 12:00 am
Question, for those smarter then me on this...  ???

How many grams of gaseous Methane and grams of gaseous O2 burned at a stoichiometric ratio in a "preburner/heater" would generate enough heat to then raise ONE whole Kilogram of either Liquid Methane OR Liquid Oxygen (likely are close but slightly different numbers) from a sub cooled at liquid nitrogen temperatures to a warm gas at say 204C (400F)...

My guess is... I don't know...  ???

On edit...
My goal with this question is to back figure how many grams of prop is needed burned to then vaporize the entire full flow staged combustion prop flow to reach the declared thrust values they have claimed...
As a percentage... I am thinking that many may be surprised how small a % number it is...  ;)

Point is... this is likely not like a typical rocket engine where a gas generator creates hot exhaust gases to drive the turbine (Merlin, F1, etc)
This is more like a power plant where heat is used to phase change something from a liquid to a gas and that then drives the turbine (much more mass at a much cooler temp)...
The nice part is the turbine exhausts can then be mixed and combusted to then make a real rocket out of it...

I may be crazy... but this is how I see Raptor... you got to forget SSME and every other rocket engine to date.
It's a different animal in many ways...  ;)

The easiest way to answer this is to use NASA's CEA online application. I have attached the outputs for for heating LOX and CH4 to about 860K. The CH4 rich gas contains about 15% LOX. The Oxygen rich gas contains about 2% CH4. Obviously to raise the temperature to 400F (860R, 478K) would reduce the amounts shown. CEA is pretty easy to run.

- I modified my plots to include a 400F case to show the effects on the LOX pre-burner. Power increased over 40% and pump exit pressure increased by over 50%.  I really don't know why you want to use such low turbine temperatures. Can you explain?

- I have to admit that I did not change the Cp and gamma for the 400F case, so the numbers are not quite right but the trend is.

John

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 06/07/2018 12:30 am
My gut feeling is Tom Mueller would design the thing to run at the lowest turbine inlet temp he could get away with... to avoid having to use any real exotic alloys on rotating parts... and to insure a long life with no need for overhauls and limited inspections.

Again, just my opinion...  ;)
He would trade needing more pump power and a higher pressure pre-turbine... if he can make the turbine out of a lower cost alloy...

It goes back to why build things to work at the ragged edge... when a bit of common sense and a few more kilo's of mass could make it last much longer...



Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/07/2018 11:01 am
1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 06/07/2018 12:46 pm
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.

Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/07/2018 01:16 pm
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.

Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.

- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.

- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.

- Is there something else about the 5000+psi environment that you are concerned about?

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TrueBlueWitt on 06/07/2018 01:37 pm
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.

Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.

- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.

- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.

- Is there something else about the 5000+psi environment that you are concerned about?

John

Let us put this in perspective.

The list price for two GE90x turbofans used on a 777 is ~$60m
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 06/07/2018 01:53 pm
Rocket turbines are usually single stage, uncooled and don't require such extended life.

That used to be the case. In a point to point scenario as described by Gwynne Shotwell they will ramp up a number of uses equivalent to a years long distanse jet flights in a month. Even assuming jet engine equivalent service time they may need to swap out one engine for major overhaul every night. Or the whole set once a month.

Edit: What about the combustion chamber and nozzle? What would be their life span with high pressure propellant pressed through their cooling channels?

It will take many years of flights and engineering advances to make the engines as long lived as needed.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: nicp on 06/07/2018 02:17 pm
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.

Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.

- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.

- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.

- Is there something else about the 5000+psi environment that you are concerned about?

John

I remember being shocked to find out that the EVA suits on Shuttle had oxygen spheres at (if I recall correctly) 6000 psi, and thought, 'wow, those guys have bombs on their backs'
So 5000+ psi at a high temperature sure impresses me, but this isn't my field. I'm really into aerospace stuff, but I'm not qualified, I'm a mere software engineer with a background in electronics.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/07/2018 07:35 pm
Rocket turbines are usually single stage, uncooled and don't require such extended life.

That used to be the case. In a point to point scenario as described by Gwynne Shotwell they will ramp up a number of uses equivalent to a years long distanse jet flights in a month. Even assuming jet engine equivalent service time they may need to swap out one engine for major overhaul every night. Or the whole set once a month.

At the pressure ratio and mass flow rate, the turbine will likely be pure impulse. On the "coking" issue, the main coking issue in with RP1 is the internal deposits that build up in the flow passages that have high heat flux. The coke can be removed on internal passages by caustic cleaning, a hassle that is best avoided.  The methane will not have that issue and if the combustion is stratified in the vicinity of the walls Either fuel rich or O2 rich to remove the LSF caused distress, the liners should take it .


Edit: What about the combustion chamber and nozzle? What would be their life span with high pressure propellant pressed through their cooling channels?

It will take many years of flights and engineering advances to make the engines as long lived as needed.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 06/07/2018 08:20 pm
1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.

John

John, thanks for this. Reading above mentioned temperatures seems high, only for the normal life of most people. The uncooled temperature limit of usual alloys gives some much needed perspective. I didnt know at all that rocket engine turbines can withstand 1700F (925 C).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/07/2018 11:22 pm
Rocket turbines are usually single stage, uncooled and don't require such extended life.

That used to be the case. In a point to point scenario as described by Gwynne Shotwell they will ramp up a number of uses equivalent to a years long distanse jet flights in a month. Even assuming jet engine equivalent service time they may need to swap out one engine for major overhaul every night. Or the whole set once a month.

At the pressure ratio and mass flow rate, the turbine will likely be pure impulse. On the "coking" issue, the main coking issue in with RP1 is the internal deposits that build up in the flow passages that have high heat flux. The coke can be removed on internal passages by caustic cleaning, a hassle that is best avoided.  The methane will not have that issue and if the combustion is stratified in the vicinity of the walls Either fuel rich or O2 rich to remove the LSF caused distress, the liners should take it .


Edit: What about the combustion chamber and nozzle? What would be their life span with high pressure propellant pressed through their cooling channels?

It will take many years of flights and engineering advances to make the engines as long lived as needed.
I was including those items in my post , the highest heat flux will be in the throat of the MCC and any failure in the cooling system in the engine itself (not the external plumbing) would be nonfatal and be noticed on later inspection (as as happened on the SSME)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 06/07/2018 11:48 pm
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.

Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.

- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.

- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.

- Is there something else about the 5000+psi environment that you are concerned about?

John
Yeah, high pressure oxidizing environment is extremely difficult to engineer. Pressure matters.

And I know jet turbines are complex and therefore expensive. In fact, I noted the big cost differential in my post.

But the high temperature, high pressure, oxidizing environment does make engines like Raptor a more extreme metallurgical challenge than jet turbines.

Additionally, rocket engines see sharper thermal spikes than jet turbines and with larger temperature gradients (due to cryogenic fuels) in the de Laval nozzle.

So absolutely “mere.” Sorry.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/08/2018 12:17 am
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/08/2018 12:55 am
1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.

John

John, thanks for this. Reading above mentioned temperatures seems high, only for the normal life of most people. The uncooled temperature limit of usual alloys gives some much needed perspective. I didnt know at all that rocket engine turbines can withstand 1700F (925 C).

That is true on the fuel rich side. I don't know about the Oxygen rich side. Probably lower. I don't have specs on Mondaloy type alloys. I am assuming lower. Trying to be conservative and starting with 1000F for the oxygen rich turbine.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 06/08/2018 12:58 am
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John
They're using a new, custom alloy that didn't exist before.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 06/08/2018 01:04 am
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John
They're using a new, custom alloy that didn't exist before.
Obtainium?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: cscott on 06/08/2018 01:49 am
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John
They're using a new, custom alloy that didn't exist before.
Obtainium?
Red matter.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/08/2018 03:54 am
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John
They're using a new, custom alloy that didn't exist before.

Do you have any thing that makes you think it is not similar to Mondaloy? My guess is that it is.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 06/08/2018 04:02 am
The referenced quote was in post 940... (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1828181#msg1828181)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 06/09/2018 10:53 am
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John

718 operate up to 1300F and industrial useful life,  this is conventional to nickel alloy, some casted nickel alloys (not single crystal yet) go even higher.

As temperatures go low,  operate life go up exponentially, 1000F could last very long.

Mondaloy should not be MUCH worse than 718.

Russians ORSC use standard cast nickel alloys that are more heat resistant than 718 but prone to oxygen fires, with fragile antioxidation coatings. When any tiny FOD got sucked into engine it explode, NK33 and RD170 with Russian/Ukranian tanks always explode thisway when they explode. US native engines never explode thisway.

Presumeably the low chamber pressure of BE4 are limited by Mondalloy?  How about AR1?

So the Mondalloy operating temperature could be calculated.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Katana on 06/09/2018 11:03 am
1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.

John

John, thanks for this. Reading above mentioned temperatures seems high, only for the normal life of most people. The uncooled temperature limit of usual alloys gives some much needed perspective. I didnt know at all that rocket engine turbines can withstand 1700F (925 C).

That is true on the fuel rich side. I don't know about the Oxygen rich side. Probably lower. I don't have specs on Mondaloy type alloys. I am assuming lower. Trying to be conservative and starting with 1000F for the oxygen rich turbine.

John
If the fuel side is limited to 1200F by soot forming (100F margin to 1300F), the oxygen side could go cooler than 1000F.

High density of LOX (compared to methane) means large oxygen turbine working mass flow and power v.s. small oxygenpump volume flow and power

Here is a major benefit of FFSC over ORSC: one kind of turbine only needs to turn one pump instead of 2 pumps, so each turbines could run cooler at same chamber pressure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/10/2018 08:17 pm
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John

I found this:
"An alloy having the weight percent composition of
72.9 nickel, 16.6 cobalt, 8.1 chromium, 1.5 aluminum and
3.9 titanium was prepared. The alloy has been tested in high
pressure gaseous oxygen environments generally more
harsh than or Similar to a full-flow Staged combustion and
oxygen-rich Staged combustion rocket engine. The alloy
exhibited both high tensile and high burn resistance.

The entire patent is at :
https://patentimages.storage.googleapis.com/58/4f/d7/4d0e60f1762cd4/US20030053926A1.pdf
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 06/10/2018 09:56 pm
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles?  If not, how low do we have to go?  I cannot find any specs on Mondaloy.

John

I found this:
"An alloy having the weight percent composition of
72.9 nickel, 16.6 cobalt, 8.1 chromium, 1.5 aluminum and
3.9 titanium was prepared. The alloy has been tested in high
pressure gaseous oxygen environments generally more
harsh than or Similar to a full-flow Staged combustion and
oxygen-rich Staged combustion rocket engine. The alloy
exhibited both high tensile and high burn resistance.

The entire patent is at :
https://patentimages.storage.googleapis.com/58/4f/d7/4d0e60f1762cd4/US20030053926A1.pdf
Maybe what SpaceX has done is an adaptation of such an alloy and made one that is 3D printable.

Excellent properties and low cost of manufacture of complex parts and shapes.

What the article shows is there are alloys that exist that will work. What SpaceX needs is one that can be easily manufactured into complex parts/shapes probably through 3D printing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/11/2018 01:48 am
Mondaloy has been 3D printed and used in AR1.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 06/11/2018 10:41 am
... What SpaceX needs is one that can be easily manufactured into complex parts/shapes probably through 3D printing.

That sounds like they are still looking for it... are you sure they aren't already in production?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: oldAtlas_Eguy on 06/11/2018 06:12 pm
... What SpaceX needs is one that can be easily manufactured into complex parts/shapes probably through 3D printing.

That sounds like they are still looking for it... are you sure they aren't already in production?
I believe they have already settled on an specific one and are into producing parts. It is just we do not know what the alloy is and what that alloy's properties are other than that it meets the pressure, temperature, and long life that SpaceX want for its Raptor engine.

Also I gamed out what costs and prices would be like depending on air frame life and Raptor max flight life. I used a spread of 100 to 1000 flights for the airframes and 20 to 100 flights for the raptor. It also uses a spread of from $150M to $300M for initial manufacture of each of the units of BFS and BFB. Cost of manufacture of Raptors is spread from $5M each to just $1M each. All other costs for launch from $5M to $7M per launch. Worst case cost and price per flight (20 flight engine life, 100 flight airframes life, $300M initial unit costs for each BFS and BFB, $5M Raptor manufacture cost, $7M operations cost, 40% profit) is $13.4M for hardware costs/flight, $20.4M for all costs/flight and  $28.56M price per flight. Best case cost and price per flight (100 flight engine life, 1000 flight airframes life, $100M initial unit costs for each BFS and BFB, $1M Raptor manufacture cost, $5M operations cost, 15% profit) is $0.53M for hardware costs/flight, $5.53M for all costs/flight and $6.36M price per flight.

The greatest variable in per flight costs prices are the life expectancy multiplied by the manufacture costs. In the cases flight life lowers costs better than a low manufacture cost.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/15/2018 12:20 am
I am a Jet engine guy and I get this same argument. They find it hard to believe that the combustion Pressure is over 100 PSI less than the defused compressor discharge :o. and that the entire flow inside the engine is subsonic (except the Fan discharge). in this case it is likely a lot lower.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 06/15/2018 01:48 am
I am a Jet engine guy and I get this same argument. They find it hard to believe that the combustion Pressure is over 100 PSI less than the defused compressor discharge :o. and that the entire flow inside the engine is subsonic (except the Fan discharge). in this case it is likely a lot lower.

I am assuming this is for a modern turbofan engine on commercial aircraft. Was/is the same true for the early pure turbojets and modern military turbojets?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/15/2018 02:43 pm
Yes, only after-burning engines have significant supersonic exhaust all others are just sonic depending on the nozzle pressure ratio.  The original post out of the blue was answering a post that now does not exist :o
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Stan-1967 on 06/15/2018 04:11 pm


...Russians ORSC use standard cast nickel alloys that are more heat resistant than 718 but prone to oxygen fires, with fragile antioxidation coatings.

Is it known what the russian anti oxidation coatings are?  The patent posted upthread was for an alloy that did not rely on a coating, which apparently is what Mondaloy also does.  What then did the Russians use?
A.  Some type of ceramic?
B.  Some type of diffused aluminide with Pt/Rh?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 06/15/2018 05:00 pm


...Russians ORSC use standard cast nickel alloys that are more heat resistant than 718 but prone to oxygen fires, with fragile antioxidation coatings.

Is it known what the russian anti oxidation coatings are?  The patent posted upthread was for an alloy that did not rely on a coating, which apparently is what Mondaloy also does.  What then did the Russians use?
A.  Some type of ceramic?
B.  Some type of diffused aluminide with Pt/Rh?

Here are some information: https://www.sto.nato.int/publications/STO%20Educational%20Notes/RTO-EN-AVT-150/EN-AVT-150-06.pdf
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Kang54 on 06/17/2018 04:32 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 06/17/2018 05:36 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.

https://en.m.wikipedia.org/wiki/Inconel
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: HMXHMX on 06/17/2018 05:58 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.

800 ATM?!  That's just a bit more than I might have expected...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 06/17/2018 06:09 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.

800 ATM?!  That's just a bit more than I might have expected...
Probably a hefty saftey margin. It's not like they can just whip up a better batch if they decide to increase the chamber pressure...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: eischei on 06/17/2018 06:27 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.

800 ATM?!  That's just a bit more than I might have expected...
Probably a hefty saftey margin. It's not like they can just whip up a better batch if they decide to increase the chamber pressure...

Or not:

http://www.buran-energia.com/energia/moteur-fusee-rocket-engine-desc.php (http://www.buran-energia.com/energia/moteur-fusee-rocket-engine-desc.php)

Quote
In the creation of the RD-170 engine, the most powerful, many problems appeared. Pumps responsible for the supply of chambers in fuels were subjected to strong constraints, pump in one stage for the oxidizer and the pump in two stages for the fuel. The pump for the oxidizer work at 14 000 trs / min under a pressure of 600 atm, the pump of fuel of the first level was at 500 atm and the second stage one at 800 atm.

I remembered something like that... the pressure drop in the RD-170 is significant, espacially over the cooling channels in the bell ( if i remember correctly)

Regards,

Sepp

€: messed up the quoting.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 06/17/2018 08:14 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.

https://en.m.wikipedia.org/wiki/Inconel
About this SX stuff: Single-crystal superalloys (https://en.m.wikipedia.org/wiki/Superalloy#Single-crystal_superalloys)

Quote
Single crystal (SX) superalloys have wide application in the high-pressure turbine section of aero and industrial gas turbine engines due to the unique combination of properties and performance.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 06/18/2018 12:32 am
About this SX stuff: Single-crystal superalloys (https://en.m.wikipedia.org/wiki/Superalloy#Single-crystal_superalloys)

SX300, SX500 will most certainly be for a proprietary SpaceX alloy, likely suitable for welding or laser sintering (like Inconel 718).  Single crystal is only of use for getting utmost temperature capability (perhaps another 50-100K) in gas turbines inlet temperatures, it is finicky to make and crazy expensive to set up for it also needs cooled disk, firtree-root inserts for the single crystal blades and internal cooling passages.  All of which are heavy and unlikely to feature in a super light weight Raptor turbine that operates at much lower temperatures.  Creep life probably won't be a significant issue in Raptor turbines given only 10 hours of engine life. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Johnnyhinbos on 06/18/2018 02:29 am
12,000 psi isn’t totally insane when looking at fault tolerance of high pressure vessels... I’ve said it before - modern diesel common rail engines run fuel at up to 36,000 psi (that 2,449 psi - to keep apples to apples). No one seems to freak out with that...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 06/18/2018 03:44 am
run fuel at up to 36,000 psi (that 2,449 psi - to keep apples to apples).

Uh....what?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 06/18/2018 04:23 am
12,000 psi isn’t totally insane when looking at fault tolerance of high pressure vessels... I’ve said it before - modern diesel common rail engines run fuel at up to 36,000 psi (that 2,449 psi - to keep apples to apples). No one seems to freak out with that...

You have an 80MPa gaseous oxygen flow area of at least 0.01m² (~500kg/s at ~250kg/m³), you will need walls at least 10mm thick (but possibly much more), and in the pre-burners to turbine inlet area you will have to accommodate <1s startup transients with metal heating up by 7-800K in that time inducing huge through-wall stresses.  Inconel718 is good for >1GPa at <850K, so it's certainly doable, but no cake-walk.

In big combined cycle gas turbine plants it is the thermal transient stresses that are generally plant life limiting.  They have to power up and down very slowly to minimise those stresses and the low-cycle fatigue.  That startup heating rate may be a significant motivator for making Raptor smaller (thinner walls to heat, possible longer life turbomachinery).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 06/18/2018 04:40 am
Think of the flow rate and volume of diesel getting pumped up to that pressure and the power required. Think of the orders of magnitude it's off from rocket engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 06/18/2018 08:05 am
About this SX stuff: Single-crystal superalloys (https://en.m.wikipedia.org/wiki/Superalloy#Single-crystal_superalloys)

SX300, SX500 will most certainly be for a proprietary SpaceX alloy, likely suitable for welding or laser sintering (like Inconel 718).  Single crystal is only of use for getting utmost temperature capability (perhaps another 50-100K) in gas turbines inlet temperatures, it is finicky to make and crazy expensive to set up for it also needs cooled disk, firtree-root inserts for the single crystal blades and internal cooling passages.  All of which are heavy and unlikely to feature in a super light weight Raptor turbine that operates at much lower temperatures.  Creep life probably won't be a significant issue in Raptor turbines given only 10 hours of engine life.

Seems you can use 3D printing: https://www.springerprofessional.de/en/additive-manufacturing-of-single-crystal-superalloy-cmsx-4-throu/10266042

Quote
The results reported here represent one of the few successes obtained in producing single-crystal epitaxial deposits through a powder bed fusion-based metal AM process and thus demonstrate the potential of SLE1 to repair and manufacture single-crystal hot section components of gas turbine systems from nickel-based superalloy powders.

1 scanning laser epitaxy
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/18/2018 02:11 pm
I would expect the Turbine would be a "Blisk" in keeping with Spacex thinking.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 06/18/2018 02:46 pm
Ok, any new updates on Raptor? 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: John Alan on 06/18/2018 03:36 pm
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)
Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

I have no idea what that really means, but it seems very relevant to this thread.

800 ATM?!  That's just a bit more than I might have expected...

Not me... I was expecting some "eye watering" number in between the pump and the turbine...

I believe the real kicker will be not the 300 Bar chamber pressure, but the much higher pre-turbine pressures that exist between the pump outlets and the turbine inlets on both sides...  :o
Likely eye watering Bar numbers... 

I mean... 800 bar = 11,600+ psi...
That is some impressive stuff right there... ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Mongo62 on 06/18/2018 04:01 pm
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Nomadd on 06/18/2018 07:20 pm
 I wouldn't think the injector drop was all that linear.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: AncientU on 06/18/2018 09:02 pm
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.


300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/18/2018 11:33 pm
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.


300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.

- sub-scale demonstrator engine  was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.

- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar

Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8   (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)

This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.





Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 06/18/2018 11:36 pm
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.


300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.

200 bar for the subscale engine, as of last year.

Quote from: Elon Musk
The test engine currently operates at 200 atmospheres, or 200 bar. The flight engine will be at 250 bar, and we believe that over time we can get that to a little over 300 bar.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 06/18/2018 11:46 pm
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.


300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.

- sub-scale demonstrator engine  was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.

- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar

Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8   (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)

This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.

I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 06/19/2018 12:16 am
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.


300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.

- sub-scale demonstrator engine  was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.

- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar

Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8   (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)

This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.

I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Build a thicker tank. 

If you're going to try things no one has done before at higher pressures then there will be changes through the system and the tanks are part of that system. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 06/19/2018 06:46 am
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 06/19/2018 10:11 pm
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.

I agree it won't be a box in a box and the tanks will obviously be an integral part of the structure, but if the tank structure they tested was not remotely like the final tank that will be used what was the point of that expensive test (with 1200 tons of liquid oxygen)? What evidence is there that the tank pressurization will be greater than 2 atmospheres?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rabidpanda on 06/19/2018 11:07 pm
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.

I agree it won't be a box in a box and the tanks will obviously be an integral part of the structure, but if the tank structure they tested was not remotely like the final tank that will be used what was the point of that expensive test (with 1200 tons of liquid oxygen)? What evidence is there that the tank pressurization will be greater than 2 atmospheres?

Well that tank was designed years ago, so I find it highly unlikely that it bears much resemblance to the current iteration of BFS (it was completely different diameter anyway).

But I imagine that testing that tank was useful for validating their manufacturing processes and analysis methods.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 06/20/2018 08:25 am
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.


In fact that test tank stood perfectly well without being pressurized.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 06/20/2018 11:03 am
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.


In fact that test tank stood perfectly well without being pressurized.

Umm...

I admit, this is after the rupture, but it still looks rather floppy.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/20/2018 11:38 am
I assumed ~3 bar because that is what tanks typically are designed for. It could be as low as ~2 but I think it will be closer to 3 for a couple of reasons. 1) makes designing the inducers and pumps easier, and 2) stabilizes the structure.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 06/20/2018 12:55 pm
I assumed ~3 bar because that is what tanks typically are designed for. It could be as low as ~2 but I think it will be closer to 3 for a couple of reasons. 1) makes designing the inducers and pumps easier, and 2) stabilizes the structure.

John

I've never thought about the net positive suction head of sub-cooled liquid oxygen and methane on turbo machinery before.

I'm curious as to how hard the autogenous system is to develop and feed back to the tanks, and control the tank pressure to not over or under shoot tank pressure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RoboGoofers on 06/20/2018 03:17 pm
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.


In fact that test tank stood perfectly well without being pressurized.

Umm...

I admit, this is after the rupture, but it still looks rather floppy.

These looks pretty floppy after rupture, too.
https://youtu.be/m6gl5vACDnQ
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 06/20/2018 06:41 pm
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.


In fact that test tank stood perfectly well without being pressurized.

Umm...

I admit, this is after the rupture, but it still looks rather floppy.



You can see that the dome is actually holding shape pretty well despite this clearly no longer being pressurized.
Remember that picture from the 2016 IAC where there was a bunch of people standing in front of the tank at the manufacturing site?
I know for a fact that the tank wasn't pressurized back then.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 06/20/2018 08:26 pm
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.

Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):

Quote
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.

There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.


In fact that test tank stood perfectly well without being pressurized.

Umm...

I admit, this is after the rupture, but it still looks rather floppy.



You can see that the dome is actually holding shape pretty well despite this clearly no longer being pressurized.
Remember that picture from the 2016 IAC where there was a bunch of people standing in front of the tank at the manufacturing site?
I know for a fact that the tank wasn't pressurized back then.
 
Or the photo of it's inside.

Nothing about the shredded remains suggests any part of it is floppy either.

Do you also consider a fallen tree floppy just because a storm blew it over? Most parts of it will remain at original structural strength even after it failed to hold up against off-nominal loads.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/21/2018 02:01 am
Autogenous pressurization is well understood technology. Shuttle used it.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 06/21/2018 07:04 am
Autogenous pressurization is well understood technology. Shuttle used it.

John

The discussion was about whether-or-not the empty tank would require constant pressurization to maintain its shape (a la the Atlas baloon tank).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 06/21/2018 11:11 am
I was responding to Wannamoonbase about autogenous pressurization.

- As far as tank design, normal practice is to make the tanks (without pressurization) stable enough to support their weight and anything that sits on top of it. Sometimes this includes the fuel, sometimes not. Also  wind loads are  considered.

- In all cases, it cannot fly without pressure stabilizing effect.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Bananas_on_Mars on 06/21/2018 12:29 pm
Autogenous pressurization is well understood technology. Shuttle used it.

John

The discussion was about whether-or-not the empty tank would require constant pressurization to maintain its shape (a la the Atlas baloon tank).
I think Elon said they will be vented/at vacuum during transit to mars, so they will act as a dewar vessel aka thermos bottle for the landing propellant stored inside the header tanks. So unpressured tanks for transit and landing on Mars, i guess.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DistantTemple on 06/21/2018 12:47 pm
The external, and gravitational forces during transit will be extremely low. During takeoff and landing the forces are enormous. An on earth and Mars intermediate. So I guess they will repressurise for EDL, I don't see how it could be otherwise. Also above pple have talked about 2 or 3 bar tank pressure for fuel delivery, and this is relevant during EDL.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: kraisee on 06/21/2018 01:16 pm
During a Mars flight I would expect the outer tanks to be vented to vacuum, for insulation purposes. But a small amount of the props from the inner tanks would be used to re-pressurise all tanks prior to any high-thrust activities, or the tanks would just not be strong enough for EDL or landing.

Internal tank pressures are likely to be determined by Raptor's startup inlet head pressure requirements, with LOX tending to be a slightly higher pressure requirement than the fuel. At liftoff the booster stage will have the benefit of head pressure created simply by the height of the propellants in such a tall rocket. But later ignition of BFR and BFS in low-g environments will depend on delivering the right inlet pressure without the benefit of gravity's influence, so that's probably going to dictate the design requirement.

Regarding ground processing, the 12m diameter one we saw previously was definitely self-supporting. It was also able to support the weight of a stairwell allowing people to access the inside of it. Given that it did not feature an airlock, that tells me that it was self-supporting in an un-pressurised state. I would expect this to be the case for the flight articles too.

Generally, composite cryo tanks like this will be built with a factor of safety higher than aluminium equivalents. The FS recommended by NASA for such tanks for human rated use is 2.0 (compared to 1.4 for alum). Assuming this is the case here too, then that should result in a tank plenty strong enough to handle ground processing activities without additional internal pressure.

A little mass could potentially be saved by pressurising the tanks permanently while on the ground, like Centaur used to do, but the endless problems and failures they experienced across that program led them away from that solution. I expect SpX don't want that sort of headache - BFR/BFS are challenging enough already - at least not in the early versions of these systems. It might be a future upgrade path, if they really need the extra performance, which I currently doubt.

Ross.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 06/21/2018 02:11 pm
This conversation is straying from Raptor.  Discussion of BFR tank design should probably go in one of the other BFR threads.  (I know it started out being related to the pressures in the propellant flow path, but the conversation has moved on from that.)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 06/21/2018 02:17 pm
I was responding to Wannamoonbase about autogenous pressurization.

- As far as tank design, normal practice is to make the tanks (without pressurization) stable enough to support their weight and anything that sits on top of it. Sometimes this includes the fuel, sometimes not. Also  wind loads are  considered.

- In all cases, it cannot fly without pressure stabilizing effect.

John

Thanks John, I'm glad it's straight forward. 

No doubt they'll enjoy not having COPVs.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 06/28/2018 09:47 pm
During a Mars flight I would expect the outer tanks to be vented to vacuum, for insulation purposes. But a small amount of the props from the inner tanks would be used to re-pressurise all tanks prior to any high-thrust activities, or the tanks would just not be strong enough for EDL or landing.

I think it would be more efficient to accumulate the gases in compressed cylinders for cold gas thrusters or GOX/CH4 thrusters.  Power is cheap, reaction mass is expensive.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: guckyfan on 06/29/2018 05:54 am
Elon Musk said they would subcool the propellant in the header tanks by venting to vacuum ahead of landing. I think they may vent to the main tanks, so the mass would not be wasted. Would the header tank fuel be cold enough before pressure inside the main tank becomes too high?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 06/29/2018 11:16 am
Elon Musk said they would subcool the propellant in the header tanks by venting to vacuum ahead of landing. I think they may vent to the main tanks, so the mass would not be wasted. Would the header tank fuel be cold enough before pressure inside the main tank becomes too high?
<snip complex answer with maths>
In short - there is basically no way of knowing.
Reasonable assumptions can lead to widely varying assumptions of in-flight temperatures and flight pressures.
You probably can't subcool this way all the way down to the coldest temperatures, if your propellant starts out very hot (at 3 bar vapour ) - but there is no particular reason to believe that the initial equilibrium temperature pressure of the tank is 3 bar, not 0.3 bar.

Pointing the BFR towards/away from the sun in interplanetary space can in principle make the walls very, very cold - enough that the 'subcooling' is done without actual venting of the inner tank at all.

I think this is consistent with what's been said.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 06/30/2018 09:53 pm
Looking back on the subject of the cause of the green startup flame , I got to thinking that there is a good chance that on this "first go" on this design, it would make sense to fit both (TEA/TEB and "Electric") and due to the unknowns of the startup behavior the tried and true use of TEA/TEB would remove an unknown in the early testing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aero on 07/01/2018 03:18 am
Looking back on the subject of the cause of the green startup flame , I got to thinking that there is a good chance that on this "first go" on this design, it would make sense to fit both (TEA/TEB and "Electric") and due to the unknowns of the startup behavior the tried and true use of TEA/TEB would remove an unknown in the early testing.

I don't think so. Removal of engine unknowns, including start-up unknowns is just what engine test stands are for. If the engines won't start very, very reliably, that model of engine will never leave the test stand.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/02/2018 03:43 pm
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/02/2018 04:21 pm
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: HMXHMX on 07/03/2018 03:34 am
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.

Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber.  For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely.  Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system.  It's coming, no doubt.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Semmel on 07/03/2018 07:18 am
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.

Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber.  For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely.  Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system.  It's coming, no doubt.

Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.

And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 07/03/2018 01:38 pm
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.

Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber.  For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely.  Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system.  It's coming, no doubt.

Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.

And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
 

I have aired the camera artefact hypothesis before, but I think far more likely than that or boron fluid, is engine rich combustion. It is testing after all. You'll notice the green flash is present soon after ignition in the September 2017 video. 
https://mobile.twitter.com/SpaceX/status/913625659947106304/video/1   
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Bananas_on_Mars on 07/03/2018 02:41 pm
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.

Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber.  For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely.  Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system.  It's coming, no doubt.

Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.

And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
With the FFSC cycle of the raptor engine, i'm not sure wether a hard start would be an issue for the combustion chamber, because both fuel and oxidiser should enter the combustion chamber as a gas, so there's should not be the possibility of fuel pooling inside the engine and then ignite when hit by liquid oxygen. On the other side, Raptor needs 3 separate ignition sources, one for each preburner and one for the combustion chamber. So they might have started one or the other with TEA-TEB and used an augmented spark ignitor at the same time...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/03/2018 03:03 pm
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.

Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber.  For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely.  Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system.  It's coming, no doubt.

Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.

And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
With the FFSC cycle of the raptor engine, i'm not sure wether a hard start would be an issue for the combustion chamber, because both fuel and oxidiser should enter the combustion chamber as a gas, so there's should not be the possibility of fuel pooling inside the engine and then ignite when hit by liquid oxygen. On the other side, Raptor needs 3 separate ignition sources, one for each preburner and one for the combustion chamber. So they might have started one or the other with TEA-TEB and used an augmented spark ignitor at the same time...

Is a MCC igniter even necessary? I would think that gaseous O2 at 800 C or thereabouts coming out of the preburner would autoignite with most fuels, especially gaseous methane.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/03/2018 03:08 pm
If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts.


That's what's called an augmented spark igniter or a torch igniter.

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120011145.pdf

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: testguy on 07/03/2018 03:43 pm
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.

How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?

I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.

Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber.  For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely.  Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system.  It's coming, no doubt.

Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.

And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.

TEA/TEB, is premixed and placed in a pressure vessel.  The vessel is pressurized with an inert gas such as nitrogen as the driving mechanism.  The plumbing for the engine is very simple, just a solenoid valve between the pressure vessel and the combustion chamber. There may also be an inert gas line near the solenoid valve, could be a three way valve, to prevent combustion products from the TEA/TEB clogging the line especially if they want restarts.  Or they could have multiple TEA/TEB systems for restarts.
The pretest preparation plumbing is more complex because you need to first get the TEA/TEB to the high pressure vessel by flowing the fluid from their low pressure DoT shipping containers, meter the amount and isolate the high pressure vessel all the while under an inert gas.
The fluid will ALWAYS ignite when it comes in contact with oxygen.  I have a great deal of experience working with TEB, not TEA/TEB, but the properties should be similar.  My experience using a spark to ignite a fuel and oxidizer have at times been problematic.  You need the correct mixture ratio and spark to be in the right location at the same time with the spark having sufficient energy.  There have been times where we have deleted the spark and just used TEB with great success.  I hope that helps.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: vulture4 on 07/05/2018 04:31 pm
Kerosene has 11 to 18 carbon atoms, roughly twice the weight of gasoline, and consequently it is not used in spark ignition piston engines where instantaneous ignition is needed. In jets, which do use kerosene, the compressor is spun up, the spark ignitor is run continuously, and the fuel flow is initiated. Ignition need not be instantaneous, and there is no LOX to keep things cold, but even so a turbine engine electrical ignitor is a lot more sophisticated and powerful than an automotive spark plug, and the cost is equally impressive. Spark ignition for a kerosene rocket engine might require starting at low fuel flow and gradual runup, as with a jet, which would not be feasible for the SpaceX booster landing where thrust is needed immediately.
http://okigihan.blogspot.com/p/turbineengine-ignition-systems-since.html

The more relevant fuel may be methane; spark ignition might require more energy than for hydrogen, but it would be easier than for kerosene or even gasoline.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/06/2018 02:37 pm
I'm a jet engine guy and I can tell you that in that application the fuel from the start nozzle must spray directly on the plug to get ignition.  In this Raptor application the spark igniter will likely ignite a "small rocket engine" which will exhaust into either the pre-burners or MCC. The the ignition must propagate to all the the injectors (assuming coaxial). The velocity of the mixed fuel exiting the injectors needs to be lower than the flame propagation velocity (in the pre-burners) to allow each injector to ignite, "nothing could go wrong"
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/06/2018 06:49 pm
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:

There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.

So one scenario would be:

Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar

In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.


300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.

- sub-scale demonstrator engine  was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.

- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar

Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8   (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)

This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.

What do you estimate for the turbine output temperature? More than 600 C?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/06/2018 10:56 pm
I'm a jet engine guy and I can tell you that in that application the fuel from the start nozzle must spray directly on the plug to get ignition.  In this Raptor application the spark igniter will likely ignite a "small rocket engine" which will exhaust into either the pre-burners or MCC. The the ignition must propagate to all the the injectors (assuming coaxial). The velocity of the mixed fuel exiting the injectors needs to be lower than the flame propagation velocity (in the pre-burners) to allow each injector to ignite, "nothing could go wrong"

The Raptor CAD model showed what looks like three ignitors. One for each of the pre-burners and one for the MCC. Probably electrically ignited torches as stated above.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/06/2018 11:04 pm
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: testguy on 07/07/2018 02:30 pm
I'm a jet engine guy and I can tell you that in that application the fuel from the start nozzle must spray directly on the plug to get ignition.  In this Raptor application the spark igniter will likely ignite a "small rocket engine" which will exhaust into either the pre-burners or MCC. The the ignition must propagate to all the the injectors (assuming coaxial). The velocity of the mixed fuel exiting the injectors needs to be lower than the flame propagation velocity (in the pre-burners) to allow each injector to ignite, "nothing could go wrong"

I started my career as a jet engine guy also.  Everything you stated is correct.  I remember the difficulty in getting early development engines to start.  BTW TEB, if I remember correctly, was first developed for the J-58 engine used in the SR-71.  The reason was that if the J-58 flamed out at altitude it could not be restarted with a spark ignition system.  TEB was used for sea level starts also.  Just in case anyone is interested TEA stands for Tri-Ethyl-Aluminum while TEB stands for Tri-Ethyl-Borane.

Sorry for the diversion.  Time to get the discussion back to Raptor.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/07/2018 04:22 pm
The cad model is the "wish"  the sub scale could be anything.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 07/07/2018 05:50 pm
The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.

.. as long as the main propellant inflow velocity isn't so fast that it extinguishes the torch, right?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/07/2018 07:38 pm
Right, luckily Methane/O2 should have a high flame propagation velocity. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/08/2018 10:59 pm
Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: deruch on 07/09/2018 12:01 am
Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.

That strikes me as the sort of Elon statement that is liable to be over parsed (are ignition fluids counted as "propellants", etc).  Or the type of engineering issue that is liable to end up changing when the final design actually makes it into production, even if it was one of the earlier goals.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/09/2018 01:07 am
Couldn't have said it better myself 8)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: marsbase on 07/09/2018 02:14 am
That strikes me as the sort of Elon statement that is liable to be over parsed (are ignition fluids counted as "propellants", etc).  Or the type of engineering issue that is liable to end up changing when the final design actually makes it into production, even if it was one of the earlier goals.
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: testguy on 07/09/2018 02:30 am
Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.

Agreed.  Every time you add a gas or liquid to the vehicle it eats into your trun around time which I thought was a major objective to minimize.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/09/2018 02:48 am
Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.

That strikes me as the sort of Elon statement that is liable to be over parsed (are ignition fluids counted as "propellants", etc).  Or the type of engineering issue that is liable to end up changing when the final design actually makes it into production, even if it was one of the earlier goals.
No, Raptor cannot use TEA/TEB, helium, or nitrogen. They add major operational headaches both on Earth and especially on Mars.

They also aren't needed. All the issues are resolvable.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 07/09/2018 08:43 am
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
Humans do not mind some nitrogen in the air. Well, it is not absolutely necessary. Helium does the job also, but... Pure oxigen is not ideal, and not only because of the fire. I am not sure if we have discussed this..
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 07/09/2018 11:12 am
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
Humans do not mind some nitrogen in the air. Well, it is not absolutely necessary. Helium does the job also, but... Pure oxigen is not ideal, and not only because of the fire. I am not sure if we have discussed this..
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.


-----
ABCD: Always Be Counting Down

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: marsbase on 07/09/2018 12:28 pm
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.

If nitrogen and argon were not available on Mars, what would be the alternative for thrusters?  Could you use cold methane or oxygen as thruster gas? At least then you would be unlikely to run out.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 07/09/2018 01:26 pm
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.

If nitrogen and argon were not available on Mars, what would be the alternative for thrusters?  Could you use cold methane or oxygen as thruster gas? At least then you would be unlikely to run out.

CO2 is used as a propellant for some applications (https://en.wikipedia.org/wiki/Air_gun#CO2), but fortunately that's not all Mars has to offer. 
Unfortunately Argon freezes near nitrogen boiling point, since SpaceX is tinkering with deep-cryo, the allure of low molecular weight in Argon is completely ruled out due to freezing. 
Nitrogen on the other hand has superior molecular weight over Oxygen and better pressurization performance (https://www.reddit.com/r/rocketry/comments/7wfe4g/vapor_pressure_graphs_for_some_substances_of/), but it's certanly not worth the added complexity in hardware and operations dealing with yet another consumable with fractional distillation complications, in your propulsion system. I fear the mass gained in lower ullage mass with N2 is more than eaten up by the extra hardware for dealing with the additional gas.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/09/2018 02:01 pm
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.

If nitrogen and argon were not available on Mars, what would be the alternative for thrusters?  Could you use cold methane or oxygen as thruster gas? At least then you would be unlikely to run out.

Gaseous pressure-fed methalox thrusters with spark or spark/torch ignition. The thruster propellant tanks will autogenously pressurize, and can be refilled by pumps compressing boiloff from the main tanks. No pneumatics or hydraulics, and no ignition fluids.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/09/2018 02:05 pm
When he says this, isn't Elon thinking of what can be made on Mars.  If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items.  Every pound of thruster nitrogen you carry to Mars is a pound less of payload.  I would think SpaceX will try very hard to avoid that.
Humans do not mind some nitrogen in the air. Well, it is not absolutely necessary. Helium does the job also, but... Pure oxigen is not ideal, and not only because of the fire. I am not sure if we have discussed this..
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.


-----
ABCD: Always Be Counting Down

They will want nitrogen for the breathable atmosphere. However this is not consumed like in a pneumatic system, only lost in small quantities to leaks and perhaps partially vented if they don't fully pump down the airlocks, so the quantities will be smaller.

More importantly, it's not a propulsion service point to get the ship re-flying. Rapid reflight means minimal servicing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: edzieba on 07/09/2018 02:07 pm
SpaceX are already developing the yet-unnamed Methalox gas/gas RCS thrusters for BFR/BFS, which will need to be rapidly and reliably igniting in order to be of much use as RCS thrusters. Could one of those be taken and used as a 'starter torch' for the Raptor in place of TEA-TEB or similar torches, or does that require a hotter or more stable flame than can be provided by the same propellants it is igniting?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/09/2018 02:14 pm
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.

John

Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/09/2018 10:02 pm
The next question is the starting method , Bootstrap or high pressure turbine impingement spinup  or, something else.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/10/2018 12:26 am
The next question is the starting method , Bootstrap or high pressure turbine impingement spinup  or, something else.
I'm betting on them using some of the RCS high pressure methane and lox  for spinning up the respective turbines.
This should yield fast starts. Its available. Why not?

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/10/2018 01:47 am
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.

John

Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?

I am currently calculating a turbine total pressure ratio of about .7-.75 and a total temperature out of about 500 C. Total temperature will not change through the injector, but static temperature drops with the pressure drop across the injectors according to adiabatic flow.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 07/10/2018 02:40 am
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.

John

Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?

I am currently calculating a turbine total pressure ratio of about .7-.75 and a total temperature out of about 500 C. Total temperature will not change through the injector, but static temperature drops with the pressure drop across the injectors according to adiabatic flow.

John

Too low to get fast autoignition with gaseous methane despite 250 bar pressures, then? The temperature is probably about right, but I'm not sure what the hold time is. Probably less than the MCC dwell time of the cold flowing props.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/11/2018 10:35 pm

I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.

John
[/quote]

Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?
[/quote]

"I am currently calculating a turbine total pressure ratio of about .7-.75 and a total temperature out of about 500 C. Total temperature will not change through the injector, but static temperature drops with the pressure drop across the injectors according to adiabatic flow.

John"


Is that The Methane or LOX turbine conditions?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 07/13/2018 01:28 am

I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.

John

Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?
[/quote]

"I am currently calculating a turbine total pressure ratio of about .7-.75 and a total temperature out of about 500 C. Total temperature will not change through the injector, but static temperature drops with the pressure drop across the injectors according to adiabatic flow.

John"


Is that The Methane or LOX turbine conditions?
[/quote]

Turbine inlet temperatures are the same for now. Also, both turbines are in the .7 to .75 pressure ratio range. Just the way its working out.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 07/13/2018 01:17 pm
Some posts back there was speculation about the ignition method for Raptor. Found this on LinkedIn:

John Bucknell (https://www.linkedin.com/in/john-bucknell-pe-7111a47/):

Quote
Created Raptor pre-burner combustor designs to achieve deep throttling as well as a spark-ignited torch igniter that was capable of starting the engine from on-board autogenous blow-down propellants.

He left SpaceX in 2012, so this might have changed...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/13/2018 04:17 pm
Some posts back there was speculation about the ignition method for Raptor. Found this on LinkedIn:

John Bucknell (https://www.linkedin.com/in/john-bucknell-pe-7111a47/):

Quote
Created Raptor pre-burner combustor designs to achieve deep throttling as well as a spark-ignited torch igniter that was capable of starting the engine from on-board autogenous blow-down propellants.

He left SpaceX in 2012, so this might have changed...

"Designed and built the subscale Raptor rocket for proof of concept testing able to test eighty-one configurations of main injector."

Oh wow, that's a lot of prototypes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: jpo234 on 07/13/2018 04:35 pm
"Designed and built the subscale Raptor rocket for proof of concept testing able to test eighty-one configurations of main injector."

Oh wow, that's a lot of prototypes.

And this was well before they really got going. Tom Mueller said recently (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/):
Quote
I’ve been working on Mars for the last four years, so I’m not going to take any credit for the Block 5 engine and all the upgrades that have happened

So, it seems the real work on Raptor started in 2014...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 07/13/2018 04:56 pm
>
So, it seems the real work on Raptor started in 2014...

They cut the ribbon on the Stennis test stand in April 2014  and did injector tests later that summer, so likely well before that.

https://www.nasa.gov/press/2014/april/nasa-spacex-cut-ribbon-to-launch-testing-partnership

http://www.nasa.gov/sites/default/files/atoms/files/septemberlagniappe2.pdf
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/13/2018 05:18 pm
They also had that LH2/LOX Raptor concept as well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/16/2018 10:03 pm
Slight change in subject.

Looking at the proximity of  McGregor  (4 miles) 45 deg. to the sub scale test stand and Oglesby  (2.72 miles)  Which is aligned with what looks like the production test stand (at least for full scale development and certification testing) I would think that the Raptor will cause some noise issues.  Any talk of some suppressed test stands for the Raptor?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 07/16/2018 10:39 pm
Slight change in subject.

Looking at the proximity of  McGregor  (4 miles) 45 deg. to the sub scale test stand and Oglesby  (2.72 miles)  Which is aligned with what looks like the production test stand (at least for full scale development and certification testing) I would think that the Raptor will cause some noise issues.  Any talk of some suppressed test stands for the Raptor?

Interesting question. Problem might remain workable until testing at a frequency requited for Booster manufacture rate is required?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/17/2018 07:22 pm
The only saving grace is the characteristics of a supersonic jet noise propagation

From "Ingeniare. Revista chilena de ingeniería, vol. 14 Nº 3, 2006"

"The maximum radiation of a jet exhaust, which is highly directional and has maximum
intensity at angles of between 30 Deg. and 45 Deg."

Which sort leaves, at least the center of town, in the cone of silence :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/17/2018 07:44 pm
The only saving grace is the characteristics of a supersonic jet noise propagation

From "Ingeniare. Revista chilena de ingeniería, vol. 14 Nº 3, 2006"

"The maximum radiation of a jet exhaust, which is highly directional and has maximum
intensity at angles of between 30 Deg. and 45 Deg."

Which sort leaves, at least the center of town, in the cone of silence :)


Except that the test stand is very close to the ground compared to jet engines in planes while flying. Sound waves reflect off the ground, and what not. It's still very very loud.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/17/2018 08:38 pm
The article was for ground test of rocket engines.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 07/17/2018 09:43 pm
The article was for ground test of rocket engines.
Says jets in the text you gave, is this a translation issue?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/17/2018 10:10 pm
it is for the supersonic jet flow from any nozzle "Jet"  Jet engine or rocket engine it makes no difference.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 07/17/2018 11:40 pm
This is the original meaning of the word, prior to the development of the turbine engine, which appropriated the word.  First three definitions from Dictionary.com:
1. a stream of a liquid, gas, or small solid particles forcefully shooting forth from a nozzle, orifice, etc.
2. something that issues in such a stream, as water or gas.
3. a spout or nozzle for emitting liquid or gas

Etymology (from Online Etymological Dictionary):
1690s, "stream of water," from French jet "a throw, a cast; a gush, spurt (of water); a shoot (of a plant)," from jeter "to throw, thrust" (from PIE root *ye- "to throw, impel"). Middle English had jet/get "a device, mode, manner, fashion, style" (early 14c.). Sense of "spout or nozzle for emitting water, gas, fuel, etc." is from 1825. Hence jet propulsion (1855, originally in reference to water)

WRT aircraft engines, not all turbines produce a jet. e.g turboprop as opposed to turbojet and turbofan. Not all jets are turbines, e.g. ramjet, scramjet

A modern granite quarry uses an ultra high pressure water jet to cut the stone as this is far more efficient than a diamond saw.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: QuantumG on 07/17/2018 11:49 pm
Yup, in Sutton you'll see a similar description.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/18/2018 01:52 pm
I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 07/18/2018 11:17 pm
I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.

Not quite sure to whom that is addressed, but my post was in response to:

Says jets in the text you gave, is this a translation issue?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/19/2018 12:55 am
Just a general reply to the subject matter...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 07/19/2018 01:01 am
I do see an new vertical test stand under construction east of the Merlin vertical stand (and "milk stool") .The flame trench and water tower as well as some small building are sprouting out of the ground. Perhaps that is the suppressed Raptor stand.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: edzieba on 07/19/2018 04:11 pm
All signs point to the new stand being an S2 test stand (to move it away from the individual Merlin test stands). The Raptor test stand is the existing Raptor test stand.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: testguy on 07/19/2018 04:49 pm
I believe the purpose is still under debate.  L2 could provide additional insight.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 07/19/2018 06:40 pm
I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.

Not quite sure to whom that is addressed, but my post was in response to:

Says jets in the text you gave, is this a translation issue?
I got it now. In my defense, the stuff at the left doesn't look very "engine-ey" in that the (apparent) bell seems smaller than the (apparent) combustion chamber.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: alang on 07/21/2018 01:48 pm
I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.

Not quite sure to whom that is addressed, but my post was in response to:

Says jets in the text you gave, is this a translation issue?
I got it now. In my defense, the stuff at the left doesn't look very "engine-ey" in that the (apparent) bell seems smaller than the (apparent) combustion chamber.

A RAH fan like you might remember when some people called rocket engines "jets" in English.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TomH on 07/21/2018 03:36 pm
This issue is known as denotation vs. connotation. The denotation is the original literal meaning of a word. The connotation is the way people use it in common vernacular. The present connotation of jet obviously is an airplane powered by turbojet or turbofan engines.

The denotation of hot is high temperature and cold is low temperature. Following a rock concert by a female performer, two young teenage boys were heard saying, Wow, she was really cool. I agree, said the other, she was so hot! A third boy, named Sheldon, exclaimed, Well hold on just a minute here; I didn't see either of you go take her temperature!

Now, back to our regularly scheduled program.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: deruch on 07/22/2018 01:37 pm
This issue is known as denotation vs. connotation. The denotation is the original literal meaning of a word. The connotation is the way people use it in common vernacular. The present connotation of jet obviously is an airplane powered by turbojet or turbofan engines.

You run into this issue a lot when engaging in technical discussions, where one participant may use one meaning and others use the other (or they may not be familiar with the technical usage, etc).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 07/25/2018 07:21 pm
Is the Isp stated is their actual test data, or the theoretical Isp that they are working toward?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 08/03/2018 01:18 am
I would think the latter...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 08/03/2018 06:26 pm
I would think the latter...

I am confused as the Wiki shows 2 different Vac Isp's. Is 356s the Isp reached, or is it due to a redesign from the previous announcement?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 08/03/2018 10:28 pm
I would think the latter...

I am confused as the Wiki shows 2 different Vac Isp's. Is 356s the Isp reached, or is it due to a redesign from the previous announcement?

SL engines have a different vac ISP than vac engines do.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hitech on 08/04/2018 02:24 pm
If they manage to operate at the MC pressures and  desired mixture ratios, then the
 chemistry and thermodynamics dictates the ISP, low MC pressure equals lower ISP.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Ultrafamicom on 09/18/2018 03:27 am
Does today's test video show a full scale Raptor running?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 09/18/2018 05:22 am
Does today's test video show a full scale Raptor running?

That’s my assumption. It looked like a different test stand (?), so I assume it was the full size engine. I hope we get some clarification.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Surgeon on 09/18/2018 06:05 am
Does today's test video show a full scale Raptor running?

That’s my assumption. It looked like a different test stand (?), so I assume it was the full size engine. I hope we get some clarification.

With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: yokem55 on 09/18/2018 06:46 am
I have a feeling that they ran into a big issue with designing the vacuum regen cooled nozzle.

It needed to be very large while still being light enough to not chew up the ISP gains, and still able to handle the thermal flux of the exhaust, which previously Elon has described as 'nuts'.

Punting on RaptorVac was probably the best way forward for now.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Cheapchips on 09/18/2018 09:13 am

If anyone missed it from the presentation, they are still planning a VacRaptor for V2 BFS. Or at least, Musk was split balling the idea.  It's hard to tell sometimes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: KelvinZero on 09/18/2018 09:37 am
I have a feeling that they ran into a big issue with designing the vacuum regen cooled nozzle.

It needed to be very large while still being light enough to not chew up the ISP gains, and still able to handle the thermal flux of the exhaust, which previously Elon has described as 'nuts'.
Could be. There were two alternate reasons given though:
(1) Get started with just one engine design instead of two (with the second only used for 4 out of 38 engines)
(2) More survivable options in the case of engine out.. can survive 4 engines failing and still land.. could be particularly useful early on while working out the bugs!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 09/18/2018 10:14 am
Does today's test video show a full scale Raptor running?

In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.

Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.

Sounds like we have a 300 bar.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aameise9 on 09/18/2018 11:50 am
Does today's test video show a full scale Raptor running?

In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.

Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.

Sounds like we have a 300 bar.

To be slightly more specific, EM called the Raptor a "200-ton class" engine.  He also stated that the chamber pressure is "300 bar, approximately".
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: niwax on 09/18/2018 12:31 pm
Does today's test video show a full scale Raptor running?

In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.

Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.

Sounds like we have a 300 bar.

To be slightly more specific, EM called the Raptor a "200-ton class" engine.  He also stated that the chamber pressure is "300 bar, approximately".

This would also mean a (very roughly) 1.1-1.2 TWR for SSTO, meaning they will have much more commonality between the early hopper and final craft. Earlier on they wanted to use a different engine configuration depending on the tests they're going to run, I could see them moving straight to final production processes and tooling now.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RoboGoofers on 09/18/2018 01:34 pm
I have a feeling that they ran into a big issue with designing the vacuum regen cooled nozzle.

It needed to be very large while still being light enough to not chew up the ISP gains, and still able to handle the thermal flux of the exhaust, which previously Elon has described as 'nuts'.
Could be. There were two alternate reasons given though:
(1) Get started with just one engine design instead of two (with the second only used for 4 out of 38 engines)
(2) More survivable options in the case of engine out.. can survive 4 engines failing and still land.. could be particularly useful early on while working out the bugs!

They might be worried about debris on unprepared ground when landing and taking off. Best to wait until there's a launch pad (which could also take the weight of increased payload) before adding the fragile Rvac nozzles.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/18/2018 01:41 pm
I updated my scale draw with 20180917 announced size.

 ;)
Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 09/18/2018 01:48 pm
Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.

This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).

(added image)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: envy887 on 09/18/2018 01:49 pm
I updated my scale draw with 20180917 announced size.

 ;)
Titus

If you go to dearmoon.earth there's a picture of the full stack with gridfins on the booster.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 09/18/2018 02:21 pm
Does today's test video show a full scale Raptor running?

That’s my assumption. It looked like a different test stand (?), so I assume it was the full size engine. I hope we get some clarification.

With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?

Quote from: Elon
 
45m28s 
So this is the Raptor engine that will power BFR, both the Ship and the Booster, it's the same engine. And this is approximately 200 tonne thrust engine. That's aiming for a roughly 300 bar or three hundred atmosphere chamber pressure. 
And - depending upon ... if you have it at a high expansion ratio - has the potential to be, to have a specific impulse above 380
And it's a staged combustion, full flow, gas-gas. [Praise for the propulsion and other teams.]

 (https://www.youtube.com/watch?v=zu7WJD8vpAQ?t=45m28s)
 
Bold and italics emphasis mine. 
TL;DR: High expansion ratio enables greater than 380 ISP 

Quote from: Tim and Elon

1h34m50s (https://youtu.be/zu7WJD8vpAQ?t=5690) 
Hey Elon, Tim Dodd the Everyday Astronaut here. I see that you changed the engine configuration for the BFS. Can you talk a little about, you know, if there's still engine out capability? Is it vacuum optimized, but still landable on sealevel? Can they function  as an abort system? Can you just kind of tell us about your new decision making on that? 

Oh yeah, actually you noticed that. That's a good thing to notice, good eye.
So in order to minimize the development risk and cost we decided to commonize the engine between the Booster and the Ship
So a future upgrade path for BFS would be to have a vacuum optimized nozzle. [pointing at diagram on wall] (https://i.imgur.com/jduFRIg.jpg) So these nozzles are kind of a sea level sized nozzle so they are able to operate well at sea level, they are essentially the booster sized nozzle. 
Where you see that cargo around the perimiter, you can actually switch out those cargo sections for a vacuum nozzle version of Raptor. And the vacuum nozzle can go all the way to the perimeter, basically the skin of the vehicle. So you can have something which has maybe 3 or 4 times the exit diameter of the raptors that you see there as engines in the perimeter, and the exchange would be you that you'd loose basically two of those cargo racks in exchange for every vacuum engine, but then your total payload performance to Mars would increase significantly. 
But we can do the 100 tonnes to surface of Mars with those engines, but I think version 2 would have the vacuum engines most likely in place of those cargo racks. 

Having those engines in that configuration with 7 engines means it's definitely capable of engine out at any time including two engine out in almost all circumstances. So you could loose two engines and still be totally safe. In fact some cases you can loose up to 4 engines and still be totally fine. It only needs 3 engines for landing, 3 out of 7.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 09/18/2018 03:53 pm
Does today's test video show a full scale Raptor running?

In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.

Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.

Sounds like we have a 300 bar.

To be slightly more specific, EM called the Raptor a "200-ton class" engine.  He also stated that the chamber pressure is "300 bar, approximately".

This would also mean a (very roughly) 1.1-1.2 TWR for SSTO, meaning they will have much more commonality between the early hopper and final craft. Earlier on they wanted to use a different engine configuration depending on the tests they're going to run, I could see them moving straight to final production processes and tooling now.

Not sure.  Bet the longer more voluminous BFS plus 3 big fins/wings/legs masses well above the old 85 tonnes
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: JamesH65 on 09/18/2018 04:07 pm
Does today's test video show a full scale Raptor running?

In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.

Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.

Sounds like we have a 300 bar.

To be slightly more specific, EM called the Raptor a "200-ton class" engine.  He also stated that the chamber pressure is "300 bar, approximately".

This would also mean a (very roughly) 1.1-1.2 TWR for SSTO, meaning they will have much more commonality between the early hopper and final craft. Earlier on they wanted to use a different engine configuration depending on the tests they're going to run, I could see them moving straight to final production processes and tooling now.

Not sure.  Bet the longer more voluminous BFS plus 3 big fins/wings/legs masses well above the old 85 tonnes

Remember they no longer need 'dedicated' landing legs, which will reduce weight.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RoboGoofers on 09/18/2018 04:10 pm
I hazard to mention this, but the tail end picture with the cargo pods made me think of a microwave feed horn (http://www.macpassradio.com/pages/15676.htm) or the Bell Labs radio telescope.

Would a vacuum nozzle extension constructed in such a way make any sense? I've seen mention of Horizontal Flow engine bells (http://www.aerospaceweb.org/design/aerospike/outflow.shtml), which would make more sense than discrete trapezoidal/curved bells, insofar as you wouldn't have corners and the walls between bells to deal with.

I thought the speculation before yesterday's presentation that the 'petals' were an articulated vacuum bell was silly, but the notion of using the entire circumference of the ship as a nozzle is not a bad idea even if a bit far fetched, especially considering they aren't working on developing a Rvac currently and won't need one for almost a decade. the vacuum engines also don't have to gimbal.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 09/18/2018 04:29 pm
Would a vacuum nozzle extension constructed in such a way make any sense? I've seen mention of Horizontal Flow engine bells (http://www.aerospaceweb.org/design/aerospike/outflow.shtml)

No.

Quote from: Elon
So you can have something which has maybe 3 or 4 times the exit diameter of the raptors that you see there as engines in the perimeter, and the exchange would be you that you'd loose basically two of those cargo racks in exchange for every vacuum engine,

Using measurements off the wall banner I can only fit diameters of 200% (or 250% if taken out beyond the edge of cargo pods). It seems when speaking of 3 to 4 fold diameter increase in the quoted answer a mistake was made and area increase was meant in stead.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: HVM on 09/18/2018 05:58 pm
Is there something odd with methlox and vacuum nozzles overall? First Blue Origin shelved development of a vacuum-optimized BE-4, and now SpaceX pushed VacRaptor, in the future...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Tomness on 09/18/2018 06:31 pm
Is there something odd with methlox and vacuum nozzles overall? First Blue Origin shelved development of a vacuum-optimized BE-4, and now SpaceX pushed VacRaptor, in the future...

Maybe because some of propulsion engineers that worked on Raptor also worked on BE-4.... coincidence i think not, JK, see consolidated raptor upperstage thread, if it not needed right now, than launch with out and bring it in later.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 09/18/2018 06:52 pm
Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.

This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).

(added image)
Looks still slightly overexpanded in image. Look at flame dia. at mach disk, it looks smaller than nozzle dia. So looks like flight spec. nozzle on Raptor being tested.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Doesitfloat on 09/18/2018 07:10 pm
At startup I didn't see the green flash.
Looks like they now have spark ignition working.  ;)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 09/18/2018 07:13 pm
Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.

This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).

(added image)
Looks still slightly overexpanded in image. Look at flame dia. at mach disk, it looks smaller than nozzle dia. So looks like flight spec. nozzle on Raptor being tested.
 

This is of great general education interest. The simple public illustrations of over versus under expansion by visual inspection overlook this detail of complex plume features. 
How does mach disc diameter relate to nozzle exit diameter and pressure?

Maybe this question and answer belong in the questions and answers forum for better discoverability?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 09/18/2018 07:23 pm
At startup I didn't see the green flash.
Looks like they now have spark ignition working.  ;)

Here is the Raptor startup in GIF form...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 09/18/2018 07:51 pm
Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.

This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).

(added image)
Looks still slightly overexpanded in image. Look at flame dia. at mach disk, it looks smaller than nozzle dia. So looks like flight spec. nozzle on Raptor being tested.
 

This is of great general education interest. The simple public illustrations of over versus under expansion by visual inspection overlook this detail of complex plume features. 
How does mach disc diameter relate to nozzle exit diameter and pressure?

Maybe this question and answer belong in the questions and answers forum for better discoverability?
Please watch the full video of the Raptor test. You see it throttle down a few seconds after start and the mach disk dia. decreases with decreasing chamber pressure. Mach disk to nozzle dia. ratio is dependent on the nozzle exit pressure if Pe does not equal Pa. Optimal expansion will yield no shock diamonds. During throttle down shock diamonds are always present indicating that the nozzle is still overexpanded at full thrust. If nozzle is underexpanded at full thrust then it will reach optimum expansion at a specified thrust setting then go overexpanded as throttle is further lowered. Going from underexpanded to optimal to overexpanded during throttle down you would very briefly see the shock diamonds disappear before returning again.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: russianhalo117 on 09/18/2018 08:41 pm
At startup I didn't see the green flash.
Looks like they now have spark ignition working.  ;)
Laser ignition was AFAIK still on the table.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: niwax on 09/18/2018 09:25 pm
At startup I didn't see the green flash.
Looks like they now have spark ignition working.  ;)

Here is the Raptor startup in GIF form...

Is that a pilot methalox torch before the vapor from the full propellant flow? That would indicate spark ignition
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: groundbound on 09/18/2018 09:32 pm
At startup I didn't see the green flash.
Looks like they now have spark ignition working.  ;)

Here is the Raptor startup in GIF form...

Is that a pilot methalox torch before the vapor from the full propellant flow? That would indicate spark ignition

Alternatively, could it be something to do with pump spin-up?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: matthewkantar on 09/19/2018 01:52 am
After Mr Musk showed the Raptor engine test and praised the SpaceX propulsion team, he said this:

"I don't think most people, even in the aerospace industry, like, know what question to ask, like it took us a long time to even frame the question correctly.  Look, once we could frame the question correctly, the answer was, I wouldn't say easy, but, the answer flowed once the question could be framed with precision. Framing that question with precision was very difficult."

Anyone here have any idea of what he was talking about? Is this about fundamentals; propellant selection, cycle choice, on what scale the engine will be built?

My guess would be it took a long time come up with a way to have the project coupled with economics size the engine correctly. Not a confident guess.

Matthew
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: titusou on 09/19/2018 05:22 am
After Mr Musk showed the Raptor engine test and praised the SpaceX propulsion team, he said this:

"I don't think most people, even in the aerospace industry, like, know what question to ask, like it took us a long time to even frame the question correctly.  Look, once we could frame the question correctly, the answer was, I wouldn't say easy, but, the answer flowed once the question could be framed with precision. Framing that question with precision was very difficult."

Anyone here have any idea of what he was talking about? Is this about fundamentals; propellant selection, cycle choice, on what scale the engine will be built?

My guess would be it took a long time come up with a way to have the project coupled with economics size the engine correctly. Not a confident guess.

Matthew
My bet: staging timing choice have direct impact on difficulty of recovering.

Look at the BFR moon trip timeline, 1st stage separation 2mins 51secs (171s). That's roughly the same with Faclon9 style booster-stage, instead of Atlas5/Delta4/H2/Arian5 style sustainer-stage design.

Falcon9 SES-10, MECO @ 160s 8,200kph 64km alt
DeltaIV WGS-9, MECO @ 240s, 17,000kph 178km alt
AtlasV WorldView-4, MECO @ 248s, 16,000kph 148km alt

That is major design choice for recovery. The higher/faster 1st stage goes, the more difficulty it will be for recovery. This just simple physics...

Then look back to rocket history... How many booster-stage design we have at core after Saturn5? Almost none other than Falcon9.

And even now... after SpaceX show that booster-stage design is the way to go for 1st stage recovery, we still have how many on-going new rocket design doing sustainer-stage design?

 8) 8) 8) 8) 8) 8) 8) 8) 8)

Titus
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: niwax on 09/19/2018 06:32 am
I think he might also be talking about having one engine type and the number of engines they ended up with. In terms of actually being able to be produced, BFR has made enormous progress since the last announcement. They can test the final engine configuration on the ground, including all the mounting points instead of producing an entirely different octaweb (heptaweb?) for hop tests.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 09/19/2018 07:14 am
Comments about which question to ask came right after extending praise from propulsion to other teams etc. Most likely referring to, business case and architecture level problem solving.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 09/19/2018 07:33 pm
The question could have been:

How to find out, how to start up such an engine, without blowing up 1000’s of engines first.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: GalacticIntruder on 09/19/2018 07:50 pm
BFS and Raptors Optimized for Point to Point Travel, and not Mars.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 09/19/2018 08:12 pm
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 09/19/2018 10:29 pm
BFS and Raptors Optimized for Point to Point Travel, and not Mars.

ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.

A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.

If P2P actually works, that means that propellant in LEO costs $10/kg. (For flights to be economic at 1000 passengers * $1000 'cheaper than economy fare').
It means fuel in solar orbit near the earth is $40/kg. (If you have a sufficient number of tankers, and a large tank).

It means that for one third or so of the cost of the BFS, you can launch nominally to Mars 2.9km/s faster (230 tons MOI)  than a vehicle with ISP 380, compared to an 'unoptimised' one with ISP=350.

The difference anyway is only some 700m/s.
One tanker with some 400 tons of fuel in the same initial orbit can make that up.

Note also, that the most recent version of the BFS has the capability to swap out vacuum for SL engines.
You only need to do one or two, and you're back at 'optimum' engine performance.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TrueBlueWitt on 09/19/2018 11:03 pm
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?

Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 09/19/2018 11:15 pm
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 09/19/2018 11:20 pm
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?

Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.

The base of the BFB flares out. The diameter is close to 10m around the bottom.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: TrueBlueWitt on 09/19/2018 11:26 pm
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?

Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.

The base of the BFB flares out. The diameter is close to 10m around the bottom.

Missed that. 42 @1.3m fit in 10m
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 09/20/2018 12:21 am
My tweak of Lars-J's image...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Rocket Surgeon on 09/20/2018 12:29 am
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?

Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.

The base of the BFB flares out. The diameter is close to 10m around the bottom.

Missed that. 42 @1.3m fit in 10m

So the Sea Level Nozzle may very well have stayed the same?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lemurion on 09/20/2018 12:39 am
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19

I wonder why they “kind of have to” add more engines later.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 09/20/2018 12:47 am
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19

I wonder why they “kind of have to” add more engines later.

EM and SpaceX itirate and evolve so much as they learn it will likely evolve more.  Best to build what you need as a minimum and start flying.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: KelvinZero on 09/20/2018 01:17 am
I wonder why they “kind of have to” add more engines later.
Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.

As one specific example, I wondered how this drop from 150 to "100+" would affect the target of being able to perform all F9 and FH missions. You need huge margins to be able to do what an expendable FH does, while remaining reusable, without refueling.

Actually, it just occurred to me, maybe this drop in performance could actually be a preparation to exceed the 150 ton figure. They did increase the volume after all. What if they had decided to oversize the upper stage, even though this would drop performance in the short term, so that more iterative performance could be delivered later while not making any future radical changes to the very difficult upper stage.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 09/20/2018 01:24 am
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk <span class="emoji-outer emoji-sizer"><span class="emoji-inner" style="background: url(chrome-extension://immhpnclomdloikkpcefncmfgjbkojmh/emoji-data/sheet_apple_32.png);background-position:97.94359576968273% 95.94594594594594%;background-size:5418.75% 5418.75%" data-codepoints="2714-fe0f"></span></span> @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19

I wonder why they “kind of have to” add more engines later.

Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: QuantumG on 09/20/2018 01:35 am
Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.

BRB. Elon used it, and I think it's cool. Big Rocket Booster but also "be right back". It's neat. Use it!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: lonestriker on 09/20/2018 01:37 am
I wonder why they “kind of have to” add more engines later.
Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.

As one specific example, I wondered how this drop from 150 to "100+" would affect the target of being able to perform all F9 and FH missions. You need huge margins to be able to do what an expendable FH does, while remaining reusable, without refueling.

Actually, it just occurred to me, maybe this drop in performance could actually be a preparation to exceed the 150 ton figure. They did increase the volume after all. What if they had decided to oversize the upper stage, even though this would drop performance in the short term, so that more iterative performance could be delivered later while not making any future radical changes to the very difficult upper stage.

Isn't Elon just referencing Hitchhiker's Guide to the Galaxy here?  31+11 = 42 engines.  If he's going to name his Mars ship Heart of Gold, that's why he has to add more engines to make it 42.  It's not a technical reason, it's an Elon reason.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 09/20/2018 01:48 am
I wonder why they “kind of have to” add more engines later.
Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.

As one specific example, I wondered how this drop from 150 to "100+" would affect the target of being able to perform all F9 and FH missions. You need huge margins to be able to do what an expendable FH does, while remaining reusable, without refueling.

Actually, it just occurred to me, maybe this drop in performance could actually be a preparation to exceed the 150 ton figure. They did increase the volume after all. What if they had decided to oversize the upper stage, even though this would drop performance in the short term, so that more iterative performance could be delivered later while not making any future radical changes to the very difficult upper stage.

Isn't Elon just referencing Hitchhiker's Guide to the Galaxy here?  31+11 = 42 engines.  If he's going to name his Mars ship Heart of Gold, that's why he has to add more engines to make it 42.  It's not a technical reason, it's an Elon reason.
Missed it originally....  Brilliant!

-----
ABCD: Always Be Counting Down

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 09/20/2018 02:34 am
Dual bell nozzles look like a pretty good solution for BFS, Can have near optimum high expansion ratio nozzle that operates after BFS separates from BFB in vacuum, with a smaller nozzle inside it with near optimum expansion for throttled landing performance (assuming need down to 20% throttling at landing to allow 3 engines to run at same time).

After early hop-tests a dual expansion nozzle should never need to transition from low expansion to high expansion nozzle mode in flight, as only operate low in atmosphere. That should help with reducing potential for asymmetric fluctuating flow separation thrust loads on dual nozzle.

A couple of good papers on dual bell designs for Vulcain2.1 on (free) Research Gate:
"Ariane 5 Performance Optimization Using Dual-Bell Nozzle Extension"
"SEA-LEVEL TRANSITIONING DUAL BELL NOZZLES"

To give a sense of nozzle masses: The Vulcain nozzle is about 450kg and Ø2.1m diameter (1350kN thrust engine).  Anyone know what SSME nozzle mass was?

Assuming mass scales with surface area scaling suggests Ø1.3m Raptor nozzles around 170kg and doubling their size from Ø1.3 to Ø2.4m would add about 360kg per engine (2.5 tonnes for 7 engines) with about 5% increase in Isp (356=>375s).  If that is so then seems like maximum sized vacuum engines will happen at some point.   using same scaling a Ø3m nozzle would mass about 900kg.  Might put one or two of those in (if possible) for a Lunar or Mars mission.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Peter.Colin on 09/20/2018 05:18 am
I kinda wondered why it was not 42 anymore (I don’t loose a bet often)

Luckily Elon is stil Elon, and obeys the unwritten rules, he has to :) :) :)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 09/20/2018 06:38 am
42 was on the larger original "ITS" version.  Larger diameter and had room for 42.  31 was chosen because of the 9m diameter, and because the total thrust could be handled by the existing launch pads 39A or B that NASA used for Saturn V and SpaceX now leases one for Falcon Heavy.  This was said in other threads on the BFR/BFS in case you missed it.  The 9m version with 31 booster engines fit most of the existing infrastructure and manufacturing capabilities of SpaceX.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 09/20/2018 08:19 am
This is starting to sound like the 9m BFR is a pathfinder Block (a) and possible EELV-2 entry, and an upcoming Block (x) returns to the 12m ITS scale.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 09/20/2018 09:20 am
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19

A taller stretched BRB is the cheapest future growth path, 31=>42 engines probably adds another 50 tonnes payload to LEO, which will be important for economics of refuelling.  Perhaps that is why they have grown the BFS cargo volume slightly.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 09/20/2018 09:46 am
BFS and Raptors Optimized for Point to Point Travel, and not Mars.

ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.

A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.

No, it cannot.

If the normal BFS LEO capacity is 100 tonnes, a tanker which is exactly 25 tonnes lighter but has exactly 25 tonnes bigger tanks lifts exactly 125 tonnes of fuel to orbit. And if the tanks are of same size, then it's ls than 125 tonnes.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 09/20/2018 09:51 am
I wonder why they “kind of have to” add more engines later.
Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.
Note 100+ includes 150.

If you assume 85 tons dry, you need close to 130 payload with sea level raptors to get VIPs round the moon without retanking.

Either BFS is 60 tons dry, or that 100+ is intended by 2023 to be closer to 130+, or refueling in orbit is part of the mission perhaps elided to conceal that it's a landing mission (as 1 retanking vs 8 is basically as hard).
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 09/20/2018 09:56 am
BFS and Raptors Optimized for Point to Point Travel, and not Mars.

ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.

A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.

No, it cannot.

If the normal BFS LEO capacity is 100 tonnes, a tanker which is exactly 25 tonnes lighter but has exactly 25 tonnes bigger tanks lifts exactly 125 tonnes of fuel to orbit. And if the tanks are of same size, then it's ls than 125 tonnes.
I am using prior generations figures, as we do not have a useful dry mass for the present version. If you calculate it strictly according to the things said in the announcement, as outlined above, you get around 60, which seems very, very optimistic.

The above refers to a tanker filled in space, not launching from earth.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rsdavis9 on 09/20/2018 01:36 pm
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk <span class="emoji-outer emoji-sizer"><span class="emoji-inner" style="background: url(chrome-extension://immhpnclomdloikkpcefncmfgjbkojmh/emoji-data/sheet_apple_32.png);background-position:97.94359576968273% 95.94594594594594%;background-size:5418.75% 5418.75%" data-codepoints="2714-fe0f"></span></span> @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19

I wonder why they “kind of have to” add more engines later.

Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.

One guess is the heat shield material weighing more than planned.

If they go for non ablative shields they are all heavier than ablative.(correct me if I am wrong).
Examples:
1. carbon-carbon
2. ceramic tiles
3. metallic x-33 style
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ThomasGadd on 09/20/2018 01:46 pm
Wasn't the 150 tons using SL and vacuum engines?  Would changing to SL engines only account for the difference?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 09/20/2018 01:47 pm
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk <span class="emoji-outer emoji-sizer"><span class="emoji-inner" style="background: url(chrome-extension://immhpnclomdloikkpcefncmfgjbkojmh/emoji-data/sheet_apple_32.png);background-position:97.94359576968273% 95.94594594594594%;background-size:5418.75% 5418.75%" data-codepoints="2714-fe0f"></span></span> @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018

https://twitter.com/elonmusk/status/1042525258899550209?s=19

I wonder why they “kind of have to” add more engines later.

Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.

One guess is the heat shield material weighing more than planned.

If they go for non ablative shields they are all heavier than ablative.(correct me if I am wrong).
Examples:
1. carbon-carbon
2. ceramic tiles
3. metallic x-33 style
x33’s TPS was pretty lightweight.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 09/20/2018 02:06 pm
BFS and Raptors Optimized for Point to Point Travel, and not Mars.

ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.

A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.

No, it cannot.

If the normal BFS LEO capacity is 100 tonnes, a tanker which is exactly 25 tonnes lighter but has exactly 25 tonnes bigger tanks lifts exactly 125 tonnes of fuel to orbit. And if the tanks are of same size, then it's ls than 125 tonnes.

I think speedevil is saying if the tanker is in orbit and was refueled by another tanker.

1800 m/s = 350 * 9.8 * ln(m0/mf)

I get m0/mf = 1.69 or mf/m0 = 0.59

41% of 1185 is 486 tons used for the first 1800 m/s maneuver.

Offload 550 tons and we are at 149 tons remaining. Solving for dry mass after another 1800 m/s gives 88 tons and the tanker should be lighter, if anything.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Cheapchips on 09/20/2018 06:34 pm

The dearmoon raptor burn was approx 1m11 in length, if that's of any interest.  They didn't show it shut off.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: deruch on 09/20/2018 08:28 pm
Folks, I know it's easy to get sidetracked as the topics are obviously related.  But this thread is supposed to be strictly for discussion of the Raptor engine and information about it's evolution.  Discussions of the vehicles (i.e. ITS, BFR, BFS, etc.) or missions/ConOps/Capacities to various orbits or locations/Materials used in heat shields/mass fractions/general, non-Raptor info discussed or released in the recent BFR event/etc. are all OFF TOPIC unless strictly related to the design and specifications of the Raptor.  In fact, we have like 4 other, very hot threads where these would all be ON topic, so please try to limit discussion of those things to there. 

If someone, in a comment that is mostly about the Raptor, includes an interesting aside about which discussion would be off topic in this thread, then you can quote them and actually post the reply in a different, and more appropriate, thread.  That will improve things in 2 ways.  First, Raptor information/discussion won't be diluted and distracted by a whole string of OT replies; and second, other threads won't miss out on interesting discussions that actually belonged there but readers didn't know to look in the Raptor thread to find it.

Thank you for your effort going forward.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: rockets4life97 on 09/20/2018 09:41 pm
Has anyone analyzed the new raptor firing video? Does it look similar to the one from last year or can you tell any differences?

I have zero expertise on rocket engines, hence the question.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 09/20/2018 09:48 pm
Has anyone analyzed the new raptor firing video? Does it look similar to the one from last year or can you tell any differences?

I have zero expertise on rocket engines, hence the question.
It appears to be the same test stand with a different engine
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 09/21/2018 02:04 am
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.
Isp is not unimportant if you are interested in whether there could be a version of the BFS that could do SSTO with a meaningful payload. Given that the market for large GEO sats is in decline and the market for small LEO sats is increasing, this could be relevant to SpaceX in the future. The 7 sea level engines would be able to get the fully loaded thing off the ground at over 1.2g. But now the vacuum Isp may end up being the limiting factor. This is why it is relevant.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: su27k on 09/21/2018 03:40 am
Public confirmation that it's a new engine: https://www.teslarati.com/spacex-raptor-engine-flight-readiness-bfr-spaceship-testing/

Quote
Most recently, photos captured earlier this summer showed that a new prototype was installed on SpaceX’s horizontal Raptor test stand in McGregor, Texas, looking nearly identical to the deep black Raptor nozzle shown in Monday’s presentation. Previous Raptor prototypes seen during testing or at the test stand appeared to have a nozzle closer to SpaceX’s silver Merlin 1Ds, whereas this newest iteration’s nozzle doesn’t seem to reflect the powerful spotlights surrounding it.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: spacenut on 09/21/2018 04:14 am
Is the new Raptor engine design similar to the SSME that can go from sea level to vacuum?  I know it looses ISP at altitude, but is that the trade off for cost and simplicity. 
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 09/21/2018 04:22 am
Is the new Raptor engine design similar to the SSME that can go from sea level to vacuum?  I know it looses ISP at altitude, but is that the trade off for cost and simplicity.
Yes, and basically every sea level engine can do that...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 09/21/2018 10:45 am
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.
Isp is not unimportant if you are interested in whether there could be a version of the BFS that could do SSTO with a meaningful payload. Given that the market for large GEO sats is in decline and the market for small LEO sats is increasing, this could be relevant to SpaceX in the future. The 7 sea level engines would be able to get the fully loaded thing off the ground at over 1.2g. But now the vacuum Isp may end up being the limiting factor. This is why it is relevant.

Sure, and I found that so interesting I started a thread about it (https://forum.nasaspaceflight.com/index.php?topic=45122).

However, it is at best tangentially relevant, unless it makes a material difference to your plans.

If you are considering the difference between $5M for BFS+BFR and somewhat under $5M for BFS alone, it is from a business perspective, marginally relevant if you're selling launches for 40M+.

ISP is similarly marginally relevant for Mars - until it comes time to do SSTO with hard-won ISRU propellant.

Being able to use soft cryogenic fuels, easing the heat gain problems on the tank, with a shared liquid range, easing the inter-tank and piping insulation concerns is the real benefit of Raptor over many engines. (LH2, or Kero fuel). Lack of TEA/TEB allowing unconstrained restarts is another nice feature.

ISP - in a system designed for rapid cheap retanking - not so much.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeM8891 on 09/21/2018 04:07 pm
Is the new Raptor engine design similar to the SSME that can go from sea level to vacuum?  I know it looses ISP at altitude, but is that the trade off for cost and simplicity.
Yes, and basically every sea level engine can do that...
Agreed. All rocket engines increase Isp with altitude. I have calculated that the exhaust pressure of the SL raptor is likely about 0.6 atm. This is equivalent to the engine being optimized for an altitude of ~4000 m or ~13,000 ft.
Title: Raptor Performance
Post by: MikeM8891 on 09/21/2018 05:32 pm
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers.  The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon.  The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design.  It is assumed that all else has remained the same.  It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.

Predicted Raptor Sea Level Engine Performance
Property20172018
Chamber Pressure250 bar300 bar
Thrust at sea level1,700 kN2,095 kN
Thrust in vacuum1,834 kN2,229 kN
Specific Impusle at sea level330 sec335 sec
Specific Impusle in vacuum356 sec356 sec

I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing.  He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS.  He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged.  For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing.  If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).
Title: Re: Raptor Performance
Post by: Elmar Moelzer on 09/21/2018 05:59 pm
I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...
Title: Re: Raptor Performance
Post by: envy887 on 09/21/2018 06:46 pm
I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...

Vac isp depends almost completely on expansion ratio. This is why throttling an engine in vacuum does not really hurt it's specific impulse.
Title: Re: Raptor Performance
Post by: Elmar Moelzer on 09/21/2018 07:30 pm
I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...

Vac isp depends almost completely on expansion ratio. This is why throttling an engine in vacuum does not really hurt it's specific impulse.
I know that the vac Isp largely depends on expansion ratio, but not solely. Chamber pressure should have at least some effect.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Bananas_on_Mars on 09/21/2018 07:49 pm
There's more metrics to a rocket engines performance when in an actual rocket than ISP alone. Going for 300 bar should optimise another metric, which is the thrust to weight ratio.
The theoretical maximum for optimising a rocket nozzle (edit: in vacuum) would be a nozzle of infinite diameter and infinite length. But that might be really bad on the thrust to weight metric.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 09/21/2018 10:19 pm
There's more metrics to a rocket engines performance when in an actual rocket than ISP alone. Going for 300 bar should optimise another metric, which is the thrust to weight ratio.
True and that is visible in the increased thrust. On the other hand, vehicle dry mass seems to have increased by at least 2 tonnes due to the increased length and increased fin size plus extra fin.
Title: Re: Raptor Performance
Post by: Slarty1080 on 09/21/2018 10:42 pm
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers.  The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon.  The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design.  It is assumed that all else has remained the same.  It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.

Predicted Raptor Sea Level Engine Performance
Property20172018
Chamber Pressure250 bar300 bar
Thrust at sea level1,700 kN2,095 kN
Thrust in vacuum1,834 kN2,229 kN
Specific Impusle at sea level330 sec335 sec
Specific Impusle in vacuum356 sec356 sec

I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing.  He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS.  He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged.  For these reasons, predictions on the vacuum Raptor would be too speculative and therefore to worth publishing.  If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).

Perhaps he meant 2-4 x the exit area?
Title: Re: Raptor Performance
Post by: MikeM8891 on 09/21/2018 11:44 pm
I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...

Vac isp depends almost completely on expansion ratio. This is why throttling an engine in vacuum does not really hurt it's specific impulse.
I know that the vac Isp largely depends on expansion ratio, but not solely. Chamber pressure should have at least some effect.
envy887 is right.  Vacuum Isp is a product of the vacuum thrust coefficient and the characteristic velocity.  The vacuum thrust coefficient reflects the exhaust expansion properties and the design of the nozzle; I assumed neither of these changed.  The characteristic velocity is basically a function of the propellants, which also did not change.  I double checked and I did actually calculate a subtle increase in vacuum Isp of 0.36 sec; this 0.1% increase got rounded off the vacuum Isp is identical.
Title: Re: Raptor Performance
Post by: livingjw on 09/22/2018 12:35 am
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers.  The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon.  The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design.  It is assumed that all else has remained the same.  It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.

Predicted Raptor Sea Level Engine Performance
Property20172018
Chamber Pressure250 bar300 bar
Thrust at sea level1,700 kN2,095 kN
Thrust in vacuum1,834 kN2,229 kN
Specific Impusle at sea level330 sec335 sec
Specific Impusle in vacuum356 sec356 sec

I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing.  He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS.  He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged.  For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing.  If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).

Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.

John
Title: Re: Raptor Performance
Post by: MikeM8891 on 09/22/2018 01:45 am
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers.  The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon.  The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design.  It is assumed that all else has remained the same.  It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.

Predicted Raptor Sea Level Engine Performance
Property20172018
Chamber Pressure250 bar300 bar
Thrust at sea level1,700 kN2,095 kN
Thrust in vacuum1,834 kN2,229 kN
Specific Impusle at sea level330 sec335 sec
Specific Impusle in vacuum356 sec356 sec

I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing.  He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS.  He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged.  For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing.  If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).

Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.

John
John,
Great work! I calculated an ER of 32.0-33.6.  My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL.  There is definitely some uncertainty in my calculation.  Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be?  My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.

Thanks, Mike

Title: Re: Raptor Performance
Post by: livingjw on 09/23/2018 12:58 am
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers.  The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon.  The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design.  It is assumed that all else has remained the same.  It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.

Predicted Raptor Sea Level Engine Performance
Property20172018
Chamber Pressure250 bar300 bar
Thrust at sea level1,700 kN2,095 kN
Thrust in vacuum1,834 kN2,229 kN
Specific Impusle at sea level330 sec335 sec
Specific Impusle in vacuum356 sec356 sec

I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing.  He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS.  He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged.  For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing.  If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).

Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.

John
John,
Great work! I calculated an ER of 32.0-33.6.  My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL.  There is definitely some uncertainty in my calculation.  Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be?  My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.

Thanks, Mike

I use NASA's CEA. It is free and on line. I have attached a typical output summary for the main chamber and pre-burners.

John
Title: Re: Raptor Performance
Post by: MikeM8891 on 09/23/2018 02:18 am

Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.

John
John,
Great work! I calculated an ER of 32.0-33.6.  My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL.  There is definitely some uncertainty in my calculation.  Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be?  My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.

Thanks, Mike

I use NASA's CEA. It is free and on line. I have attached a typical output summary for the main chamber and pre-burners.

John
Wow! Thank you, this is amazing! It appears you provided the output for the 2016 raptor. I honestly do not understand why the output Isp does not match the Isp SpaceX advertised, but I am excited to learn more about NASA's CEA. Thank you.
Title: Re: Raptor Performance
Post by: livingjw on 09/24/2018 01:06 am

Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.

John
John,
Great work! I calculated an ER of 32.0-33.6.  My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL.  There is definitely some uncertainty in my calculation.  Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be?  My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.

Thanks, Mike

I use NASA's CEA. It is free and on line. I have attached a typical output summary for the main chamber and pre-burners.

John
Wow! Thank you, this is amazing! It appears you provided the output for the 2016 raptor. I honestly do not understand why the output Isp does not match the Isp SpaceX advertised, but I am excited to learn more about NASA's CEA. Thank you.

CEA just gives you straight chemistry. There are no losses or inefficiencies in it. You have to add those in your model. I estimate combustion efficiency, and nozzle efficiency based on correlation and a few CFD runs.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Zed_Noir on 09/29/2018 07:48 pm
@livingjw

Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: tdperk on 09/29/2018 08:07 pm
@livingjw

Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?

I believe that within a certain small % range of pressure increase, thrust and Isp all go up quite linearly with pressure, that does not whole true for example for a 100% increase.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: aero on 09/30/2018 12:08 am
@livingjw

Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?

I believe that within a certain small % range of pressure increase, thrust and Isp all go up quite linearly with pressure, that does not whole true for example for a 100% increase.

Linearly OK, but at what rate of change, isp/delta-bar? I hope it is 42!
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeM8891 on 09/30/2018 01:30 am
@livingjw

Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?

I believe that within a certain small % range of pressure increase, thrust and Isp all go up quite linearly with pressure, that does not whole true for example for a 100% increase.

Linearly OK, but at what rate of change, isp/delta-bar? I hope it is 42!

Here are some rough estimates. I expect these are accurate to within 2%. Basically the vacuum thrust scales linearly with chamber pressure, something like ~7.4 kN/bar. The sea level thrust is going to be 134 kN less than the vacuum thrust; this is based on the 1.3 m nozzle exit diameter. Vacuum Isp is not affected by chamber pressure.

Chamber Pressure250 bar300 bar315bar330bar
Thrust at sea level1,700 kN2,095 kN2,206 kN2,318 kN
Thrust in vacuum1,834 kN2,229 kN2,340 kN2,452 kN
Specific Impusle at sea level330.0 sec334.6 sec335.6 sec336.5 sec
Specific Impusle in vacuum356 sec356 sec356 sec356 sec
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 09/30/2018 05:17 pm
Looks about about right to me. The only way to increase the vacuum Isp is to increase the expansion ratio. You need about 200:1 er to get close to 380 sec.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: IainMcClatchie on 10/02/2018 02:58 am
Jon, how did you estimate the fuel:oxy mass ratios for the preburners?  Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 10/02/2018 07:03 am
. Vacuum Isp is not affected by chamber pressure.

Chamber Pressure250 bar300 bar315bar330bar
Thrust at sea level1,700 kN2,095 kN2,206 kN2,318 kN
Thrust in vacuum1,834 kN2,229 kN2,340 kN2,452 kN
Specific Impusle at sea level330.0 sec334.6 sec335.6 sec336.5 sec
Specific Impusle in vacuum356 sec356 sec356 sec356 sec

What about improvements in combustion efficiency when running the engine further still from stoichiometric? Does higher chamber pressure enable running even more methane-rich?

Other question: does increased chamber pressure mean decreased throat area? For a theoretical const-mdot engine?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/02/2018 12:15 pm
Jon, how did you estimate the fuel:oxy mass ratios for the preburners?  Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....

I chose 1000 F to minimize the pressure drop across the turbines needed to drive the pumps. The higher the turbine inlet temperature, the lower the pressure drop, the lower the overall pressure rise required of the pumps. Turbine materials are able to handle 100's of degrees higher temperatures, uncooled.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 10/02/2018 12:27 pm
. Vacuum Isp is not affected by chamber pressure.

Chamber Pressure250 bar300 bar315bar330bar
Thrust at sea level1,700 kN2,095 kN2,206 kN2,318 kN
Thrust in vacuum1,834 kN2,229 kN2,340 kN2,452 kN
Specific Impusle at sea level330.0 sec334.6 sec335.6 sec336.5 sec
Specific Impusle in vacuum356 sec356 sec356 sec356 sec

What about improvements in combustion efficiency when running the engine further still from stoichiometric? Does higher chamber pressure enable running even more methane-rich?

Other question: does increased chamber pressure mean decreased throat area? For a theoretical const-mdot engine?

- Once you distance your mixture ratio from stoichiometric (3.7 should do it), high combustion efficiencies are easier to achieve. More distance from stoichiometric will probably not help.
 
- If you want to maintain constant mass flow with increasing pressure, you will have to decrease throat size. You usually wouldn't have to do this since increased pressure from the pumps (spinning faster) would normally increase mass flow.

John
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: acsawdey on 10/02/2018 02:22 pm
Jon, how did you estimate the fuel:oxy mass ratios for the preburners?  Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....

I chose 1000 F to minimize the pressure drop across the turbines needed to drive the pumps. The higher the turbine inlet temperature, the lower the pressure drop, the lower the overall pressure rise required of the pumps. Turbine materials are able to handle 100's of degrees higher temperatures, uncooled.

John

Evolution of Rolls-Royce air-cooled turbine blades and feature analysis (https://ac.els-cdn.com/S1877705814038065/1-s2.0-S1877705814038065-main.pdf?_tid=2a885133-3a4d-4cc5-b8ea-0f5c8b4d6a3b&acdnat=1538489242_c8deeea0cade4657c4a01d8d68b7ac9e)

Even the materials in the very earliest turbojets could handle 1000K (1340F) turbine inlet temperature. To me this feels like one of the big advantages of full-flow -- the temperatures are low compared to gas turbines.


Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 10/07/2018 09:04 pm
Here are some rough estimates. I expect these are accurate to within 2%. Basically the vacuum thrust scales linearly with chamber pressure, something like ~7.4 kN/bar. The sea level thrust is going to be 134 kN less than the vacuum thrust; this is based on the 1.3 m nozzle exit diameter. Vacuum Isp is not affected by chamber pressure.

Chamber Pressure250 bar300 bar315bar330bar
Thrust at sea level1,700 kN2,095 kN2,206 kN2,318 kN
Thrust in vacuum1,834 kN2,229 kN2,340 kN2,452 kN
Specific Impusle at sea level330.0 sec334.6 sec335.6 sec336.5 sec
Specific Impusle in vacuum356 sec356 sec356 sec356 sec
What would 300, 315 and 330 bar look like with 2.4 m and 1.7 m exit diameter bells?
I assume that the higher chamber pressure would improve the performance of larger ER nozzles at SL as well? Or am I wrong here?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Hominans Kosmos on 10/09/2018 12:23 pm
Isn't 1.7 m too large?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Elmar Moelzer on 10/10/2018 05:14 am
Isn't 1.7 m too large?
I thought they would have nozzles with up to 2.4 meters diameter for the vac version?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: RobLynn on 10/11/2018 09:53 am
Isn't 1.7 m too large?

SSME has a huge Ø2.3m nozzle with 1860kN thrust at 363s SL ISp and a huge 77.5 expansion ratio.  Mass flow not that much different from the Raptor, and chamber pressure only 20.5MPa

SSME used a hooked trailing edge to increase pressure near nozzle rim to 30-40kPa far above the 14kPa found in centre of the outlet flow.

There is obviously a lot of scope for further optimisation of Raptor nozzles for both booster and spaceship.  I think the Ø9m booster suffers badly from being too small a diameter.  If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.  Or they might prefer to run even more (61 engines) at significantly lower pressures to improve thrust chamber life and minimize lifetime costs.  But at the moment getting it flying as cheaply as possible is for sure the right approach.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/11/2018 10:09 am
There is obviously a lot of scope for further optimisation of Raptor nozzles for both booster and spaceship.  I think the Ø9m booster suffers badly from being too small a diameter.
The booster tail seems to not be 9m.
Most of the flare visible in some images is due to the fins, but not all.
(https://forum.nasaspaceflight.com/assets/46395.0/1510970.jpg)
It's around 10m.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 10/11/2018 04:25 pm
I think the Ø9m booster suffers badly from being too small a diameter.  If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.

I agree on the problems with the smaller diameter.  There appears to be advantages to increasing the diameter beyond 9 meters.

Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.

Maybe make the 1.0 model of the BFS at 9 meters and get flying.  But why not order another tool.

That's my 2 cents from my armchair.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/11/2018 05:13 pm
I think the Ø9m booster suffers badly from being too small a diameter.  If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.

I agree on the problems with the smaller diameter.  There appears to be advantages to increasing the diameter beyond 9 meters.

Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.

Maybe make the 1.0 model of the BFS at 9 meters and get flying.  But why not order another tool.

That's my 2 cents from my armchair.

This is SpaceX.  The minute the 9 m flies, they'll be working on the 12, or the 15.

Why would it be "decades" if even the first design didn't take even a single decade?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 10/11/2018 05:53 pm
I think the Ø9m booster suffers badly from being too small a diameter.  If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.

I agree on the problems with the smaller diameter.  There appears to be advantages to increasing the diameter beyond 9 meters.

Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.

Maybe make the 1.0 model of the BFS at 9 meters and get flying.  But why not order another tool.

That's my 2 cents from my armchair.

This is SpaceX.  The minute the 9 m flies, they'll be working on the 12, or the 15.

Why would it be "decades" if even the first design didn't take even a single decade?

Did SpaceX immediately start working on 4 and 5m diameter Falcon rockets after F9 flew in 2010?

Why would they when it is far easier to stretch the 9m? There is no real benefit for just going to 12m from 9m.

There are ground infrastructure issues that makes going larger than 9-10m difficult. If they plan on using existing launch sites, like 39A. And perhaps ANY land launch site in the continental US. I don't think they will be able to go bigger without a large floating launch infrastructure.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: hkultala on 10/12/2018 01:53 pm
Is the planned expansion ratio still the same as in 2017 version? Or did they make the nozzle of the "sea level optimized" raptor slightly bigger in the 2018 version? (optimizing it for slightly higher than sea level to be better comphromize for vacuum, or optimizing it for higher chamber pressure even on sea level)

Increasing the nozzle size might explain the width increase in the base of the rocket?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/12/2018 02:22 pm
Is the planned expansion ratio still the same as in 2017 version? Or did they make the nozzle of the "sea level optimized" raptor slightly bigger in the 2018 version? (optimizing it for slightly higher than sea level to be better comphromize for vacuum, or optimizing it for higher chamber pressure even on sea level)

Increasing the nozzle size might explain the width increase in the base of the rocket?

I think it's mostly just a protective skirt for the engine bells. The size does not look to have changed much, perhaps a slight increase.

if you look at the 2017 BFR the outer ring of engine bells extends outside the diameter of the rocket, now it is inside.

That being said, I would not be surprised given the increase in pressure if the throat was narrowed somewhat and the ER pushed from ~35 to more like 40.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/12/2018 02:34 pm
Here are my own personal guesses on the changes from Raptor 2017 to Raptor 2018:

Bell diameter: ~1.30m -> ~1.33m
Throat diameter: ~0.22m -> ~0.21m
Expansion ratio: ~35:1 -> ~40:1
chamber pressure: 250bar -> 300bar
ISP (sea level): ~330 -> ~334
ISP (vacuum): ~356 -> ~358
Thrust (sea level): ~1,700kN -> 1,975kN
Thrust (vacuum): ~1,900kN -> 2,117kN
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/12/2018 04:13 pm
I think the Ø9m booster suffers badly from being too small a diameter.  If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.

I agree on the problems with the smaller diameter.  There appears to be advantages to increasing the diameter beyond 9 meters.

Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.

Maybe make the 1.0 model of the BFS at 9 meters and get flying.  But why not order another tool.

That's my 2 cents from my armchair.

This is SpaceX.  The minute the 9 m flies, they'll be working on the 12, or the 15.

Why would it be "decades" if even the first design didn't take even a single decade?

Did SpaceX immediately start working on 4 and 5m diameter Falcon rockets after F9 flew in 2010?

Why would they when it is far easier to stretch the 9m? There is no real benefit for just going to 12m from 9m.

There are ground infrastructure issues that makes going larger than 9-10m difficult. If they plan on using existing launch sites, like 39A. And perhaps ANY land launch site in the continental US. I don't think they will be able to go bigger without a large floating launch infrastructure.
Because their Mars plans are insanely large and they only went to 9 since the leap was too large, required extra infrastructure, etc.

Also because it's in their nature. They can't help it. I don't think they'll ever sit for "decades" on the same launch vehicle.  F9/H lasted exactly one decade.

The 9m is not what they wanted to build, it's what they could practically build.  Some parallels to F1.

I think you'll see a progressive increase in capabilities over 3-4 launch campaigns, and then they'll move on.




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ABCD: Always Be Counting Down

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/12/2018 05:50 pm
The 9m is not what they wanted to build, it's what they could practically build.  Some parallels to F1.
If you are looking at a rapid build up, launching more and more per synod, you cannot assume steady state economics - amortising vehicles over 10 synods is a nonsense if you're doubling every synod.

This means that capital cost is a major limit - and while in principle a 14m (say) rocket might have half the capital cost per ton, this does not make it meaningful in the context of launches costing sub $10/kg to LEO.
(P2P cost of 9m).

The vast majority of supplies to Mars do not (after the first several launches) need to go in 14m rockets, at half the capital cost per ton.
They are just fine in a 8.5m Aluminium tank, at perhaps 5PSI, with passive thermal control and no ECLSS, and just enough attitude control to remain nice and stable for a few hours once near Mars. (launched from, perhaps even herded by, and caught by a 'barn door' BFS.)

You may want large vessels for bulk passenger transport at some time in the future, but these are going to remain a lot more expensive than bulk cargo on-off delivery tanks.

(This is all post 2026, and first ISRU crew return on the nominal plan)
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/12/2018 06:31 pm
The 9m is not what they wanted to build, it's what they could practically build.  Some parallels to F1.
If you are looking at a rapid build up, launching more and more per synod, you cannot assume steady state economics - amortising vehicles over 10 synods is a nonsense if you're doubling every synod.

This means that capital cost is a major limit - and while in principle a 14m (say) rocket might have half the capital cost per ton, this does not make it meaningful in the context of launches costing sub $10/kg to LEO.
(P2P cost of 9m).

The vast majority of supplies to Mars do not (after the first several launches) need to go in 14m rockets, at half the capital cost per ton.
They are just fine in a 8.5m Aluminium tank, at perhaps 5PSI, with passive thermal control and no ECLSS, and just enough attitude control to remain nice and stable for a few hours once near Mars. (launched from, perhaps even herded by, and caught by a 'barn door' BFS.)

You may want large vessels for bulk passenger transport at some time in the future, but these are going to remain a lot more expensive than bulk cargo on-off delivery tanks.

(This is all post 2026, and first ISRU crew return on the nominal plan)
I guess we'll see... Chat again about this 4 years after first BSR Mars-bound flight?

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ABCD: Always Be Counting Down

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: gongora on 10/12/2018 07:07 pm
This isn't the general BFR discussion thread.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DigitalMan on 10/12/2018 07:53 pm
SpaceX current priorities are to get the initial Raptor out of testing and into production and then to certify engines for flight. 

1) I wonder if the initial production engines for test configurations will be different from the engines in the first operational BFS/R vehicles?

2) Given various possible priorities for SpaceX, when would SpaceX begin to work on improvements for Raptor?

3) What will the first Raptor improvements be?  Vacuum version perhaps?  Increased reliability/reduced strain from reuse?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/12/2018 09:11 pm
3) What will the first Raptor improvements be?  Vacuum version perhaps?  Increased reliability/reduced strain from reuse?
I think one of the first long poles in the tent would be adjustments for rapid production.
They are going to want a hundred of these in the nearish future, both for flight, and for testing, and maybe as many as a thousand in the next decade.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 10/14/2018 03:51 pm
(Sorry, if discussed already.) Do we know, how raptors will be gimbaled? Elon was very proud of the kerosene hydraulics of Merlin. It will not work with methane.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Robotbeat on 10/14/2018 04:02 pm
... It will not work with methane.
Says who?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 10/14/2018 04:12 pm
... It will not work with methane.
Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Gliderflyer on 10/14/2018 04:23 pm
... It will not work with methane.
Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?

It also appears to be a compressible liquid at traditional hydraulic system pressures:
https://www.engineeringtoolbox.com/methane-d_1420.html
I haven't looked up how squishy it is yet, but it might make for a bouncy hydraulic system even if you could keep it liquid.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: MikeM8891 on 10/14/2018 04:26 pm
... It will not work with methane.
Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?
The methane in the tanks will be liquid. A hydraulic pump can bring the the methane up to a pressure that should prevent boiling. It is not ideal, but I think it is workable.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/14/2018 04:35 pm
... It will not work with methane.
Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?
The methane in the tanks will be liquid. A hydraulic pump can bring the the methane up to a pressure that should prevent boiling. It is not ideal, but I think it is workable.
Yeah, but can you guarantee that all hydraulic fluid will circulate fast enough?  One stagnant region near an edge of a cylinder, and you have  gas pocket.

Not even stagnant - just a pocket of fluid that didn't circulate back to the tanks.

Yeah, it's not 1:1 from F9 here.  But I see nothing wrong with a dedicated hydraulic system.  It's a long-lived vehicle, no different than an airplane in that respect.

F9 does carry with it some old rocket DNA...

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ABCD: Always Be Counting Down

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 10/14/2018 04:40 pm
... It will not work with methane.
Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?
The methane in the tanks will be liquid. A hydraulic pump can bring the the methane up to a pressure that should prevent boiling. It is not ideal, but I think it is workable.
There are a pile of interrelated problems in this area that smell like they might be amenable to a really clever combined solution.

Autogenous pressurisation heating/vapourisation in the low megawatt range.
Compression and heating of gas for the RCS to ~room temperature and 3000PSIish.
Actuation for the fins and gimbals.
Possible spot-cooling of critical points at reentry.
Rapid start high pressure gas reservoir for the turbopumps. (being able to start up very very rapidly would be nice for contingencies)

Of course, really clever combined solutions have fun failure modes that dumb ones don't.

Gimbal requirements need to also bear in mind the RCS that can 'cope with 60km/h winds' on landing.
RCS at the top, even quite small, has a massive lever arm.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: geza on 10/14/2018 05:02 pm
Actually, I asked about gimballing, because we discussed actuation of the leg-fins in another thread. Long time ago I read a hint/guess that engine gimballing would be made by electromechanic actuator. However, fin-actuation was envisaged via hydraulics by everybody. Sure, the electric solution would not solve your other problems...
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: deruch on 10/14/2018 05:33 pm
Instead of hydraulics, could they use a pneumatic system and use the methane in gaseous form?
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Gliderflyer on 10/14/2018 09:53 pm
Instead of hydraulics, could they use a pneumatic system and use the methane in gaseous form?
Not easily, because you want the TVC system to be stiff. Hydraulic fluid is incompressible, so when you put a load on the end of an actuator the pressure will increase, but the volume (and therefore position) stays the same. In a pneumatic system the actuator will move until the actuator pressure balances the applied load. Theoretically, if you had an extremely fast active system balancing the pressure on each side of the actuator piston using valves that have all the flow area and instant response time, it might be possible. It would also be the embodiment of "all you have to do is just...".

This is also why using liquid methane won't work. Even with magic insulation to keep things from boiling, the cylinders are dead-headed volumes where the fluid will not be circulated out constantly. The methane will go supercritical if it gets too hot and then the density becomes very nonlinear with respect to pressure and the system loses stiffness.

Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: meekGee on 10/14/2018 09:57 pm
Instead of hydraulics, could they use a pneumatic system and use the methane in gaseous form?
Not easily, because you want the TVC system to be stiff. Hydraulic fluid is incompressible, so when you put a load on the end of an actuator the pressure will increase, but the volume (and therefore position) stays the same. In a pneumatic system the actuator will move until the actuator pressure balances the applied load. Theoretically, if you had an extremely fast active system balancing the pressure on each side of the actuator piston using valves that have all the flow area and instant response time, it might be possible. It would also be the embodiment of "all you have to do is just...".

This is also why using liquid methane won't work. Even with magic insulation to keep things from boiling, the cylinders are dead-headed volumes where the fluid will not be circulated out constantly. The methane will go supercritical if it gets too hot and then the density becomes very nonlinear with respect to pressure and the system loses stiffness.
Yup.  But as Geza says, electro-mechanical has potential.  Also electro-mechanical with local closed hydraulic amplification.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 10/17/2018 12:27 am
Jon, how did you estimate the fuel:oxy mass ratios for the preburners?  Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....

I chose 1000 F to minimize the pressure drop across the turbines needed to drive the pumps. The higher the turbine inlet temperature, the lower the pressure drop, the lower the overall pressure rise required of the pumps. Turbine materials are able to handle 100's of degrees higher temperatures, uncooled.

John

Evolution of Rolls-Royce air-cooled turbine blades and feature analysis (https://ac.els-cdn.com/S1877705814038065/1-s2.0-S1877705814038065-main.pdf?_tid=2a885133-3a4d-4cc5-b8ea-0f5c8b4d6a3b&acdnat=1538489242_c8deeea0cade4657c4a01d8d68b7ac9e)

Even the materials in the very earliest turbojets could handle 1000K (1340F) turbine inlet temperature. To me this feels like one of the big advantages of full-flow -- the temperatures are low compared to gas turbines.

The other big advantage of full flow is using the full flow. While Raptor has a higher chamber pressure than both the RD-170/180/190 series and SSME, since it uses the full flow I would not be surprised if pressure at the preburner/compressor outlet was actually lower in the Raptor engine.

Also, the fuel side of the Raptor is probably almost a hybrid expander cycle, could put that regenerative heat flux for good use.

The Raptor is an extremely impressive design in many different ways....
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 11/18/2018 09:40 pm
I wonder how shops/vendors are coping with the increased demand from these megalithic engines. I'm talking about forging, and casting shops that already have typically long lead times. Has SpaceX bought up forges like they have done for machine shops? SpaceX is talking about putting a dozen engines per BFR, and I presume they will still be getting business making the Merlins.

I see a lot of the turbomachinery being castings, and forgings, especially with their proprietary alloys I presume only select few forges will be producing.

Blue's also producing their rocket engines, Aerojet Rocketdyne is as well (well, not so much with the announcement, but SLS RS-25 is still going on)

Forges don't pop up overnight either, and they have demand from other industries as well.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 11/19/2018 10:01 am
I wonder how shops/vendors are coping with the increased demand from these megalithic engines. I'm talking about forging, and casting shops that already have typically long lead times. Has SpaceX bought up forges like they have done for machine shops? SpaceX is talking about putting a dozen engines per BFR, and I presume they will still be getting business making the Merlins.

I see a lot of the turbomachinery being castings, and forgings, especially with their proprietary alloys I presume only select few forges will be producing.

Blue's also producing their rocket engines, Aerojet Rocketdyne is as well (well, not so much with the announcement, but SLS RS-25 is still going on)

Forges don't pop up overnight either, and they have demand from other industries as well.

I'm fairly certain that SpaceX builds the Raptor engine inhouse and a lot of parts are 3D printed
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: speedevil on 11/19/2018 02:26 pm
I'm fairly certain that SpaceX builds the Raptor engine inhouse and a lot of parts are 3D printed
The nozzle is only the large part, in addition, and it's probably not going to be cast.
Title: Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
Post by: DigitalMan on 11/19/2018 03:15 pm
I'm fairly certain that SpaceX builds the Raptor engine inhouse and a lot of parts are 3D printed
The nozzle is only the large part, in addition, and it's probably not going to be cast.

Indeed.  Here is a video of how SpaceX have previously made a nozzle.

https://www.popularmechanics.com/space/rockets/a24245/spacex-rocket-nozzles-musk-tweet/ (https://www.popularmechanics.com/space/rockets/a24245/spacex-rocket-nozzles-musk-tweet/)
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 12/22/2018 10:17 pm
Wow

https://twitter.com/nasaspaceflight/status/1076615264043757569

Quote
While we have you, Elon.... How well is Raptor performing during test stand firings at McGregor? On track to support your Super Heavy/Starship schedule?

https://twitter.com/elonmusk/status/1076616737020231681

Quote
Yes. Radically redesigned Raptor ready to fire next month.

Edit to add:

https://twitter.com/elonmusk/status/1076618077301665793

Quote
Yes, full flow, gas-gas, staged combustion. Will take us time to work up to 300 bar. That is a mad level.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 12/22/2018 10:23 pm
 :o

https://twitter.com/elonmusk/status/1076618886932353024

Quote
You def don’t want electric pumps on a rocket engine! Raptor turbopumps alone need 100,000 horsepower per engine. That’s not a typo.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 12/22/2018 10:33 pm
Well what can I say - I'm getting the redesign blues here.  :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 12/22/2018 10:35 pm
Well what can I say - I'm getting the redesign blues here.  :(

I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Davidthefat on 12/22/2018 10:40 pm
Gotta stay "agile" as they say. If it doesn't work, why bother keep bandaging it? Move on to a design that works. Don't get emotionally attached to a design.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Jcc on 12/23/2018 12:03 am
They must have something to bolt onto the BFS "hopper" by March/April. Could be "Raptor 1A" to be followed by 1B that never flies, then 1C for the first flight hardware, then 1D  and 1D "full thrust". 😄
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 12/23/2018 12:06 am
Possiblities:

 - relocation of the lox pump
 - addition of boost pumps
 - use of lox for some of the cooling. This would remove the need for a separate lox autogenous heat exchanger

Any other ideas?

John
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: DigitalMan on 12/23/2018 12:36 am
Possiblities:

 - relocation of the lox pump
 - addition of boost pumps
 - use of lox for some of the cooling. This would remove the need for a separate lox autogenous heat exchanger

Any other ideas?

John

One of the previous statements was that thrust to weight optimized lower than expected.  Which resulted in the previously shown Raptor.  Perhaps something here has changed to a non-optimal configuration?
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Cinder on 12/23/2018 01:25 am
Well what can I say - I'm getting the redesign blues here.  :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\
Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later?  That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it.  No room for complacency.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: wannamoonbase on 12/23/2018 01:43 am
100,000 HP Whoa 😮
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lar on 12/23/2018 02:25 am
100,000 HP Whoa 😮
I don't think that's the highest ever..  I was pretty sure the RS-25 was higher, but ... "The fuel pump alone delivers as much as 71,000 horsepower, the oxygen pump delivers about 23,000"[1]

1 - https://www.nasa.gov/missions/highlights/webcasts/shuttle/sts111/ssme-qa.html
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: hamerad on 12/23/2018 02:59 am
Crossposting since i think i got the wrong thread
https://twitter.com/elonmusk/status/1076684059827302400

Quote
SpaceX metallurgy team developed SX500 superalloy for 12000 psi, hot oxygen-rich gas. It was hard. Almost any metal turns into a flare in those conditions.

https://twitter.com/elonmusk/status/1076686201061404672

Quote
Our superalloy foundry is now almost fully operational. This allows rapid iteration on Raptor.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 12/23/2018 04:57 am
The Raptor change is hopefully equivalent to the M1C to M1D jump. Which was simpler, cheaper, and more reliable, and also more thrust. :)
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 12/23/2018 05:37 am
Well what can I say - I'm getting the redesign blues here.  :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\
Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later?  That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it.  No room for complacency.
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most, but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.

On the other hand if the breakthrough radically changes the internal layout and geometry at the heart of the engine as well as introducing new materials then they will need to start testing again to gain experience. The more radical the redesign the greater the loss of experience with that design. Still here's hoping, I guess it’s just the SpaceX way.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Lars-J on 12/23/2018 05:50 am
Well what can I say - I'm getting the redesign blues here.  :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\
Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later?  That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it.  No room for complacency.
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most, but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.

On the other hand if the breakthrough radically changes the internal layout and geometry at the heart of the engine as well as introducing new materials then they will need to start testing again to gain experience. The more radical the redesign the greater the loss of experience with that design. Still here's hoping, I guess it’s just the SpaceX way.

They are well aware of any such risks.

I'm not sure if you were around back when SpaceX was upgrading the F9v1.0 to the v1.1. New engines, basically a brand new rocket. "They are changing too much!", was the constant complaint on this forum. Over and over. Some posters insisted that they needed to fly out the CRS contract with the v1.0 - anything else was just breach on contract and the sky was falling too.

But time has proved SpaceX right. It was the new design that allowed them to actually reach the performance and flight rate they originally hoped for, to reach the point now where they have caught up with their order backlog.

So I wouldn't worry. Assume they know what they are doing. And surely it is better to make drastic(?) changes before production has already started?
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: joek on 12/23/2018 05:55 am
... Still here's hoping, I guess it’s just the SpaceX way.

Yes, it is the SpaceX way.  Their previous experience is not lost or forgotten.  They tried.  They learned something... how to make it better.  Now they are making it better.  Absolutely no reason for SpaceX to pause that cycle.  Why should they?[1]  There is no impending NASA or DoD demand or deadline for design freeze or X flights of the same configuration.


[1] Your personal angst notwithstanding.  Put a fork in it.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 12/23/2018 05:56 am
Well what can I say - I'm getting the redesign blues here.  :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\
Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later?  That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it.  No room for complacency.
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most, but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.

On the other hand if the breakthrough radically changes the internal layout and geometry at the heart of the engine as well as introducing new materials then they will need to start testing again to gain experience. The more radical the redesign the greater the loss of experience with that design. Still here's hoping, I guess it’s just the SpaceX way.

They are well aware of any such risks.

I'm not sure if you were around back when SpaceX was upgrading the F9v1.0 to the v1.1. New engines, basically a brand new rocket. "They are changing too much!", was the constant complaint on this forum. Over and over. Some posters insisted that they needed to fly out the CRS contract with the v1.0 - anything else was just breach on contract and the sky was falling too.

But time has proved SpaceX right. It was the new design that allowed them to actually reach the performance and flight rate they originally hoped for, to reach the point now where they have caught up with their order backlog.

So I wouldn't worry. Assume they know what they are doing. And surely it is better to make drastic(?) changes before production has already started?
True and somewhat reassuring, I hope your right.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Coastal Ron on 12/23/2018 06:22 am
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most...

Be careful what you say - those could be fighting words...  ;)

Quote
...but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.

Wasn't it about a year ago that Elon Musk said that they only had about 5% of their staff on BFR/BFS?

It occurs to me that in the meantime they may have shifted more staff to SH & SS, and the additional staff has allowed them to reassess and optimize what had been done by a relatively smaller staff. That morphing we're seeing today is the result of them putting a LOT of additional resources toward SH/SS.

My $0.02
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: woods170 on 12/23/2018 10:34 am
Well what can I say - I'm getting the redesign blues here.  :(

I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.

This.

From what I hear from sources: The previous version of Raptor, which was extensively tested over the past two years was a DEVELOPMENT engine.
During the extensive testing campaign Tom's team learned many, many new things. They also found dozens upon dozens of items to improve on the next, closer-to-flight-design.

That next, closer-to-flight-design is what will be on the test-stand soon.

But the basic principle behind the engine is still the same: full flow, gas-gas methane-lox rocket engine. And yes: the improvements - as confirmed by Elon - are also designed to reach that magical 300 bar number.

This is what Agile rocket engine development is all about: build a minimal viable product (the initial Raptor we have seen) and test the hell out of it. Learn all you can and build-in all the improvements. Test the hell out of it. Learn some more and build an even better one. Than test the hell out of that as well.


Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because it takes the most amount of time to get to something that is ready to do the job.
Rocket Engines Are Hard (TM).

Don't expect the current version of Jeff's BE-4 to be the one that will be on the first New Glenn and Vulcan. It will be a thoroughly improved and enhanced BE-4 as well.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Slarty1080 on 12/23/2018 01:15 pm
Well what can I say - I'm getting the redesign blues here.  :(

I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.

This.

From what I hear from sources: The previous version of Raptor, which was extensively tested over the past two years was a DEVELOPMENT engine.
During the extensive testing campaign Tom's team learned many, many new things. They also found dozens upon dozens of items to improve on the next, closer-to-flight-design.

That next, closer-to-flight-design is what will be on the test-stand soon.

But the basic principle behind the engine is still the same: full flow, gas-gas methane-lox rocket engine. And yes: the improvements - as confirmed by Elon - are also designed to reach that magical 300 bar number.

This is what Agile rocket engine development is all about: build a minimal viable product (the initial Raptor we have seen) and test the hell out of it. Learn all you can and build-in all the improvements. Test the hell out of it. Learn some more and build an even better one. Than test the hell out of that as well.


Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because it takes the most amount of time to get to something that is ready to do the job.
Rocket Engines Are Hard (TM).

Don't expect the current version of Jeff's BE-4 to be the one that will be on the first New Glenn and Vulcan. It will be a thoroughly improved and enhanced BE-4 as well.
You may well be right in which case excellent news! It’s just that scaling a third scale development engine to a full sized engine doesn't sound like a radical redesign, especially when IIRC it was stated that this was always the plan and would not be "that difficult".
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Owlon on 12/23/2018 01:24 pm
You may well be right in which case excellent news! It’s just that scaling a third scale development engine to a full sized engine doesn't sound like a radical redesign, especially when IIRC it was stated that this was always the plan and would not be "that difficult".

It's not just scaling. The development engine(s) we've seen had a very much not flight-like thrust to weight ratio and had a lot of optimization in that department yet to be done. There was some comment about that by Elon within the last year or so, but I'm too lazy to go dig it up.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 12/23/2018 01:57 pm
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.

John
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: RotoSequence on 12/23/2018 02:00 pm
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.

John

If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 12/23/2018 02:40 pm
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.

John

If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.

Are you talking about boost pumps?
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: RotoSequence on 12/23/2018 02:45 pm
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.

John

If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.

Are you talking about boost pumps?

More than two, parallel turbopumps.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: livingjw on 12/23/2018 03:04 pm
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.

John

If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.

Are you talking about boost pumps?

More than two, parallel turbopumps.

They currently use a two stage methane pump and a single stage lox pump. Do you expect that to change?
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: niwax on 12/23/2018 03:06 pm
You may well be right in which case excellent news! It’s just that scaling a third scale development engine to a full sized engine doesn't sound like a radical redesign, especially when IIRC it was stated that this was always the plan and would not be "that difficult".

It's not just scaling. The development engine(s) we've seen had a very much not flight-like thrust to weight ratio and had a lot of optimization in that department yet to be done. There was some comment about that by Elon within the last year or so, but I'm too lazy to go dig it up.

I could see that fit with with a radical redesign. A clean sheet design based on the architecture and learnings of the subscale model. Similar to how you always throw away a software prototype.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: philw1776 on 12/23/2018 03:28 pm
I'd said it previously but I always expected SpaceX to fly prototype upper stages with less Raptor engines in order to catch initial real world flight integration bugs and to factor in learned improvements before building lots of engines such as for populating a full first stage.  I expected the 1st stage to fly early on with ~half the engines or so.  When the design stabilizes enough to build a bunch, they'll build the workhorse orbital rocket ships.

I never, ever expected a flying water tower boilerplate as the engine proof of concept vehicle.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: docmordrid on 12/23/2018 03:41 pm
>
I never, ever expected a flying water tower boilerplate as the engine proof of concept vehicle.

After a dozen years of following SpaceX, and more "W-T-F!?!" moments than I can remember, very little  surprises me anymore - but this one did. Scratch another sprayed keyboard.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: ZachF on 12/23/2018 06:45 pm
100,000 HP Whoa 😮
I don't think that's the highest ever..  I was pretty sure the RS-25 was higher, but ... "The fuel pump alone delivers as much as 71,000 horsepower, the oxygen pump delivers about 23,000"[1]

1 - https://www.nasa.gov/missions/highlights/webcasts/shuttle/sts111/ssme-qa.html

I would bet the RD-170 is the king in terms of Turbopump power since it has both the highest thrust and chamber pressure.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 12/23/2018 07:53 pm
100,000 HP Whoa 😮
= 74.6MW. From this looks like Raptor thrust is back towards 2016 ITS design levels. Would allow Super Heavy engine no. to be reduced to 19.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 12/23/2018 07:54 pm
100,000 HP Whoa 😮
I don't think that's the highest ever..  I was pretty sure the RS-25 was higher, but ... "The fuel pump alone delivers as much as 71,000 horsepower, the oxygen pump delivers about 23,000"[1]

1 - https://www.nasa.gov/missions/highlights/webcasts/shuttle/sts111/ssme-qa.html

I would bet the RD-170 is the king in terms of Turbopump power since it has both the highest thrust and chamber pressure.
You are right, 170MW but a bit OT here.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: billh on 12/24/2018 04:39 pm
One of the challenges of optimization is getting stuck in a local minimum, which might be far from the global minimum. I suspect conservative engineering organizations might have the same problem in design space. Big changes are only a problem if you are not gaining big improvements...or you run out of development funding!
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: FutureSpaceTourist on 12/25/2018 07:18 am
https://twitter.com/andrewsamoylich/status/1077477123911225344

Quote
Are the nozzles of the Raptors built taking into account the need to fly engines forward when entering the atmosphere at the first or second space velocity? That is, will the Raptors be a kind of heat shield?

https://twitter.com/elonmusk/status/1077477602091257857

Quote
No, Raptors must be shielded during atmospheric entry. Although, maybe not ...
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Jcc on 12/25/2018 03:48 pm

"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."

Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 12/25/2018 04:06 pm

"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."

Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
That's the plan- "Raptors must be shielded during reentry."
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Jcc on 12/25/2018 04:36 pm

"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."

Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
That's the plan- "Raptors must be shielded during reentry."
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.

But can they take the dynamic pressure?
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: HMXHMX on 12/25/2018 04:42 pm

"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."

Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
That's the plan- "Raptors must be shielded during reentry."
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.

But can they take the dynamic pressure?

Entry dynamic pressure behind the shock is quite low, far less than what the bell sees during ascent.  The issue is unsteady flow and differential heating.  This can be mitigated by running a fuel bleed thought the cooling channels then out the chamber to lightly pressurize the nozzle.  Even for base on entry, this is a feasible method to limit bell temperatures and relieve differential loads, so it is probably overkill for side entry.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: rakaydos on 12/25/2018 05:09 pm

"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."

Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
That's the plan- "Raptors must be shielded during reentry."
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.

But can they take the dynamic pressure?
That's one of MANY questions for the engineers, when he isnt busy with twitter.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Tuna-Fish on 12/26/2018 12:39 am
That's one of MANY questions for the engineers, when he isnt busy with twitter.

I suspect the  reason he's so busy with twitter is that all his engineers are at home with their families and he's bored. Bored Musk = lots of info for us, so yay.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: cferreir on 12/28/2018 04:06 pm
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: Cinder on 12/28/2018 04:21 pm
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.
Quoting in full instead of linking, because twitter embeds can confuse which post is being linked by scrolling your browser view.
Well what can I say - I'm getting the redesign blues here.  :(

I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.

This.

From what I hear from sources: The previous version of Raptor, which was extensively tested over the past two years was a DEVELOPMENT engine.
During the extensive testing campaign Tom's team learned many, many new things. They also found dozens upon dozens of items to improve on the next, closer-to-flight-design.

That next, closer-to-flight-design is what will be on the test-stand soon.

But the basic principle behind the engine is still the same: full flow, gas-gas methane-lox rocket engine. And yes: the improvements - as confirmed by Elon - are also designed to reach that magical 300 bar number.

This is what Agile rocket engine development is all about: build a minimal viable product (the initial Raptor we have seen) and test the hell out of it. Learn all you can and build-in all the improvements. Test the hell out of it. Learn some more and build an even better one. Than test the hell out of that as well.


Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because it takes the most amount of time to get to something that is ready to do the job.
Rocket Engines Are Hard (TM).

Don't expect the current version of Jeff's BE-4 to be the one that will be on the first New Glenn and Vulcan. It will be a thoroughly improved and enhanced BE-4 as well.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: DJPledger on 12/28/2018 04:21 pm
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.
Star Hopper will most likely be using the dev. version of Raptor not the new version that SpaceX plans to fire next month. Not possible to go from 1st test fire to flight qualification within 3 months.
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: 50_Caliber on 12/28/2018 04:22 pm
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.

Tom Meuller stated back in May that flight-ready Raptor engines were already "in work".

https://www.teslarati.com/spacex-raptor-engine-flight-readiness-bfr-spaceship-testing/ (https://www.teslarati.com/spacex-raptor-engine-flight-readiness-bfr-spaceship-testing/)
Title: Re: Super Heavy (ITS-BFR) Propulsion – The evolution of the SpaceX Raptor engine
Post by: RedLineTrain on 12/28/2018 04:27 pm
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.
Star Hopper will most likely be using the dev. version of Raptor not the new version that SpaceX plans to fire next month. Not possible to go from 1st test fire to flight qualification within 3 months.

In certain circumstances that I'm not sure apply here, it can be done in 3 months, as was done for the Merlin 1C.

https://www.nasa.gov/offices/c3po/home/spacex_merlin_qual_f1.html
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cferreir on 12/29/2018 06:28 am
Woa, there is a huge difference between firing a 2/3 scale engine and a full scale. Then there is a huge jump  from firing an engine, learning something, then building a modified engine and testing that and so forth. Difference between a Merlin 1C and 1D was minor. Has anyone seen/heard a full scale Raptor firing?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 12/29/2018 08:16 am
Woa, there is a huge difference between firing a 2/3 scale engine and a full scale. Then there is a huge jump  from firing an engine, learning something, then building a modified engine and testing that and so forth. Difference between a Merlin 1C and 1D was minor. Has anyone seen/heard a full scale Raptor firing?
Full scale Raptor not fired yet. EM tweeted new Raptor which I assume to be full scale to be fired next month.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Ronsmytheiii on 12/29/2018 03:03 pm
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/

Also when awarded, the contract was for an upperstage for Falcon 9 and Falcon Heavy. Later, SpaceX decided to use it on BFR/Spaceship only:

Quote
This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.

https://dod.defense.gov/News/Contracts/Contract-View/Article/642983/
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JamesH65 on 12/29/2018 04:21 pm
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/

Also when awarded, the contract was for an upperstage for Falcon 9 and Falcon Heavy. Later, SpaceX decided to use it on BFR/Spaceship only:

I don't think that is right, Didn't SpaceX have the rapter under development well before the DOD contract? And I'm pretty sure they have known from the start what it was going to be used for....and its not a DoD upper stage.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 12/29/2018 04:34 pm
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>

The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.

April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)

Which quotes an announcement of intent 6 months earlier, so Q4 2013.

Long before Air Force funding.

Edit: typo, clarity
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 12/29/2018 04:37 pm
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>

That story link was 2016. SpaceX was working on the powerhead at Stemmis in 2014, so decisions were made well before then.

April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)

Which quotes an announcement of intent 6 months earlier, so Q4 2013.

Long before Air Force funding.

And a 6 year development cycle for such an engine is really fast!

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: deruch on 12/29/2018 08:05 pm
Plus there was a period in the early work where Raptor was going to be a Hydralox engine instead of Methalox.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Ronsmytheiii on 12/29/2018 09:04 pm
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>

The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.

April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)

Which quotes an announcement of intent 6 months earlier, so Q4 2013.

Long before Air Force funding.

Edit: typo, clarity

But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. Raptor enabled BFR development, but BFR wasnt the cause of Raptor development. Air force funding and SpaceX matching investment turned the engine from a TRL 3 design into a TRL 7 functioning prototype, but the BFR system only crystallized as Raptor engine development accelerated, not starting at the same point.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: oiorionsbelt on 12/29/2018 09:22 pm

But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. (snip)
Mars colonization vehicle/BFR was ALWAYS the goal. Raptor was simply the first part of that vehicle to be developed. It was always intended for the Mars vehicle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Ronsmytheiii on 12/29/2018 09:32 pm

But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. (snip)
Mars colonization vehicle/BFR was ALWAYS the goal. Raptor was simply the first part of that vehicle to be developed. It was always intended for the Mars vehicle.

It was first meant for a hydrolox upperstage engine, but yes a large rocket has been in consideration for awhile. There is a difference though between a generic push for a big rocket (Falcon X using Merlin 2s) and the BFR architecture that was unveiled at IAC 2016, and that difference was (edit: the AF contract for Falcon family upperstage) Raptor development.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 12/29/2018 09:55 pm
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>

The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.

April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)

Which quotes an announcement of intent 6 months earlier, so Q4 2013.

Long before Air Force funding.

Edit: typo, clarity

But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. Raptor enabled BFR development, but BFR wasnt the cause of Raptor development. Air force funding and SpaceX matching investment turned the engine from a TRL 3 design into a TRL 7 functioning prototype, but the BFR system only crystallized as Raptor engine development accelerated, not starting at the same point.

Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>

The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.

April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)

Which quotes an announcement of intent 6 months earlier, so Q4 2013.

Long before Air Force funding.

Edit: typo, clarity

But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. Raptor enabled BFR development, but BFR wasnt the cause of Raptor development.
>

Musk has been talking about a Mars architecture since the mid 2000's. Initially the architecture was named Mars Colonial Transporter (MCT), and the engine Merlin 2 - then a hydrogen engine.

Merlin 2 was renamed before this 2013 Space News  article a few months after the Stennis ribbon cutting. Link... (http://www.spacenews.com/article/launch-report/37859spacex-could-begin-testing-methane-fueled-engine-at-stennis-next-year)

Quote
The current Raptor concept “is a highly reusable methane staged-combustion engine that will power the next generation of SpaceX launch vehicles designed for the exploration and colonization of Mars,”

The architecture itself was renamed BFR in this 2015 GQ interview. Link... (https://www.gq.com/story/elon-musk-mars-spacex-tesla-interview)

Both of these predate the 2016 USAF contract award.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 12/30/2018 04:00 am
No, it wasn't the 2016 US Air Force ask/contract for SpaceX to do an upper stage methalox engine for Falcon 9 (in 2016) that led SpaceX to start development of a new methalox FFSC engine in 2012.

SpaceX was doing that Raptor development since at least October 2012 (https://www.flightglobal.com/news/articles/spacex-aims-big-with-massive-new-rocket-377687/) ("... won't use Merlin's RP-1 fuel"), and specified it was methalox one month later, in November 2012 (https://web.archive.org/web/20160611083349/http://seradata.com/SSI/2012/11/musk_goes_for_methane-burning/).

As noted above by docmordrid, SpaceX were already doing fairly large scale engine subsystem tests on the powerhead alone by 2014.

The USAF came along in 2016 and hopped on the bandwagon, trying to get in on a little of that methalox Raptor action (gotta have "US engines" for "US national security reasons..." and all that), offering US$33 million to SpaceX, so that SpaceX would continue (after that point) to put double that amount in on the project.  I assume, then, that it is up to USAF contracting officers to monitor the progress of that contract and only pay out if the contract terms are met.  And USAF apparently threw more $$$ on that contract in October 2017 (https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/), to accomplish something else the Air Force/US government ostensibly wanted to achieve. 

(NOTE:  this is a small fraction of the sort of $$$ the US government throws at other rocket engine dev efforts like Aerojet Rocketdyne and Orbital ATK NG.]

It is historical revisionism to say that the answer to woods170's question
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
was merely 'cause
Because the USAF gave them $61.4 million to develop it.

SpaceX started developing that engine in 2012 'cause they have a good propulsion team, and know it typically takes 7 years or so before first flight to develop a major new orbital-class rocket engine, even with Uncle Sam big bucks funding from the beginning, and even when the dev team isn't pushing into new-ish tech (in US terms) for FFSC methalox engines.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: woods170 on 12/30/2018 10:28 am
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?

Because the USAF gave them $61.4 million to develop it:

https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/

I'm afraid you are mistaken. By the time USAF stepped into Raptor development (sometime in early 2016) SpaceX had already spent several years on Raptor development. They have been known to be working on Raptor (and it's earlier designs) from 2011 forward.
USAF was therefore NOT the one to ask SpaceX to start Raptor development.


Also when awarded, the contract was for an upperstage for Falcon 9 and Falcon Heavy. Later, SpaceX decided to use it on BFR/Spaceship only:

Quote
This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.

https://dod.defense.gov/News/Contracts/Contract-View/Article/642983/

Again: incorrect. Raptor, and it's earlier design iterations, have been intended for a Mars-capable vehicle from the get-go. That is: from 2011 forward. You clearly have never heard of the Falcon X, Falcon X Heavy and Falcon XX concepts from 2010. Those preliminary design concepts were for SHLVs capable of putting humans on Mars. SpaceX understood, as early as 2010, that continued upgrades of the Merlin architecture were never going to be sufficient to make the Falcon X-series a reality.
It was that realization that led to what we now know as Raptor engine development.

USAF, in 2014, became aware of what SpaceX was trying to do. Around the same time the ban on military use of RD-180 was beginning to loom. To hedge their bets USAF put out an RFI for replacement of the RD-180.
SpaceX, which was now actively working on Raptor, responded to this RFI by offering the Raptor for USAF use.
But, there was a caveat: SpaceX was not going to allow USAF to use Raptor as a booster stage engine on anything other than a SpaceX rocket. So, the only purpose that Raptor would serve, on a relatively short term, was to improve performance of the upper stage of that other SpaceX vehicle, which was than under development: Falcon Heavy.

Stating that USAF funding took Raptor to TRL-7 and thus leading to BFR/BFS is categorically false. Although the first tangible concept (MCT/ITS) was revealed by Musk in 2016 it had in fact been in development for at least 4 years.
Yes, you read that correctly: a very small team of SpaceX folks have been working on concepts for ITS/MCT/BFR/BFS since 2012. Musk has been teasing details about ITS/MCT/BFR/BFS - and its Raptor engine - since late 2013. That's almost three years before USAF decided to fund Raptor development:

https://spacenews.com/37859spacex-could-begin-testing-methane-fueled-engine-at-stennis-next-year/
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mnelson on 12/30/2018 02:28 pm
And a 6 year development cycle for such an engine is really fast!

In 1988 I had the great privilege of attending an "Introduction to Rocket Engines" lecture by a retired rocket engineer in Huntsville. He was German and worked on the V2 before being brought to the United States after the war to work on engines for the US. Really fascinating as he walked us around to several different engines in the museum and could talk of the issues and fixes they worked through.

I can still hear him saying in his German accent: "It takes 10 years to develop a new engine. People always want it faster but it takes 10 years."
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 12/30/2018 06:36 pm
And a 6 year development cycle for such an engine is really fast!

In 1988 I had the great privilege of attending an "Introduction to Rocket Engines" lecture by a retired rocket engineer in Huntsville. He was German and worked on the V2 before being brought to the United States after the war to work on engines for the US. Really fascinating as he walked us around to several different engines in the museum and could talk of the issues and fixes they worked through.

I can still hear him saying in his German accent: "It takes 10 years to develop a new engine. People always want it faster but it takes 10 years."
New superalloy foundry could allow the next version of Raptor to be dev. in as little as one year. EM said it would allow for rapid iteration of Raptor.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Ronsmytheiii on 12/30/2018 06:53 pm
You clearly have never heard of the Falcon X, Falcon X Heavy and Falcon XX concepts from 2010. Those preliminary design concepts were for SHLVs capable of putting humans on Mars.

And you have clearly not read my posts where I talked about Falcon X. I have been on this forum for over a decade, and heard the first rumblings about a SpaceX SHLV since 2010.

I consider the Falcon X to be almost a completely different design to the BFR architecture unveiled in 2016, and that was developed on the basis of Raptor.

I could keep on, but I think that this has gone far enough. I think there is a fundamental disagreement here on the difference between concepts and actual system development, and will leave it at that.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 12/30/2018 07:13 pm

And you have clearly not read my posts where I talked about Falcon X. I have been on this forum for over a decade, and heard the first rumblings about a SpaceX SHLV since 2010.

I consider the Falcon X to be almost a completely different design to the BFR architecture unveiled in 2016, and that was developed on the basis of Raptor.

I could keep on, but I think that this has gone far enough. I think there is a fundamental disagreement here on the difference between concepts and actual system development, and will leave it at that.

We should explore this Ronsmytheiii.  I am genuinely curious what you think is the "engine system development" that SpaceX is doing here.

Since you put forth the argument that USAF contract $$$ to SpaceX in Jan 2016 (just US$33 million) is what got SpaceX to develop the Raptor engine, and since many serious propulsion engineers/managers have said it takes 7 to 10 years to develop a new rocket engine, ...

how do you account for the simple reality that SpaceX is currently planning on flying a multi-engine suborbital test flight on three Raptor engines in approx. April 2019?   :o  Just three years and a few months after the USAF gave them a contract for a set of upper stage engine milestones.

Is SpaceX just special, and they can develop new orbital class rocket engines on paltry US government funding in a bit over three years?

Or, Occam's razor, do we just accept that SpaceX started quite serious development of the Raptor methalox full-flow staged combustion engine on their own dime, with private capital, in 2012, as all sources seem to indicate?  And that they needed, like most others, at least seven years to get the new engine from start of development to maiden flight?  But also that, in order to get 'er done in just 7 years, they certainly could not wait 'til the government showed up with a few tens of millions of $$ in year 4, and had to basically put substantial development resources on it at the beginning of the effort.

Edit: fixed a typo

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DistantTemple on 12/30/2018 07:39 pm
It might be a wrong assumption, but there seems to a general acknowledgement that woods170 is close to SX, and has an awareness of SX unavailable through only NSF and public media. I  read his comments as if they are likely better informed. However he does a great job of carefully quoting what IS public in careful detail.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: John Alan on 12/30/2018 09:05 pm
Lets also not forget, that Tom Mueller (head engine guy) has been in the Rocket Engine R & D field going way back...

If you don't think he hasn't been been thinking and doodling on FFSC engines since like his college days, you likely don't understand how good engineers think and do with their spare time...

A good working FFSC engine is like THE Holy Grail in that field...
I bet he has spent lots of time thinking on it over the years...

 ;)

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jpo234 on 12/30/2018 09:47 pm
From Tom Mueller himself https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/

Quote
“I’ve been working on Mars for the last four years, so I’m not going to take any credit for the Block 5 engine and all the upgrades that have happened,” he said. “I’ll take credit for developing the team that developed the Merlin 1D engine.”

Quote
Mueller told GeekWire that he’s been mulling over the Raptor for about a decade. The engine doesn’t make use of the Merlin design, but goes instead with a full-flow, staged-combustion system that requires a clean-sheet design.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RedLineTrain on 12/30/2018 10:26 pm
I can still hear him saying in his German accent: "It takes 10 years to develop a new engine. People always want it faster but it takes 10 years."

Note that not all engine development efforts are created equal -- they can differ in the number of seconds fired by two or more orders of magnitude.  These efforts can also be paced by the manpower efficiency of the tests.

It's pretty amazing that Mueller and Co. developed an engine with a more-or-less novel combustion cycle and fuel utilizing novel metallurgy in a half dozen years or less on a comparatively shoe-string budget.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: indaco1 on 12/31/2018 11:30 am
Merlin developement has been quite fast.  We could assume Raptor will take about the same.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Rabidpanda on 12/31/2018 04:26 pm
Merlin developement has been quite fast.  We could assume Raptor will take about the same.

Why? Raptor is much much more complex than Merlin.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 04:40 pm
Do we have a solid, official explanation of why SpaceX is going for both LOX/methane and full-flow staged combustion without easier intermediate steps, such as gas-generator LOX-methane and/or fuel-rich staged combustion?

There have been rumours and a lot of speculation that LOX/methane is needed for reuse without a lot of maintenance, since kerolox suffers from coking problems. Full-flow staged combustion is of course intended to give maximum Isp and T/W.

But how necessary is LOX/methane for reusability and how important are staged combustion and full-flow staged combustion in particular for performance? Hasn't Musk said that current Merlins can be used up to ten times without a lot of maintenance? I'm getting the feeling that people are skeptical about this, but do we have any hard facts?

It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 12/31/2018 04:43 pm
It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.
But probably wouldn't be good enough for a Mars architecture -- not at a size that SpaceX can afford to build.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 12/31/2018 05:50 pm
Do we have a solid, official explanation of why SpaceX is going for both LOX/methane and full-flow staged combustion without easier intermediate steps, such as gas-generator LOX-methane and/or fuel-rich staged combustion?

There have been rumours and a lot of speculation that LOX/methane is needed for reuse without a lot of maintenance, since kerolox suffers from coking problems. Full-flow staged combustion is of course intended to give maximum Isp and T/W.

But how necessary is LOX/methane for reusability and how important are staged combustion and full-flow staged combustion in particular for performance? Hasn't Musk said that current Merlins can be used up to ten times without a lot of maintenance? I'm getting the feeling that people are skeptical about this, but do we have any hard facts?

It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.

I remember seeing a Musk presentation a while back where he was comparing different fuels against a list of criteria he had drawn up for the new engine. Methane won hands down as (from memory) it was easier to handle and store long term than hydrogen, didn't coke like RP1, but above all it could be relatively easily manufactured on Mars unlike RP1.

As for the full flow staged combustion cycle, I believe this is the most efficient cycle so an obvious choice for a billionaire who has hired some of the best smarts in the world and wants the highest performance engine for his Mars spaceship.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 05:59 pm
As for the full flow staged combustion cycle, I believe this is the most efficient cycle so an obvious choice for a billionaire who has hired some of the best smarts in the world and wants the highest performance engine for his Mars spaceship.

Sure, that must be the reason to want it eventually. But given that it's such a long journey to get there, why not do something simpler first? The answer might be that coking isn't as urgent a problem as people are suggesting, because the current Merlins can indeed already handle ~10 launches. On the other hand, there seems to be an urgent desire to get SH and BFS operational as soon as possible.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: AC in NC on 12/31/2018 06:04 pm
he answer might be that coking isn't as urgent a problem as people are suggesting, because the current Merlins can indeed already handle ~10 launches.

IIRC coking is an urgent problem as Merlins can handle ~10 launches but are cleaned between launches.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 12/31/2018 06:34 pm
As for the full flow staged combustion cycle, I believe this is the most efficient cycle so an obvious choice for a billionaire who has hired some of the best smarts in the world and wants the highest performance engine for his Mars spaceship.

Sure, that must be the reason to want it eventually. But given that it's such a long journey to get there, why not do something simpler first? The answer might be that coking isn't as urgent a problem as people are suggesting, because the current Merlins can indeed already handle ~10 launches. On the other hand, there seems to be an urgent desire to get SH and BFS operational as soon as possible.

Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 06:43 pm
Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.

You're not answering my question. I didn't say full-flow staged combustion was impossible, just that it was hard and (for whatever reason) they appear to be in a hurry. We already know why LOX/methane was chosen and why full-flow staged combustion is awesome. The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.

Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious what that reason is.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 12/31/2018 07:05 pm
Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.

You're not answering my question. I didn't say full-flow staged combustion was impossible, just that it was hard and (for whatever reason) they appear to be in a hurry. We already know why LOX/methane was chosen and why full-flow staged combustion is awesome. The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.

Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious what that reason is.

They are working incrementally, though. They don't have to do all the possible increments. They are starting with a smaller, lower pressure version of Raptor and improving it.

Also, they want to get to FFSC eventually, and doing methalox GG doesn't really get them there any faster.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 12/31/2018 07:13 pm

The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry?


'Cause he hired, and then built over time, a super top-notch propulsion engineering team, and they thought they could build a methalox FFSC engine in about 7 years if adequately resourced.  They obviously had that argument put forward by October 2012, when Musk announced the next-gen engine would not use Merlin's fuel, and confirmed it as methane in Nov 2012.   Based on the testing of the Raptor dev engine(s) since first fire in Sept 2016, and it's now Dec 2018, I'd say they are right on target.  Improvements to the dev Raptor test engine, ostensibly, the initial production engine, hits the test stand in early 2019.  First flight using three (3) Raptors by 2Q2019 on the Starship BFH1.
This is a competent propulsion team!

Why methane?  Because Mars!  No oil on Mars to make RP-1 for the return flight.  Musk explained this in detail in his IAC talk in Sept 2017.

Why FFSC?  Because Efficiency!  Higher ISP, which is needed in the long run for the sorts of objectives SpaceX has for themselves.

Why these choices, and not the choices you want? (or anybody else on NSF wants, or the US government might "want" if they were to have been asked?)  Because private capital, and choices that can be made by actors in a world of economic incentives rather than that political incentives that drove all nation state/government space efforts for the past six+ decades.

This is not your father's, nor your government's, space technology development effort.

So why do we keep comparing what a private company like SpaceX can do to these legacy ways of thinking, and then being surprised by the results?  ::)

Edit:  fixed a typo
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 12/31/2018 07:14 pm
Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.

You're not answering my question. I didn't say full-flow staged combustion was impossible, just that it was hard and (for whatever reason) they appear to be in a hurry. We already know why LOX/methane was chosen and why full-flow staged combustion is awesome. The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.

Merlin 1D made several breakthroughs over M1C. Most of them were not required but desirable to lower production cost and make it simpler and more reliable. Some short term challenges are worth if it the long term benefits more then make up for it.

We simply don't know the entire story. But this is obviously something that Muellers team feel that they can pull off.


Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious what that reason is.

This idea that Musk only (or even mostly) only works incrementally has little or no basis in fact. Do you think that the BFR architecture is just an incremental step?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mgeagon on 12/31/2018 07:16 pm
The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.
Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious to know what that reason is.
It is unlikely any CH4 engine would ever be made “fuel-rich” as it would still be susceptible to some coking in the pre-burner. Blue Origin’s BE-4, for example, is oxidizer rich for this reason. The time to develop a ORSC engine is neither shorter nor less expensive. The FFSC engine offers slower and cooler running turbo pumps, ideal for reliability and reuse. In addition, they do not suffer from seal failure and introduction of hot, oxygen rich gas being injected into the fuel plumbing via the shared turbine axis.

Developing an interim methane engine using a gas generator would also cost a lot of time and money that would be of little short term use and would vastly increase the overall budget, further hindering a company looking for ways to reduce research outlays to the extent possible. No part of developing a GG would inform the eventual FFSC. The development cost would simply be wasted. If the goal was to simply get a super heavy into orbit as soon as possible, stick with the FH architecture, strap on a couple more boosters and fly (judging from the time and money it took to “strap-on” two boosters, I just hand-waved, sorry). If one wishes to go to Mars and return, it seems reasonable to just build the engine that is required. This is in fact the cheapest and fastest way.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 07:19 pm
This idea that Musk only (or even mostly) only works incrementally has little or no basis in fact. Do you think that the BFR architecture is just an incremental step?

No, you're right, the whole BFR project, not just Raptor, appears to be following a different philosophy. Tesla Model 3 too, come to think of it. And, as you say, we don't know the full story. I'm sure that if and when we ever hear it, it will make sense. It's just fun to speculate what the reasons might be.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: lonestriker on 12/31/2018 07:28 pm

The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry?


'Cause he hired, and then built over time, a super top-notch propulsion engineering team, and they thought they could build a methalox FFSC engine in about 7 years if adequately resourced.  They obviously had that argument put forward by October 2012, when Musk announced the next-gen engine would not use Merlin's fuel, and confirmed it as methane in Nov 2012.   Based on the testing of the Raptor dev engine(s) since first fire in Sept 2016, and it's now Dec 2018, I'd say they are right on target.  Improvements to the dev Raptor test engine, ostensibly, the initial production engine, hits the test stand in early 2019.  First flight using three (3) Raptors by 2Q2019 on the Starship BFH1.
This is a competent propulsion team!

Why methane?  Because Mars!  No oil on Mars to make RP-1 for the return flight.  Musk explained this in detail in his IAC talk in Sept 2017.

Why FFSC?  Because Efficiency!  Higher ISP, which is needed in the long run for the sorts of objectives SpaceX has for themselves.

Why these choices, and not the choices you want? (or anybody else on NSF wants, or the US government might "want" if they were to have been asked?)  Because private capital, and choices that can be made by actors in a world of economic incentives rather than that political incentives that drove all nation state/government space efforts for the past six+ decades.

This is not your father's, nor your government's, space technology development effort.

So why do we keep comparing what a private company like SpaceX can do to these legacy ways of thinking, and then being surprised by the results?  ::)

Edit:  fixed a typo

And one other big benefit:

Methane is cheap.  Elon mentioned this in one of his talks:  When you're flying expendable or only partially reusable, the cost of fuel is almost a rounding error.  When you're fully fully reusable, fuel costs become significant, especially when you're doing in-orbit fueling with many flights.

It's all a brave new world.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 07:40 pm
Why these choices, and not the choices you want? (or anybody else on NSF wants, or the US government might "want" if they were to have been asked?)

No, my question was a different one, why no intermediate steps given that they were (or at least appear now to be) in a hurry? People have given several useful potential explanations: perhaps it would not in fact have gained them much time, or perhaps it would have but only at a cost that made the idea unattractive. Or Musk doesn't always prefer a philosophy of incrementalism. I can think of several other potential explanations: maybe circumstances changed, new information came to light, maybe some things might have been done differently in hindsight, but wouldn't improve the timeline at this late stage.

Quote
This is not your father's, nor your government's, space technology development effort. So why do we keep comparing what a private company like SpaceX can do to these legacy ways of thinking, and then being surprised by the results?  ::)

NASA's legacy approach to LV development has been big-bang. Incrementalism would in fact be the be the more modern and agile way to do it, more in line with private-sector thinking.

On a more philosophical note: asking people to speculate about SpaceX's motives is not the same thing as criticising SpaceX. I'm sure there are good reasons for their choices, I'm just curious what they are. Several people have made useful suggestions what they might be, maybe one day we'll hear the full story.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RedLineTrain on 12/31/2018 07:43 pm
Maybe they aren't really in a hurry.  Raptor has earned some flight time, so there's no reason to wait around before flying it.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: AC in NC on 12/31/2018 07:45 pm
No, my question was a different one, why ...

You gracefully stepped down at #1303.  You should've left it there.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: GreenShrike on 12/31/2018 08:12 pm
Do we have a solid, official explanation of why SpaceX is going for both LOX/methane and full-flow staged combustion without easier intermediate steps, such as gas-generator LOX-methane and/or fuel-rich staged combustion?

What is a GG or FRSC engine going to teach you that just outright building a FFSC won't?

Let's look at this logically. Assume you build a first prototype -- it will either be a non-FFSC (your GG or FRSC engines) or FFSC (SpaceX's test Raptor) configuration. In building that first prototype, there will be two types of experience gained to be used for a follow-on FFSC project: that which is common between the non-FFSC and the FFSC prototypes, in which case you gain no time building a non-FFSC since the issues need to be solved in either case; or that which is uncommon between non-FFSC and FFSC prototypes, in which case you gain no time (in fact, you actively waste it) since building the non-FFSC engine isn't teaching you something necessary for the follow-on FFSC engine.

In addition, it will be easier to iterate from a FFSC prototype to a follow-on FFSC production engine, than from a non-FFSC prototype to a FFSC production engine, which would necessitate *another* generation of engines before the final production FFSC engine.

As such, if your goal is a FFSC engine, then building anything else is just a waste of time and money.

It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.

What makes you think a FF methalox engine would have a substantially better ISP than Merlin 1D?

Methane and kerosene have very close theoretical max ISPs, and while methane's is indeed a couple percent higher, methane's low density compared to RP-1 also means rather larger tankage. As the Bruce Dunn Fuel Table says, "Going from RP-1 to methane gains 3.8 % in Isp, but costs about 22% in density."  With identical engine cycles, the overall *system* result is pretty much a wash, performance-wise.

For an outright win over kerolox, you need sub-cooled propane -- slightly higher ISP, and slightly better bulk density. ;-)

Make no mistake: the only reason Raptor outperforms Merlin is because its FFSC cycle a) enables a ridiculous increase of chamber pressure and b) eliminates those losses incurred by Merlin's open cycle gas generator.

SpaceX didn't develop a GG methalox or FRSC methalox engine because they have no need of one. Merlin gets the job done for Falcon -- which will be retired soonest anyway -- and they need FFSC Raptors for BFR.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 08:12 pm
[...]

Those reasons sound plausible. One thing jumped out at me though: the idea that a fuel-rich preburner would also involve coking, and would therefore not be a solution to the affordable reusability problems some people speculate Merlins suffer from. If that's true, and it does sound logical, how does that work with full flow? Is it possible to get full flow without having both an oxygen-rich preburner for the oxygen turbopump and a fuel-rich one for the fuel? Or will Raptor only have oxygen-rich preburners as I believe RD-180 does?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 12/31/2018 08:15 pm
As such, if your goal is a FFSC engine, then building anything else is just a waste of time and money.

Yeah, that's basically the argument that mgeagon made, and it sounds plausible.

Quote
Methane and kerosene have very close theoretical max ISPs, and while methane's is indeed a couple percent higher,

Thanks, I didn't know that, I thought the Isp difference was much bigger.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 12/31/2018 09:00 pm
Thanks, I didn't know that, I thought the Isp difference was much bigger.

For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.

Methalox needs staged combustion to beat kerolox on performance.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mgeagon on 12/31/2018 09:12 pm
Is it possible to get full flow without having both an oxygen-rich preburner for the oxygen turbopump and a fuel-rich one for the fuel? Or will Raptor only have oxygen-rich preburners as I believe RD-180 does?
This very thread has the answers. The first few pages go into great detail. In short summary, using pure CH4 in the fuel rich pre-burner of a FFSC engine produces little coking because the pressures and temperatures are kept low enough. If one were to run methane or in Blue Origin’s case, LNG, through a FRSC cycle, it would run high and hot enough to produce coking. The greater mass of O2 also makes ORSC more desirable than FRSC. The SSME is FRSC because H2 does not coke and NASA didn’t believe it was possible to reuse an engine that had been exposed to high pressure, hot O2.
By definition, a FFSC engine contains a fuel-rich preburner and an oxidizer-rich preburner. Since all propellant molecules go through both the pre and main combustion processes, the cycle is the most efficient, garnering the highest ISP possible given the fuel.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 12/31/2018 09:17 pm
Upthread it said raptor has the same shaft for methane preburner and oxygen reburner. I think they have different shafts?

Also coking in the methane preburner:
What temperature does coking happen at?
I thought I remember reading that the methane preburner was relatively cool and could almost be just be an expander cycle. Which I assume to mean is that just vaporizing the methane is enough to provide the horsepower.

Do I have that right?

EDIT: Sounds like point 2 is already answered.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mgeagon on 12/31/2018 09:27 pm
For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.

Methalox needs staged combustion to beat kerolox on performance.
Absolutely. The only advantages of CH4 out of Earth’s dense atmosphere and gravity well is its potential to be carbon neutral and its effect on reusability, including clean burning, spark ignition and autogenous tank pressurization.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 12/31/2018 09:43 pm
For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.

Methalox needs staged combustion to beat kerolox on performance.
Absolutely. The only advantages of CH4 out of Earth’s dense atmosphere and gravity well is its potential to be carbon neutral and its effect on reusability, including clean burning, spark ignition and autogenous tank pressurization.
You can also use the super efficient engine cycles that hydrogen can use: FFSC and expander.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mwood on 12/31/2018 09:44 pm
For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.

Methalox needs staged combustion to beat kerolox on performance.
Absolutely. The only advantages of CH4 out of Earth’s dense atmosphere and gravity well is its potential to be carbon neutral and its effect on reusability, including clean burning, spark ignition and autogenous tank pressurization.
Another advantage, at least for Boca Chica, is the ability to tap directly into nearby CNG feed line and refine their own methane. That's got to save a few bucks.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 12/31/2018 09:53 pm
Upthread it said raptor has the same shaft for methane preburner and oxygen reburner. I think they have different shafts?

Also coking in the methane preburner:
What temperature does coking happen at?
I thought I remember reading that the methane preburner was relatively cool and could almost be just be an expander cycle. Which I assume to mean is that just vaporizing the methane is enough to provide the horsepower.

Do I have that right?

EDIT: Sounds like point 2 is already answered.

I found one paper that listed the coke limit ~950 K.

Preburner output temperature needs to be less than that so that the methane can be used to cool the chamber.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Hominans Kosmos on 12/31/2018 11:23 pm
why not do something simpler first?
perhaps it would not in fact have gained them much time
Absolutely certainly it would have lost them time. It's not a good question to ask.

https://en.wikipedia.org/wiki/Busy_work
See also: Boondoggle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Zpoxy on 01/01/2019 12:41 am
This idea that Musk only (or even mostly) only works incrementally has little or no basis in fact. Do you think that the BFR architecture is just an incremental step?

No, you're right, the whole BFR project, not just Raptor, appears to be following a different philosophy. Tesla Model 3 too, come to think of it. And, as you say, we don't know the full story. I'm sure that if and when we ever hear it, it will make sense. It's just fun to speculate what the reasons might be.

I know it's fun to speculate here on NSF but you're overthinking this. The reason Spacex is following this route is because they can! Just that simple.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: AC in NC on 01/01/2019 12:45 am
I know it's fun to speculate here on NSF but you're overthinking this. The reason Spacex is following this route is because they can! Just that simple.

Not that I really want this guy's sub-thread to live but what am I missing?  What the heck is the purpose of discussing development alternatives for a design that was fired up and working two years ago?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rubicondsrv on 01/01/2019 01:21 am
Upthread it said raptor has the same shaft for methane preburner and oxygen reburner. I think they have different shafts?

Also coking in the methane preburner:
What temperature does coking happen at?
I thought I remember reading that the methane preburner was relatively cool and could almost be just be an expander cycle. Which I assume to mean is that just vaporizing the methane is enough to provide the horsepower.

Do I have that right?

EDIT: Sounds like point 2 is already answered.


that lower temperature is an advantage of FFSC on both the fuel and oxidizer pumps.  That is the other major advantage, a less damaging operating environment for the turbomachinery compered with say a standard ORSC cycle, and due to coking concerns that would be the likely alternate choice.   
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 01/01/2019 02:27 pm
Pixel measuring the latest pictures of the testbed shows that the Raptor bell diameter is 1/6.89 the diameter of the test article as a whole... That works out to 1.3 meters if it is 9m in diameter.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 01/01/2019 02:29 pm
Pixel measuring the latest pictures of the testbed shows that the Raptor bell diameter is 1/6.89 the diameter of the test article as a whole... That works out to 1.3 meters if it is 9m in diameter.

and remind me of what the different size estimates of the bell have been over the years?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: baldusi on 01/01/2019 03:24 pm
Let me clarify the reasons for going oxidizer or fuel rich (besides cocking issues). Turbopumps are thermal machines that provide power. You want as much power as hassle-freeas possible, to i crease chamber pressure and thus increase isp and reduce weight.
Simplifying, available power on a TP would be massflow * specific heat.
In a hydrolox engine, even though the mass relationship is 6:1 (or so) for the oxidizer, the specific heat difference is so much that fuel rich gives more than 70% (IIRC) of power of going FF.
In the RP-1 case, you simply can’t go fuel rich, so it is your only solution.
But in the mathalox case, the O/F ratio is around 3.5, and the specific heat difference is not that high. In fact, I think to remember, the oxidizer-rich give something like 60% of the FF power.
Overall, just for the methalox case, while oxi-rich is the best choice for staged combustion, going full flow gives you an extra 70% power. So it makes a lot more sense than for hydrolox.
Also have to compare the power balance on the oxidizer and fuel pump. If you have most power on the fuel turbine, but require most of it on the oxidizer, you won’t get so much advantages of going full flow.
Again, methalox full flow balances very nicely there.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 01/01/2019 03:54 pm
Which helps underscore the point that SpaceX and propulsion head Tom Mueller have assembled a really top-notch team of rocket engine designers, technicians, testers, and the ops team that keeps it all running.  They had very good reasons to go the way they chose to plan to go, and by all accounts, are achieving on their development milestones.

Probably the best and most experienced rocket engine development team today globally, given their achievements and recent iterative and incremental real-world experience in many rocket engines with many propellant mixes.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nomadd on 01/01/2019 04:04 pm
 Nice Baldusi. Now, do want to show up more than once a year? You've been seriously missed here. And I still owe you dinner.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 01/01/2019 04:10 pm
So do you already know what these are:

"♫ Raptors marching three by three
Some from the land and some from the sea ♫"
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/01/2019 04:13 pm
Pixel measuring the latest pictures of the testbed shows that the Raptor bell diameter is 1/6.89 the diameter of the test article as a whole... That works out to 1.3 meters if it is 9m in diameter.

and remind me of what the different size estimates of the bell have been over the years?

1.3 m. Has anyone else attempted to estimate the engine bell diameter?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/01/2019 04:19 pm
Could be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/01/2019 04:25 pm
Could be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.

John

It's not the wrong size if the purpose is deep throttle at SL, instead of vacuum optimization for which the dual bell was traditionally proposed.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 01/01/2019 04:31 pm
Which helps underscore the point that SpaceX and propulsion head Tom Mueller have assembled a really top-notch team of rocket engine designers, technicians, testers, and the ops team that keeps it all running.  They had very good reasons to go the way they chose to plan to go, and by all accounts, are achieving on their development milestones.

Probably the best and most experienced rocket engine development team today globally, given their achievements and recent iterative and incremental real-world experience in many rocket engines with many propellant mixes.
They have suffered personnel losses to several other engine development teams (mostly NewSpace) ... significant ones. Enough to make those other teams quite good (that's a good thing, IMHO). And yet I suspect they are still the best.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 01/01/2019 04:41 pm
If we assume that it's not a mockup, then those need to be sub scale development version of the Raptor? Like the first Grasshopper used development version of the Merlin 1D.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Peter.Colin on 01/01/2019 06:51 pm
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 01/01/2019 07:12 pm
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.
No, the new radically redesigned Raptor will have it's 1st test fire next month. I think the Raptors on Star Hopper may be the 2017 spec. ones with 1.3m dia. nozzles which are likely the 1st revision from the original dev. Raptor. The new radically redesigned Raptor will most likely be revision 2 and likely have somewhat greater thrust than the 2017 spec. judging from EM's tweet about it's pump power.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 01/01/2019 08:16 pm
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.
No, the new radically redesigned Raptor will have it's 1st test fire next month.
>

Not trying to be too picky, but Musk didn't say 1st/first test fire,

Elon Musk ✔ @elonmusk
Yes. Radically redesigned Raptor ready to fire next month.
6:13 PM - Dec 22, 2018

which could mean the hopper with radically redesigned Raptors doing it's first hop.  RR Raptor's tests could have been at Stennis.

https://www.nasa.gov/centers/stennis/about/stennis/index.html

Quote
Stennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 01/01/2019 08:45 pm
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.
No, the new radically redesigned Raptor will have it's 1st test fire next month.
>

Not trying to be too picky, but Musk didn't say 1st/first test fire,

Elon Musk ✔ @elonmusk
Yes. Radically redesigned Raptor ready to fire next month.
6:13 PM - Dec 22, 2018

which could mean the hopper with radically redesigned Raptors doing it's first hop.  RR Raptor's tests could have been at Stennis.

https://www.nasa.gov/centers/stennis/about/stennis/index.html

Quote
Stennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.

What are the chances they use the hopper as the test stand? We were speculating about testing multiple engines next to each other on a test stand before integration.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 01/01/2019 09:10 pm
  --- snip ---
which could mean the hopper with radically redesigned Raptors doing it's first hop.  RR Raptor's tests could have been at Stennis.

https://www.nasa.gov/centers/stennis/about/stennis/index.html

Quote
Stennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.

Do you have a source for any testing by SpaceX at Stennis after 2015?  The only sources I've seen say that SpaceX did test a Raptor powerhead (turbines and preburners but no combustion chamber) testing at Stennis, starting in 2014, and that they wrapped up that Stennis testing in 2015. 

I know for a fact that Stennis would like to have SpaceX back; but I've never seen a source since very early 2016 that even discusses SpaceX and the US government rocket engine test facilities at Stennis.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 01/01/2019 10:01 pm
  --- snip ---
which could mean the hopper with radically redesigned Raptors doing it's first hop.  RR Raptor's tests could have been at Stennis.

https://www.nasa.gov/centers/stennis/about/stennis/index.html

Quote
Stennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.

Do you have a source for any testing by SpaceX at Stennis after 2015? 
>
I know for a fact that Stennis would like to have SpaceX back; but I've never seen a source since very early 2016 that even discusses SpaceX and the US government rocket engine test facilities at Stennis.

It's part of the Oct. 2017 USAF Raptor funding. Also noticed something in L2.

DoD announcement...  (https://dod.defense.gov/News/Contracts/Contract-View/Article/1348379/)

Quote
Space Exploration Technologies Corp., Hawthorne, California, has been awarded a $40,766,512 modification (P00007) for the development of the Raptor rocket propulsion system prototype for the Evolved Expendable Launch Vehicle program.  Work will be performed at NASA Stennis Space Center, Mississippi; Hawthorne, California; McGregor, Texas; and Los Angeles Air Force Base, California; and is expected to be complete by April 30, 2018.
>
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 01/01/2019 11:28 pm
Thanks, doc.  That's super helpful.

I've not read the contract details, but that makes good sense.  Since USAF added that US$40 million to the Raptor engine contract in late 2017, which was originally (early 2016, $33 million) to develop a prototype Raptor upper stage methalox engine for a Falcon 9, it makes perfect sense that USAF would want performance demonstration and testing on that contract to occur at a US government rocket engine test facility. 

I'll get around to reading whatever contract detailz I can find in the coming weeks.  What we can say for now is that we don't have any good source that says SpaceX is needing/wanting to buy Stennis test facility services for their own commercial engine development purposes.  As NSFers know, SpaceX has a super state-of-the-art rocket engine test facility at McGregor, Texas, averaging >1 rocket engine firing per day, and that by all accounts seems entirely adequate to provide all testing to support their engine development objectives.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: CT Space Guy on 01/04/2019 09:58 pm
I was wondering how come I haven't heard any discussion here about the possibility of a dual-bell design for the Raptor engine?

https://forum.nasaspaceflight.com/index.php?topic=46795.msg1896944#msg1896944

https://forum.nasaspaceflight.com/index.php?topic=46795.msg1896932#msg1896932
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Kenp51d on 01/04/2019 10:08 pm
I get the ideal of a duel bell. Any notion of how effective this is.
Has there ever been an operational duel bell used before?

Ken


Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 01/04/2019 10:11 pm
@ CT SpaceGuy
Could be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.

John

It's not the wrong size if the purpose is deep throttle at SL, instead of vacuum optimization for which the dual bell was traditionally proposed.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: CT Space Guy on 01/04/2019 11:14 pm
@ CT SpaceGuy
Could be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.

John

It's not the wrong size if the purpose is deep throttle at SL, instead of vacuum optimization for which the dual bell was traditionally proposed.

Thanks I guess I didn't understand the dual bell post until I read more about it elsewhere
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: 50_Caliber on 01/04/2019 11:16 pm
I get the ideal of a duel bell. Any notion of how effective this is.
Has there ever been an operational duel bell used before?

Ken
The space shuttle used a de Laval nozzle which helped with flow separation at sea level.

https://en.wikipedia.org/wiki/Space_Shuttle_main_engine (https://en.wikipedia.org/wiki/Space_Shuttle_main_engine)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/04/2019 11:35 pm
I get the ideal of a duel bell. Any notion of how effective this is.
Has there ever been an operational duel bell used before?

Ken
The space shuttle used a de Laval nozzle which helped with flow separation at sea level.

https://en.wikipedia.org/wiki/Space_Shuttle_main_engine (https://en.wikipedia.org/wiki/Space_Shuttle_main_engine)

A de Laval nozzle is any contracting-expanding nozzle.

The SSME used a slight reduction from optimum expansion to inhibit flow separation at SL. But it could not throttle at SL as Raptor needs to do.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/05/2019 12:28 am
Cross posting this:

What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/05/2019 02:32 am
Cross posting this:

What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.

John

Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: MikeAtkinson on 01/05/2019 03:15 am
What we see could just be where the raptor armour joins the bell. We know raptor has armour for RUD protection, this armour proably joins onto the bell, it may also take some of the thrust.
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/05/2019 03:16 am
Cross posting this:

What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.

John

Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.

Yep. Full regen. And radiative nozzles really only work well for single engines. (Otherwise they radiate heat to each other) And Starship will have 7 relatively closet packed engines.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: su27k on 01/05/2019 03:21 am
Cross posting this:

What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.

John

Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.

The production version will surely have regenerative cooled nozzle, but this may not be the production version or anything close to that. I had the same thought as livingjw when I first saw the nozzle images, it is as if they welded a MVac nozzle extension to the nozzle of the dev Raptor. This could be just a temporary test setup, radiation cooling should work with just 3 of them, similar to Delta-IV Heavy.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/05/2019 03:37 pm
Cross posting this:

What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.

John

Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.

The production version will surely have regenerative cooled nozzle, but this may not be the production version or anything close to that. I had the same thought as livingjw when I first saw the nozzle images, it is as if they welded a MVac nozzle extension to the nozzle of the dev Raptor. This could be just a temporary test setup, radiation cooling should work with just 3 of them, similar to Delta-IV Heavy.

The RS-68 nozzle is primarily ablatively cooled, and has a much larger view of the sky on DIVH than these Raptors do.

Anyway, Musk said "Engines currently on Starship hopper are a blend of Raptor development & operational parts. First hopper engine to be fired is almost finished assembly in California. Probably fires next month."

I read that as meaning the currently mounted hopper engines are only for fit checks, are not fully operational but are the same size and geometry as the operational Raptor. They probably have operating TCV to check range of motion, which explains the gimbal.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 01/05/2019 05:46 pm
Except the claim is it is the outside bell (the extension) that is SL optimized.

So the inner bell would would be attached at above Earth SL pressure.

Which makes no sense, unless people are mis estimating the sizes of both and/or the operating pressure of the engine.

It makes a lot of sense, when you consider that the engine needs to be able to deep throttle, and expansion ratio is one of the hard limits on deep throttling. Assuming the numbers put forwards by others in this thread, it provides 1 to 50 expansion ratio when going full throttle, and yet allows throttling down to 15% or even less. That's something that a standard bell really couldn't do.

It also sort of explains how they intend to deal with some of the hard problems of steeped nozzles: The nozzle is never intended to spend any real time in the dangerous zone where it's overflowing the inner nozzle but not really fully filling up the outer. On startup at full power, the full nozzle is always in use. When returning, the 3 engines are spun up and fired at a throttle level that just barely fills out the inner one.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/05/2019 07:07 pm
Except the claim is it is the outside bell (the extension) that is SL optimized.

So the inner bell would would be attached at above Earth SL pressure.

Which makes no sense, unless people are mis estimating the sizes of both and/or the operating pressure of the engine.

It makes a lot of sense, when you consider that the engine needs to be able to deep throttle, and expansion ratio is one of the hard limits on deep throttling. Assuming the numbers put forwards by others in this thread, it provides 1 to 50 expansion ratio when going full throttle, and yet allows throttling down to 15% or even less. That's something that a standard bell really couldn't do.

It also sort of explains how they intend to deal with some of the hard problems of steeped nozzles: The nozzle is never intended to spend any real time in the dangerous zone where it's overflowing the inner nozzle but not really fully filling up the outer. On startup at full power, the full nozzle is always in use. When returning, the 3 engines are spun up and fired at a throttle level that just barely fills out the inner one.

- The claim is that the inside bell is SL optimized, not the outer.

- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.

- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.

- In the case of the hopper, the lower bell can be radiation cooled, even with the some radiation blockage, because it would not be exposed to attached flow at low hopper altitudes.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 01/05/2019 07:33 pm
Except the claim is it is the outside bell (the extension) that is SL optimized.

So the inner bell would would be attached at above Earth SL pressure.

Which makes no sense, unless people are mis estimating the sizes of both and/or the operating pressure of the engine.

It makes a lot of sense, when you consider that the engine needs to be able to deep throttle, and expansion ratio is one of the hard limits on deep throttling. Assuming the numbers put forwards by others in this thread, it provides 1 to 50 expansion ratio when going full throttle, and yet allows throttling down to 15% or even less. That's something that a standard bell really couldn't do.

It also sort of explains how they intend to deal with some of the hard problems of steeped nozzles: The nozzle is never intended to spend any real time in the dangerous zone where it's overflowing the inner nozzle but not really fully filling up the outer. On startup at full power, the full nozzle is always in use. When returning, the 3 engines are spun up and fired at a throttle level that just barely fills out the inner one.

- The claim is that the inside bell is SL optimized, not the outer.

- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.

- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.

- In the case of the hopper, the lower bell can be radiation cooled, even with the some radiation blockage, because it would not be exposed to attached flow at low hopper altitudes.

John
The sizes of the bells and the raptor's known characteristics dont line up with your claim of a vac-optimized outer bell. They DO line up with Tuna's claim of a SL optimised outer bell, deep throttle optimized inner bell.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 01/05/2019 07:52 pm
- The claim is that the inside bell is SL optimized, not the outer.

I think that claim is wrong, and does not line up with published performance numbers.

Quote
- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.

For the latest performance numbers, 1.3m would be rather small for a vacuum nozzle, and since the engines are on the rocket apparently to perform fit tests, it would not make sense for the final articles to have larger ones.

Quote
- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.

That is one of the things you can get by designing dual nozzles. However, it is not the only possible design goal.

Raptor engines differ from most other rocket engines by having an additional requirement: The ability to deep throttle down when right at sea level altitude. This raises from the need to support multi-engine-out vertical landing. (Engine startup is much slower than raising throttle on an already running engine, so it's necessary to be able to land on at least two less engines than are typically lit, and preferably still at much less than 100% throttle so you have reserve for the "oh crap" situation where you lose two engines and only realize it when their thrust falls to 0 and you need to very rapidly compensate.)

Traditionally designed engine nozzles cannot support this, unless they are very low expansion ratio and therefore very low performance. Using a dual nozzle where the outer nozzle is sea level optimized for full throttle and at maximum possible expansion ratio, and the inner nozzle is sea level optimized but for a very low power output is an elegant solution to their problem.

And if feel that the inner nozzle is sea level optimized and the outer one is vacuum optimized, how do you believe Raptor will achieve <20% throttle with the inner nozzle?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DigitalMan on 01/05/2019 08:10 pm
- The claim is that the inside bell is SL optimized, not the outer.

I think that claim is wrong, and does not line up with published performance numbers.

Quote
- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.

For the latest performance numbers, 1.3m would be rather small for a vacuum nozzle, and since the engines are on the rocket apparently to perform fit tests, it would not make sense for the final articles to have larger ones.

Quote
- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.

That is one of the things you can get by designing dual nozzles. However, it is not the only possible design goal.

Raptor engines differ from most other rocket engines by having an additional requirement: The ability to deep throttle down when right at sea level altitude. This raises from the need to support multi-engine-out vertical landing. (Engine startup is much slower than raising throttle on an already running engine, so it's necessary to be able to land on at least two less engines than are typically lit, and preferably still at much less than 100% throttle so you have reserve for the "oh crap" situation where you lose two engines and only realize it when their thrust falls to 0 and you need to very rapidly compensate.)

Traditionally designed engine nozzles cannot support this, unless they are very low expansion ratio and therefore very low performance. Using a dual nozzle where the outer nozzle is sea level optimized for full throttle and at maximum possible expansion ratio, and the inner nozzle is sea level optimized but for a very low power output is an elegant solution to their problem.

And if feel that the inner nozzle is sea level optimized and the outer one is vacuum optimized, how do you believe Raptor will achieve <20% throttle with the inner nozzle?

I'm a little unclear on how your description works, hopefully you can shed some light on this:

some potential facts:

1) Thrust is produced when it interacts with the nozzle, which is attached to the rocket engine, producing a force.

2) The atmosphere interacts with the plume, causing expansion to occur as the atmosphere nears a vacuum.

3) In order to obtain a higher force in more expanded environments (vacuum) a larger nozzle is necessary to provide more surface area for the plume to interact with in order to capture the force since the interactive area of the plume is much wider in vacuum

So, here is the question:

If the outer part of the nozzle is sea level it would not be wider than the plume at sea level and since the inner part of the nozzle will have less surface area than the outer part, how does your interpretation of this work?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/05/2019 08:54 pm
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.

But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wolfram66 on 01/05/2019 09:27 pm
Used BocaChicaGals image cropped and enhanced to better show the frankenRaptor boiler plate engine/bells dual bell shape.

Edit: attaching 2014 NASA study on dual bell performance
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140011268.pdf (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140011268.pdf)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 01/05/2019 09:45 pm
I have a dumb question.

So, assuming the double bell is sea level/deep throttle, what pressures could the 19:1 expansion ratio and 300 bar full throttle work at? Is a shiny 1-way trip to Venus a possibility?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/05/2019 10:12 pm
....
Raptor engines differ from most other rocket engines by having an additional requirement: The ability to deep throttle down when right at sea level altitude. This raises from the need to support multi-engine-out vertical landing. (Engine startup is much slower than raising throttle on an already running engine, so it's necessary to be able to land on at least two less engines than are typically lit, and preferably still at much less than 100% throttle so you have reserve for the "oh crap" situation where you lose two engines and only realize it when their thrust falls to 0 and you need to very rapidly compensate.)

Traditionally designed engine nozzles cannot support this, unless they are very low expansion ratio and therefore very low performance. Using a dual nozzle where the outer nozzle is sea level optimized for full throttle and at maximum possible expansion ratio, and the inner nozzle is sea level optimized but for a very low power output is an elegant solution to their problem.

And if feel that the inner nozzle is sea level optimized and the outer one is vacuum optimized, how do you believe Raptor will achieve <20% throttle with the inner nozzle?
.... blew my quote !

- Could be. That implies that  the nozzles would have to be regenerative cooled. Which implies that these are the 2017 sized engines and not the demo engines. I have attached the 2017 sized engine dimensions. This engine could have been designed with a deep throttling dual bell.

- For the actual SS one might contemplation a triple bell. ;^)

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DigitalMan on 01/06/2019 02:46 am
Thanks, I hadn't been thinking about it in terms of capability to throttle down.  It has been an interesting discussion.
 A triple nozzle would be quite a thing to see! 

This does make me wonder what the target thrust is going to be.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RotoSequence on 01/06/2019 02:57 am
The conceivable complexity of a deep throttling, hybrid bell that is efficient in vacuum, but also works decently at sea level, probably explains why they'd forego vacuum Raptors for early versions of Starship.
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/06/2019 03:27 am
People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?

I don’t expect such a nozzle on Raptor. But I could be wrong.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/06/2019 05:17 am
People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?

I don’t expect such a nozzle on Raptor. But I could be wrong.
Mish mash doesn't mean they concatenated two nozzles they had lying around the scrapyard.

These nozzles came.from somewhere.  They look dual-belled, and they look hi fidelity.  Why would mock-up bells be so complex?  Mock-up bells would be conical...

The strongest argument for these being temporary engines is that it doesn't look like there are engine covers on them..  but they are probably interface-compatible, and might support initial testing.

Unless Musk wasn't clear, they will put production engines on this rocket, and according to the schedule, test them at McGregor and on BFH almost concurrently... (With McGregor leading, no doubt)

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ABCD: Always Be Counting Down

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/06/2019 05:45 am
People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?

I don’t expect such a nozzle on Raptor. But I could be wrong.
Mish mash doesn't mean they concatenated two nozzles they had lying around the scrapyard.

These nozzles came.from somewhere.  They look dual-belled, and they look hi fidelity.  Why would mock-up bells be so complex?  Mock-up bells would be conical...

Indeed they would be - the upper part on these nozzles is exactly that - conical. The bottom part appears to be a real nozzle (or shaped like a real nozzle), but the upper part is the “fake bit” IMO. This  part is what people see and think dual/hybrid. But it is strictly conical, and thus IMO just a conical cover over where real engine detail/plumbing will be.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/06/2019 08:06 am
People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?

I don’t expect such a nozzle on Raptor. But I could be wrong.
Mish mash doesn't mean they concatenated two nozzles they had lying around the scrapyard.

These nozzles came.from somewhere.  They look dual-belled, and they look hi fidelity.  Why would mock-up bells be so complex?  Mock-up bells would be conical...

Indeed they would be - the upper part on these nozzles is exactly that - conical. The bottom part appears to be a real nozzle (or shaped like a real nozzle), but the upper part is the “fake bit” IMO. This  part is what people see and think dual/hybrid. But it is strictly conical, and thus IMO just a conical cover over where real engine detail/plumbing will be.
We can only see a tiny bit of it - can't tell if conical or not.

Joining such nozzles is a lot of work - and to what end?  Everyone knows the rocket's under construction, so adding the engines in a few weeks would have been no different...

Plus, Musk said they are a mixture of parts from development and production engines - he did not say anything was fake.  Remember the FH second stage, described with similar language?  That two was a hybrid, but it certainly wasn't a mock-up.

The simple explanation is that they took development engines, added what's necessary to mate them to a production baseplate, and this is what we're seeing here, to be used for simple bring-up testing.  Meanwhile they're working on finishing the first set of production engines, which will actually hop the hopper.

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ABCD: Always Be Counting Down

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 01/06/2019 01:18 pm
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.

But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.

The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.

Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 01/06/2019 02:30 pm
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Kaputnik on 01/06/2019 02:41 pm
What are the disadvantages of a dual-bell nozzle?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nomadd on 01/06/2019 02:51 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HMXHMX on 01/06/2019 02:54 pm
What are the disadvantages of a dual-bell nozzle?

They are:

(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;

(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 01/06/2019 03:19 pm
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?

I would assume there are 3 states.

Attached flow. No turbulence.

Not attached. Turbulence is separated from nozzle.

Somewhere inbetween where some sections are attached and some not. Very time dependent attachment. This is what creates the large rapidly changing stresses.

So the smaller nozzle guarantees complete separation of flow.

EDIT: One of my favorite videos showing the twang on the SSME on startup.
https://imgur.com/gallery/W7UZ1

https://i.imgur.com/uEYMGMW.mp4
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DigitalMan on 01/06/2019 03:37 pm
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.

But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.

The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.

Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

I wonder if this will give them another advantage:

On launch from an unprepared surface (Mars) you could potentially ignite all the engines at a low throttle setting and allow them a short time to stabilize before moving to full thrust. 
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/06/2019 05:15 pm
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?

The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/06/2019 05:28 pm
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?

The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.
And from the looks of it, that separation ring is naturally reinforced

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ABCD: Always Be Counting Down

Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/06/2019 06:16 pm
The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.
And from the looks of it, that separation ring is naturally reinforced


Huh? Wha... where do you get that from?

These “engine” pictures really are the ultimate Rorschach test. People see whatever the heck they want. Double/triple/quintuple nozzles, natural reinforced points, you name it. Pretty soon someone will argue that this is proof of Raptor being an aerospike engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/06/2019 06:35 pm
The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.
And from the looks of it, that separation ring is naturally reinforced


Huh? Wha... where do you get that from?

These “engine” pictures really are the ultimate Rorschach test. People see whatever the heck they want. Double/triple/quintuple nozzles, natural reinforced points, you name it. Pretty soon someone will argue that this is proof of Raptor being an aerospike engine.
Because, irrespective of your angst over these specific nozzles, any location where the curvature of the generatrix of a bell is discontinuous will be naturally reinforced, as you'd like it to be if it is where you want the flow to separate.

Again folks - with regards to this prototype, I'd urge people to take a more relaxed attitude...  We have a bunch of opinions here, and the water tower has its own opinions, and the water tower will do as it will...

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ABCD: Always Be Counting Down

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: biosehnsucht on 01/06/2019 07:37 pm
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.

But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.

The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.

Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

I wonder if this will give them another advantage:

On launch from an unprepared surface (Mars) you could potentially ignite all the engines at a low throttle setting and allow them a short time to stabilize before moving to full thrust.

I wonder if you could combine retro-thrust from RCS with what gravity Mars has to provide hold down forces during Raptor startup (and thus if the engines don't come up to a low throttle correctly, you can abort without taking off). This assumes large RCS thrusters are already available for landing in "most" wind conditions on Earth, and possibly RCS as a backup to Raptors for touchdown on Mars / for landing on unprepared surfaces.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/06/2019 07:42 pm
What are the disadvantages of a dual-bell nozzle?

They are:

(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;

(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).

I believe there’s one more:

(3) there’s a transient instability spike at a certain atmospheric pressure where it transitions from bell to the other. This could damage or destroy the bell. However this may not be an issue because during ascent the booster will transition this atmospheric pressure and it won’t (I assume) have these bells. The Starship will ignite in vacuum so avoids the issue. On descent there may be concern for the Starship, but careful engine shutdown and reignition selection could avoid this issue, as could throttling.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: John Alan on 01/06/2019 08:15 pm
What are the disadvantages of a dual-bell nozzle?

They are:

(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;

(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).

I believe there’s one more:

(3) there’s a transient instability spike at a certain atmospheric pressure where it transitions from bell to the other. This could damage or destroy the bell. However this may not be an issue because during ascent the booster will transition this atmospheric pressure and it won’t (I assume) have these bells. The Starship will ignite in vacuum so avoids the issue. On descent there may be concern for the Starship, but careful engine shutdown and reignition selection could avoid this issue, as could throttling.

My take is, with proper use of the throttle, you stay away from the transition points...
Taper off as your getting close to a critical altitude, then quickly adjust to "step across the issue"...

My 2 cent opinion...  ;)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: indaco1 on 01/06/2019 09:36 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 01/06/2019 10:04 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/06/2019 10:07 pm
Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.

And they only need 3 engines that work at sea level for landing.  The rest can (and ultimately probably will) be optimised entirely for vacuum.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: indaco1 on 01/06/2019 10:30 pm
The 3 dual mode engines could be used for not much time after MECO, just to keep gravity losses low, and cut off as the veichle has lost mass enough.   Most of SS burn phase during launch and most of delta V could be made by vacuum optimized engines only.
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/07/2019 02:12 am
Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.

And they only need 3 engines that work at sea level for landing.  The rest can (and ultimately probably will) be optimised entirely for vacuum.

No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.

The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.

Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 02:21 am
Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.

And they only need 3 engines that work at sea level for landing.  The rest can (and ultimately probably will) be optimised entirely for vacuum.

No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.

The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.

Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: AC in NC on 01/07/2019 02:40 am
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?

Pretty sure 31 is readily doable based on past discussion.  https://forum.nasaspaceflight.com/index.php?topic=41363.msg1729558#msg1729558

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 02:41 am
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?

Pretty sure 31 is readily doable based on past discussion.  https://forum.nasaspaceflight.com/index.php?topic=41363.msg1729558#msg1729558
I think these bells are a lot larger tho...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 01/07/2019 02:44 am
Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.

And they only need 3 engines that work at sea level for landing.  The rest can (and ultimately probably will) be optimised entirely for vacuum.

No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.

The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.

Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?
It is quite possibe to fit 31 1.3m wide engines on the bottom of a 9m wide booster. Outer ring of 12, middle ring of 12 and inner ring of 6 with 1 in the middle or example.
9m booster cross sectional area = 64 sqm 31 raptos at 1.3m cross sectional area 41 sqm. Of course whether this is a good idea or not is another matter. I assume it is as they tend to be quite a smart bunch.
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 02:48 am
Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.

And they only need 3 engines that work at sea level for landing.  The rest can (and ultimately probably will) be optimised entirely for vacuum.

No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.

The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.

Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?
It is quite possibe to fit 31 1.3m wide engines on the bottom of a 9m wide booster. Outer ring of 12, middle ring of 12 and inner ring of 6 with 1 in the middle or example.
9m booster cross sectional area = 64 sqm 31 raptos at 1.3m cross sectional area 41 sqm. Of course whether this is a good idea or not is another matter. I assume it is as they tend to be quite a smart bunch.
Agreed - however are we positive the three engines we’re seeing on the BFH are 1.3 meters in diameter? With the advent of a dual bell design are we sure the diameter hasn’t increased? Someone cut out one of the workers in an image with the engines showing and turn him sideways...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: AC in NC on 01/07/2019 02:52 am
Agreed - however are we positive the three engines we’re seeing on the BFH are 1.3 meters in diameter?

Including some spare pixel "padding", I've got 6.8 engines across the 9m diameter for 1.32m.  I'd bet the pixel padding brings that right in at 1.3m.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 02:56 am
Agreed - however are we positive the three engines we’re seeing on the BFH are 1.3 meters in diameter?

Including some spare pixel "padding", I've got 6.8 engines within 9m for 1.32m
Thanks - looks pretty spot on.

(And now your picture will show up on Facebook claiming that SpaceX has added a second stage to their BFH...)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JH on 01/07/2019 03:01 am
Also, renders of the booster show that it widens to ~10 meters near the base.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/07/2019 03:05 am
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?

- Because it is a clean separation at the angle change. Separation in a normal bell is random and that is what causes the damage.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/07/2019 03:21 am
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.

But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.

The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.

Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

- A nozzle optimized for SL will produce more thrust at SL than a larger nozzle optimized for a higher altitude, whether it is separated or not. This is due to atmospheric pressure miss match at the nozzle exit.

John
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/07/2019 03:27 am
Also, renders of the booster show that it widens to ~10 meters near the base.

Exactly. And with 10 meters diameter the fit is more generous. Based on some images it is also possible that only the central 7 will gimbal. The outer ones are more tightly packed.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/07/2019 05:02 am
Given how many things have changed, I wouldn't rule the vacuum raptor out just yet.
It was a temporary compromise even as presented...  So with the switch to metal and change in time-table - who knows.

Maybe the aforementioned "Acceleration" also includes more work on the engine?

Me, I'd love to see a "cross" arrangement, with 3 SL engines inline, and 2 large diameter bells on two sides of it.
Or maybe 5 in the center, and then two large ones:

    . .
O   .   O
    . .

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/07/2019 07:15 am
Given how many things have changed, I wouldn't rule the vacuum raptor out just yet.
It was a temporary compromise even as presented...  So with the switch to metal and change in time-table - who knows.

Their big constraint was cost limiting them to Ø9m composite tooling, original ITS was much larger diameter.

But now the tooling costs with stainless steel construction have probably eased up considerably and the TPS requirement for stainless steel superheavy is probably vanished so that a higher deltaV bigger superheavy becomes more optimal (lots of other benefits too).  We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 01/07/2019 09:02 am
>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)

A few months ago Musk  tweeted something about 42 engines as a BFR upgrade.

Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.

Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for  Starship.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 01/07/2019 09:37 am
What are the disadvantages of a dual-bell nozzle?

They are:

(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;

(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).

I believe there’s one more:

(3) there’s a transient instability spike at a certain atmospheric pressure where it transitions from bell to the other. This could damage or destroy the bell. However this may not be an issue because during ascent the booster will transition this atmospheric pressure and it won’t (I assume) have these bells. The Starship will ignite in vacuum so avoids the issue. On descent there may be concern for the Starship, but careful engine shutdown and reignition selection could avoid this issue, as could throttling.

Ok, that fits sooo well together that I am going out and speculate: This is one of the reasons for the BFH. The flow separation and instability is very hard to model. Whatever software they use, its probably not taking ALL effects that matter in reality into account. So validating the flow instability for Raptor and throttled engine fireing is a very valid reason to do BFH. They cant test it on the test stand due to the constant air pressure there. They probably cant model it to perfection. They absolutely need to know the boundary conditions for operating Raptor, both on the Starship and on Super Heavy. That might also be the reason we see the dual nozzle bells. And thats also the reason why its not important to have the latest Raptor engines in the ship, the uncertainty is in the nozzle design, not the engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 01/07/2019 10:48 am
Seems we have got to a situation where we need another big update from uncle Elon to clarify where things stand. After the Dear Moon announcement things seemed to be clarified, but since then there have been a number of changes that have clouded the picture.

The move to Stainless steel has given freedom from the RCF winding mandrel size requirements and opened up the possibility of a different diameter Super Heavy or at least the possibility of a 10m skirt which in turn would give more engine options.

The “radically redesigned” Raptor is another confusing issue. This could mean new alloys, it could mean the engines are more powerful (larger scale) or might mean the nozzles have been redesigned to the dual bell shape or all three. We don’t know for sure.

So this leaves things a bit muddled. What will the lower diameter of the Super Heavy be? How many raptor engines of what size will it include and what thrust rating will they have? We can see our way through some of this, but will need further announcements to properly clarify the situation.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 01/07/2019 12:40 pm
>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)

A few months ago Musk  tweeted something about 42 engines as a BFR upgrade.

Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.

Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for  Starship.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/07/2019 12:46 pm
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.

But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.

The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.

Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.

I wonder if this will give them another advantage:

On launch from an unprepared surface (Mars) you could potentially ignite all the engines at a low throttle setting and allow them a short time to stabilize before moving to full thrust.

Mars is essentially a vacuum. Even a very highly expanded vacuum engine could run at low throttle without flow sep.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: philw1776 on 01/07/2019 01:05 pm
>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)

A few months ago Musk  tweeted something about 42 engines as a BFR upgrade.

Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.

Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for  Starship.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.

This ^^^
Long felt that 31 engines even with the add on mass of the 10m skirt was a design patch not an enhancement.  Whether it's 19, 21 or whatever, I also look for raptor thrust upscaling.
What's interesting about this is let's say it's 19 engines.  Even if initial Super Heavy's have interim not quite full thrust raptors, the full stack could easily loft ~60+ tonnes to LEO.  Launches of Starlink could commence with gradual payload upratings, just like with the Merlin family.  Earlier useful flight starts and experience gained while learning and enhancing.
Schedule acceleration.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nomadd on 01/07/2019 02:19 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.

It is an SSFO.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/07/2019 02:44 pm
&gt;
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)

A few months ago Musk  tweeted something about 42 engines as a BFR upgrade.

Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.

Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for  Starship.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.
That's reasonable.

The other option is that the extra push will be used for larger payloads and heavier tankers.

Or they'll wait for the return of the 12m rocket for that...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nomadd on 01/07/2019 02:55 pm
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.
Higher chamber pressure equals lower thrust?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: launchwatcher on 01/07/2019 03:59 pm
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.
Higher chamber pressure equals lower thrust?
If you hold pump power constant.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 01/07/2019 04:19 pm
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.
Higher chamber pressure equals lower thrust?
If you hold pump power constant.

But to vary chamber pressure during development, assuming you're not modifying the chamber geometry - wouldn't you be changing pressure by varying pump power?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 01/07/2019 05:38 pm
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.
Higher chamber pressure equals lower thrust?
If you hold pump power constant.

But to vary chamber pressure during development, assuming you're not modifying the chamber geometry - wouldn't you be changing pressure by varying pump power?
EM is looking at 25MPa Pc for Raptor at IOC then gradually raising it to at least 30MPa later. If the new Raptor produces 3.25MN SL thrust at 25MPa Pc, it will produce around 3.9MN SL thrust at 30MPa Pc assuming everything else is constant.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Peter.Colin on 01/07/2019 05:53 pm
>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)

A few months ago Musk  tweeted something about 42 engines as a BFR upgrade.

Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.

Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for  Starship.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.

This ^^^
Long felt that 31 engines even with the add on mass of the 10m skirt was a design patch not an enhancement.  Whether it's 19, 21 or whatever, I also look for raptor thrust upscaling.
What's interesting about this is let's say it's 19 engines.  Even if initial Super Heavy's have interim not quite full thrust raptors, the full stack could easily loft ~60+ tonnes to LEO.  Launches of Starlink could commence with gradual payload upratings, just like with the Merlin family.  Earlier useful flight starts and experience gained while learning and enhancing.
Schedule acceleration.




The 19 vs 42 debate...


Suppose if 19 big Raptor engines give the same total thrust as 42 small engines, the total weight of the engines is going to decide what makes more sense for Superheavy.


Many small engines or a few big ones, obviously 42 engines will win.. :-)



Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Spindog on 01/07/2019 06:02 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.

So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile? This transition zone with associated flow separation would have to be avoided, correct? It would seem that even if it was quick, the rapid transitions of all of the many engines would cause serious vibration. So what is the solution? The only one I could think if is throttling down significantly, which I understand is already done late in the booster burn on an F9 to avoid over-acceleration with an almost empty rocket.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 06:03 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.

So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile? This transition zone with associated flow separation would have to be avoided, correct? It would seem that even if it was quick, the rapid transitions of all of the many engines would cause serious vibration. So what is the solution? The only one I could think if is throttling down significantly, which I understand is already done late in the booster burn on an F9 to avoid over-acceleration with an almost empty rocket.
Why would the booster have dual bell Raptors?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Spindog on 01/07/2019 06:16 pm
I was assuming (as others have?) that there's now a single raptor design for both the booster and the starship. I can see how the dual bell would work for the Starship but it would make more sense to me if there was a sea  level bell on the booster version ...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 06:22 pm
I was assuming (as others have?) that there's now a single raptor design for both the booster and the starship. I can see how the dual bell would work for the Starship but it would make more sense to me if there was a sea  level bell on the booster version ...
I think perhaps there is a common engine design but the bells are different - kind of like the Merlin 1D SL and Vac (although I think this is an oversimplification as I believe the M1D Vac has other optimizations / modifications from the sea level version).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 01/07/2019 06:28 pm
Having them all(booster and spaceship) regeneratively cooled chambers and bells probably makes them more the same then different. Different bell shapes for different ambient air pressure regimes. Probably stepped bell for spaceship and "normal" SL bell for booster.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JonathanD on 01/07/2019 06:29 pm
I think perhaps there is a common engine design but the bells are different - kind of like the Merlin 1D SL and Vac (although I think this is an oversimplification as I believe the M1D Vac has other optimizations / modifications from the sea level version).

Maybe the dual-bell config allows them to have fewer design differences between the variants compared to Merlin 1D vs Vac.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/07/2019 06:31 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.

So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile?

This depends on the exact chamber pressure and expansion ratios, but in all likelihood, no. Any Raptor at full chamber pressure is going to have attached flow all the way out any expansion ratio up to 75:1 or thereabouts.

The smaller inner bell, (if that's what it is), is only ever relevant at SL AND at deep throttle. It has to be at BOTH conditions, or flow won't detach from the larger nozzle. The small bell (if that's what it is) plays no part at all in ascent, which is all at either wide-open throttle or reduced pressure.

Why would the booster have dual bell Raptors?

Because it has to deep throttle to land at SL on Earth. That's the only flight regime where a 15 or 20 ER nozzle on Raptor plays any role at all. In all other flight regimes the flow will be fully attached out to an ER of 40 or greater.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 01/07/2019 06:44 pm
What are the disadvantages of a dual-bell nozzle?
From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow  expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.

Just a note about requirements.

They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn.   No other altitudes and no transitions, this is not a SSTO.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.

So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile?

This depends on the exact chamber pressure and expansion ratios, but in all likelihood, no. Any Raptor at full chamber pressure is going to have attached flow all the way out any expansion ratio up to 75:1 or thereabouts.

The smaller inner bell, (if that's what it is), is only ever relevant at SL AND at deep throttle. It has to be at BOTH conditions, or flow won't detach from the larger nozzle. The small bell (if that's what it is) plays no part at all in ascent, which is all at either wide-open throttle or reduced pressure.

Why would the booster have dual bell Raptors?

Because it has to deep throttle to land at SL on Earth. That's the only flight regime where a 15 or 20 ER nozzle on Raptor plays any role at all. In all other flight regimes the flow will be fully attached out to an ER of 40 or greater.
So you're saying the booster, which doesn't need a vac optimized engine (as apposed to the Starship), needs to have a dual bell nozzle, rather than just the smaller diameter one? I'm not questioning your statement, rather trying to understand for myself. It would seem that unless absolutely required that the inherent risks of a dual bell over a single diameter design would dictate that the booster has a sea level optimized bell (like the M1D SL) and the Starship would have the dual bell (analogous to the F9 Upper Stage using the M1D VAC).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: abaddon on 01/07/2019 06:53 pm
I think the idea is an optimal SL expansion nozzle would have instability attachment issues when deep throttling.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/07/2019 07:12 pm
So you're saying the booster, which doesn't need a vac optimized engine (as apposed to the Starship), needs to have a dual bell nozzle, rather than just the smaller diameter one? I'm not questioning your statement, rather trying to understand for myself. It would seem that unless absolutely required that the inherent risks of a dual bell over a single diameter design would dictate that the booster has a sea level optimized bell (like the M1D SL) and the Starship would have the dual bell (analogous to the F9 Upper Stage using the M1D VAC).

The ~1.3 meter dual-bells on the BFH are not analogous to MVac. They are MUCH smaller than MVac's ~2.5 meter nozzle. We haven't seen a true vac-optimized nozzle (ER 120:1 to 200:1, diameter near 4 meters) on Raptor since 2017 IAC, and we have no information that SpaceX is building one anytime soon. Musk said at the dearmoon presentation that they might build one later, but going to a single configuration helps they get flying much faster.

The SH booster could run about a 40:1 standard single nozzle and be fine for deep throttle at SL. So to answer your first question, no it doesn't NEED a dual-bell nozzle, and there are reasons it would be better with a 40:1 single shorter bell. Lower ER is also helpful for packing and gimbal range. They might sacrifice some ISP to use a shorter nozzle and pack more engines to generate more thrust. Once the ER gets above 50:1 it will have trouble deep throttling.

I can't tell if this ~1.3 meter dual bell is 40:1 or 50+:1. But I can tell that if it is indeed 40:1, there's no point in it having the smaller inner bell, while at 50:1 or greater the smaller inner is perfectly sized for deep throttle landings at SL on Earth.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/07/2019 07:18 pm
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.
I don't think so.  2000kN thrust at 330s Isp gives 617kg/s propelleant flow, with 930kg/m³ densified 3.8:1 methalox mix that is 0.66m³/s, and with 80MPa outlet pressure get 53MW ideal pumping power.  But as I recall rocket turbopumps sacrifice a lot of efficiency for weight savings, and are generally only in 70-80% range.  With 70% pump efficiency would have about 74MW shaft power.  That is close enough to quoted figure that I don't think we are looking at a massive change in thrust.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: launchwatcher on 01/07/2019 07:38 pm
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.
Higher chamber pressure equals lower thrust?
If you hold pump power constant.
But to vary chamber pressure during development, assuming you're not modifying the chamber geometry - wouldn't you be changing pressure by varying pump power?
That's likely what SpaceX is doing.  We, on the other hand, are estimating Raptor thrust based on fixed numbers (like turbopump power) tweeted by Elon that are probably still being tweaked...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 01/07/2019 07:47 pm
So you're saying the booster, which doesn't need a vac optimized engine (as apposed to the Starship), needs to have a dual bell nozzle, rather than just the smaller diameter one? I'm not questioning your statement, rather trying to understand for myself. It would seem that unless absolutely required that the inherent risks of a dual bell over a single diameter design would dictate that the booster has a sea level optimized bell (like the M1D SL) and the Starship would have the dual bell (analogous to the F9 Upper Stage using the M1D VAC).

You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/07/2019 07:57 pm
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.

SpaceX don't yet have a vacuum test stand do they?  Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/07/2019 08:26 pm
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.

SpaceX don't yet have a vacuum test stand do they?  Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.

1) This is NOT a vacuum nozzle, and doen't need a vac stand for testing.
2) There is no detachment nor re-attachment during ascent, for a 1.3 meter nozzle on Raptor at full throttle. This nozzle will have fully attached flow for all flight regimes except terminal landing. This is assured by the amount of mass Raptor has to flow.

The only time the nozzle flow will be detached from the wall is during startup (see SSME ringing) and terminal landing. Both can be tested at sea level.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: philw1776 on 01/07/2019 08:28 pm
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.

SpaceX don't yet have a vacuum test stand do they?  Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.

Thus a major rationale for The hopper to fly ASAP so they can get on with 1st generation engine pre-production
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/07/2019 08:38 pm
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.

SpaceX don't yet have a vacuum test stand do they?  Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.

Thus a major rationale for The hopper to fly ASAP so they can get on with 1st generation engine pre-production

There are many things they need to test, but flow separation transients on ascent are not one of them since that won't ever happen.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 01/07/2019 08:47 pm
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 01/07/2019 08:54 pm
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.
At that point, it's less "deep throttling" and more "prelighting" backup engines in case there is an engine failure during descent. By having a couple spare engines idling at >5%, they can be rapidly throttled up if needed, more easilly than doing a fresh turbopump ignition>Main chamber ignition>full power landing.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/07/2019 08:58 pm
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.

You don't need deep throttling, but you do need to rapidly go from ~6 times the vehicles weight in thrust down to 1.5 times or less. Anything else leads to longer landing burns and wastes a lot of heavy fuel that you just accelerated to Mach 8 and back.

You could fire 5 engines at max thrust, then cut out 4 of them. But you better have a really good handle on the shutdown transients, and have a high confidence that the single engine won't die on you at the worst possible moment.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: 1 on 01/07/2019 09:34 pm
You could fire 5 engines at max thrust, then cut out 4 of them. But you better have a really good handle on the shutdown transients, and have a high confidence that the single engine won't die on you at the worst possible moment.

I think this is quite likely. SpaceX already has a lot of experience doing 1-3-1 landing burns with F9.  They're clearly comfortable with betting the fate of their boosters on their confidence in understanding the shutdown (and startup!) transients of multiple Merlins. That said, I can see them using deep throttling on the SH booster in the early days, especially if it's less risky, and then switching to a more aggressive landing profile as they accrue similar confidence in raptor. Starship will need deep throttle capability regardless, so data from those early flights will still be very valuable to them.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/07/2019 11:07 pm
1) This is NOT a vacuum nozzle, and doen't need a vac stand for testing.
2) There is no detachment nor re-attachment during ascent, for a 1.3 meter nozzle on Raptor at full throttle. This nozzle will have fully attached flow for all flight regimes except terminal landing. This is assured by the amount of mass Raptor has to flow.

The only time the nozzle flow will be detached from the wall is during startup (see SSME ringing) and terminal landing. Both can be tested at sea level.

Good points.

Are there any scenarios where we might see large throttle changes at altitude?
- Throttling back for max q on superheavy before throttling back up?
- An engine failing at startup on final landing burn (from 1-2km up - admittedly not too much different from Sea Level), another engine failing from a nozzle transient then could be rather bad.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/08/2019 12:14 am
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.

You still need throttling. Because while turning off engines can be reasonably predictable, turning them on is not. And so if you ever hope to increase thrust during a landing, all the necessary engines need to be already running.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: John Alan on 01/08/2019 12:28 am
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.

You still need throttling. Because while turning off engines can be reasonably predictable, turning them on is not. And so if you ever hope to increase thrust during a landing, all the necessary engines need to be already running.

Three engines for landing also gives then a rather powerful roll axis (pivot on the rocket centerline) tool they can use to line up direction it will face once down...
One engine on centerline can't do that (as recent CRS "event" pointed out)

2 cents on subtopic...  ;)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: corneliussulla on 01/08/2019 06:07 am
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.

You don't need deep throttling, but you do need to rapidly go from ~6 times the vehicles weight in thrust down to 1.5 times or less. Anything else leads to longer landing burns and wastes a lot of heavy fuel that you just accelerated to Mach 8 and back.

You could fire 5 engines at max thrust, then cut out 4 of them. But you better have a really good handle on the shutdown transients, and have a high confidence that the single engine won't die on you at the worst possible moment.

I am not sure landing on one engine would be compatible with the risk profile associated with intra planetary travel such as airlines which is Musks stated goal.

Edit/Lar: Fix quotes. Use the preview button people!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Crispy on 01/08/2019 11:40 am
It gets nowhere near. T/W at landing is around 2.
Blue Origin's New Shepard can hover before landing, but that has engines capable of deeper throttling, and a heavier mass fraction.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 01/13/2019 12:02 pm
How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/13/2019 02:40 pm
How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?

I think they are hydrodynamic bearings which use the cryo propellants.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 01/13/2019 03:12 pm
How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?

I think they are hydrodynamic bearings which use the cryo propellants.

John
I suspected that too but then I have more questions:
1. How do they start? Hydrodynamic bearing has to have some static pressure untill it  it transits to hydrodinamic mode.
2. Cryo proppelants are fluid, but also low pressure. How's the sealing to between the bearing and hot, high pressure side of pump? How do cryo fluids in bearings don't start to boil?

Does anyone have any rocket turbopump cross section? I dissassembled large turbines with additional pumps for hydrostatic or hydrodynamic bearings but how does this work without additional pump?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/14/2019 04:20 pm
Moved my post here from the update thread:

Pics of the bells and the bottom of a fin. Taken on 1-8-19 during the fit test excitement.

Using BocaChicaGal's photo, focusing on left 2 Raptors: Enhanced to show additional Raptor components through gap in the leg.

Note**  that the left bell and expander section are mismatched in O.D. ... whereas center is much smoother transition.

Peeking around leg, we can see part of the thrust chamber and lack of plumbing. This confirms EM's statement that thee were not flight ready units but development test articles being used as boiler plate fit check and alignment units...

Thanks for the great work BCG/Nomadd/Austin... and many others..

Yep, great images. Those clearly show (IMO) that the top part is a real engine part. But the larger nozzle looks welded on.

Could it still be a dual nozzle engine? Eh.. It's possible, but the likelihood just dropped even further. The rough look of that transition (and how it differs from engine to engine) makes it seem like a temporary thing.

EDIT: Attached image being discussed.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/14/2019 04:41 pm
How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?

I think they are hydrodynamic bearings which use the cryo propellants.

John
I suspected that too but then I have more questions:
1. How do they start? Hydrodynamic bearing has to have some static pressure untill it  it transits to hydrodinamic mode.
2. Cryo proppelants are fluid, but also low pressure. How's the sealing to between the bearing and hot, high pressure side of pump? How do cryo fluids in bearings don't start to boil?

Does anyone have any rocket turbopump cross section? I dissassembled large turbines with additional pumps for hydrostatic or hydrodynamic bearings but how does this work without additional pump?

The pumps are bathed in propellants prior to start (pre-chill), bearings probably also. Remember, you have 3 bar of pressure in the tanks.

This little paper outlines some of the problems and solutions:
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910009666.pdf

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/14/2019 05:35 pm
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:

https://twitter.com/blueorigin/status/950365085091811330

Quote
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 01/14/2019 06:07 pm
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:

https://twitter.com/blueorigin/status/950365085091811330

Quote
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Hydrodynamic bearings have rubbing until the speed is high enough, thus they have a lot of wear with start/stop cycles. Using a hydrostatic bearing pump means you can eliminate the rubbing before start-up, eliminating that source of wear and tear from start/stop cycles, increasing the cycle life.

Could be that the bearings are also hydrodynamic, it's just that they use the hydrostatic pump during start-up and spin-down.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 01/14/2019 06:10 pm
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:

https://twitter.com/blueorigin/status/950365085091811330

Quote
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Hydrostatic and hydrodynamic bearings are different.
https://en.m.wikipedia.org/wiki/Fluid_bearing

I thought that Raptor uses hydrodynamic ones because I couldn't see the piping for hydrostatic bearings in that Raptor render from IAC.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: IntoTheVoid on 01/14/2019 07:04 pm
At that point, it's less "deep throttling" and more "prelighting" backup engines in case there is an engine failure during descent. By having a couple spare engines idling at >5%, they can be rapidly throttled up if needed, more easilly than doing a fresh turbopump ignition>Main chamber ignition>full power landing.
My recollection (which admittedly could be wrong) was that Elon's justification for landing on 3 engines at low throttle was to be able to land on 2 at higher throttle, if there was an engine issue. I therefore, wouldn't expect any 'idling' engines.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Oersted on 01/14/2019 07:19 pm
Maybe the Super Heavy could have different-sized engine bells on different engines and direct fuel to them accordingly, to optimize performance durent ascent and descent. Just throwing the idea out there...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 01/14/2019 08:11 pm
More on bearings:

Quote
Yes, plenty of work has been done in that area. The best primer is probably the paper that kicked off the whole hydrostatic bearing discussion (this is what you mean by "some high pressure gas"), "Reddecliff and Vohr, 1969". A lot of work has been done since then. The French had a hydrostatic bearing test stand for the never-completed European staged combustion engine (TPX/TPTech). Pratt & Whitney wanted to replace the pump side bearing in one of the Space Shuttle turbopumps with a hydrostatic one, but nothing came of this either.

IHI, the Japanese rocket turbopump manufacturer, has released information about a turbopump that they want to sell overseas that uses hydrostatic bearings, so they are probably the furthest along. Unless of course SpaceX uses hydrostatic, but for all we know their bearings run on fairy dust.

Superconducting magnetic bearings have been studied as well.

Here are the three largest problems for these bearings:

    The stiffness is tiny compared to rolling element bearings. Hydrostatic bearings have at most a tenth the stiffness, magnetic bearings maybe only a hundredth or a thousandth. This means that rotordynamic instability is a huge development risk, since it can't be well predicted.
    They require an additional supply of propellant, which lowers the efficiency of the turbopump and increases the complexity.
    Rolling element bearings have made huge strides, so expendable and even some reusable rockets have really no need. Silicon nitride bearing balls together with induction-harded Cronidur steel races have made some other components in the Space Shuttle look really bad in terms of life time.

https://space.stackexchange.com/questions/14848/has-any-work-been-done-on-alternatives-to-rolling-element-bearings-in-rocket-tur

So with more reading from other sources I don't see any definite conclusion that Raptor has fluid bearings, and with compactness of bearing positions and absence of some  visible pressure source for fluid bearings I believe that it's more probable that Raptor uses advanced roller bearings especially as they don't need that high level of clean fluid as fluid bearings.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: philw1776 on 01/14/2019 09:08 pm
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:

https://twitter.com/blueorigin/status/950365085091811330

Quote
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn

Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 01/14/2019 09:12 pm
Snip...
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
I suspect it probably does. Although this is probably moderated by the fact that Blue have not reached orbit yet.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: gongora on 01/14/2019 09:15 pm
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?

BE-4's thrust has nothing to do with Raptor.  If SpaceX cared about having a bigger engine they'd build one.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: matthewkantar on 01/14/2019 10:08 pm
SpaceX set out to build a cheap reliable Mars transportation system. The size of the engines was determined by the engineering trades made in the (apparently ongoing) design process. If they were dumb enough to get into a pissing match over engine size, they would have the biggest engine powering a more expensive or less capable or incapable spacecraft.

Matthew
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 01/14/2019 10:19 pm
SpaceX set out to build a cheap reliable Mars transportation system. The size of the engines was determined by the engineering trades made in the (apparently ongoing) design process. If they were dumb enough to get into a pissing match over engine size, they would have the biggest engine powering a more expensive or less capable or incapable spacecraft.

Matthew

Of course they are! And I have every hope that they will succeed. Of course they wouldn't be so dumb as to have some stupid engine size competition. But that still wouldn't stop Elon being irked by the size of Jeff's rocket engine at some level.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/14/2019 10:24 pm
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?

Cost to orbit and ISRU on Mars are the only metrics Space X cares about.  High Isp and high thrust to weight with low maintenance are the major goals.  Bigger engines have worse thrust to weight due to longer fluid flow paths.  They also reduce failure redundancy and mass-production advantages, and get more expensive and difficult to manufacture and assemble (larger heavier components that can't be easily man-handled).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lemurion on 01/14/2019 10:46 pm
Any degree of irk that Elon may feel from the size of BE-4 will likely vanish should Raptor fly first, which it should do within the next 60-90 days at the outside-- especially since the reference to deep throttling tests indicates that Blue may not have achieved full design thrust yet anyway.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/15/2019 03:31 am
Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor?  Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: matthewkantar on 01/15/2019 05:11 am
Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor?  Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?

Yes. The other facility will be economical soon, the entirety of outer space.

Edited 01.15.19, added "be"
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mme on 01/15/2019 05:26 am
Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor?  Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?
My recollection is that modeling showed that using more smaller engines counterintuitively resulted in a better T/W even including all of the plumbing.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: TrevorMonty on 01/15/2019 06:07 am


Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:

https://twitter.com/blueorigin/status/950365085091811330

Quote
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Hydrostatic and hydrodynamic bearings are different.
https://en.m.wikipedia.org/wiki/Fluid_bearing

I thought that Raptor uses hydrodynamic ones because I couldn't see the piping for hydrostatic bearings in that Raptor render from IAC.

Hydrostatic have less wear so better for multiple starts the downside is lower performance. For RLV the lower performance is worthwhile trade for longer engine live.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/15/2019 06:32 am
Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor?  Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?
Yes. The other facility will economical soon, the entirety of outer space.
A really high stakes way of testing a nozzle, with more limited instrumentation potential.  In particular long term vac nozzle reliability will be hard to prove out without an altitude test facility.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 01/15/2019 08:40 am
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?

BE-4's thrust has nothing to do with Raptor.  If SpaceX cared about having a bigger engine they'd build one.
EM tweeted that Raptor's TP power will be 74.6MW (100,000HP) so the new radically redesigned Raptor may have more thrust than BE-4. We will not know until EM announces the specs. for the new version of Raptor. I doubt EM bothers about what size engines BO develop.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/15/2019 03:43 pm


Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:

Quote
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Hydrostatic and hydrodynamic bearings are different.
https://en.m.wikipedia.org/wiki/Fluid_bearing

I thought that Raptor uses hydrodynamic ones because I couldn't see the piping for hydrostatic bearings in that Raptor render from IAC.

Hydrostatic have less wear so better for multiple starts the downside is lower performance. For RLV the lower performance is worthwhile trade for longer engine live.

Speculation on my part. Could be either.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/15/2019 08:47 pm
Looks like the fake Raptors have been removed. (See attached image, from bocachicagal in the update thread: https://forum.nasaspaceflight.com/index.php?topic=47120.msg1901640#msg1901640)

Looks like the Raptor parts were even more minimal that it appeared at first. And the lower nozzle part wasn't even welded on, it looks like it was just a loose piece that was welded around it to just hang from the small nozzle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 01/15/2019 09:05 pm
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 01/15/2019 09:29 pm
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!

Well to be fair, a dual nozzle is still a possibility. (and I say that as someone who very much doubts it) It's just that the supporting evidence for it has withered away.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 01/15/2019 09:34 pm
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?

BE-4's thrust has nothing to do with Raptor.  If SpaceX cared about having a bigger engine they'd build one.
EM tweeted that Raptor's TP power will be 74.6MW (100,000HP) so the new radically redesigned Raptor may have more thrust than BE-4. We will not know until EM announces the specs. for the new version of Raptor. I doubt EM bothers about what size engines BO develop.

If it has increased thrust from the 2017 version it'll likely be because chamber pressure was upped from 250bar to 300 bar.

99% chance it's getting a 1.3m Raptor. If it wasn't they wouldn't have bothered with the 1.3m mockup bells.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tyrred on 01/16/2019 07:29 am
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!

Well to be fair, a dual nozzle is still a possibility. (and I say that as someone who very much doubts it) It's just that the supporting evidence for it has withered away.

Still a possibility, my question being "Why are these placeholder engines equipped with a nozzle skirt that isn't even attached?"

Photoshoot optics of the Hopper for how it will *approximately look with the upcoming production engines?

Protection of the placeholder engines from the elements?  (Why protect an engine that most likely will not be flown on this test article? And how would they be more protective than the usual blue coverings used on Merlins?)

Some type of test of a bell extension for gritty Boca Chica conditions?

UFO DefenceTM?

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: edzieba on 01/16/2019 08:26 am
Don't worry, speculation will now move from OMG dual bell! to OMG retractable nozzle!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 01/16/2019 09:30 am
Don't worry, speculation will now move from OMG dual bell! to OMG retractable nozzle!
Calm down, the real flight Raptors may well just have conventional nozzles after all. The loosely fitted nozzle extensions on the placeholders gave the visual appearance of dual bells. We should find out within the next couple of months when the flight Raptors get installed. The nozzle extensions on the placeholders may have only been held by friction.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: edzieba on 01/16/2019 11:06 am
They were just held on by gravity, the inner diameter of the lower bell segment being slightly smaller than the outer diameter of the upper bell segment attached to the thrust chamber.
They could be a close match to the desired final Raptor configuration, (dual bell and/or extendable nozzle section), or they could just be reject Merlin bell segments tin-snipped to length as 'close enough' boilerplate stand-ins.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cppetrie on 01/16/2019 02:32 pm
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!

Well to be fair, a dual nozzle is still a possibility. (and I say that as someone who very much doubts it) It's just that the supporting evidence for it has withered away.

Still a possibility, my question being "Why are these placeholder engines equipped with a nozzle skirt that isn't even attached?"

Photoshoot optics of the Hopper for how it will *approximately look with the upcoming production engines?

Protection of the placeholder engines from the elements?  (Why protect an engine that most likely will not be flown on this test article? And how would they be more protective than the usual blue coverings used on Merlins?)

Some type of test of a bell extension for gritty Boca Chica conditions?

UFO DefenceTM?

I suspect the placeholder engines are wrapped more to protect from prying eyes than the elements. They are actual test hardware after all.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 01/20/2019 09:18 pm
Lacking a vacuum testing facility could a Raptor vacuum expansion nozzle be adequately tested for mechanical durability and cooling adequacy using a partial conical base filling plug held inside the nozzle?  The (ablatively covered or transpiration cooled?) cone would deflect flow outwards to hug the nozzle walls without separation that would otherwise occur due to over-expansion.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 01/21/2019 12:58 pm
Lacking a vacuum testing facility could a Raptor vacuum expansion nozzle be adequately tested for mechanical durability and cooling adequacy using a partial conical base filling plug held inside the nozzle?  The (ablatively covered or transpiration cooled?) cone would deflect flow outwards to hug the nozzle walls without separation that would otherwise occur due to over-expansion.

SpaceX will soon have cheap access to the biggest vacuume facility ever- Orbit. Testing isnt going to be an issue.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 01/27/2019 12:05 pm
There has been some discussion on the engineering thread suggesting that the Raptor engine over all burns fuel rich, although pre-burners are heavily oxygen rich on the oxygen side and fuel rich on the fuel side. I had assumed that overall the raptor burned oxygen rich? Is fuel rich the norm in rocket engines?

This lead on to the topic of the Raptor engine mixture ratio. There seems to be a degree of uncertainty over the exact ratio that will be used as well as the nozzle expansion ratio and bell geometry. So is this just lack of public knowledge of the specific number that SpaceX have settled on? Or is it possible that SpaceX are thinking to change these figures as part of the optimisation program? This then leads to the question of overall optimisation if they change the mixture ratio won't they then need to change the proportions of LOX and Methane carried and move the LOX-methane bulkhead?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 01/27/2019 12:35 pm
Is fuel rich the norm in rocket engines?

It is.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 01/27/2019 02:49 pm
Is fuel rich the norm in rocket engines?

It is.

At least in the west. Russia invested heavily in oxygen-resistant alloys so they have some fairly advanced engines running oxygen rich.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 01/27/2019 02:58 pm
Is fuel rich the norm in rocket engines?

It is.

Correct me if I am wrong but I think the reason for "all" rocket engines being fuel rich is that they run cooler that way and minimal impact to the isp.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 01/27/2019 03:30 pm
At least in the west. Russia invested heavily in oxygen-resistant alloys so they have some fairly advanced engines running oxygen rich.

RD-180 uses oxygen-resistant alloys for its oxygen-rich preburner, but as AFAIK the overall cycle is still fuel-rich. I've never heard of overall oxidiser-rich engines, but maybe they exist. There certainly do exist oxygen-rich thrusters, for instance those using hydrogen peroxide monopropellant, but that's a different situation.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mmeijeri on 01/27/2019 03:34 pm
Correct me if I am wrong but I think the reason for "all" rocket engines being fuel rich is that they run cooler that way and minimal impact to the isp.

Yeah, and in fact fuel-rich typically has slightly better Isp than stoichiometric, in addition to being cooler. Also, fuel-rich is more compatible with ordinary alloys. At least, that's my understanding.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: b.lorenz on 01/28/2019 08:35 pm
Correct me if I am wrong but I think the reason for "all" rocket engines being fuel rich is that they run cooler that way and minimal impact to the isp.

Yeah, and in fact fuel-rich typically has slightly better Isp than stoichiometric, in addition to being cooler. Also, fuel-rich is more compatible with ordinary alloys. At least, that's my understanding.

So I have heard too. Since most (though not all) rocket fuel are hydro(carbons), while oxidizer typically consists of compounds of oxygen and nitrogen, fuel rich means trading a bit of power for a lower mean molecular mass of the exhaust, allowing optimization of Isp.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: guckyfan on 01/29/2019 09:47 am
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 01/30/2019 11:53 am
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.

Someone may correct me, but I understand it as this:
Water has a worse molecular mass (18) than methane (16). So I would assume that water does not increase ISP as much as methane. Also, you never get a stochiometric combustion anyway because the combustion process is not perfectly uniform within the engine. As a result, you would get oxygen rich local environments in the exhaust. I can imagine that this would potentially damage the engine bell. By burning fuel rich, the non-uniformity of combustion will always (or say to much much higher likelihood) stay fuel rich, which is non-corrosive.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 01/30/2019 12:20 pm
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.

Someone may correct me, but I understand it as this:
Water has a worse molecular mass (18) than methane (16). So I would assume that water does not increase ISP as much as methane. Also, you never get a stochiometric combustion anyway because the combustion process is not perfectly uniform within the engine. As a result, you would get oxygen rich local environments in the exhaust. I can imagine that this would potentially damage the engine bell. By burning fuel rich, the non-uniformity of combustion will always (or say to much much higher likelihood) stay fuel rich, which is non-corrosive.

Yep, anything containing oxygen will likely be your heaviest combustion product, that's why you're running fuel rich in the first place. You can do that up to the limit where going richer would mean exhausting too much solid C which reduces efficiency while not increasing isp. High efficiency tripropellant concepts typically use hydrogen as the working fluid for that reason.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: guckyfan on 01/30/2019 12:25 pm
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.

Someone may correct me, but I understand it as this:
Water has a worse molecular mass (18) than methane (16). So I would assume that water does not increase ISP as much as methane. Also, you never get a stochiometric combustion anyway because the combustion process is not perfectly uniform within the engine. As a result, you would get oxygen rich local environments in the exhaust. I can imagine that this would potentially damage the engine bell. By burning fuel rich, the non-uniformity of combustion will always (or say to much much higher likelihood) stay fuel rich, which is non-corrosive.

Yep, anything containing oxygen will likely be your heaviest combustion product, that's why you're running fuel rich in the first place. You can do that up to the limit where going richer would mean exhausting too much solid C which reduces efficiency while not increasing isp. High efficiency tripropellant concepts typically use hydrogen as the working fluid for that reason.

I came up with the idea mostly because of propellant production on Mars. Propellant is valuable there because of energy requirement to produce it. Using water as a third propellant would reduce that cost. If nothing else the added complexity of a three propellant engine would probably be a reason not to go that way.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: edzieba on 01/30/2019 12:32 pm
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.
If instead of water you inject more fuel and oxidiser, you get the Thrust Augmented Nozzle (https://selenianboondocks.com/2007/11/thrust-augmented-nozzles/). As far as I am aware, no examples have been flown.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Chris Bergin on 01/30/2019 02:05 pm
Report to mods about off topic posts, so drag this back and I'll start a thread 2 in the near future (long threads tend to drift).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: spacenut on 01/30/2019 02:49 pm
Ok, at this point, does anyone know at what stage Raptor is at?  Have they increased thrust to max yet?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Taxidermista on 01/31/2019 11:31 am
Teslarati: SpaceX sends “radically redesigned” Starship engine to Texas for hot-fire tests (https://www.teslarati.com/spacex-radically-redesigned-starship-engine-shipped-texas-hot-fire-testing/)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 01/31/2019 02:21 pm
Teslarati: SpaceX sends “radically redesigned” Starship engine to Texas for hot-fire tests (https://www.teslarati.com/spacex-radically-redesigned-starship-engine-shipped-texas-hot-fire-testing/)
Quote
Most notable was an obvious secondary preburner/turbopump stack and the lack of any exhaust port, whereas M1D relies on a single turbopump and exhausts the gases used to power it.
So does Raptor no longer have the oxygen tubopump integrated above the MCC?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 01/31/2019 02:26 pm
Teslarati: SpaceX sends “radically redesigned” Starship engine to Texas for hot-fire tests (https://www.teslarati.com/spacex-radically-redesigned-starship-engine-shipped-texas-hot-fire-testing/)
Quote
Most notable was an obvious secondary preburner/turbopump stack and the lack of any exhaust port, whereas M1D relies on a single turbopump and exhausts the gases used to power it.
So does Raptor no longer have the oxygen tubopump integrated above the MCC?

Thats the design of FFSC. Full Flow Staged Combustion.
Basically a preburner for oxygen and methane.
Both preburners on separate shafts power their own fuel/oxygen pump. So the output of the preburners goes directly to the combustion chamber already heated and partly combusted. This should make the mixing of the fuel and oxidizer much better. Also having separate shafts for both sides of the fuel/oxidizer means no seal with oxygen on one side and fuel on the other.

EDIT: So I misunderstood the question. I believe the design I saw had oxy in line and integrated with the combustion chamber and the fuel turbopump off to the side. I remember someplace a picture of the engine and annotations labelling everything. I think livingjw did it.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 01/31/2019 02:44 pm
EDIT: So I misunderstood the question. I believe the design I saw had oxy in line and integrated with the combustion chamber and the fuel turbopump off to the side. I remember someplace a picture of the engine and annotations labelling everything. I think livingjw did it.
That picture of Raptor is in the article.

edit: the labeled version is somewhere in this thread.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 01/31/2019 05:07 pm
Are there any pictures of the Raptor mentioned in the Teslarati article?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: abaddon on 01/31/2019 06:50 pm
I believe this (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1700053#msg1700053) is the post with the labelled version mentioned above.

Also this flow diagram (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1826715#msg1826715) might be helpful in visualizing how it works.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 01/31/2019 07:49 pm
I believe this (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1700053#msg1700053) is the post with the labelled version mentioned above.

Also this flow diagram (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1826715#msg1826715) might be helpful in visualizing how it works.

The one difference from that flow diagram is that I expect is for the fuel line for regenerative cooling to start from after the first pump stage and return to the second stage. Given their insane chamber pressures, pushing the coolant through the entire cooling circuit at full pressure would require making the structure of the rocket a lot heavier than it needs to be.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 01/31/2019 08:00 pm
I believe this (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1700053#msg1700053) is the post with the labelled version mentioned above.

Also this flow diagram (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1826715#msg1826715) might be helpful in visualizing how it works.

The one difference from that flow diagram is that I expect is for the fuel line for regenerative cooling to start from after the first pump stage and return to the second stage. Given their insane chamber pressures, pushing the coolant through the entire cooling circuit at full pressure would require making the structure of the rocket a lot heavier than it needs to be.

Could be, there is some hidden plumbing, but the regen out seems to go directly to the preburner.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/01/2019 01:03 am
And - it's on!

https://twitter.com/elonmusk/status/1091154222173712384
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/01/2019 01:14 am
For the Twitter/Tapatalk-less

Elon Musk ✔ @elonmusk
 Preparing to fire the Starship Raptor engine at @SpaceX Texas
9:00 PM - Jan 31, 2019


Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JBF on 02/01/2019 01:14 am
No wonder they had to beef up the test stand.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/01/2019 01:17 am
Can anyone superimpose a full-up M1D SL for comparison?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nehkara on 02/01/2019 01:19 am
https://twitter.com/elonmusk/status/1091156245132673024
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureMartian97 on 02/01/2019 01:24 am
WOW is that thing massive!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JonathanD on 02/01/2019 01:24 am
Wow....that's a piece of kit.  Can't wait to see it fire up.  Is it just me or are some of the pipes hanging pretty far out?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JBF on 02/01/2019 01:32 am
Ok, here is an eyeball comparison assuming the height of the people is about the same. I do not guarantee the accuracy.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wolfram66 on 02/01/2019 01:35 am
Tweaked Elon’s twitPic
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: MichaelBlackbourn on 02/01/2019 02:08 am
I love that we're seeing a Mars rocket engine just hit the stand... And it will likely be flying in a month or two. Methane FFSC here we come!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 02:25 am
That rocket engine is a piece of art.  8)

Finally we can take the crown from the Russians!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: acsawdey on 02/01/2019 02:29 am
That rocket engine is a piece of art.  8)

Finally we can take the crown from the Russians!

The turbo machinery is arranged so it all fits above the footprint of the bell, probably so they can cram 31 of these things under the "Super Heavy" booster.

I'm curious what the large metal piece with the slots in the side at the back is though.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 02:30 am
250 tonnes of thrust in the 'high thrust' version must mean the 300+ bar variant.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jak Kennedy on 02/01/2019 03:36 am
So what type of nozzle is that? An inverse duel bell? It looks reversed compared to a duel bell nozzle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cuddihy on 02/01/2019 03:51 am
It looks like a more graduated dual bell nozzle. Is that a thing or is SpaceX inventing that now?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Okie_Steve on 02/01/2019 04:12 am
I keep remembering how the Merlin up through C had lots of pipes valves bells whistles and spangles on it. Supposedly Elon asked Tom how they could make it simpler and cheaper which resulted in the much cleaner looking D.

I wonder if something similar will happen with Raptor in a future update. 1st make it work ...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 04:16 am
If the nozzle is 1.3m, here is a rough size comparison using pixel measuring between Raptor, Merlin, and a 1.78m man.

(https://i.imgur.com/osNnhWQ.jpg)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Joseph Peterson on 02/01/2019 04:18 am
Would anyone care to speculate on how long it will take to set up for the first test fire?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: yokem55 on 02/01/2019 04:19 am
Just to make the units a little more clear:
'200 metric tons' == 1.96 meganewton = 440,000 foot lbs.

Or in simple terms its got almost double the umph of the first development Raptors.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nomadd on 02/01/2019 04:24 am
 Am I the only one who can't tell where that engine is in relationship to the people in the picture?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Scylla on 02/01/2019 04:35 am
Am I the only one who can't tell where that engine is in relationship to the people in the picture?
Nope.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Cinder on 02/01/2019 04:36 am
Assuming the man is holding the engine from a peripheral component, that would put him maybe half an arm's length away from below the edge of the engine.  Otherwise it's hard to pick out the shadows' perspective lines.
Tentative arm in pink.

Otherwise the man's side (leg and up) coincides with the edge of the bell shadow.  That'd make the engine seem larger by the foreshortening of the actual distance between engine and the blue-shirt man behind it.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 04:46 am
Am I the only one who can't tell where that engine is in relationship to the people in the picture?

It's in between the two of them
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/01/2019 05:06 am
Ok, here is an eyeball comparison assuming the height of the people is about the same. I do not guarantee the accuracy.

It's way off, here are my quick measurements:
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/01/2019 05:10 am
Wow, cool!  8)

One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.

And it has to be said... [sarcasm]Wow, look at that glorious dual bell nozzle.[/sarcasm]   ::) ;D

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/01/2019 05:11 am
So what type of nozzle is that? An inverse duel bell? It looks reversed compared to a duel bell nozzle.

It's not a dual bell nozzle, it never was. Sigh...

Its amusing to see the number of people on twitter bending over backwards to try to see the dual nozzle where one isn't. Even our own Chris G is getting in on the act.

EDIT... Here is an enhanced/leaned up image someone posted in the follow up to Elons tweet. Source: https://twitter.com/John_Gardi/status/1091161496506257408

EDIT 2 - I added an even brighter version to make the bell outline clearer.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tyrred on 02/01/2019 05:21 am
There appears to be a large blue graphics card attached.  To make the engine look bigger?  ;)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/01/2019 05:40 am
It is an interesting coincidence that this first version of Raptor will have roughly the same thrust as RS-25/SSME. Although in a much more compact form. (nozzle diameter 1.3m vs SSME's 2.3m)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/01/2019 06:07 am
For the record (and to bound just how radical the raptor redesign is):

https://twitter.com/tritexan/status/1091158260726546435

Quote
Full flow staged combustion cycle still?

https://twitter.com/elonmusk/status/1091158310261125120

Quote
Yes
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Cheapchips on 02/01/2019 06:17 am
Wow, cool!  8)

One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.


This is the first production engine, according to Musk's December tweet.  I wonder how much scope they have for simplification, since all the lessons from Merlin's will be incorporated? 
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/01/2019 06:21 am
Wow, cool!  8)

One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.


This is the first production engine, according to Musk's December tweet.  I wonder how much scope they have for simplification, since all the lessons from Merlin's will be incorporated?

While true, the first engines on a test stand would always have extra instrumentation and margins. Compare some of the first M1Ds on the test stand (image 1) to a recent production version (image 2). I think they already have plans to simplify it. But this first version may be good enough to mount on the Hopper.

There is precedent for this - Grasshopper also used a pre-production M1D (similar to picture 1), before they optimized it for the engine clustering in F9v1.1.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 02/01/2019 07:48 am
Wow, that engine looks chunky compared to Merlin. Since thrust is 200t, maybe 250t later I doubt TWR will increase much over Merlin (~100t), but it might be equal.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jpo234 on 02/01/2019 07:58 am
Wow, that engine looks chunky compared to Merlin. Since thrust is 200t, maybe 250t later I doubt TWR will increase much over Merlin (~100t), but it might be equal.

From Tom Mueller (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/):

Quote
“The Merlin holds the thrust-to-weight record for now,” he said. “But the Raptor’s coming.”
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/01/2019 08:45 am
Wow, that engine looks chunky compared to Merlin. Since thrust is 200t, maybe 250t later I doubt TWR will increase much over Merlin (~100t), but it might be equal.

Forget the plumbers nightmare aspect, the weight is mostly down to how short they have managed to keep the high pressure flow paths (minimising the containing pressure vessel weight), and those turbopumps and pre-burners look very compact, LOX turbopump still between gimbal and thrust chamber for shortest flowpath for 60% of volume flow, and methane turbopump with the other 40% of flow seems very efficiently integrated - not big either.

NK33 staged combustion managed 137:1 T/W 50 years ago at a lower 15MPa chamber pressure.  RD191 with similar 26MPa chamber pressure has T/W of 89:1, but bulky legacy turbopump layout.  http://www.russianspaceweb.com/images/rockets/angara/rd191/infograph_1.jpg
RD193 is supposedly improved to 103:1 T/W 10 years back (not flown)

Won't be too surprising if 40-50 years improvement in analysis tools sees a jump to 150-200:1 T/W for Raptor as hinted at.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cebri on 02/01/2019 12:09 pm
Wow, that's a pretty cool looking engine.

From a (relatively) simple GG to an SC full flow engine. A think the term "generational leap" falls short to describe what SpaceX has done. I'm still a bit sceptical that the engine will be as "reusable" as SpaceX wants it to be, at least for the first iterations. I wonder how it compares to the SSME in complexity and refurbishment needs after each flight.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: First Mate Rummey on 02/01/2019 12:18 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).

Raptor source:
https://twitter.com/elonmusk/status/1091156245132673024

BE-4 source:
BE-4 is capable of producing 2,400 kN (550,000 lbf) thrust
https://www.blueorigin.com/engines/be-4
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/01/2019 12:39 pm
If the nozzle is 1.3m, here is a rough size comparison using pixel measuring between Raptor, Merlin, and a 1.78m man.

(https://i.imgur.com/osNnhWQ.jpg)
Nice drawing! Thanks for the comparison.

(Looks like the gimbal mount was a late add in your drawing )
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Darth Pathetic on 02/01/2019 12:45 pm
Poll: When will full-scale hot-fire testing of Raptor begin?

You can (almost) stop the poll now.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 02/01/2019 12:48 pm
The teslarati article wasn't kidding about the mass of plumbing all around the engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/01/2019 12:48 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing! (edit not so big after all)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 01:46 pm
Looking at the picture, I think you can see the fuel compressor/turbine, but it's located behind the combustion chamber in this picture.
(https://i.imgur.com/HZjV9HG.jpg)

i think the major difference between this and the old raptor is the new fuel rich turbine assembly is at an angle relative to the line of thrust.

it still looks like the oxygen line is integrated into the top of the combustion chamber.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 01:46 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

I'm pretty sure BE-4 has a 1.8m nozzle
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jpo234 on 02/01/2019 01:49 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

Could you add SSME?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: acsawdey on 02/01/2019 02:02 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

I'm pretty sure BE-4 has a 1.8m nozzle

Yes, see here: https://forum.nasaspaceflight.com/index.php?topic=45518.0 (https://forum.nasaspaceflight.com/index.php?topic=45518.0)

Still a big difference between 1.3m and 1.8m.

Seems that the ratio of nozzle area almost exactly matches the inverse ratio of chamber pressures:

(1.8^2)/(1.3^2) ~= 250/130

The packaging of Raptor is so much more compact than BE-4.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/01/2019 02:05 pm
That rocket engine is a piece of art.  8)

Finally we can take the crown from the Russians!

The turbo machinery is arranged so it all fits above the footprint of the bell, probably so they can cram 31 of these things under the "Super Heavy" booster.

I'm curious what the large metal piece with the slots in the side at the back is though.

Looks like a combination of gimbal bearing and thrust structure. The top half of the TCV is connected to it, see the right side of the test stand picture.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/01/2019 02:07 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

I'm pretty sure BE-4 has a 1.8m nozzle

~ = +- 60cm (I updated thö picture tho')

(google: how tall is Jeff Bezos (http://naplesherald.com/wp-content/uploads/Private-Space-Travel-NH.jpg))
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jpo234 on 02/01/2019 02:28 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

I'm pretty sure BE-4 has a 1.8m nozzle

Using  ImageJ  (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.


Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/01/2019 03:35 pm
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

I'm pretty sure BE-4 has a 1.8m nozzle

Using  ImageJ  (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.

That's a manlet. (1.30 m)

So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).
...
Yes that is bit insane... Look at the size of that thing!

Could you add SSME?

You need to ask ZachF to draw it (https://forum.nasaspaceflight.com/assets/42585.0/1542151.jpg).

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jpo234 on 02/01/2019 03:45 pm
Using  ImageJ  (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.

That's a manlet.

Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: acsawdey on 02/01/2019 03:55 pm
Using  ImageJ  (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.

That's a manlet.

Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).

No but a 2.4m nozzle does not match up with the BE-4 reverse engineering thread that I linked above. Those calculations get you something in the neighborhood of 1.8m. I think there is a bit less uncertainty in that than in the man shaped outline somebody stuck on that picture.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 04:00 pm
Using  ImageJ  (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.

That's a manlet.

Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).

No but a 2.4m nozzle does not match up with the BE-4 reverse engineering thread that I linked above. Those calculations get you something in the neighborhood of 1.8m. I think there is a bit less uncertainty in that than in the man shaped outline somebody stuck on that picture.

There are now New Glenn cutaways you can compare the width of the BE-4 against the width of the rocket as a whole (7m), you get somewhere between 1.8-1.9m.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/01/2019 04:09 pm
Using  ImageJ  (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.

That's a manlet.

Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).

No but a 2.4m nozzle does not match up with the BE-4 reverse engineering thread that I linked above. Those calculations get you something in the neighborhood of 1.8m. I think there is a bit less uncertainty in that than in the man shaped outline somebody stuck on that picture.

Right. This image from Blue Origin is more accurate (no perspective distortion), it supports a ~1.7-1.8m diameter.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ulm_atms on 02/01/2019 04:16 pm
Man, SpaceX played a mean game of Tetris with that plumbing.  It is much more compact then I was expecting considering the combustion cycle type.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/01/2019 04:19 pm
Man, SpaceX played a mean game of Tetris with that plumbing.  It is much more compact then I was expecting considering the combustion cycle type.

If they are going to stick so many of them close to each other on the booster, the compact volume is a necessity.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: alhenry1231 on 02/01/2019 04:21 pm
The thrust per cubic meter is for fun.
it is the volume calculated from the diameter of the nozzle by the height.

of note I calculate the nozzle of the BE4 closer to 1.9m than 1.8m based on Blue's 6' scale.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wargrim on 02/01/2019 04:35 pm
Good to see this progress. Hope to see the test being successful, this has to be the number one item on the checklist for the Hopper program to start flying.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/01/2019 04:51 pm
It hits my eye as if I'm looking at a V12 engine sat vertically on a display stand.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: acsawdey on 02/01/2019 05:32 pm
The thrust per cubic meter is for fun.
it is the volume calculated from the diameter of the nozzle by the height.

of note I calculate the nozzle of the BE4 closer to 1.9m than 1.8m based on Blue's 6' scale.

Nice.

I think thrust per square meter is more like it for the purposes of seeing how much you can put on a booster. I assumed a square the size of the nozzle diameter:


           w     th    a  th/a
BE4        1.9  2.45 3.61 0.68
Raptor 250 1.3  1.96 1.69 1.16
Raptor 300 1.3  2.45 1.69 1.45
M1D        0.92 0.85 0.85 1.00
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cuddihy on 02/01/2019 05:48 pm
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/01/2019 06:10 pm
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)

Fuel line leading from the fuel turbopump to the regen cooling system.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/01/2019 06:19 pm
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)

Fuel line leading from the fuel turbopump to the regen cooling system.

- I think that might be coming from a fuel valve downstream of the pump. Hard to tell, but I believe the turbo-pump and pre-burner is on the backside and not shown on either photo.

- black coating might a little insulation, or not.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/01/2019 06:32 pm
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)

Fuel line leading from the fuel turbopump to the regen cooling system.

- I think that might be coming from a fuel valve downstream of the pump. Hard to tell, but I believe the turbo-pump and pre-burner is on the backside and not shown on either photo.

- black coating might a little insulation, or not.

John

Could be. Looks to me like the fuel pump is mounted transversely now, and that line is coming done off the end of it. Most of the pump and all of the preburner would still be hidden behind the engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/01/2019 06:38 pm
I think the nozzle outline does appear a little odd, though it may be nothing. The shape of the nozzle might be due to trying to limit separation at part throttle at sea level similar to this discussion taken from "Nozzle Design"
by R.A. O'Leary and J. E. Beck, Spring 1992


....
The first choice for an SSME nozzle contour would obviously be one that maximized
nozzle thrust; that is, a Rao optimum contour. However, if a Rao optimum contour was
used, the wall pressure at the nozzle exit, (pw(exit)) would be much lower than the
ambient pressure at sea level. Even at 100 percent power levels, corresponding to a
chamber pressure equal to 3,000 psia, (pw(exit)) would be 4.6 psia or 31 percent of the
ambient pressure at sea level. Past experience showed that nozzle flow separation would
likely occur if the wall pressure approached this level. Since nozzle flow separation is
dependent upon a number of variables (boundary-layer thickness, pressure gradient,
Mach number, etc.) and is thus difficult to accurately predict, additional margin in exit
pressure was sought. Some margin was also required to permit sea-level testing of engine
throttling capability.

A Rao design resulted in a wall angle of 7.5° at the nozzle exit. By reducing this angle,
additional flow turning is produced, and then, an increase in nozzle wall pressure is
created. A study was performed by Pratt & Whitney Rocketdyne engineers in which a large
number of parabolic-shaped contours, with a variety of different initial wall angles (qmax)
and exit wall angles (qe), were analyzed. After careful analysis of these contours, it was
determined that a parabolic contour with qmax=37° and qe=5.3° would produce the
desired wall pressure increase with the least amount of performance loss. The wall exit
pressure was raised 24 percent (from 4.6 psia to 5.7 psia) at a cost of only 0.1 percent in
nozzle efficiency. Validation of the design approach was provided by subsequent testing of
the SSME which demonstrated that the engine can be throttled to below 80 percent power
level at sea level without nozzle flow separation.
....

This technique could allow the Raptor to throttle much deeper since it is not already over expanded like the SSME due to its exit diameter constraints.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/01/2019 06:41 pm
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)

Fuel line leading from the fuel turbopump to the regen cooling system.

- I think that might be coming from a fuel valve downstream of the pump. Hard to tell, but I believe the turbo-pump and pre-burner is on the backside and not shown on either photo.

- black coating might a little insulation, or not.

John

Could be. Looks to me like the fuel pump is mounted transversely now, and that line is coming done off the end of it. Most of the pump and all of the preburner would still be hidden behind the engine.

At first I thought that as well, but it doesn't look big enough. I'm betting we are seeing a fuel valve. Reading tea leaves is fun. ;^)

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: drzerg on 02/01/2019 06:42 pm
is it actually both turbo pumps stacked on each other on top? or inside each other?

look at the ropes. they hold engine close to the center. if there was huge massive machinery it will be below engine on that photo
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/01/2019 06:46 pm
is it actually both turbo pumps stacked on each other on top? or inside each other?

look at the ropes. they hold engine close to the center. if there was huge massive machinery it will be below engine that photo

There are two lift ropes, and the CG could be anywhere between them.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dglow on 02/01/2019 06:54 pm
It hits my eye as if I'm looking at a V12 engine sat vertically on a display stand.

Well-said. I am awestruck by the pure aesthetic beauty of this engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: wannamoonbase on 02/01/2019 06:58 pm
Man, SpaceX played a mean game of Tetris with that plumbing.  It is much more compact then I was expecting considering the combustion cycle type.

If they are going to stick so many of them close to each other on the booster, the compact volume is a necessity.

That was my first impression too.  They stacked all the plumbing into a tight profile so they can cram those things under a big booster easier.  Ease of plumbing connection appears to be factored in from the start.

Looks like a very impressive tight package.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/01/2019 07:10 pm
I am pretty sure there is a fair size turbopump/preburner unit on the far side. Look at the other photo. You can see the fuel intake next to the inline Lox intake.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Oersted on 02/01/2019 07:32 pm
First fire would probably be something less than a one-second burp, wouldn't it?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/01/2019 07:34 pm
Wow, cool!  8)

One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.


This is the first production engine, according to Musk's December tweet.  I wonder how much scope they have for simplification, since all the lessons from Merlin's will be incorporated?

While true, the first engines on a test stand would always have extra instrumentation and margins. Compare some of the first M1Ds on the test stand (image 1) to a recent production version (image 2). I think they already have plans to simplify it. But this first version may be good enough to mount on the Hopper.

There is precedent for this - Grasshopper also used a pre-production M1D (similar to picture 1), before they optimized it for the engine clustering in F9v1.1.

I suspect a lot of that piping is to supply boost pumps. They probably need boost pumps for both the fuel and oxidizer, and perhaps pumps going into the preburners (Ox into the fuel rich loop, and fuel into the ox-rich loop).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RoboGoofers on 02/01/2019 08:11 pm
I think the nozzle outline does appear a little odd, though it may be nothing. The shape of the nozzle might be due to trying to limit separation at part throttle at sea level similar to this discussion taken from "Nozzle Design"
by R.A. O'Leary and J. E. Beck, Spring 1992


....
The first choice for an SSME nozzle contour would obviously be one that maximized
nozzle thrust; that is, a Rao optimum contour. ...
Here's the link for that Rao (https://en.wikipedia.org/wiki/G._V._R._Rao) paper from 1958 "Exhaust Nozzle Contour for Optimum Thrust":
https://doi.org/10.2514/8.7324

(I don't know if it's available anywhere for free.)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/01/2019 09:06 pm
I think the nozzle outline does appear a little odd, though it may be nothing. The shape of the nozzle might be due to trying to limit separation at part throttle at sea level similar to this discussion taken from "Nozzle Design"
by R.A. O'Leary and J. E. Beck, Spring 1992


....
The first choice for an SSME nozzle contour would obviously be one that maximized
nozzle thrust; that is, a Rao optimum contour. ...
Here's the link for that Rao (https://en.wikipedia.org/wiki/G._V._R._Rao) paper from 1958 "Exhaust Nozzle Contour for Optimum Thrust":
https://doi.org/10.2514/8.7324

(I don't know if it's available anywhere for free.)

I haven't looked for the paper recently, but found this which gives you the general idea:
http://seitzman.gatech.edu/classes/ae6450/bell_nozzle.pdf

 The Rao paper outlines how he came up with the quadratic approximation to the Method Of Characteristics, MOC, optimum nozzle shape.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/01/2019 10:15 pm
Does anyone have any information as to whether Elon's quoted thrust was sea level or vacuum?

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Spindog on 02/02/2019 04:36 am
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/02/2019 04:48 am
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?

In 2012 DLR estimated 25 uses for an LH2 full-flow staged combustion, but with SpaceX's new alloys and CH4...??

Section 3.2

http://elib.dlr.de/78208/1/Prop2012-2.pdf
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lemurion on 02/02/2019 04:51 am
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?

That’s one of the things that will be fascinating to find out. Having said that, if you look at projected flight rates it definitely appears that SpaceX is looking to reuse Raptors both more times and more frequently.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tea monster on 02/02/2019 07:54 am
I tried to do my own processing of these images. I went into the page source to recover the direct images, then did a a basic 200% blowup (bicubic smoother) and smart-sharpen in Photoshop. The blow up images are a tradeoff between sharpening artifacts and visible detail.

The first two images are the untouched JPEG originals off twitter (in case anyone wants to experiment themselves). The second pair are the enlarged and enhanced. I've saved them as PNG files without compression to reduce any artefacts. 

EDIT: I did a googly search on the image and found that one guy on imgur has already tried to blow one of the images up, with seemingly good results. Just beware that any kind of sharpening and enhancing procedure can introduce stuff that isn't in the original image.

Link to Imgur page of CKStones: https://imgur.com/gallery/z5OAMfV
I've included his image as a direct attachment as well.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/02/2019 12:22 pm
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?

Hard to know with so little info about Raptor. From a theoretical perspective, Raptor is FFSC (should lower temps and pressures needed in preburners and pumps), while BE-4 uses an oxygen rich staged combustion cycle (in theory worse for reuse) but at almost half the final chamber pressure of Raptor. Raptor doesn't have to worry about finicky turbopump seals like BE4 does, on the other hand.

Edit: Forgot to add that BE4's turbopump uses hydrostatic bearings, which would give it a reusability edge over Raptor. But for all we know, operationalized Raptor could have hydro bearings too.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/02/2019 12:56 pm
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?

That’s one of the things that will be fascinating to find out. Having said that, if you look at projected flight rates it definitely appears that SpaceX is looking to reuse Raptors both more times and more frequently.
I predict even though Raptor is more compact and higher thrust, it will do better at this, because this is not their first reusable engine.  Yes, Blue has some experience from New Shepard, to be sure. But not quite the same, with dozens of reused Merlin engines flown now.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 02/02/2019 04:31 pm
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?

That’s one of the things that will be fascinating to find out. Having said that, if you look at projected flight rates it definitely appears that SpaceX is looking to reuse Raptors both more times and more frequently.
I predict even though Raptor is more compact and higher thrust, it will do better at this, because this is not their first reusable engine.  Yes, Blue has some experience from New Shepard, to be sure. But not quite the same, with dozens of reused Merlin engines flown now.

And the Raptors will get tested rapidly. After the first two demo missions, Merlin had already been fired twenty times. In just one SH/SS launch, there will be at least 50 Raptor ignitions!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Rei on 02/02/2019 11:58 pm
Elon's plane just landed in Waco.  Looks like he wants to be there for the raptor test firing  ;)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/03/2019 01:15 am
Cross posting this Raptor schematic from the autogenous pressurization thread. Better place for it. Started with the SSME schematic and changed it to reflect Raptor as I currently understand it.
Modified to add a heat exchanger in the methane pre-burner turbine exit flow. After adiabatic expansion pressurant gas was still too cold.

Modified (4 Feb 2019) changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/03/2019 01:42 am
Lovely diagram. Where's the heat exchanger to gasify the methane for tank pressurant? Or is it going to gasify on its own when introduced since it is hot (from the engine main combustion and nozzle heat exchange) and high pressure liquid which when released to a lower pressure environment, gasifies?

Edit: Keep reading though, the diagram just above this post has changed from when I asked the question...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/03/2019 01:54 am
Lovely diagram. Where's the heat exchanger to gasify the methane for tank pressurant? Or is it going to gasify on its own when introduced since it is hot (from the engine main combustion and nozzle heat exchange) and high pressure liquid which when released to a lower pressure environment, gasifies?

Yes, methane is hot after cooling MCC. Will gasify during pressure drop.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/03/2019 06:23 am
https://twitter.com/elonmusk/status/1091958352513425408

Quote
At @SpaceX Texas with engineering team getting ready to fire new Raptor rocket engine
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/03/2019 06:25 am
For the Twitterless,

Elon Musk ✔ @elonmusk
At @SpaceX Texas with engineering team getting ready to fire new Raptor rocket engine
2:15 AM - Feb 3, 2019
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/03/2019 06:59 am
Is that a pen in his hand?  Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/03/2019 07:09 am
Is that a pen in his hand?  Are the team all signing the nozzle perhaps? I am s ure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.

And yet trauma surgeons operate on people with 30+ hours under their belt; those 'a plane hit a train, they hit a bus  and it rolled on a car' days. 
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: su27k on 02/03/2019 07:12 am
Is that a pen in his hand?  Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.

Hear, hear, but come to think about it all the Raptor footage we've seen is at night, so this may be par for the course for them.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/03/2019 07:22 am
Is that a pen in his hand?  Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.

This is concern trolling of the highest degree, sir. But perhaps you can convince us otherwise by assuring us that you have deep contacts at McGregor or SpaceX HR job schedules? Surely you know when their job shift started and can tell us?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/03/2019 07:30 am
Some observations about the image.

#1 - I guess this is the final nail in the "dual bell" theory. I know some where speculating that "the interior could have a funky shape" (loosely translated). Nope.

#2 - Note the hexagonal plate behind the engine, that protects the sensitive parts of the engine. Could this be something added just for this engine test cell.... or, is this the load bearing plate for the engine hexagonal bay? The equivalent of the top plate in this M1D image:
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ejb749 on 02/03/2019 07:32 am
Aren't there strict sound limits at McGregor this time of night?  I wouldn't expect this to light up until daylight.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/03/2019 07:36 am
Aren't there strict sound limits at McGregor this time of night?  I wouldn't expect this to light up until daylight.

I don't think they're supposed to test after 2200 Local. Betting they'll try for Sunday so it'll make the Monday news cycle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/03/2019 07:40 am
Aren't there strict sound limits at McGregor this time of night?  I wouldn't expect this to light up until daylight.

I don't think they're supposed to test after 2200 Local. Betting they'll try for Sunday so it'll make the Monday news cycle.

SpaceX most certainly does not schedule test fires based on hitting news cycles. Musk flew in tonight, so they're almost certainly aiming for a static fire tonight (nominally).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/03/2019 07:44 am
Aren't there strict sound limits at McGregor this time of night?  I wouldn't expect this to light up until daylight.

I don't think they're supposed to test after 2200 Local. Betting they'll try for Sunday so it'll make the Monday news cycle.

SpaceX most certainly does not schedule test fires based on hitting news cycles. Musk flew in tonight, so they're almost certainly aiming for a static fire tonight (nominally).

They won't fire tonight if it's after midnight Local. Testing at a decent hour tomorrow IF THEY'RE READY won't rile the locals & Church Ladies and still makes the Monday news cycle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: EngrDavid on 02/03/2019 08:42 am
Is that a pen in his hand?  Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.

Elon is holding a torch light.  You can see the light shining in the bell.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: drzerg on 02/03/2019 09:06 am
i really want to see where is methane turbo pump :(
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/03/2019 09:07 am
...
Note the hexagonal plate behind the engine, that protects the sensitive parts of the engine. Could this be something added just for this engine test cell.... or, is this the load bearing plate for the engine hexagonal bay? The equivalent of the top plate in this M1D image:
Yes, it is part of the test stand, engine (with sensitive parts) is front of it:
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 02/03/2019 10:00 am
Cross posting this Raptor schematic from the autogenous pressurization thread. Better place for it. Started with the SSME schematic and changed it to reflect Raptor as I currently understand it.

John
Shouldn't the O2 for CH4 precombustion be taken after the GO2 heat exchanger and not after the LO2 pump?
O2 would be gasseous when injected to CH4 precombustion chamber which is better for combustion.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Bananas_on_Mars on 02/03/2019 10:17 am
Lovely diagram. Where's the heat exchanger to gasify the methane for tank pressurant? Or is it going to gasify on its own when introduced since it is hot (from the engine main combustion and nozzle heat exchange) and high pressure liquid which when released to a lower pressure environment, gasifies?

Yes, methane is hot after cooling MCC. Will gasify during pressure drop.

John

Actually for most of the flow through the engine, the methane should be supercritical, right?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/03/2019 10:19 am
Cross posting this Raptor schematic from the autogenous pressurization thread. Better place for it. Started with the SSME schematic and changed it to reflect Raptor as I currently understand it.

John
Shouldn't the O2 for CH4 precombustion be taken after the GO2 heat exchanger and not after the LO2 pump?
O2 would be gasseous when injected to CH4 precombustion chamber which is better for combustion.
IIUC, only the CH4 has a heat exchanger (the nozzle jacket)

However, could they have used "oxygen rich gas"?  I think the answer lies in the starting sequence of the engine.

Maybe after everything is running, this can be done, or maybe there is a heat exchanger not shown in the diagram.

Remember again this is based on the Wikipedia diagram for SSME, not on an actual Raptor diagram.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/03/2019 10:29 am
Is that a pen in his hand?  Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.

Elon is holding a torch light.  You can see the light shining in the bell.
That's flashlight for you Yanks, not a pen.

...IIUC, only the CH4 has a heat exchanger (the nozzle jacket)
...
How they then pressurize GO2 tank, electric heater? With Webasto?

(They have 14-62 ICE preheaters for use, you know?)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: alhenry1231 on 02/03/2019 01:10 pm
https://twitter.com/elonmusk/status/1091958352513425408

Quote
At @SpaceX Texas with engineering team getting ready to fire new Raptor rocket engine

I’m currently not around my computer, would someone be so kind as to get the nozzle diameter from Elon’s height of 1.828m [6’2”].

Edit- just realized Elon’s whole body is not in the picture. So this is not as straightforward as I thought. Still a nice nozzle shot for reference to diameter
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wolfram66 on 02/03/2019 01:40 pm
Enhanced twitpic to see inside bell and show that Elon is holding a penlight not a pen
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: kevinof on 02/03/2019 01:40 pm
1.8 m = 5 Feet, 11 (ish)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/03/2019 01:40 pm
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.

But you can try to estimate the size by this palm print here:

; P
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wolfram66 on 02/03/2019 01:41 pm
Enhanced twitpic to see inside bell and show that Elon is holding a penlight not a pen
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: alhenry1231 on 02/03/2019 01:55 pm
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.

But you can try to estimate the size by this palm print here:

; P

Perfect! Is the hand what you used to set your scale in your Merlin/BE4 image?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/03/2019 04:46 pm
Modified Raptor schematic to add a heat exchanger in the methane pre-burner turbine exit flow. After adiabatic expansion pressurant gas was still too cold.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/03/2019 10:06 pm
The pictures above show the nozzle to be smooth on the inside and smooth on the outside.  How happens cooling channels within that?  Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough).  Does that imply that its 3D printed?  Not liquid cooled?  Other?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/03/2019 10:11 pm
The pictures above show the nozzle to be smooth on the inside and smooth on the outside.  How happens cooling channels within that?  Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough).  Does that imply that its 3D printed?  Not liquid cooled?  Other?

The same way it is done for Merlin 1D (see image)... with internal channels. There is basically three layers... the inside, middle, and outside. The middle copper(?) or other alloy layer is milled to have lots of small channels.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/03/2019 10:43 pm
The pictures above show the nozzle to be smooth on the inside and smooth on the outside.  How happens cooling channels within that?  Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough).  Does that imply that its 3D printed?  Not liquid cooled?  Other?

https://en.wikipedia.org/wiki/SpaceX_rocket_engines#/media/File:Making_a_SpaceX_Engine.jpg
"Making a SpaceX Merlin 1D engine. The shiny nozzle here has channels etched vertically in it for the fuel to run down and back up, keeping the nozzle from melting during use. 2012 photo by Steve Jurvetson."
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/03/2019 11:44 pm
The pictures above show the nozzle to be smooth on the inside and smooth on the outside.  How happens cooling channels within that?  Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough).  Does that imply that its 3D printed?  Not liquid cooled?  Other?

The same way it is done for Merlin 1D (see image)... with internal channels. There is basically three layers... the inside, middle, and outside. The middle copper(?) or other alloy layer is milled to have lots of small channels.

They just use a copper alloy inner channeled wall with an outer high strength wall. Bell has alternating channels. See photo of milling test coupon below.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: octavomaran on 02/04/2019 12:49 am
If EverydayAstronaut’s sources are correct and if we’re extrapolating accurately from them, we may have just had a test fire of the raptor currently on the stand:

https://mobile.twitter.com/erdayastronaut/status/1092220360727187457?s=21

Full text of the tweet:
Reports coming out from people in the town of McGregor about a “big boom”... now we just have to wait to find out if it was a good big boom or a bad big boom 🤞😂 rocket engines can easily make either 😂
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jketch on 02/04/2019 01:05 am
<blockquote class="twitter-tweet" data-lang="en"><p lang="en" dir="ltr">Another 1 second test, 1 minute ago, sounded same, rattled my house.</p>&mdash; Xobmo (@xobmo) February 4, 2019 (https://twitter.com/xobmo/status/1092242089285439488?ref_src=twsrc%5Etfw)
<script async src="https://platform.twitter.com/widgets.js" charset="utf-8"></script>

Sounds like there was another firing.

Full text:

Another 1 second test, 1 minute ago, sounded same, rattled my house.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/04/2019 01:08 am
Video:

https://twitter.com/wjordaniv/status/1092242263177084929

Quote
We have a (brief) test fire!!! Only caught the audio  @Erdayastronaut @johnkrausphotos
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/04/2019 02:50 am
https://twitter.com/elonmusk/status/1092268892339273730

Quote
First firing of Starship Raptor flight engine! So proud of great work by @SpaceX team!!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/04/2019 02:53 am
https://twitter.com/elonmusk/status/1092268892339273730

Quote
First firing of Starship Raptor flight engine! So proud of great work by @SpaceX team!!
A) You quick!

B) If I said I wasn't waiting for this for like 2 hours now, I'd be lying.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mclumber1 on 02/04/2019 02:57 am
What can we glean from the exhaust plume in relation to the nozzle? 
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/04/2019 02:58 am
https://twitter.com/elonmusk/status/1092270756715737088
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/04/2019 03:00 am
What can we glean from the exhaust plume in relation to the nozzle?

Not much, really. It doesn't look a whole lot different than Be-4 tests.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ThePonjaX on 02/04/2019 03:02 am
We need someone to synchronize the video from Elon with the sound from B.J. Jordan.
I'd love the video from Elon has sound.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wolfram66 on 02/04/2019 03:04 am
And is the green from the copper in the cooling channels? Was there some burn-through to the inner layers?

Either way it was way cool!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Orbiter on 02/04/2019 03:05 am
https://twitter.com/elonmusk/status/1092272377889738753
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DigitalMan on 02/04/2019 03:05 am
Wow, 31 of those is going to be incredible
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/04/2019 03:06 am
I'd say that taking it from assembly to >1 sec burn in 3 days is a sign of very high confidence.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/04/2019 03:06 am
Wow - that was some bright ignition! What do you suppose that vapor cloud was behind the bell just before ignition. If you step thought the video I thought it was Going to be a RUD...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureMartian97 on 02/04/2019 03:07 am
https://twitter.com/elonmusk/status/1092272377889738753

Was that supposed to happen...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/04/2019 03:09 am
2nd raptor video attached
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: x15_fan on 02/04/2019 03:11 am
And is the green from the copper in the cooling channels? Was there some burn-through to the inner layers?

Either way it was way cool!

Same observation on original Raptor firing back in 2016. I think most people settled on the sensor in the camera being saturated and shifting green. I believe someone discussed this phenomena with digital camera sensors back then.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/04/2019 03:12 am
https://twitter.com/elonmusk/status/1092272377889738753

Was that supposed to happen...
Judging by the response of the observers, I'd say yes.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Orbiter on 02/04/2019 03:19 am
The video looks to be filmed from a distance of ~650-700 meters away judging from the sound delay and the engine was still powerful enough to rock the camera. Just shows how much more powerful of an engine the Raptor is compared to the Merlin. Congratulations to everyone in SpaceX for getting the first Raptor engine assembled.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: TrueBlueWitt on 02/04/2019 03:26 am
Anyone else wondering about the greenish color in the flame at the end of the burn? Curious how well their advanced materials are holding up.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: S.Paulissen on 02/04/2019 03:30 am
Wow - that was some bright ignition! What do you suppose that vapor cloud was behind the bell just before ignition. If you step thought the video I thought it was Going to be a RUD...

Before ignition is usually high pressure helium to spin up the turbopumps.  I don't know what they're using in raptors for spin up.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/04/2019 03:33 am
Anyone else wondering about the greenish color in the flame at the end of the burn? Curious how well their advanced materials are holding up.
Engine-rich combustion? Copper?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/04/2019 03:34 am
Saw the same thing with the development engine and IIRC the consensus was it's an oversaturated CCD in the camera.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/04/2019 03:36 am
Saw the same thing with the development engine and IIRC the consensus was it's an oversaturated CCD in the camera.
People keep abusing the word "consensus."

Consensus means that basically everyone agreed.

There DEFINITELY wasn't consensus there.

It's green.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lemurion on 02/04/2019 03:38 am
So now they have had the first fire, how long will it take to reach full thrust— a week?

At this rate it will hit full size and full thrust well before BE-4.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Navier–Stokes on 02/04/2019 03:39 am
https://twitter.com/elonmusk/status/1092280273599979520

Edit: fixed embedding

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/04/2019 03:40 am
https://twitter.com/elonmusk/status/1092280273599979520?s=09
Dan Guisinger

 
@dguisinger
 17m17 minutes ago
More
Discussion from the SpaceX reddit suggests it is the camera sensor unable to handle the brightness

1 reply 0 retweets 11 likes
Reply 1   Retweet   Like 11

Elon Musk

Verified account
 
@elonmusk
Follow Follow @elonmusk
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Replying to @dguisinger @Erdayastronaut @DanielDavisA
Engine use methox torch igniters. Green tinge is either camera saturation or a tiny bit of copper from the chamber.

8:34 PM - 3 Feb 2019
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/04/2019 03:42 am
https://twitter.com/elonmusk/status/1092282107639021569

Quote
Gaseous CH4/O2 & heavy duty spark plugs. Basically, a 💨 of insane power 😀
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: x15_fan on 02/04/2019 03:52 am
Saw the same thing with the development engine and IIRC the consensus was it's an oversaturated CCD in the camera.
People keep abusing the word "consensus."

Consensus means that basically everyone agreed.

There DEFINITELY wasn't consensus there.

It's green.

Fair enough, that was my recollection from several years ago. One thing that is weird, two tests were reported around 2 hours apart. If the jacket was being eaten up you wouldn't think they would go back for more in the same night. And if they had some quick adjustment to the flow-rate cooling the jacket for second test, you'd think they would release that video instead (unless the quick change was unsuccessful). In any case I don't see how this adds up to two tests in one night.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/04/2019 04:01 am
A small amount of engine-rich exhaust at shutdown is not necessarily a showstopper for initial hops.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/04/2019 04:12 am
BTW, I find it odd that Elon doesn't seem to know whether or not it was engine-rich combustion or saturation (and this doesn't look like saturation, FWIW).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/04/2019 04:16 am
If you want to know why I’m being kind of a pain in the butt about this, then read this thread by actual propulsion engineers who have tested and destroyed several rocket engines.

There’s even a term of art for it: “unintentional green.”

(and there were similar comments about a previous Raptor firing video)

https://mobile.twitter.com/unrocket/status/1092273797015625728


...this reply in particular:

https://mobile.twitter.com/wikkit/status/1092278314734112768
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/04/2019 04:29 am
BTW, I find it odd that Elon doesn't seem to know whether or not it was engine-rich combustion or saturation (and this doesn't look like saturation, FWIW).

Maybe because the amount of copper necessary to create the tint (if that's what it was) is small enough that it's not a problem during start-up / shut-down, or maybe that it's not a problem during initial tests, or maybe that the term you used above is more appropriate during steady state firing or during major anomalies when it is of more consequence.

Either way, maybe we should worry about it if engines prove short-lived because of it?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JonathanD on 02/04/2019 04:42 am
It is a test after all...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dnavas on 02/04/2019 05:37 am
Either way, maybe we should worry about it if engines prove short-lived because of it?

Well, to my completely, utterly, untrained eyes, that didn't look like a healthy progression of colors.  But I don't know if this is actually a fault to worry about, normal during bring up tests, or a complete red herring.  So, while I'm not going to worry about it, and I'll just be happy with the vids, I sure would appreciate the opinions of people with more experience at this sort of thing than I have.  No worries, no concerns, just interest.

edit to add [as apparently I was misunderstood]:
I am interested in the process of bringing a rocket engine up.  What things typically go wrong, what things typically go right, how long does it take to run through a test program, what things are you looking at and looking for, etc.  Creating a new rocket engine isn't the sort of thing that's done every day, so it's interesting to me.  Apologies if that makes me seem like a weirdo  :shrug:
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/04/2019 05:41 am
If I can’t concern troll myself, can someone else do it please? Is that what you are asking for?

Geez, this is BEYOND ridiculous.

Would Elon post it so quickly if the result wasn’t what they expected and hoped for? C’mon.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 02/04/2019 06:04 am
Would Elon post it so quickly if the result wasn’t what they expected and hoped for? C’mon.
Sure, because he's Elon Musk and he can do that.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tyrred on 02/04/2019 06:23 am
I'm not an expert, but...  Great scott!  What a beast.  I almost can't stand the magnificence.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FinalFrontier on 02/04/2019 06:24 am
Under certain mix conditions methane in combustion can produce a greenish tint or a green flash. Depends on temperature pressure % O2 in the burn and also other atmospheric gases interfering. The wide field video with sound shows much less green than the close in shot. Still good possibility exists it could also be due to internal erosion of some kind in the chamber. Really easy way to find out stick a camera or scope up there after it's cooled down and see if there is damage.

If no visible damage continue test campaign see if engine RUD's itself after multiple firings. Reminder the last SSME test saw a large hydrogen flame shooting out of part of the power-head or test stand and that was not really a massive show stopper, without the government shutdown test campaign would probably still be proceeding. Also engine did not undergo RUD despite failure. Chamber erosion is very different however, but again, if no visible damage is found with camera/scope then just keep firing and see what happens.

Easy way to find the flaws in a piece of machinery is push the test article until it breaks and then see where it broke and why. Like with the windblown starhopper fairing these things are test articles and are meant to be broken. Do not be surprised if they explode break fall over or otherwise fail.

What you don't want is to be on a flight to the moon and discover a flaw the hard way then, as Apollo 13 did. Or worse during launch or re-entry.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: woods170 on 02/04/2019 06:27 am
With regards to the green:
Over-analysis is running wild here (again).

And for all supposed rocket engine engineers here: unless you actually work for SpaceX, and specifically on Raptor, I suggest you keep your "facts" to yourselves for it is nothing but speculation.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FinalFrontier on 02/04/2019 06:28 am
With regards to the green:
Over-analysis is running wild here (again).
I agree. Folk's should just refer to Musk's tweet. And if it was actually chamber erosion that's not hard to find or to fix really, adjusting the burn/mix ratio can fix things like that.
BTW, I find it odd that Elon doesn't seem to know whether or not it was engine-rich combustion or saturation (and this doesn't look like saturation, FWIW).

Engine probably still cooling down+people need to sleep. They can come back in the morning and stick a camera in there to find out.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: SciNews on 02/04/2019 06:41 am
Edited Raptor test
https://www.youtube.com/watch?v=MAAzbjG_Duc
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FinalFrontier on 02/04/2019 06:43 am
If I can’t concern troll myself, can someone else do it please? Is that what you are asking for?

No.  Please re-read the actual words I wrote.  I think I was very explicit about what I was not interested in.  I wasn't terribly helpful about what I was interested in, though, so let me correct that.

I am interested in the process of bringing a rocket engine up.  What things typically go wrong, what things typically go right, how long does it take to run through a test program, what things are you looking at and looking for, etc.  Creating a new rocket engine isn't the sort of thing that's done every day, so it's interesting.  I'm interested in precisely the sort of thing one might expect someone that reads "the evolution of the spacex raptor engine" to be interested in.  So yes, I'm interested in "green", but I'm not interested in "oh gods, everything is coming apart at the seams."

I hope that's really clear now.

If not, well, don't worry, this is the last I'll write about it.
There is no defined graph or table or exact timeline on how long it takes to develop a new rocket engine from scratch. Aside from nuclear reactor design, nuclear submarine design, supersonic and hyper-sonic aircraft design, and sub-sea infrastructure design, this is basically the hardest thing to create and build from scratch in the human world right now. It takes as long as it takes depending on a wide variety of factors including but not limited to LV requirements first, cash available second, materials third, and breakage in testing fourth. With that said certain kinds of breakage are worse than others, burning up your combustion chamber liner during startup is definitely not good, but it's a test article. Also, it is often *semi easy* to fix problems like this in early test engines by changing things like mixture ratio or making other small tweaks. Not at all a show stopper.

You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's. This motor fired and did not explode or destroy itself on the first run, and its being developed at rapid pace, with many rapid changes, and probably not much cash. Doing pretty good so far, as far as anyone knows the sub-scale raptor never had an RUD during last year's campaign and this variant didn't on the first test either.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dnavas on 02/04/2019 06:51 am
\Engine probably still cooling down+people need to sleep. They can come back in the morning and stick a camera in there to find out.
I was actually surprised they ran this at all on Superbowl Sunday (in TX?!)

I figure they have a lot of instrumentation already on the important bits, unless there are parts that are difficult to monitor, in which case I suppose you live with post-analysis.  Is it actually likely that they would need to stick a camera down the guts during testing, or does it seem more likely they'd rely on instrumentation to tell them what's going on?  I assume post analysis after a testing suite would be normal, but I assume that this would not be that time?

You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's.

More familiar with the stories of the former than the latter.  So, yes, I'm *really* impressed with Merlin, Raptor, BE* engine development!

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/04/2019 06:54 am
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.

But you can try to estimate the size by this palm print here:

; P

Perfect! Is the hand what you used to set your scale in your Merlin/BE4 image?

Isn't that bit rich, coming from a *rebellious colony which still uses dead king's feet for measurements. Anyway here more of The True Size of Things (https://forum.nasaspaceflight.com/index.php?action=dlattach;topic=42585.0;attach=1542450;image):

* : P
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 02/04/2019 07:28 am
I dont want to speculate on any implications. I am ooking at the 8s version here and it does not look like saturation or any other camera issue.

1. The lamp in the background is brighter than the flame it self, if there was saturation, the lamp would show a green tint as well. Also, this is likely a CMOS detector, they dont bleed in the same way when saturated as CCD detectors. It also wouldnt effect just the green pixels because the color is determined by a micro-mask on top of the pixels, the charge information inside the pixels is independent of color. We would see blooming in all colors if any pixel saturates.
2. The green tint starts at the top of the flame and follows the mach diamonds. Then a few frames later it also starts at the bottom of the flame. Both sides have roughly equal brightness. This could be a localized white balance change from the camera, but I dont know any camera that does local white balance as opposed to global.
3. It could be an image ghost, but that doesnt sound right with the time staged effect that we see. An image ghost would correlate with brightness. And brightness was pretty constant before and after the green tint occurred.

I cant think of any other camera effect that would cause this.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/04/2019 07:32 am
An engine that can be transported in a cargo van that produces 7GW, enough power to run Switzerland, with all that power passing through a hole the size of a toilet seat.  Rocket engineering is best engineering. :)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JamesH65 on 02/04/2019 08:52 am
As someone on arocket said, it was a first test of a new rocket engine, and nothing blew up. Therefor its a success. I agree. There might be some issues to sort out, as would be entirely expected, no-on expects things to be perfect the first time, but they still have the motor and they still have the test stand. Success by anyones standards.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/04/2019 09:14 am
SpaceX must have parts for at least 2 more Raptors either assembled or nearing completion for the hopper, so if they do suffer any issues that don't require significant design modifications then they can probably swap in another engine or parts thereof for continued testing within a few days.  Here's hoping for a noisy month at McGregor.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: CapitalistOppressor on 02/04/2019 09:17 am
So during the burn video there are 2 green flame episodes, one near the beginning, one at shutdown. Since these are not actually at startup, and supposedly Raptor uses spark ignition anyway, the only explanation I can think of is a bit of copper chamber or bell vaporizing.

The first frame of the early incident at 5:44 you just see a little streak by the bell, the next frame it's partway down the jet, then it's at the end of the jet. For some reason this frame is doubled. Then the next frame, no more green. However the vapor patterns along the ground don't suggest that any video is missing, they change from frame to frame in a consistent way.

Also, is it possible to tell the exhaust velocity from the spacing of the mach diamonds? They are extremely consistent right until shutdown starts around 6:20. You can see from one frame to the next that they are sliding to the right at that point.

This is just an example of this exact discussion starting up after previous video releases earlier in the Raptor development process. People probably should relax.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/04/2019 09:19 am
Under certain mix conditions methane in combustion can produce a greenish tint or a green flash. Depends on temperature pressure % O2 in the burn and also other atmospheric gases interfering. The wide field video with sound shows much less green than the close in shot. Still good possibility exists it could also be due to internal erosion of some kind in the chamber. Really easy way to find out stick a camera or scope up there after it's cooled down and see if there is damage.

If no visible damage continue test campaign see if engine RUD's itself after multiple firings. Reminder the last SSME test saw a large hydrogen flame shooting out of part of the power-head or test stand and that was not really a massive show stopper, without the government shutdown test campaign would probably still be proceeding. Also engine did not undergo RUD despite failure. Chamber erosion is very different however, but again, if no visible damage is found with camera/scope then just keep firing and see what happens.

Easy way to find the flaws in a piece of machinery is push the test article until it breaks and then see where it broke and why. Like with the windblown starhopper fairing these things are test articles and are meant to be broken. Do not be surprised if they explode break fall over or otherwise fail.

What you don't want is to be on a flight to the moon and discover a flaw the hard way then, as Apollo 13 did. Or worse during launch or re-entry.
I am sure that the throat diameter is measured to a um after each firing, plus microscopy shots to examine the microstructure, plus any number of techniques to ascertain surface composition...  It's so easy these days, and those are such obvious questions...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tyrred on 02/04/2019 09:21 am
Wonder when gimbal testing firing will happen, or if we'll be privy to video.  Does anybody recall video of gimbal testing in previous dev raptor videos?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/04/2019 09:23 am
Quote from: @psalman03
What made you decide to not use TEA-TEB for ignition? Is the idea that during engie chill, CH4/O2 are released to ignite for liftoff? I don't think I've ever heard of spark plugs being used for rockets.

https://twitter.com/elonmusk/status/1092320321229643776

Quote
Spark plugs ignite dual blow torches that ignite preburners & main chamber
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 02/04/2019 09:31 am
With that said certain kinds of breakage are worse than others, burning up your combustion chamber liner during startup is definitely not good, but it's a test article. Also, it is often *semi easy* to fix problems like this in early test engines by changing things like mixture ratio or making other small tweaks. Not at all a show stopper.

You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's. This motor fired and did not explode or destroy itself on the first run, and its being developed at rapid pace, with many rapid changes, and probably not much cash. Doing pretty good so far, as far as anyone knows the sub-scale raptor never had an RUD during last year's campaign and this variant didn't on the first test either.

That's why I don't think this is a massive issue. This is the first firing of a new engine, not a final design they have built a hundred of already. Making slight changes to the mixture or bell isn't the worst thing that could have happened and the hopper might be fine to run as is or with an added ablative liner.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tyrred on 02/04/2019 09:37 am
With that said certain kinds of breakage are worse than others, burning up your combustion chamber liner during startup is definitely not good, but it's a test article. Also, it is often *semi easy* to fix problems like this in early test engines by changing things like mixture ratio or making other small tweaks. Not at all a show stopper.

You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's. This motor fired and did not explode or destroy itself on the first run, and its being developed at rapid pace, with many rapid changes, and probably not much cash. Doing pretty good so far, as far as anyone knows the sub-scale raptor never had an RUD during last year's campaign and this variant didn't on the first test either.

That's why I don't think this is a massive issue. This is the first firing of a new engine, not a final design they have built a hundred of already. Making slight changes to the mixture or bell isn't the worst thing that could have happened and the hopper might be fine to run as is or with an added ablative liner.

Nitpick: This is not a new engine.  It is an iteration of the dev raptor, which was fired multiple times in testing.  Of course it looks different, but compare Merlin 1A to 1D.  Agree with the rest.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/04/2019 09:40 am
This is also the first firing of this physical engine. Any difference between the nozzle surface of a pristine and a fired engine will show up in the flame. Any artifacts left over from manufacturing, etc.

We'll see what happens after a few long duration burns.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: StuffOfInterest on 02/04/2019 10:47 am
Quote from: @psalman03
What made you decide to not use TEA-TEB for ignition? Is the idea that during engie chill, CH4/O2 are released to ignite for liftoff? I don't think I've ever heard of spark plugs being used for rockets.

https://twitter.com/elonmusk/status/1092320321229643776

Quote
Spark plugs ignite dual blow torches that ignite preburners & main chamber

This seems good from the standpoint of not needing additional consumables (TEA-TEB) on board.  Makes sense if the Starship is taking off from another celestial body as you don't have to haul a tank of something around.  Simpler is better.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wargrim on 02/04/2019 11:37 am
The green is a reflection of the envy that other rocket engines felt right that very moment.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: 50_Caliber on 02/04/2019 11:46 am
The green is a reflection of the envy that other rocket engines felt right that very moment.
In my HIGHLY amateur opinion, I though the green exhaust looked cool.  8)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/04/2019 12:28 pm
The subscale Raptor also burnt some copper during starts and stops, and seems to run just fine:
https://www.youtube.com/watch?v=r_OM3WPijqs
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: glennfish on 02/04/2019 12:56 pm
Under certain mix conditions methane in combustion can produce a greenish tint or a green flash. Depends on temperature pressure % O2 in the burn and also other atmospheric gases interfering. The wide field video with sound shows much less green than the close in shot. Still good possibility exists it could also be due to internal erosion of some kind in the chamber. Really easy way to find out stick a camera or scope up there after it's cooled down and see if there is damage.

If no visible damage continue test campaign see if engine RUD's itself after multiple firings. Reminder the last SSME test saw a large hydrogen flame shooting out of part of the power-head or test stand and that was not really a massive show stopper, without the government shutdown test campaign would probably still be proceeding. Also engine did not undergo RUD despite failure. Chamber erosion is very different however, but again, if no visible damage is found with camera/scope then just keep firing and see what happens.

Easy way to find the flaws in a piece of machinery is push the test article until it breaks and then see where it broke and why. Like with the windblown starhopper fairing these things are test articles and are meant to be broken. Do not be surprised if they explode break fall over or otherwise fail.

What you don't want is to be on a flight to the moon and discover a flaw the hard way then, as Apollo 13 did. Or worse during launch or re-entry.

I don't think you can draw much conclusion from the colors in the video.  Maybe with a spectrograph.  Methane Oxygen combustion has a wide variety of spectra depending on the pressure and mix:  see the attached for spectra under different combustion conditions:  https://www.researchgate.net/publication/224989842_LOXMethane_Technology_Efforts_for_Future_Liquid_Rocket_Engines

for reference, green is 560 nm on most days.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: alhenry1231 on 02/04/2019 01:01 pm
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.

But you can try to estimate the size by this palm print here:

; P

Perfect! Is the hand what you used to set your scale in your Merlin/BE4 image?

Isn't that bit rich, coming from a *rebellious colony which still uses dead king's feet for measurements. Anyway here more of The True Size of Things (https://forum.nasaspaceflight.com/index.php?action=dlattach;topic=42585.0;attach=1542450;image):

* : P

Oh, very rich :)
No sarcasm from me, I was thrilled at the ingenuity.

I'm happy to see our scale references so far holding up.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/04/2019 01:38 pm
Another possible source of 'green' may be the ignition torch sparkplug electrodes.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 02/04/2019 01:57 pm
If "engine rich combustion" is a predictable startup/shutdown transient, what does that mean for the PREDICATABILITY of service life? Obvously it's not going to last as long in absolute terms, but if you know that the chamber ablates x micrometers per activation, does that mean you could simply have a "engine must be serviced every XX firings for chamber ablation" and skip actual per-firing checks?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: abaddon on 02/04/2019 02:06 pm
This is just an example of this exact discussion starting up after previous video releases earlier in the Raptor development process. People probably should relax.
And, that engine ran for over a thousand seconds of burn time during the test campaign, if I recall correctly.  So A) we've seen this before and B) the engine ran for a really long time.  And it is literally the first firing of the first production engine.  What we saw was a great success.  It doesn't mean it's ready to fly tomorrow.

The hand-wringing around here has become really tiresome.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wolfram66 on 02/04/2019 02:10 pm
So people, this was equivalent to SpaceX doing a 1-3 sec Static Fire of a Merlin 1-D. so Chill-ax...
It's their process to check to make sure everything is set up right in the plumbing, controllers and telemetry channels. The did just ship this unit across the country... just working the use case / test case check-list like always. now. back to monday  :-\
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: BunkerTheHusky on 02/04/2019 03:58 pm
Will anybody with the video editing know how make a shake-stabilized version of these videos? I'm curious to see how violent that shutdown appears without the camera going ham.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Vemaster on 02/04/2019 05:30 pm
SpaceX: Completed a two-second test fire of the Starship Raptor engine that hit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power.
https://www.instagram.com/p/BteDxoUFUOX/
So Raptor test was 2 seconds, 170 bar, 116 metric tons force, ~60% power...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: aaminotti on 02/04/2019 05:31 pm
Somebody in reddit upload this synched video: https://www.reddit.com/r/SpaceXLounge/comments/an406g/tried_my_best_to_sync_both_raptor_test_videos_and/?utm_source=reddit-android

Enviado desde mi SM-N950F mediante Tapatalk

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/04/2019 06:07 pm
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/04/2019 06:11 pm
Quote
hit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power

What does "power" mean in this context?

I'll go out on a bit of a hopeful limb here until someone that knows corrects me back to reality.  If we're at 60% of max thrust and 60% of Pch then wow, it would be;

   Pressure @ 100% = 283 bar, 4104 psi

   Thrust @ 100% = 193.2 mt, 426,000 lbf

Don't run with these numbers, they are more of a question than a statement and are likely to be stricken from the record when someone with real knowledge sets me straight.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Davidthefat on 02/04/2019 06:17 pm
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.

John

John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/04/2019 06:20 pm
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.

John
To assist folks finding John's schematic, here's a link to the post:

https://forum.nasaspaceflight.com/index.php?topic=41363.msg1907500#msg1907500

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/04/2019 06:21 pm
...
No sarcasm from me, I was thrilled at the ingenuity.

I'm happy to see our scale references so far holding up.
Here's  the scale references:
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/04/2019 06:36 pm
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.

John

John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?

? The Raptor has never been shown with boost pumps.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/04/2019 06:38 pm
Quote
hit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power

What does "power" mean in this context?

I'll go out on a bit of a hopeful limb here until someone that knows corrects me back to reality.  If we're at 60% of max thrust and 60% of Pch then wow, it would be;

   Pressure @ 100% = 283 bar, 4104 psi

   Thrust @ 100% = 193.2 mt, 426,000 lbf

Don't run with these numbers, they are more of a question than a statement and are likely to be stricken from the record when someone with real knowledge sets me straight.

And vacuum thrust would be around 208 tonnes if the ISPs are the same (330sl/356vac)... I think these numbers are more right than wrong.

...This could also mean that the "Raptor Full Thrust" may be well over 300 bar in chamber pressure!

If chamber pressure increases with the thrust increase Elon was talking about (200t to 250t), We could be talking about ~350 bar chamber pressure if the throat doesn't change.  :o
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Davidthefat on 02/04/2019 06:38 pm
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.

John

John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?

? The Raptor has never been shown with boost pumps.

John

So you are saying it's a two stage pump. The first stage would be akin to a "boost pump" in my books.

Edit: What was the thought process behind having the preburner be downstream the regen? What if the regen dumped back into the low pressure side. Then you don't have your concern about the pressure margin in the preburners.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/04/2019 06:48 pm
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.

Also... 350/280 (bar) = 250/200 (tonnes force)

 ;)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: punder on 02/04/2019 06:50 pm
An engine that can be transported in a cargo van that produces 7GW, enough power to run Switzerland, with all that power passing through a hole the size of a toilet seat.  Rocket engineering is best engineering. :)

Mods delete if considered frivolous, but... that was one heck of a mental image.  :)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/04/2019 06:57 pm
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.

John

John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?

? The Raptor has never been shown with boost pumps.

John

So you are saying it's a two stage pump. The first stage would be akin to a "boost pump" in my books.

To the best of our knowledge from statements and CAD pictures:

- Methane has a two stage pump with a single stage turbine.

- Lox has a single stage pump with a single stage turbine.

- Number of turbine stages may not be accurate.

- Pumps may have inducers on their front ends to help with cavitation.

- If a rocket has boost pumps they are usually powered by tapping off matching high pressure fluid to drive a boost pump turbine. Look up a schematic of the SSME to see the idea. Boost pumps are low pressure ratio pumps used to guard against cavitation.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/04/2019 07:13 pm
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.

Also... 350/280 (bar) = 250/200 (tonnes force)

 ;)

I am more inclined to think the current rating of 200-ish tonnes is 100% rated at 250 bar. That would make the 250-ish tonne version 100% rated around 310 bar. Obviously a small increase in throat size would be needed which would reduce the expansion ratio a little. I am getting ready to remodel the Raptor with the latest data.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/04/2019 07:43 pm
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.

Also... 350/280 (bar) = 250/200 (tonnes force)

 ;)

I am more inclined to think the current rating of 200-ish tonnes is 100% rated at 250 bar. That would make the 250-ish tonne version 100% rated around 310 bar. Obviously a small increase in throat size would be needed which would reduce the expansion ratio a little. I am getting ready to remodel the Raptor with the latest data.

John

I am honestly too, just think it's an interesting coincidence of math at this point...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: matthewkantar on 02/04/2019 08:12 pm
Does anyone know what very loud bark is at the end of the run?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Gotorah on 02/04/2019 08:16 pm
Fuel rich exhaust is yellow, sufficient or oxygen rich exhaust is blue. Mix that amount of them and you get a green flash at the switch from one condition to the other.  In my humble opinion !
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: testguy on 02/04/2019 08:30 pm
Does anyone know what very loud bark is at the end of the run?

Not unusual. It may happen with a rapid decrease in pressure be it the engine or feed lines.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RoboGoofers on 02/04/2019 08:36 pm
I've never noticed the, for lack of a better term, flashlight or headlight effect that the engine is casting on the ground. It's as if the engine were a Klieg light.

Does anyone know what very loud bark is at the end of the run?
Not unusual. It may happen with a rapid decrease in pressure be it the engine or feed lines.
Not that it's relevant but it sounds like the barking dog chemistry experiment.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/04/2019 08:37 pm
Quote
hit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power

What does "power" mean in this context?

I'll go out on a bit of a hopeful limb here until someone that knows corrects me back to reality.  If we're at 60% of max thrust and 60% of Pch then wow, it would be;

   Pressure @ 100% = 283 bar, 4104 psi

   Thrust @ 100% = 193.2 mt, 426,000 lbf

Don't run with these numbers, they are more of a question than a statement and are likely to be stricken from the record when someone with real knowledge sets me straight.

At sea level, thrust is not linear with chamber pressure because of atmospheric backpressure. I would expect that 200 metric tonnes at 250 bar is the goal.

As chamber pressure goes up at SL, both exhaust velocity and mass flow go up. Since thrust is mass flow * exhaust velocity, thrust increases faster than chamber pressure.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/04/2019 08:48 pm
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.

Also... 350/280 (bar) = 250/200 (tonnes force)

 ;)

I am more inclined to think the current rating of 200-ish tonnes is 100% rated at 250 bar. That would make the 250-ish tonne version 100% rated around 310 bar. Obviously a small increase in throat size would be needed which would reduce the expansion ratio a little. I am getting ready to remodel the Raptor with the latest data.

John

What expansion ratio are you assuming? I can't get a 50:1 to fit with the numbers SpaceX released and a 1.3 meter nozzle. 35:1 with a 1.41 m nozzle fits much better.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/04/2019 09:05 pm
I watched BE-4's and even XCOR's XR's starts and shut downs, and did not see any greens there.

Fuel rich exhaust is yellow, sufficient or oxygen rich exhaust is blue. Mix that amount of them and you get a green flash at the switch from one condition to the other.  In my humble opinion !

And you just can't mix them; colors depend on the chemical mix of the flame. The electrons "jump" from their ground state to a higher energy level in high temperature. As they return to their ground state they emit visible light. There needs to something that emits green...

Here are measurements on spectra of the methane-air flame. There are some spikes on green (525 nm) with high fuel rations, but it's out classed by cyan and blue.

Images from: http://lup.lub.lu.se/luur/download?func=downloadFile&recordOId=4695221&fileOId=4695231 (http://lup.lub.lu.se/luur/download?func=downloadFile&recordOId=4695221&fileOId=4695231)

May 2014 Spectral Analysis of Flame Emission for Optimization of Combustion Devices on Marine Vessels 

Thesis of Panagiota Stamatoglou
Supervisor: Mattias Richter Division of Combustion Physics Department of Physics Lund University
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 02/04/2019 09:49 pm
And you just can't mix them; colors depend on the chemical mix of the flame. The electrons "jump" from their ground state to a higher energy level in high temperature. As they return to their ground state they emit visible light. There needs to something that emits green...

What about different portions of the plume having different chemical mixes? is it reasonable to have a "skin" of yellow fuel rich combustion surrounding a core of blue combustion, at least during startup and shut down?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: marsbase on 02/04/2019 10:25 pm
And you just can't mix them; colors depend on the chemical mix of the flame. The electrons "jump" from their ground state to a higher energy level in high temperature. As they return to their ground state they emit visible light. There needs to something that emits green...

What about different portions of the plume having different chemical mixes? is it reasonable to have a "skin" of yellow fuel rich combustion surrounding a core of blue combustion, at least during startup and shut down?
Because we are talking about light emission, yellow plus blue makes white.  In the pigment model, yellow plus blue makes green.  Since green is a primary light color, it is correct that to get green something must be emitting in the green band.  http://learn.leighcotnoir.com/artspeak/elements-color/primary-colors/
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dubya on 02/04/2019 11:22 pm
Because we are talking about light emission, yellow plus blue makes white.  In the pigment model, yellow plus blue makes green.  Since green is a primary light color, it is correct that to get green something must be emitting in the green band.  http://learn.leighcotnoir.com/artspeak/elements-color/primary-colors/


Don't tell the RGB screen you are looking at that emitted colors don't mix. You will lose all those wonderful yellow, oranges, purples and whites. /snark

Edit/Lar: fixed quotes
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jketch on 02/04/2019 11:53 pm
Because we are talking about light emission, yellow plus blue makes white.  In the pigment model, yellow plus blue makes green.  Since green is a primary light color, it is correct that to get green something must be emitting in the green band.  http://learn.leighcotnoir.com/artspeak/elements-color/primary-colors/


Don't tell the RGB screen you are looking at that emitted colors don't mix. You will lose all those wonderful yellow, oranges, purples and whites. /snark

The point is that when an RGB wants to emit green, it just emits green light, as green is a primary color. There's no such thing as a RYB TV which makes green by mixing yellow and blue.

Edit/Lar: Fixed quotes
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/05/2019 12:14 am
(mod) I think that we've all red enough on basic color, yeah?. More off topic would turn me blue. Don't throw shade on my bad puns.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/05/2019 02:37 am
Just started looking at my Raptor model. Things match really well if I assume the quoted Thrusts are vacuum thrusts.

- Starting assumption is that the 2019 Raptor has an exit diameter of 1.3 meters. If anyone knows different, please let me know.

- I kept the 2017 Raptor (er=35) dimensions and increased the chamber pressure from 250 to 270 bar. This takes the engine from  1.7 sl MN / 1.83 vac MN to 1.83 / 1.96 MN (200 tonnes vacuum).

- I then did similar with the 2017 vacuum Raptor 119 which was assumed to have the same throat diameter as the Raptor (er=35) , but an expansion ratio of 119.  I increased the chamber pressure from 250 to 320 bar. The vacuum engine went from 1.93 MN to 245 MN (250 tonnes vacuum).

- I haven't done the Raptor (er=35) with 320 bar yet. That's next.

- All ISPs are within a couple of seconds of previously stated results. I will post a graph of the engines performance when I am done.

This is only one approach but it does seem to work out. I attached the 2017 engine model summary.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: hkultala on 02/05/2019 03:23 am
Has anybody made an analysis where he/she takes the published  raptor picture and annotates with speculation, which part of the picture is which part of the engine? Which pipes are which etc. ?

The oxygen pump should just at the end of the chamber. The methane pump.. is it inline forward from it?

(https://images.livemint.com/img/2019/02/04/600x338/Musk2_1549262740050.jpg)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: su27k on 02/05/2019 04:32 am
Poll: When will full-scale hot-fire testing of Raptor begin?

You can (almost) stop the poll now.

Now we can conclude the poll, 31.1% got it correctly (Integrated tests -  2019), 51.5% voted 2018 which is pretty close, looks like wisdom of crowds worked well this time.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 02/05/2019 06:53 am
What drives me crazy about these images of Raptor... where is the Methane preburner and turbopump? Did they take the images deliberately so it is on he other side of the engine? Or is it in line, like the LOX preburner now?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/05/2019 01:07 pm
Does anyone know what very loud bark is at the end of the run?

If you are referring to the video in Musk's tweet, the audio and video are not synced because it was filmed from some distance away. The loud noise just before shutdown is actually the ignition.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/05/2019 03:10 pm
Hard telling what pipes are for, but here is some speculation. We really need to see the other side of the engine in a little more detail.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Wargrim on 02/05/2019 04:00 pm
This tweet chain made me look like  :o

https://twitter.com/bluemoondance74/status/1092615576210550786

But there is no way that was a Raptor burn already, or is it? Must be a Merlin test?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: matthewkantar on 02/05/2019 04:02 pm
Does anyone know what very loud bark is at the end of the run?

If you are referring to the video in Musk's tweet, the audio and video are not synced because it was filmed from some distance away. The loud noise just before shutdown is actually the ignition.

The noise between 6 and 7 seconds in this video? https://streamable.com/nl6y7

On edit: maybe "honk" is a better description of the noise than "bark."
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ugordan on 02/05/2019 04:11 pm
This tweet chain made me look like  :o

https://twitter.com/bluemoondance74/status/1092615576210550786

But there is no way that was a Raptor burn already, or is it? Must be a Merlin test?

No way it's the Raptor. They're not going from 2 seconds to 6 minutes just like that, especially given the green flame from the first firing (yeah, I'm in the "it's not a camera artifact" camp).

6 minutes is a dead giveaway for either an MVac or integrated 2nd stage test.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Hitech on 02/05/2019 08:04 pm
Does anyone know what very loud bark is at the end of the run?
Anyone that has worked with pulse jets especially large ones will recognize that noise.
The result of "ambient pressure" ignition of residual fuel in a "pipe". On this engine, as it was shutting down "rich" at some point the "fresh" air enters the nozzle and combustion chamber till the mixture leans to the point where rapid combustion occurs and induces an resonant oscillation in the "pipe" (not unlike the "bark" of the hydrogen in the test tube)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 02/05/2019 08:18 pm
How would that schematic change if you put the fuel preburner and turbopump ontop of the LOX one? I ask becaue in none of the new images, the Methane preburner is visible. So if the Methane preburner were to sit on top of the LOX preburner, how do you get the hot methane to the combustion chamber? I dont see fat pipes snaking from one end to the other either. The only large pipes that are visible seem to be for cooling the nozzle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/05/2019 08:31 pm
This tweet chain made me look like  :o

https://twitter.com/bluemoondance74/status/1092615576210550786

But there is no way that was a Raptor burn already, or is it? Must be a Merlin test?

Barring substantial upgrades to the Raptor test cells, they weren't even able to support test fires longer than 100 seconds as of late 2017 (and probably 2018, too). That will undoubtedly change if it hasn't already, given the need for mission-duration tests. The new stand under construction at McGregor would be a prime candidate, unless a 200t Raptor can be realistically static-fired for 2+ minutes without a regen-cooled trough or deflector.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JamesH65 on 02/06/2019 12:52 pm
Does anyone know what very loud bark is at the end of the run?
Anyone that has worked with pulse jets especially large ones will recognize that noise.
The result of "ambient pressure" ignition of residual fuel in a "pipe". On this engine, as it was shutting down "rich" at some point the "fresh" air enters the nozzle and combustion chamber till the mixture leans to the point where rapid combustion occurs and induces an resonant oscillation in the "pipe" (not unlike the "bark" of the hydrogen in the test tube)

So...a backfire?  ;-)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/06/2019 04:51 pm
Maybe we need to hive off a subthread to talk about image editing and improvements and things to keep this focused on the engine itself?  Let me have a think.

(I blame me for this in part although clearly it's not actually me because I'm never wrong)

That said I love these diagrams because they help laymen understand how this beast works. I know they're conjecture but it's very informed conjecture and I really appreciate  the effort that goes into them.

Edit: And so I did.
https://forum.nasaspaceflight.com/index.php?topic=47383
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: WormPicker959 on 02/06/2019 05:34 pm
In regards to the green emissions seen in raptor test vids, is it possible that the first burn of the rocket does indeed "burn off" some initial residue, or a first layer of copper, or something to that effect, as part of a completely normal process? If it's a residue left from manufacture, it'd be burnt off and not present later. If it's a "first layer" of copper, this could be bad, but could it also be normal - as in, after heating the alloy in this specific way, it changes the structure or something somewhat, making it more resistant to this kind of torture in the future? It seemed a pretty bright green and highly localized to be merely one of the emission lines of methalox combustion, but I could obviously be wrong about that.

I know nothing about metallurgy, so please be kind to my dumb speculations ;)

In any case I'm not trying to amplify any sense or worry or anxiety, as I'm fairly confident these guys know what they're doing. I'm just curious! I want to know more about how rockets and stuff work!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Prettz on 02/06/2019 06:07 pm
livingjw's schematic as vector graphic:
This is really great, and it's reminded me of a question I've had about the oxygen preburner since the first reveal of the Raptor: what goes on in a torus-shaped combustion chamber? What's the injector have to be like?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/06/2019 06:12 pm
In regards to the green emissions seen in raptor test vids, is it possible that the first burn of the rocket does indeed "burn off" some initial residue, or a first layer of copper, or something to that effect, as part of a completely normal process? If it's a residue left from manufacture, it'd be burnt off and not present later. If it's a "first layer" of copper, this could be bad, but could it also be normal - as in, after heating the alloy in this specific way, it changes the structure or something somewhat, making it more resistant to this kind of torture in the future? It seemed a pretty bright green and highly localized to be merely one of the emission lines of methalox combustion, but I could obviously be wrong about that.

I know nothing about metallurgy, so please be kind to my dumb speculations ;)

In any case I'm not trying to amplify any sense or worry or anxiety, as I'm fairly confident these guys know what they're doing. I'm just curious! I want to know more about how rockets and stuff work!

If it is a bit of the combustion chamber wall vaporizing, it might be a little too thick which reduces heat flux into the coolant and increases the hot side temperature. As it vaporizes off, the thickness drops and the heat flux increases, which reduces the hot side temps and stops the vaporizing. It's a negative feedback cycle that stabilizes, just as long as it doesn't get so thin that it goes kablooey.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ellindsey on 02/06/2019 06:55 pm
livingjw's schematic as vector graphic:
This is really great, and it's reminded me of a question I've had about the oxygen preburner since the first reveal of the Raptor: what goes on in a torus-shaped combustion chamber? What's the injector have to be like?
The design of the oxygen turbopump on the Raptor actually reminds me of the early centrifugal compressor type jet engines, with the compressor, combustion chamber, and turbine all in a compact line like that.  Those had individual compressor cans, but some modern jet engines have toroidal combustion chambers instead, and I suspect that those have some design similarity to the Raptor's oxygen preburner.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/06/2019 08:10 pm
There are a lot of advantages to making engines smaller (higher thrust to weight and lower manufacturing cost per unit of thrust) but there is one disadvantage - using very fuel-rich mixture [edit, next to wall] or transpiration cooling of wall costs you more performance owing to lower area/circumference ratios.

Transpiration cooling of the combustion chamber and nozzle throat region would be one possible means for greatly reducing wall heat fluxes to increase the life of the nozzle and create a higher cycle life engine.  Wonder if they have done something like that (or plan to) - laser drilling is a very mature process now from gas turbine manufacture.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/06/2019 08:30 pm
livingjw's schematic as vector graphic:
This is really great, and it's reminded me of a question I've had about the oxygen preburner since the first reveal of the Raptor: what goes on in a torus-shaped combustion chamber? What's the injector have to be like?
The design of the oxygen turbopump on the Raptor actually reminds me of the early centrifugal compressor type jet engines, with the compressor, combustion chamber, and turbine all in a compact line like that.  Those had individual compressor cans, but some modern jet engines have toroidal combustion chambers instead, and I suspect that those have some design similarity to the Raptor's oxygen preburner.

- Any burner requires a region of near stoichiometric mixing to get good combustion.

- In turbine engines this is done in a recirculation region near the front of the burner. Additional air is mixed downstream of the stoichiometric region. This dilutent air reduces the temperature to a temperature that the turbine blades can handle.

- The Raptor Lox pre-burner must work in an analogous manner. I suspect there is a series of localized stoichiometric recirculating regions surrounding the front of the toroidal combustion chamber. Additional Lox is mixed downstream of this region to vaporize it and reduce the temperature before entering the turbine.

- There is a torch igniter for each pre-burner, as well as one for the main chamber.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/07/2019 02:27 am
I'll go lazy here and use someone else's words to save some typing...

In regards to the green emissions seen in raptor test vids,
I'm not understanding how copper is exposed unless there is a significant erosion of another material.  Isn't it true that the copper is clad on both sides with another material?  Perhaps nickel and cobalt if I understand right?  Or maybe I got that material wrong but looking into the nozzle its certainly not copper colored inside and [?]that material extends through the throat and into the combustion chamber[/?]
In any case I'm not trying to amplify any sense or worry or anxiety, as I'm fairly confident these guys know what they're doing. I'm just curious! I want to know more about how rockets and stuff work!
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/07/2019 04:17 am
People are stuck on the whole “green equals copper” assertion. Don’t make that assumption - or if you do, know that it is an assumption and not fact.

(I would only consider people with experience in *methane* engines as any sort of authority on the issue. And even then the alloys used by SpaceX are a well kept secret.)

Then there is the 2nd assumption, that “green equals BAD”. Also not proven.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/07/2019 07:17 am
https://twitter.com/elonmusk/status/1093423297130156033

Quote
Raptor just achieved power level needed for Starship & Super Heavy
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 02/07/2019 07:22 am
https://twitter.com/elonmusk/status/1093423297130156033

Quote
Raptor just achieved power level needed for Starship & Super Heavy

And not a hint of green on it :) wow.. that was soon! Look at the flame surrounding the central column of exhaust with mach diamonds. What does this, film cooling of the nozzle with methane?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/07/2019 07:23 am
https://twitter.com/elonmusk/status/1093424663269523456

Quote
Design requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/07/2019 07:43 am
Previous green was the engine:

Quote from: @AndreTI
Did you folks track down the source of the green hue?

https://twitter.com/elonmusk/status/1093428938871779328

Quote
Vaporized some copper
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OhYeah on 02/07/2019 07:45 am
250 bar pressure already? I vaguely remember that 300 was the target for future iterations of the engine and I got the impression that even reaching 250 will be a challenge. Things are moving pretty quickly.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Zed_Noir on 02/07/2019 07:46 am
Elon's tweet about 10% to 20% more with cryo propellants. Does that meant thrust, chamber pressure or both?

If it is chamber pressure than the Raptor can do about 300 bars with cryo propellant right now.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/07/2019 07:47 am
250 bar pressure already? I vaguely remember that 300 was the target for future iterations of the engine and I got the impression that even reaching 250 will be a challenge. Things are moving pretty quickly.
Your member name says it all...

(curious as to the duration of this test)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/07/2019 07:53 am
https://twitter.com/elonmusk/status/1093424663269523456

Quote
Design requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.

Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?

Also, if chamber pressure grows linearly with that 10-20% thrust increase, this particular Raptor would experience chamber pressures of at least ~280 bar (and perhaps more than 300 bar) at full thrust. In mid-December, he suggested that it would "take [SpaceX] time to work up to 300 bar...that is a mad level".

https://twitter.com/elonmusk/status/1076618077301665793
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: kevinof on 02/07/2019 08:00 am
That was my first thought also. Could still be just a short duration of a couple of seconds.

..
(curious as to the duration of this test)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/07/2019 08:01 am
https://twitter.com/elonmusk/status/1093424663269523456

Quote
Design requires at least 170 metric tons of force. Engine reached 172 mT &amp; 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.

Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?

Also, if chamber pressure grows linearly with that 10-20% thrust increase, this particular Raptor would experience chamber pressures of at least ~280 bar (and perhaps more than 300 bar) at full thrust. In mid-December, he suggested that it would "take [SpaceX] time to work up to 300 bar...that is a mad level".

https://twitter.com/elonmusk/status/1076618077301665793
Isn’t therefore the energy density higher? Which in turn leads to higher chamber pressure?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Semmel on 02/07/2019 08:01 am
https://twitter.com/elonmusk/status/1093424663269523456

Quote
Design requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.

Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?

Also, if chamber pressure grows linearly with that 10-20% thrust increase, this particular Raptor would experience chamber pressures of at least ~280 bar (and perhaps more than 300 bar) at full thrust. In mid-December, he suggested that it would "take [SpaceX] time to work up to 300 bar...that is a mad level".

https://twitter.com/elonmusk/status/1076618077301665793

Someone might correct me, but here is how I understand it.

The turbopumps are limited in their rotation speed by cavitation. They cant run any faster than that maximum speed. Densified propellant will pump more mass for the same rotation speed of the pump, which means one can bring more propellant into the combustion chamber before cavitating. I dont know if the cavitation speed is higher or lower for densified propellant. That depends on the speed of sound in the medium. Expert input required. But I bet that the net effect is an increased in propellant flow by mass for colder temperatures.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/07/2019 08:19 am
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!

The gist, assuming thrust is roughly proportional to chamber pressure:

~60% power   = 116t thrust, 170 bar
~90% power   = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 02/07/2019 08:23 am
https://twitter.com/elonmusk/status/1093424663269523456

Quote
Design requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.

Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?

I assumed it's related to mass flow. The same machinery can transport 10-20% more propellant if it's denser. Your point about chamber pressure might work the opposite way - they would theoretically reach 280 bar from increased pump capability alone but might have to dial it back because the chamber can't take it.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JamesH65 on 02/07/2019 08:31 am
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.

Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/07/2019 08:38 am
https://twitter.com/elonmusk/status/1093423297130156033

Quote
Raptor just achieved power level needed for Starship & Super Heavy

And not a hint of green on it :) wow.. that was soon! Look at the flame surrounding the central column of exhaust with mach diamonds. What does this, film cooling of the nozzle with methane?
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: guckyfan on 02/07/2019 08:42 am
The copper may have just been traces from production and burned off in the first firing. Elon seemed not to worry about it at all.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/07/2019 08:53 am
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.

You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

[Edit]"Engines like SSME, F-1, J-2, RS-27, Vulcain 2, RD-171 and RD-180 use film cooling technique for combustion chamber cooling" https://www.sciencedirect.com/science/article/pii/S2212540X1830004X
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Pete on 02/07/2019 08:58 am
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.

Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.

Remember that you are dealing with the company that, when a second stage engine had a cracked nozzle, fixed it by simply snipping off the cracked bit *by hand*, and launching the next day.

The same problem at Nasa, or even Roscosmos, would have required a vehicle disassembly and complete engine replacement.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Cheapchips on 02/07/2019 09:00 am
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.

Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.

Tom Mueller talked about testing Merlin's gas generator with the XCOR guys. 

After their first successful test Mueller said, "Lets do it again."
XCOR guys, "We didn't plan on doing more than one a day."
Mueller, "Its 11 in the morning. What are we going to do?"

At least some of the slow pace of testing is due habit.  Obviously a full up engine isn't the same as testing a gas generator, but it does show that they won't hang around. 

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/07/2019 09:30 am
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.

You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/07/2019 09:51 am
You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.

A little googling shows that film cooling is ubiquitous for high chamber pressure engines on top of the regenerative cooling.  Doesn't cost that much isp when using a low molecular weight high specific heat fuels like (in particular) hydrogen or methane as the film coolants.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Johnnyhinbos on 02/07/2019 10:33 am
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.

You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
Technically speaking, the whole Starship will be film cooled
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: philw1776 on 02/07/2019 01:11 pm
https://twitter.com/elonmusk/status/1093424663269523456

Quote
Design requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.

Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?

I assumed it's related to mass flow. The same machinery can transport 10-20% more propellant if it's denser. Your point about chamber pressure might work the opposite way - they would theoretically reach 280 bar from increased pump capability alone but might have to dial it back because the chamber can't take it.

The formula for thrust is roughly

Force (Kg-m)/sec2 = (mass flow Kg/sec) X (exhaust velocity m/sec)

So higher mass flow from densified propellants yields more thrust.  The cooler propellants may reduce exhaust velocity (related to ISP) a tiny bit but I'm not a rocket propulsion engineer. 

Color me surprised (Again!) at how quick a pace SpaceX's engine testing is running at!

Raptor on schedule on its quest to achieve Yarak.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/07/2019 01:18 pm
An approx duration for the latest test, from an 'ear'-witness:

https://twitter.com/abbygarrettX/status/1093509953267122177

Quote
I heard the test from 30 miles away!

https://twitter.com/abbygarrettX/status/1093511737331064832

Quote
Maybe a little longer. My guess is under 10 seconds.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dnavas on 02/07/2019 01:28 pm
You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.

A little googling shows that film cooling is ubiquitous for high chamber pressure engines on top of the regenerative cooling.  Doesn't cost that much isp when using a low molecular weight high specific heat fuels like (in particular) hydrogen or methane as the film coolants.

Interesting.  Tom definitely used this technique on the Merlin engines and talked about it here https://www.youtube.com/watch?v=0xWRhKB3JTM?t=882 (14:40ish) so I was wondering if they might use a similar strategy for Raptor.  Same design engineer, but given the difference in engines and fuels it wasn't clear whether that strategy was even available.  If you could link some citations for the newbs among you (raises hand) I sure would appreciate it!

I found https://www.sciencedirect.com/science/article/pii/S2212540X1830004X which talks about ISP reduction, but without figures for methane.  It also discusses coking problems using RP-1 as the coolant, which is applicable to Merlin.  That article links to https://arc.aiaa.org/doi/pdf/10.2514/6.2010-6721, but I sadly don't have institutional access.  If someone does and indicates it's worth it, I'll be happy to create an account (:sigh:) and purchase it.  I can't even find the price without an account apparently though.

[the really interesting pointer appears to be to https://iris.unicampania.it/handle/11591/172095#.XFxBW89KiGR, but it wasn't immediately clear to me how to even get access to that one.]

BTW, for anyone interested in the story Cheapchips mentioned above, Tom's XCor story ("let's go again!") is related at 11:22ish in this same video.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/07/2019 02:33 pm
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!

The gist, assuming thrust is roughly proportional to chamber pressure:

~60% power   = 116t thrust, 170 bar
~90% power   = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar

I suggested earlier ~280bar for initial Raptor and ~350bar for full thrust version.

Honestly, given everything we know so far the math works out the best with those numbers, and they so far seem to fit better and better as we learn more... But they seem crazy. :o
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 02/07/2019 02:35 pm
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!

The gist, assuming thrust is roughly proportional to chamber pressure:

~60% power   = 116t thrust, 170 bar
~90% power   = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar

I suggested earlier ~280bar for initial Raptor and ~350bar for full thrust version.

Honestly, given everything we know so far the math works out the best with those numbers, and they so far seem to fit better and better as we learn more... But they seem crazy. :o

Is there any chance that when elon said 250 ton he was referring to the vac version?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/07/2019 02:42 pm
Here's the time of the Raptor test and confirmation of short duration:

https://twitter.com/bluemoondance74/status/1093376766754914304

Quote
.@SpaceX employees appear to be working over-time tonight @ the McGregor, TX facility, as the burst and brief, low rumble of an engine test was heard @ 10:51 pm CST. Note: There are Noise Ordinances in neighboring towns. I checked, and 11pm is the cut-off. Just in-time! 🔥👍

Quote from: @bluemoondance74
(Amendment: The most recent Ordinance I found for McGregor was a few yrs old: 9 pm limit, except for prior-approved cases, w/ 115-decibel max. Surrounding towns have later cut-off times, much lower decibel limits. *McGregor is the only town which can enforce an Ord. on SpaceX.)🧐

https://twitter.com/bluemoondance74/status/1093463566093611008

Quote
Only a few seconds
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ugordan on 02/07/2019 02:49 pm
So, what's up with the weird-looking exhaust plume? It's like there's a detached sheath of methane flowing out of the nozzle, igniting and creating a secondary, much slower plume. I wonder whether this image was taken during steady operation or ignition/shutdown transient.

I can't notice a similar detached plume developing at any point in any the previously released video.

Crazy thought: Is there any reason whatsoever you'd ever want to drill some small holes at the nozzle bottom on a test engine and release some of the regen cooling methane out into the void? I can't think of one, but this almost looks like that to me.  ???
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/07/2019 02:56 pm
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!

The gist, assuming thrust is roughly proportional to chamber pressure:

~60% power   = 116t thrust, 170 bar
~90% power   = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar

My model to date (work in progress):
- Keeping the 2017 Raptor (er=35, Aexit=1.3m) dimensions and increased the chamber pressure from 250 to 270 bar. This takes the engine from  1.7 sl MN / 1.83 vac MN to 1.83 / 1.96 MN (200 tonnes vacuum).

 pressure, thrust
 ----------------------
- 170 bar, 113 mt SL
- 257 bar, 177 mt SL
- 270 bar, 183 mt SL  200 mt vacuum

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cppetrie on 02/07/2019 03:02 pm
Curious what the informed opinion is on next test steps. They have apparently reached their required thrust targets. Do they continue increasing thrust to see what they can get to or now begin increasing test duration at target thrust levels to simulate full flight profile?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: MikeAtkinson on 02/07/2019 03:05 pm
Curious what the informed opinion is on next test steps. They have apparently reached their required thrust targets. Do they continue increasing thrust to see what they can get to or now begin increasing test duration at target thrust levels to simulate full flight profile?

Throttling response next I think.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: [email protected] on 02/07/2019 06:14 pm
My model to date (work in progress):
- Keeping the 2017 Raptor (er=35, Aexit=1.3m) dimensions and increased the chamber pressure from 250 to 270 bar. This takes the engine from  1.7 sl MN / 1.83 vac MN to 1.83 / 1.96 MN (200 tonnes vacuum).

 pressure, thrust
 ----------------------
- 170 bar, 113 mt SL
- 257 bar, 177 mt SL
- 270 bar, 183 mt SL  200 mt vacuum

John

Does anyone think they will achieve 300+ bars?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/07/2019 06:28 pm
>
Does anyone think they will achieve 300+ bars?

I've lost count of the "SpaceX will/could never..." proclamations which ended up as roadkill.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: intelati on 02/07/2019 06:37 pm
>
Does anyone think they will achieve 300+ bars?

I've lost count of the "SpaceX will/could never..." proclamations which ended up as roadkill.

At this point I'm just watching their shock and awe...

I'm reminded of Intel's tick-tock ding-dong model (https://en.wikipedia.org/wiki/Tick%E2%80%93tock_model) from the mid 2000's. SpaceX is so far forward of the rest of the launchers (ULA/BO/ESA...) in their research and development it's not even funny
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/07/2019 06:47 pm
So, what's up with the weird-looking exhaust plume? It's like there's a detached sheath of methane flowing out of the nozzle, igniting and creating a secondary, much slower plume. I wonder whether this image was taken during steady operation or ignition/shutdown transient.

I can't notice a similar detached plume developing at any point in any the previously released video.

Crazy thought: Is there any reason whatsoever you'd ever want to drill some small holes at the nozzle bottom on a test engine and release some of the regen cooling methane out into the void? I can't think of one, but this almost looks like that to me.  ???
Your question about Raptor's exhaust plume was answered in previous posts. They increased the film cooling with methane to avoid overheating the thrust chamber allowing them to test at full thrust with current thrust chamber design. They will learn from this and make the next iteration of Raptor with redesigned cooling channels to reduce the amount of film cooling required.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: matthewkantar on 02/07/2019 06:48 pm
300 bar is the stated goal, that number wasn’t pulled from a hat. I would imagine it was at the intersection of what is achievable at X engine weight, Y safety factor, Z material toughness, etc. Mr Musk starts with what is physically possible and aims for that.

For an outfit that has professed a desire for economy and not performance, SpaceX  is attempting a no compromise engine, near the top of all performance categories, reusable and cheap.

Raptor looks like it is hitting its marks, I don’t see why they won’t hit 300 bar take or give a few percent. 
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/07/2019 06:52 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/07/2019 06:59 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.

Indeed, I think for maximum thrust chamber life (low-cycle fatigue is the limitation) they will run at lowest peak engine pressure they can get away with to reduce peak heat-flux induced thermal strain per firing cycle.  Turn it up to 11 only in emergencies.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rsdavis9 on 02/07/2019 07:00 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.

Probably after they get their 3(or 4 one spare) raptors for boca chica.
Don't want to slow down the hopper people. :)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: scr00chy on 02/07/2019 07:03 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/07/2019 07:08 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.

Agreed. At the end of the day, the true test of Raptor's design is not simply reaching 300 bar or 200 tons of thrust (that has never been the challenge AFAIK) but doing so with extreme reliability and reusability. I can't imagine that the propulsion team has even scratched the surface of determining whether Raptor's updated design can satisfy those end goals.

I'm optimistic but am definitely reserving judgement until we have additional information.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/07/2019 07:20 pm
I found https://www.sciencedirect.com/science/article/pii/S2212540X1830004X which talks about ISP reduction, but without figures for methane.  It also discusses coking problems using RP-1 as the coolant, which is applicable to Merlin.  That article links to https://arc.aiaa.org/doi/pdf/10.2514/6.2010-6721, but I sadly don't have institutional access.  If someone does and indicates it's worth it, I'll be happy to create an account (:sigh:) and purchase it.  I can't even find the price without an account apparently though.

[the really interesting pointer appears to be to https://iris.unicampania.it/handle/11591/172095#.XFxBW89KiGR, but it wasn't immediately clear to me how to even get access to that one.]

In the spirit of democratizing research, here are your mentioned papers.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Norm38 on 02/07/2019 07:25 pm
For an outfit that has professed a desire for economy and not performance, SpaceX  is attempting a no compromise engine, near the top of all performance categories, reusable and cheap.

It's targeted performance obviously.  The F9 is optimized for cost, not performance, but the Merlin 1D still has the highest thrust to weight ratio.  That is the performance that allowed them to not have to obsess over every gram on the rocket, to add legs, grid fins, beef up where they had to, and land the first stage.

By making Raptor be the most powerful, most advanced engine they possibly can, they give the overall system the performance margins so that they can optimize on cost everywhere else.

Then they loop round again and use one variant of the Raptor everywhere (to start) to further reduce cost. And they will do what they can in manufacturing to reduce cost further.  By pushing the performance envelope as far as they have, they can probably make small sacrifices in production if it gives a big cost savings.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: intelati on 02/07/2019 07:26 pm
I found https://www.sciencedirect.com/science/article/pii/S2212540X1830004X which talks about ISP reduction, but without figures for methane.  It also discusses coking problems using RP-1 as the coolant, which is applicable to Merlin.  That article links to https://arc.aiaa.org/doi/pdf/10.2514/6.2010-6721, but I sadly don't have institutional access.  If someone does and indicates it's worth it, I'll be happy to create an account (:sigh:) and purchase it.  I can't even find the price without an account apparently though.

[the really interesting pointer appears to be to https://iris.unicampania.it/handle/11591/172095#.XFxBW89KiGR, but it wasn't immediately clear to me how to even get access to that one.]

In the spirit of democratizing research, here are your mentioned papers.

Appreciate it. I lost access to my account (YES!), but I don't have a problem with the sharing (https://twitter.com/hwitteman/status/1015049411276300289) within forums or stockpiling on offline access. Thanks.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 02/07/2019 07:34 pm

Does anyone think they will achieve 300+ bars?

Does anyone honestly think that they will not?

One thing that SpaceX really likes is continuous, incremental improvement. The chamber pressure of the Merlin engine almost doubled during it's life. Right now, the very first Raptor is clocking in at pressures >250 in the first week of testing. I would be genuinely surprised if a few years from now they were still running on engines that have seen minimal improvement from today.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JonathanD on 02/07/2019 07:50 pm

Oh they'll hit 300 bar.  Might pop the test article, but if stuff ain't breakin' you ain't tryin'  8)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Star One on 02/07/2019 07:55 pm
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.

Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.

I imagine it’s very dedicated, highly skilled and motivated people willing to work 24/7. Not besmirching people working on other engines who I am sure have equal qualities.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/07/2019 08:01 pm
Once again, "reaching" 300 bar or certain thrust targets is a pittance compared to what Musk actually means when he says things like "Will take us time to work up to 300 bar, that is a mad level." The challenge is "working up to 300 bar" while ensuring that Raptor's thrust to weight ratio, reusability, and reliability remain absolutely exceptional.

Starhopper and Starship and Super Heavy can continue development almost unimpeded even without reaching those targets now, but the BFR program will have failed to fulfill its purpose (i.e. enabling sustainable access to orbit and interplanetary transport) if Raptor is not eventually able to meet them.

The way I lately think about it is through the analogy of car engines or batteries. If you want immense power output in a short period of time and a compact package, that typically requires major tradeoffs. Capacitors can store a fraction of the charge of batteries, while top-fuel dragsters often need their engines entirely inspected and/or rebuilt after just a few full-power runs. Raptor at 300+ bar is the equivalent of SpaceX trying to build a top-fuel dragster engine that can do 10 or 100 top-speed runs back-to-back with little to no repairs or inspections and minimal performance drop-off.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/07/2019 08:09 pm
Two things I've got to get out of my system;

1) We're used to 'Elon time', but both the hopper and the recent Raptor arrival to first test timing then the first test to second test timing are exactly the opposite of the usual Elon timing plan, they're surprising us with their rapidity, rapidity beyond not behind what Elon's words would have had us expecting.

2) 100,000 Hp just to run the fuel pumps.  For one individual engine.  100,000 Hp to power the fuel pumps.  There might be 31 of them going off at once.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/07/2019 08:23 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
SpaceX may test Raptor briefly at 30MPa soon although it will likely be a few years before Raptor is routinely running at that during normal operational missions.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: launchwatcher on 02/07/2019 08:29 pm
I'm reminded of Intel's tick-tock model (https://en.wikipedia.org/wiki/Tick%E2%80%93tock_model) from the mid 2000's. SpaceX is so far forward of the rest of the launchers (ULA/BO/ESA...) in their research and development it's not even funny
hmm, tick-tock is not directly applicable here - it's holding one thing constant while changing another (in intel's case, semiconductor manufacturing process vs chip design).    In chip design -- where you can have relatively clean interfaces that let you change one thing without changing the other -- you can pull it off, and it lets you debug your unproven manufacturing process using a chip design that's known to work, and your unproven chip designs with a manufacturing process known to work.

But you don't have that in rockets - you can't trivially replace the Merlin engines on a Falcon 9 with Raptors without changing a bunch of other things as well..
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: intelati on 02/07/2019 08:32 pm
I'm reminded of Intel's tick-tock model (https://en.wikipedia.org/wiki/Tick%E2%80%93tock_model) from the mid 2000's. SpaceX is so far forward of the rest of the launchers (ULA/BO/ESA...) in their research and development it's not even funny
hmm, tick-tock is not directly applicable here - it's holding one thing constant while changing another (in intel's case, semiconductor manufacturing process vs chip design).    In chip design -- where you can have relatively clean interfaces that let you change one thing without changing the other -- you can pull it off, and it lets you debug your unproven manufacturing process using a chip design that's known to work, and your unproven chip designs with a manufacturing process known to work.

But you don't have that in rockets - you can't trivially replace the Merlin engines on a Falcon 9 with Raptors without changing a bunch of other things as well..

Tock as in the engine families. Tick as in the thrust upgrades. Instead of starting with a clean sheet everytime you want to make an improvement (AKA SSME vs RS-68)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RotoSequence on 02/07/2019 08:36 pm
Forgive this question from ignorance, but I've recently come off of a misunderstanding of how fuel mixture ratios are measured (by molecular weight, as it turns out); Raptor's 3.81 mix ratio should mean a fuel rich engine cycle. The bright violet flame from Raptor should, to my knowledge, reflect oxygen-rich combustion. What am I missing?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/07/2019 08:54 pm
100,000 Hp just to run the fuel pumps.  For one individual engine.  100,000 Hp to power the fuel pumps.  There might be 31 of them going off at once.

Back to my bogglement with the power of the Raptor fuel pumps (propellant pumps actually, both sides)...
100,000 hp x 31 engines = 3,100,000 hp total for a SH.

Now some stats for Hoover Dam:
Flow through turbines - 906 m^3/s
Head - 158.5 m
Max power - 3,000,000 hp or 2,000,000 kW.

So about the same. 

Keep in mind I'm not talking about the power of the Raptor engines themselves, only the liquid pumps that keep them fed.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lemurion on 02/07/2019 09:21 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: gongora on 02/07/2019 09:23 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Rocket Surgeon on 02/07/2019 09:46 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.

BE-4 is fairly far along.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HeartofGold2030 on 02/07/2019 09:53 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.

BE-4 is fairly far along.

The duration argument is valid, but Raptor has already fired at a higher thrust than BE-4 ever has. The current version of the BE-4 is only able to throttle up to 70% (1.68MN) due to "setbacks", while Raptor has already fired at 1.72MN, albeit only for a short time.

https://twitter.com/jeff_foust/status/1088088915771342848
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Doesitfloat on 02/07/2019 10:00 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.

BE-4 is fairly far along.

Respectfully disagree on the state of BE-4.

Blue and X use different terms to describe their development engines.
BE-4 has long firing of "full scale engine". This enging was full size but incapable of more than 70% of full thrust.
X has long firing of "sub-scale raptor". This engine was full size but only able to reach 250 of the 300 bar pressure goal.


These two development engines appear the same   full size not full thrust.

BE-4 is designing a new engine that will be able to fire at full thrust.

X just unveiled a "flight engine" that has briefly firied at lowest acceptable thrust level for system to work.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jketch on 02/07/2019 10:14 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.

BE-4 is fairly far along.

Respectfully disagree on the state of BE-4.

Blue and X use different terms to describe their development engines.
BE-4 has long firing of "full scale engine". This enging was full size but incapable of more than 70% of full thrust.
X has long firing of "sub-scale raptor". This engine was full size but only able to reach 250 of the 300 bar pressure goal.


These two development engines appear the same   full size not full thrust.

BE-4 is designing a new engine that will be able to fire at full thrust.

X just unveiled a "flight engine" that has briefly firied at lowest acceptable thrust level for system to work.

Agree regarding BE-4s lower thrust testing, but AFAIK the Raptor tests that happened before this week were done with an engine that was physically smaller, not just lower thrust. I think this is the first time that a full-size Raptor has been fired and it hasn't done any long duration burns yet.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: hplan on 02/07/2019 10:45 pm
Agree regarding BE-4s lower thrust testing, but AFAIK the Raptor tests that happened before this week were done with an engine that was physically smaller, not just lower thrust. I think this is the first time that a full-size Raptor has been fired and it hasn't done any long duration burns yet.

First time it has been publicly announced. But do we really know there were no full-sized test engines in the last year and a half? It's hard to imagine that SpaceX has just now produced the first full-size "radically-redesigned" test article and fired it at near full power a couple of days later, or that Musk would so confidently predict the hopper would be flying on these in the next two months, unless this were a smaller revision of an earlier full-size development engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lemurion on 02/07/2019 10:55 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.

BE-4 is fairly far along.

BE-4 is pretty far along, but Blue has yet to fire the full-thrust version of the engine. SpaceX has fired their flight thrust version of Raptor.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Jcc on 02/07/2019 10:59 pm
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.

BE-4 is fairly far along.

BE-4 is pretty far along, but Blue has yet to fire the full-thrust version of the engine. SpaceX has fired their flight thrust version of Raptor.

Let's see which one flies first!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/07/2019 11:08 pm
BE-4 is pretty far along, but Blue has yet to fire the full-thrust version of the engine. SpaceX has fired their flight thrust version of Raptor.
Let's see which one flies first!
Unless Blue is planning a suborbital test vehicle, that will be no contest.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: abaddon on 02/07/2019 11:10 pm
How about we not turn this into yet another BO vs SoaceX thread?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: wannamoonbase on 02/07/2019 11:18 pm
Elon is stressing the importance of being fast and cheap on this development. 

It maybe hard to press engine development but I bet we see an insane charge at getting this version of Raptor ready. 

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/07/2019 11:43 pm
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
Okay, but a month before that, they were still seemingly using carbon fiber for Starship, and they called it BFR.

Now, they've switched to stainless, built a mockup/fairing, had a photo-op, had it blow over, and are now fitting out the mockup to do flights.

A heck of a lot changes in a months' time nowadays.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ulm_atms on 02/08/2019 12:11 am
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
Okay, but a month before that, they were still seemingly using carbon fiber for Starship, and they called it BFR.

Now, they've switched to stainless, built a mockup/fairing, had a photo-op, had it blow over, and are now fitting out the mockup to do flights.

A heck of a lot changes in a months' time nowadays.

But you have to admit...  It's fun to watch!  ;D
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/08/2019 12:17 am
How about we not turn this into yet another BO vs SoaceX thread?
(mod) Yes, let us not do that, thanks
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/08/2019 02:29 am
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.

Raptor has fired for a few seconds.

I imagine that with the tremendous flow rates, thin and high thermal conductivity materials and heat fluxes at play within the Raptor that at 2s everything is basically steady state.  If it survives for 2s most likely it will survive for 200s
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DanielW on 02/08/2019 02:37 am
100,000 Hp just to run the fuel pumps.  For one individual engine.  100,000 Hp to power the fuel pumps.  There might be 31 of them going off at once.

Back to my bogglement with the power of the Raptor fuel pumps (propellant pumps actually, both sides)...
100,000 hp x 31 engines = 3,100,000 hp total for a SH.

Now some stats for Hoover Dam:
Flow through turbines - 906 m^3/s
Head - 158.5 m
Max power - 3,000,000 hp or 2,000,000 kW.

So about the same. 

Keep in mind I'm not talking about the power of the Raptor engines themselves, only the liquid pumps that keep them fed.
]

So what you are saying is, that the turbopumps on the Super heavy could push the Colorado river back up into lake meade.

Edit: somehow quoted the wrong post.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Herb Schaltegger on 02/08/2019 02:41 am
Agree regarding BE-4s lower thrust testing, but AFAIK the Raptor tests that happened before this week were done with an engine that was physically smaller, not just lower thrust. I think this is the first time that a full-size Raptor has been fired and it hasn't done any long duration burns yet.

First time it has been publicly announced. But do we really know there were no full-sized test engines in the last year and a half? It's hard to imagine that SpaceX has just now produced the first full-size "radically-redesigned" test article and fired it at near full power a couple of days later, or that Musk would so confidently predict the hopper would be flying on these in the next two months, unless this were a smaller revision of an earlier full-size development engine.


We know because SpaceX’s McGregor test facility has too many eyes on the ground (and in the sky - See L2 content) watching every move to have missed a full size Raptor test, while the effects of plenty of sub-scale testing has been documents (again, see L2 content for reference).

The only facilities in the country capable of testing flight-size Raptors are NASA facilities (and we know about power pack tests already), Blue Origin’s (presumably, since that’s where BO tests its own), or McGregor. Do you really think SpaceX could do secret tests at a NASA site, with federal and contractor workers all over the place, and not have it leak out? Do you really think Blue would let SpaceX test engines in its secretative site far from watching eyes?

That leaves McGregor. Which only recently was upgraded and modified to support full-size Raptor testing (again, see L2).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 02/08/2019 03:24 am
Here's a paper on "transpiration cooling" in rocket nozzles.  Same technique as SpaceX now say they will use to cool the skin during atmospheric reentry.j . From 1995; not that new a cooling technique.  Well studied.

Transpiration cooling effects on nozzle heat transfer and performance (https://arc.aiaa.org/doi/10.2514/3.26718) :  D. Keener, J. Lenertz, R. Bowersox, and J. Bowman.  "Transpiration cooling effects on nozzle heat transfer and performance", Journal of Spacecraft and Rockets, Vol. 32, No. 6 (1995), pp. 981-985.
https://doi.org/10.2514/3.26718 (https://doi.org/10.2514/3.26718)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/08/2019 03:29 am
Everyday Astronaut illustrates transpiration nicely, with modern analogs (aircraft deicing). Starts at about 06:55.

https://www.youtube.com/watch?v=LogE40_wR9k
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HMXHMX on 02/08/2019 04:51 am
100,000 Hp just to run the fuel pumps.  For one individual engine.  100,000 Hp to power the fuel pumps.  There might be 31 of them going off at once.

Back to my bogglement with the power of the Raptor fuel pumps (propellant pumps actually, both sides)...
100,000 hp x 31 engines = 3,100,000 hp total for a SH.

Now some stats for Hoover Dam:
Flow through turbines - 906 m^3/s
Head - 158.5 m
Max power - 3,000,000 hp or 2,000,000 kW.

So about the same. 

Keep in mind I'm not talking about the power of the Raptor engines themselves, only the liquid pumps that keep them fed.
]

So what you are saying is, that the turbopumps on the Super heavy could push the Colorado river back up into lake meade.

Edit: somehow quoted the wrong post.

Decades ago when I thought I might actually run a profitable space transportation company one day, I fantasized about a nice company campus with a lake in the middle, equipped with a fountain.  Nothing too odd about that, you say.

But I had my eye on a  surplus J2 centrifugal turbopump to be the fountain pump...the top of the stream would have been about 750 foot high...needless to say, you'd only run it on special occasions!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: kkattula on 02/08/2019 05:53 am
I remember following Armadillo Aerospace back in the day.  They rapidly tried a LOT of different engines cooling techniques, and eventually settled on a simple stainless steel chamber and nozzle with film cooling.  Stainless is a terrible choice for regen cooling because it conduct sheat poorly, but the film cooling (ethanol) worked so well they didn't need anything else.

I believe Exos is still using the same technology
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: alhenry1231 on 02/08/2019 08:07 pm
What makes the Raptor hardware special?

I would prefer not to get into the social, budgetary or time of development etc. aspect.

Specifically hardware.
By special lets try and keep it in a class of its own or maybe only one other engine with a similar capability.

I'll kick off with something likely to be special about the Raptor:
Thrust to weight ratio.
Mueller has stated that Merlin currently holds the record but Raptor is coming.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tulse on 02/08/2019 08:24 pm
It will be the first full flow staged combustion engine to fly (presuming it actually does).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/08/2019 08:35 pm
To get the ball rolling.

What makes Raptor special:

- Full flow cycle allows higher chamber pressures with lower peak pump pressures and lower turbine temperatures.

- Use of Methane as fuel improves cooling and reduces coking relative to RP1.

- Inline Lox pump and pre-burner reduces plumbing weight.

- 3D printing of complex parts reduces weight and cost.

- Designed from the outset for long life.


Also What is not special about Raptor:

- Copper alloy MCC with brazed steel outer shell. Brazing MCC outer shell developed in USSR.

- Full flow researched and partially developed by USSR and USA.

- Use of high temperature and pressure oxygen compatible material for turbo-pumps and pre-burners.


I am sure I left out a lot. Feel free to add to.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mead on 02/08/2019 09:39 pm
Where did the term gas-gas come from? The Wikipedia article on FFSC references an NSF article, and Elon tweeted that the Raptor is gas-gas, but I can't find other references. I have a very poor understanding of chemistry, but aren't both the fluids coming out of the preburners supercritical? The pressure should be well over 3000psi.

On a related not, what would be the upper and lower bounds for O/F ratios on the preburners? I'm curious how the individual preburners' environments differ.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/08/2019 09:54 pm
Where did the term gas-gas come from? The Wikipedia article on FFSC references an NSF article, and Elon tweeted that the Raptor is gas-gas, but I can't find other references. I have a very poor understanding of chemistry, but aren't both the fluids coming out of the preburners supercritical? The pressure should be well over 3000psi.

On a related not, what would be the upper and lower bounds for O/F ratios on the preburners? I'm curious how the individual preburners' environments differ.

They become gasses (you're probably technically right that they are closer to supercritical fluids though) when they go through the preburners. On FRSC or ORSC half goes through preburners so it's gas-liquid. On FFSC, both go through preburners so they mix as gas-gas. Most others are liquid-liquid.

The chamber pressure is over 3,700 psi (250bar), so the preburner pressure should be much higher.

The Preburners in the SSME run at 361 bar, and the turbine inlets are 326 bar, the fuel compressor outlet pressure is 405 bar. This is for an engine with a chamber pressure of about 210 bar.

For the RD-170, compressor outlet on the Ox line is 602(!) bar preburner is 535 bar going to 509 bar at the turbine inlet. This is for an engine with a chamber pressure of about 250 bar.

Because of the efficiencies of full flow, I wouldn't be surprised if Raptor has lower pressure (than the RD-170) at the Oxidizer compressor/preburner, even on the 300+bar version.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/08/2019 10:21 pm
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD). 

Also interesting parts about IPD reports "Hydrostatic bearings are applicable to any nonpressure fed rocket engine using oxygen [and hydrogen]" and "The most stressing case of all the potential MSP concepts incorporated a vertical landing vehicle. This concept would require a propulsion system with an ability to throttle down to low power levels (to allow for a soft landing). The engine would also have to be restarted in order to land the vehicle safely. Accordingly, the IPD contractors were asked to develop restartable designs capable of throttling down to 20% of sea level design thrust."
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mead on 02/08/2019 11:01 pm
I guess gas-gas is easier to say than supercritical fluid-supercritical fluid?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: billh on 02/08/2019 11:19 pm
Perhaps the distinction being made by calling it gas-gas is that it's all single phase flow downstream of the preburners - even though it technically is probably supercritical.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: NH22077 on 02/08/2019 11:48 pm
I've show the Raptor 2 second test fire to a few people & some were put off by the audio delay.
We on NSF know the camera was at a safe distance. 600+ feet, sound takes time to travel that far. ;)
So I added some black frames to the beginning & shifted the audio. Till I was happy with it.
Here are the results.

Ned
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: 1 on 02/09/2019 12:49 am
Where did the term gas-gas come from?

Answering this somewhat more generally, I believe the [phase of matter]-[phase of matter] terminology is how the Russians normally classify general engine types. Some other posters here have mentioned that elsewhere.

While you're at it, you should check out the wikipedia page on Raptor. livingjw's Raptor schematic is now on both English and Russian-language pages. :-)

https://en.wikipedia.org/wiki/Raptor_(rocket_engine_family)#Design
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: su27k on 02/09/2019 01:03 am
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD). 

Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/09/2019 01:13 am
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/09/2019 02:18 am
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John

Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".

https://twitter.com/elonmusk/status/1008385171744174080

Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/09/2019 02:32 am
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John

Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".

https://twitter.com/elonmusk/status/1008385171744174080

Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: gongora on 02/09/2019 02:33 am
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD). 

Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.

Well, they hired Jeff Thornburg:
Quote
Jeff was stationed at Edwards AFB, CA where he joined the liquid rocket engine branch at the Air Force Research Laboratory and worked several component and engine technology programs that included his leadership of the joint Air Force-NASA Integrated Powerhead Demonstration engine which was the world’s first hydrogen full-flow staged combustion cycle rocket engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OneSpeed on 02/09/2019 02:40 am
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John

That's great. Do you have updated Isp figures to go with the new thrust and pressure values?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mead on 02/09/2019 03:23 am
What mass flow rate are you using for those numbers?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/09/2019 03:27 am
>
Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.

Musk said in a Dec 22, 2018 tweetstorm,

Quote
....Raptor turbopumps alone need 100,000 horsepower per engine. That’s not a typo.

100k HP = 74.57 MW

https://twitter.com/elonmusk/status/1076618886932353024
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: mme on 02/09/2019 04:32 am
>
Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.

Musk said in a Dec 22, 2018 tweetstorm,

Quote
....Raptor turbopumps alone need 100,000 horsepower per engine. That’s not a typo.

100k HP = 74.57 MW

https://twitter.com/elonmusk/status/1076618886932353024
100,000 horsepower is about 75 MW.  And he says per engine, so could that be 37.5 MW per pump? All this unit willy-nilliness hurts my head.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Brunberg on 02/09/2019 05:05 am
I would imagine that oxygen pump is more powerful.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: azinyk on 02/09/2019 07:48 am
I would imagine that oxygen pump is more powerful.
Can you elaborate?  Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.

On the shuttle, the fuel pumps were about 3x as powerful.  Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Star One on 02/09/2019 08:58 am
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD). 

Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.

Well, they hired Jeff Thornburg:
Quote
Jeff was stationed at Edwards AFB, CA where he joined the liquid rocket engine branch at the Air Force Research Laboratory and worked several component and engine technology programs that included his leadership of the joint Air Force-NASA Integrated Powerhead Demonstration engine which was the world’s first hydrogen full-flow staged combustion cycle rocket engine.

So acquired means hired the guy with the knowledge in his head, to put it crudely.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/09/2019 12:48 pm
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD). 

Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.
Nope

...
Either way, HVM and livingjw, I'll be happy to help get a properly licensed and up-to-date version of your graphic up on Wikipedia....
Do you know what happens people who post traced art online?

Also it's not ready.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/09/2019 12:59 pm
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John

That's great. Do you have updated Isp figures to go with the new thrust and pressure values?

I will dig them out for you. They didn't change more than a second or two. Since I am using the same ER=35. Sea level Isp goes up a couple of seconds.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/09/2019 02:02 pm
I would imagine that oxygen pump is more powerful.
Can you elaborate?  Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.

On the shuttle, the fuel pumps were about 3x as powerful.  Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.

The volumetric flow of methane is a little lower, but the pressure is higher since it has to flow through the regen system while the oxygen goes straight to the MCC. I would expect the fuel pump to need slightly more power, but they should be roughly similar.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/09/2019 02:06 pm
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John

Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".

https://twitter.com/elonmusk/status/1008385171744174080

Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.

Here are a few possible guesses/suggestions that may help reconcile the numbers:

-The peak pressure may be much higher on the fuel side because it will be sized for a pressure drop to feed a much larger vacuum nozzle.

-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it.

-The Oxidizer line into the fuel rich preburner may need a boost pump. I think the RD-180 boosts fuel pressure to 700+bar in order to feed into the preburners. SSME has to boost ox pressure to 472 bar to feed into the preburners. Fuel line may not need one if my first suggestion is true.

(https://pbs.twimg.com/media/CL2pTAAUYAAuQGf.jpg)

Boost pumps in general might be the culprit for the apparent increase in plumbing in the new Raptor model.

-Liquid-gas preburners seem to need much larger pressure drops on the liquid side to feed well.

Just some guesses.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/09/2019 02:07 pm
I would imagine that oxygen pump is more powerful.
Can you elaborate?  Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.

On the shuttle, the fuel pumps were about 3x as powerful.  Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.

The volumetric flow of methane is a little lower, but the pressure is higher since it has to flow through the regen system while the oxygen goes straight to the MCC. I would expect the fuel pump to need slightly more power, but they should be roughly similar.

The fuel pump is probably also sized for an eventual vacuum nozzle.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/09/2019 02:23 pm
I would imagine that oxygen pump is more powerful.
Can you elaborate?  Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.

On the shuttle, the fuel pumps were about 3x as powerful.  Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.

Methane pump takes more power due to low density and higher pressure needed because of cooling pressure drop.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/09/2019 02:50 pm
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.

John

Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".

https://twitter.com/elonmusk/status/1008385171744174080

Quote
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.

Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.

Here are a few possible guesses/suggestions that may help reconcile the numbers:

-The peak pressure may be much higher on the fuel side because it will be sized for a pressure drop to feed a much larger vacuum nozzle.

-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it.

-The Oxidizer line into the fuel rich preburner may need a boost pump. I think the RD-180 boosts fuel pressure to 700+bar in order to feed into the preburners. SSME has to boost ox pressure to 472 bar to feed into the preburners. Fuel line may not need one if my first suggestion is true.

(https://pbs.twimg.com/media/CL2pTAAUYAAuQGf.jpg)

Boost pumps in general might be the culprit for the apparent increase in plumbing in the new Raptor model.

-Liquid-gas preburners seem to need much larger pressure drops on the liquid side to feed well.

Just some guesses.

All good suggestions. I am looking into all of those use both the SSME and RD180 for guidance. Thanks.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/09/2019 02:52 pm
Here are a few possible guesses/suggestions that may help reconcile the numbers:
...
-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it....

Another guess: 800 bar is the very local dynamic pressure due to pressure waves generated by the turbine blades, which is chemically relevant in local areas of the turbine, but is not the mean pressure at a larger scale.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/09/2019 04:32 pm
Here are a few possible guesses/suggestions that may help reconcile the numbers:
...
-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it....

Another guess: 800 bar is the very local dynamic pressure due to pressure waves generated by the turbine blades, which is chemically relevant in local areas of the turbine, but is not the mean pressure at a larger scale.

Pressures could also be higher during startup and/or shutdown.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: alang on 02/10/2019 12:52 pm
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.

You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.

If the amount of film cooling is so easily tunable that it could be altered for a nearly immediate subsequent test, and if they had this problem at the start of testing of the previous Raptor, then I wonder why they didn't start with more film cooling and then tune it back to a point where a hint of copper started to show in the spectrum?
Maybe the film cooling theory is wrong.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Hominans Kosmos on 02/10/2019 03:27 pm
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.

You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.

If the amount of film cooling is so easily tunable that it could be altered for a nearly immediate subsequent test, and if they had this problem at the start of testing of the previous Raptor, then I wonder why they didn't start with more film cooling and then tune it back to a point where a hint of copper started to show in the spectrum?
Maybe the film cooling theory is wrong.
 

The copper may have just been traces from production and burned off in the first firing. Elon seemed not to worry about it at all.
 

I mean, piston engines (while more complex in terms of moving parts) have a break-in period... why wouldn't a brand new rocket engine have some settling-in phenomena? /notarocketscientist
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureMartian97 on 02/10/2019 08:00 pm
Possible Raptor test?

https://twitter.com/DJSnM/status/1094701660390121479
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/10/2019 08:23 pm
Possible Raptor test?

https://twitter.com/DJSnM/status/1094701660390121479
There are regular rocket engine tests in McGregor so the test heard may not necessarily be a Raptor test. Perhaps someone can find out if the noise was a Raptor test or not. A regular M1D test can sometimes sound a lot louder than normal due to a temp. inversion.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/10/2019 08:25 pm
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.

You think that thin halo around the core exhaust stream is near-pure methane?  From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?

That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.

If the amount of film cooling is so easily tunable that it could be altered for a nearly immediate subsequent test, and if they had this problem at the start of testing of the previous Raptor, then I wonder why they didn't start with more film cooling and then tune it back to a point where a hint of copper started to show in the spectrum?
Maybe the film cooling theory is wrong.

Maybe a hint of copper isn't actually a problem. We're probably talking tenths of a gram of vaporized metal.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: CardBoardBoxProcessor on 02/10/2019 09:34 pm
So what's the benefit or the lox hardware being directly atop The chamber?
And why is it but done before?

Is the idea directly from the integrated power head demonstrator?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/10/2019 09:51 pm
So what's the benefit or the lox hardware being directly atop The chamber?
And why is it but done before?

Is the idea directly from the integrated power head demonstrator?

- lighter weight, but requires tight integration.

- Integrated power head demonstrator never integrated parts together in a flight weight engine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/10/2019 10:08 pm
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.

- Added pressures, temperatures and pump powers to engine schematic.

- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.

- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power

- At 320 bar I get  2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.

Added some more flow detail.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: xyv on 02/10/2019 11:19 pm
This engine schematic just keeps getting better.  I see you are still in the 'no boost pumps' camp - I do not know enough high pressure fluid dynamics to have an informed opinion.  One slight suggestion: you show both autogenous heat exchanger loops with the liquid exiting them.  Without adding more colors (which is hard with an easy on the eyes pastel palette...) maybe use polka dots for the gas?

It's not actually confusing but the purist in me is bothered.  Again, really appreciate the effort here.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 02/10/2019 11:25 pm
So what's the benefit or the lox hardware being directly atop The chamber?

One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.

Quote
And why is it but done before?

That's a good question.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/10/2019 11:50 pm
So what's the benefit or the lox hardware being directly atop The chamber?

One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.

Quote
And why is it but done before?

That's a good question.

I know of one small research engine that was constructed that way. I think it was done by a university.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: docmordrid on 02/11/2019 12:06 am
This engine schematic just keeps getting better.  I see you are still in the 'no boost pumps' camp - I do not know enough high pressure fluid dynamics to have an informed opinion. 
>

Not sure if this applies; boost pumps = multi-stage? but from 2016...

David Ki Sun Yoon v@DavidKYoon 
@elonmusk Sweet Jesus, that means you are pumping to 45-50 MPa... Surely this will be using multiple stage pumps?
|
Elon Musk ✓ @elonmusk
@DavidKYoon yes
11:41 PM - 25 Sep 2016

https://twitter.com/elonmusk/status/780296159315173376
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/11/2019 12:51 am
So what's the benefit or the lox hardware being directly atop The chamber?

One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.

Quote
And why is it but done before?

That's a good question.

Meanwhile, the RD-180 oxidizer line takes a pretty snaking route. You're talking about a half meter of travel from the Oxidizer turbine inlet into the combustion chamber for the Raptor. On the RD-180, you're talking about like 2.5 meters of travel from the Compressor outlet to the preburner at 602-535 bar, about 2 meters of travel at 535-509 bar to the turbine, then about 3 meters of travel to the combustion chamber at 262-245 bar. This is probably the primary culprit for the RD-180's TWR of 80:1 vs possibly 200:1 on Raptor.

(NOTE: Pressure values are for the RD-170, so the RD-180 will differ slightly)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Mongo62 on 02/11/2019 01:28 am
https://twitter.com/elonmusk/status/1094782854007910400
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/11/2019 01:40 am
https://twitter.com/elonmusk/status/1094782854007910400

Is anyone aware of other flight-grade engines that have reached 269 bar or higher in ground testing? I feel like it's not exactly fair to compare Raptor with the chamber pressure of an engine that's been flying for decades.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Robotbeat on 02/11/2019 01:42 am
I'm sure other engines have been tested on the test stand to higher pressures. But among flight engines, none have exceeded Raptor's chamber pressure here.

We probably can't definitively give the crown to Raptor until it flies, but this is incredibly impressive.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Scylla on 02/11/2019 01:43 am
https://twitter.com/elonmusk/status/1094788148230406146
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Scylla on 02/11/2019 01:48 am
https://twitter.com/elonmusk/status/1094785437216829440
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Scylla on 02/11/2019 01:52 am
https://twitter.com/elonmusk/status/1094788783331897344
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/11/2019 01:53 am
https://twitter.com/elonmusk/status/1094789114258313216

Quote
Yes, aiming for 380 sec Isp with vac nozzle. Maybe 382 if we get lucky.

https://twitter.com/elonmusk/status/1094792462436990976

Quote
But not an extendable nozzle though, as that just saves length. Nozzle diameter is limited by body diameter.

https://twitter.com/elonmusk/status/1094790663646760961

Quote
Propellant was not deep cryo. CH4 & O2 were just barely below liquid temp at 1 bar. In theory, Raptor should do ~300 bar at deep cryo, provided everything holds together, which is far from certain. However, only 250 bar is needed for nominal operation of Starship/Super Heavy.

https://twitter.com/elonmusk/status/1094791972944920582

Quote
Much above 300 bar main chamber pressure means extreme oxygen preburner pressure of 700 to 800+ bar. Definitely pushing the limit of known physics.

Edit to add: pressure graph from Elon’s first tweet tonight posted earlier (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1910143#msg1910143) in this thread
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/11/2019 02:00 am
Quote from: @13ericralph31
Is the SL engine's expected vac ISP relatively unchanged from IAC 2017?

https://twitter.com/elonmusk/status/1094792729865797632

Quote
Close
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/11/2019 02:03 am
https://twitter.com/elonmusk/status/1094793664809689089

Quote
This will sound implausible, but I think there’s a path to build Starship / Super Heavy for less than Falcon 9

Quote from: @SpaceXFan97
Wow! I assume the switch to Stainless Steel is a big factor in this?

https://twitter.com/elonmusk/status/1094794147980931073

Quote
Yes

Edit to add: borderline for this raptor thread, as Elon’s answer was much broader. I’ve moved subsequent tweet about launch costs to tweet summary thread (https://forum.nasaspaceflight.com/index.php?topic=47352.msg1910163#msg1910163)

Further addition: missed another follow-up on build cost

Quote from: @John_Gardi
Stainless steel is CHEAP! Welding it is easy!
Super Heavy & Starship will eventually use gaseous methane & oxygen to pressurize the tanks so no helium is needed. They'll use the same gasses for the maneuvering thrusters as well, so no nitrogen or hydrazine will be needed either.

https://twitter.com/elonmusk/status/1094849332841304064

Quote
True
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/11/2019 02:26 am
My model show that thrust at 269 bar to be about ~1.83 MN, ~1.87 tonne. The model shows peak pressure is the methane pump output at ~570 bar, Lox pump at ~563 bar. This is lower than RD180's Lox pump at 606 bar, but higher than its RP1 pump 509 bar.  RD180 does have a small RP1 pump 809 bar which feeds the burner.

At 300 bar chamber pressure, model predicts a Lox pump pressure of 633 bar. I don't know if Elon mean the 700 bar to go with the 300 bar chamber pressure.

At 320 bar chamber pressure, model predicts a Lox pump pressure of 683 bar. CH4 pump 698 bar. 320 chamber pressure vacuum engine, ER=119, matched stated 250 tonne max value.

SX data coming fast and furious!

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/11/2019 03:17 am
268.9 Bar equates to 3900.0 psi.  I think someone at McGregor is working in English units and Elon had to translate before tweeting.  I don't think those are Elon's favorite units.

-----------------

So what's the benefit or the lox hardware being directly atop The chamber?

One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.

Quote
And why is it but done before?

That's a good question.
Could this be something that was made (more) possible by 3D metal printing?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: meekGee on 02/11/2019 03:18 am
268.9 Bar equates to 3900.0 psi.  I think someone at McGregor is working in English units and Elon had to translate before tweeting.  I don't think those are Elon's favorite units.

-----------------

So what's the benefit or the lox hardware being directly atop The chamber?

One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both &gt;500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.

Quote
And why is it but done before?

That's a good question.
Could this be something that was made (more) possible by 3D metal printing?

Well Bar isn't an SI unit either...

And they tried to get metric gauges, but couldn't fit the metric NPT threads to anything :)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/11/2019 03:25 am
268.9 Bar equates to 3900.0 psi.  I think someone at McGregor is working in English units and Elon had to translate before tweeting.  I don't think those are Elon's favorite units.

-----------------

So what's the benefit or the lox hardware being directly atop The chamber?

One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.

Quote
And why is it but done before?

That's a good question.
Could this be something that was made (more) possible by 3D metal printing?

It's probably a factor, but the most likely culprit is modern CAD.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Mark K on 02/11/2019 03:26 am
268.9 Bar equates to 3900.0 psi.  I think someone at McGregor is working in English units and Elon had to translate before tweeting.  I don't think those are Elon's favorite units.

Just look at the graph he showed - it is just the value reached, not a translation.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/11/2019 03:34 am
This engine schematic just keeps getting better.... One slight suggestion:
How's about showing a pintle injector in the main chamber?  I'm not certain but I think we know that to be the case?  and if it is pintle was that what you assumed in your pressure drop calculations?  How's about the injector type in the preburners, are they pintle a well or is that not applicable in that application?  I know what you (livingjw) said about having local areas of stoichiometric combustion in a sea of oxygen or a sea of methane but does that mean it has to be multiple areas or would they go for just one if the preburners are pintle-ly injected?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: PM3 on 02/11/2019 04:26 am
So what's the benefit or the lox hardware being directly atop The chamber?

Besides of the weight advantage: Hot oxygen-rich gas is agressive and needs special material to contain it. May save some bucks to minimize the usage of that special metal ...?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 02/11/2019 04:27 am
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.

- Added pressures, temperatures and pump powers to engine schematic.

- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.

- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power

- At 320 bar I get  2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.

Added some more flow detail.

John
LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: WormPicker959 on 02/11/2019 04:37 am
This engine schematic just keeps getting better.... One slight suggestion:
How's about showing a pintle injector in the main chamber?  I'm not certain but I think we know that to be the case?  and if it is pintle was that what you assumed in your pressure drop calculations?  How's about the injector type in the preburners, are they pintle a well or is that not applicable in that application?  I know what you (livingjw) said about having local areas of stoichiometric combustion in a sea of oxygen or a sea of methane but does that mean it has to be multiple areas or would they go for just one if the preburners are pintle-ly injected?
I'm not sure where this "knowledge" comes from, but my understanding is that pintle-type injectors are excellent at mixing liquid propellants - something that would be required in the preburners, but not the main combustion chamber. If the propellants are in the gas phase, they should mix well enough not to need a complicated injector. I have no idea if this distinction still applies with supercritical fluids, as I know next to nothing about them and their characteristics relative to liquids or gasses, so on that I must defer to someone with more expertise.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/11/2019 04:39 am
How's about showing a pintle injector in the main chamber?  I'm not certain but I think we know that to be the case?  ...
I'm not sure where this "knowledge" comes from, but my understanding is that pintle-type injectors are excellent at mixing liquid propellants - something that would be required in the preburners, but not the main combustion chamber. If the propellants are in the gas phase, they should mix well enough not to need a complicated injector.

That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 02/11/2019 04:48 am
How's about showing a pintle injector in the main chamber?  I'm not certain but I think we know that to be the case?  ...
I'm not sure where this "knowledge" comes from, but my understanding is that pintle-type injectors are excellent at mixing liquid propellants - something that would be required in the preburners, but not the main combustion chamber. If the propellants are in the gas phase, they should mix well enough not to need a complicated injector.

That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.
Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you 😁
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/11/2019 06:11 am
That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.
Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you

Yes but that does not have to mean it uses a pintle injector. There are many kinds.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/11/2019 07:33 am
https://twitter.com/elonmusk/status/1094782854007910400
I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 02/11/2019 10:20 am
That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.
Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you

Yes but that does not have to mean it uses a pintle injector. There are many kinds.

Are there any informed opinions on here regarding the best injectors to use for gas-gas?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: kemen on 02/11/2019 10:32 am
https://twitter.com/elonmusk/status/1094782854007910400
I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).

I think based on the graph’s legend, Musk is showing the RD180 pressures. So not space cows harmed in the making of that graph.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Vemaster on 02/11/2019 10:58 am
At 320 bar chamber pressure, model predicts a Lox pump pressure of 683 bar. CH4 pump 698 bar. 320 chamber pressure vacuum engine, ER=119, matched stated 250 tonne max value.

250 tonne thrust was stated for a sea-level: https://twitter.com/elonmusk/status/1091156245132673024
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Crispy on 02/11/2019 11:02 am
https://twitter.com/elonmusk/status/1094782854007910400
I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).

I think based on the graph’s legend, Musk is showing the RD180 pressures. So not space cows harmed in the making of that graph.
Seems to me the white horizontal line is RD180 operating pressure and the wiggly line is data from the Raptor test fire, showing a ramp up over time
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: kemen on 02/11/2019 11:11 am
https://twitter.com/elonmusk/status/1094782854007910400
I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).

I think based on the graph’s legend, Musk is showing the RD180 pressures. So not space cows harmed in the making of that graph.
Seems to me the white horizontal line is RD180 operating pressure and the wiggly line is data from the Raptor test fire, showing a ramp up over time

Yes, I think your explanation makes more sense than what I was thinking the graph meant.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/11/2019 12:37 pm
This engine schematic just keeps getting better.... One slight suggestion:
How's about showing a pintle injector in the main chamber?  I'm not certain but I think we know that to be the case?  and if it is pintle was that what you assumed in your pressure drop calculations?  How's about the injector type in the preburners, are they pintle a well or is that not applicable in that application?  I know what you (livingjw) said about having local areas of stoichiometric combustion in a sea of oxygen or a sea of methane but does that mean it has to be multiple areas or would they go for just one if the preburners are pintle-ly injected?

I think the shape is wrong for a pintle nozzle on the main chamber, but I will be changing the CH4 pre-burner to a pintle like central lox nozzle forming a stoichiometric recirculation region at the head of the chamber.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/11/2019 12:42 pm
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.

- Added pressures, temperatures and pump powers to engine schematic.

- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.

- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power

- At 320 bar I get  2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.

Added some more flow detail.

John
LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?

Could be, or there might be a boost pump hidden somewhere. I have not finished my pump design model yet, but I think you can do it with a single stage. Same goes for the turbine. We will see.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: garcianc on 02/11/2019 12:42 pm
Looking at the recent Twitter thread, and at the risk of rehashing some of the arguments from pages 72-74 of this thread, I was wondering about this question on twitter and whether the poster misspoke:

https://twitter.com/John_Gardi/status/1094784514503659522


and Elon Musk simply responded as the question was asked:

https://twitter.com/elonmusk/status/1094792462436990976


I have not seen much recent discussion about extendable nozzles in relation to Raptor, only Blue Origin. I believe (perhaps wishfully) that the real question was meant to be about a dual bell design, or has that question been settled? I looked through the thread and I didn't see anything to indicate so.


Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/11/2019 12:53 pm
At 320 bar chamber pressure, model predicts a Lox pump pressure of 683 bar. CH4 pump 698 bar. 320 chamber pressure vacuum engine, ER=119, matched stated 250 tonne max value.

250 tonne thrust was stated for a sea-level: https://twitter.com/elonmusk/status/1091156245132673024

I am looking into that.  So far I have maintained the 2017 geometry (Dexit=1.30m, ER=35). This matched so well I went with it. Chamber pressure will have to go up to well beyond 320 bar to achieve 250 tonne.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DusanC on 02/11/2019 01:04 pm
That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.
Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you

Yes but that does not have to mean it uses a pintle injector. There are many kinds.
I agree and I tried to say the same but it wasn't obvious from my comment.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tulse on 02/11/2019 01:23 pm
This may be a newbie question, but when the engine is throttled, is it controlled simply by changing flow via valves from the tank inlets?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 02/11/2019 02:43 pm
This may be a newbie question, but when the engine is throttled, is it controlled simply by changing flow via valves from the tank inlets?

Doing that quickly would reduce pressure in the pumps, possibly causing cavitation which would destroy them.

A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: matthewkantar on 02/11/2019 02:58 pm
The X axis in the graph is time,  represents 2 tenths of a second?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tulse on 02/11/2019 03:07 pm
A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.
Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/11/2019 03:56 pm
A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.
Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?

Basically yes.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/11/2019 04:01 pm
A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.
Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
Can't be exactly similar since they are different cycles, but one could throttle the gas generator maybe?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 02/11/2019 06:00 pm
So I note that the schematics listed seem to have very tight seals between the sides of the turbopump. Historically, those seals have been a source of failure if it meant getting oxygen in the fuel feed or vice versa.

But that's not an issue for a FFSC, because it segregates the fuel pump from the oxidizer flow, and vice versa.

So how much does that tight seal affect performance? Is it worth letting some small amount of pressurized fluid leak across as a lubricant?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/11/2019 06:13 pm
A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.
Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
Can't be exactly similar since they are different cycles, but one could throttle the gas generator maybe?

The gas generator is essentially a partial flow preburner, it throttles the same way as Raptor: reducing GG fuel/ox flow which reduces turbine power.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/11/2019 06:25 pm
Just of the record, this was from a Teslarati Jan 31 article on the Raptor engine. It confirms original layout with the methane turbopump-preburner on the side. We just haven't seen a picture yet.

.... the Raptor engine spotted departing SpaceX’s Hawthorne, CA factory last week was reportedly immense in person, towering over an M1D engine. Raptor also featured a mass of spaghetti-like plumbing (complexity necessary for its advanced combustion cycle), with a significant fraction of the metallic pipes and tubes displaying mirror-like finishes. Most notable was an obvious secondary preburner/turbopump stack and the lack of any exhaust port, ....

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JonathanD on 02/11/2019 06:28 pm
lack of any exhaust port

Well, no secondary exhaust port aside from the big one  ;D
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/11/2019 06:46 pm
LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?

That's what inducers (the helical screw shaped part of the impeller at the front) are for - to gently accelerate the inlet fluid (which is already flowing at ~15m/s) into the main part of the pump impeller.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Davidthefat on 02/11/2019 07:24 pm
This engine schematic just keeps getting better.  I see you are still in the 'no boost pumps' camp - I do not know enough high pressure fluid dynamics to have an informed opinion. 
>

Not sure if this applies; boost pumps = multi-stage? but from 2016...

David Ki Sun Yoon v@DavidKYoon 
@elonmusk Sweet Jesus, that means you are pumping to 45-50 MPa... Surely this will be using multiple stage pumps?
|
Elon Musk ✓ @elonmusk
@DavidKYoon yes
11:41 PM - 25 Sep 2016

https://twitter.com/elonmusk/status/780296159315173376

It's a matter of terminology IMO. A "stage" in a pump is a distinct pump in itself. In multistage pumps, each stage is fed by the preceding stage. To call the pump that takes the propellant from tank pressure to the pressure for the main pump a "boost pump" vs just the first stage of the 2 stage pump is just terminology. They are both doing the same thing. I suppose since if it's driven off of the same turbine, it would make it a "first stage" pump, but it's still a "boost pump" IMO. As John pointed out (and what I asked about) is that boost pumps historically had a separate turbine that powered the pumps that used a tap off of the high pressure outlet of the main or an even higher pressure "kick pump" that is used to power the preburner.

I know a specific ORSC engine ran the kick pump off of the same turbine as the main pump, technically making it a "second stage" of the main->kick pump assembly... So in the same logic, I just call the pump that "boosts" the propellant from the tank pressure to inlet of the main pump a "boost pump" regardless of what's driving it.

edit: I know there are distinctly different designs for the outlets of a single stage in a multi stage pump vs a single stage pump, like the cross over vanes.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/12/2019 01:11 pm
LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?

That's what inducers (the helical screw shaped part of the impeller at the front) are for - to gently accelerate the inlet fluid (which is already flowing at ~15m/s) into the main part of the pump impeller.

I am analyzing a Lox inducer+axial stage in front of main centrifugal pump similar to pictures below.

John
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: PM3 on 02/12/2019 06:35 pm
Suggest that a mod moves post #1892 and all replies to a separate Thread.

Edit/Lar:
Good suggestion. Better suggestion? Report to Mod rather than posting in the thread. Less clutter.
Your suggestion has been implemented. Here's the thread:
https://forum.nasaspaceflight.com/index.php?topic=47420
Further on that subtopic here may get aetherized, or sent out the airlock... depending on mood.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Rondaz on 02/13/2019 04:59 pm
THE “IMPOSSIBLE” TECH BEHIND SPACEX’S NEW ENGINE

by: Tom Nardi  Posted February 13, 2019 

https://hackaday.com/2019/02/13/the-impossible-tech-behind-spacexs-new-engine/
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tulse on 02/13/2019 06:28 pm
https://hackaday.com/2019/02/13/the-impossible-tech-behind-spacexs-new-engine/
That's a very clear and thorough description of the point of staged-combustion engines, and why full-flow is so beneficial.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DanielW on 02/13/2019 10:15 pm
https://hackaday.com/2019/02/13/the-impossible-tech-behind-spacexs-new-engine/
That's a very clear and thorough description of the point of staged-combustion engines, and why full-flow is so beneficial.

However, unless I am misunderstanding something badly, the author implies that Fuel-rich or Oxygen rich staged combustion engines are still somehow leaking efficiency. They aren't as the full flow still goes through the combustion chamber. Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/13/2019 10:47 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Hominans Kosmos on 02/15/2019 08:10 am
Tweaked CH4 pump to clear turbine exhaust path. Changed thrust to 2.05 MN. Changed Title to 12 Feb 2019.
 

Quoting from the schematic thread. Where do you think they would fit the main LOX valve? upstream from the pump? Squished between pump and gas generator?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/15/2019 01:47 pm
Tweaked CH4 pump to clear turbine exhaust path. Changed thrust to 2.05 MN. Changed Title to 12 Feb 2019.
 

Quoting from the schematic thread. Where do you think they would fit the main LOX valve? upstream from the pump? Squished between pump and gas generator?

Upstream of the LOX pump appears to be the only logical place. I probably should add one.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: vaporcobra on 02/16/2019 09:58 am
"SpaceX foundry casting Raptor engine manifold out of Inconel"

https://twitter.com/elonmusk/status/1096722535217917952
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Oersted on 02/16/2019 11:28 am
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dubya on 02/16/2019 12:55 pm
I assume that is a vacuum casting rig? If so is vacuum casting of Inconel common these days?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/16/2019 04:12 pm
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.

Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ppb on 02/16/2019 04:25 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: WormPicker959 on 02/16/2019 05:29 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?

My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.

I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!

Edit: I'm wrong about all this! :) But CorvusCorvax has provided a clarification/explanation (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1912510#msg1912510), and now we're all better for it!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rashomon on 02/16/2019 05:50 pm
People need to get closer to manufacturing roots if they think investment casting is medieval. (Though the basic technique is truly ancient.) How do you make near-net parts affordably? It usually involves melting metal.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: dglow on 02/16/2019 05:52 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?

My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.

I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!

There exist several staged-combustion engines which utilize twin turbines but are not full-flow, and SSME is among these.

Isn't there an additional aspect – beyond dual turbines alone – of running all your propellant through the preburners that leads to lower turbopump stress?   
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RedLineTrain on 02/16/2019 07:02 pm
What are the approximate dimensions of the Raptor engine manifold?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/16/2019 08:13 pm
What are the approximate dimensions of the Raptor engine manifold?

Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/16/2019 08:33 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?

- It doesn't. For a main chamber pressure of 300 bar (4351 psi), the CH4 pump outputs around 650 bar (9428 psi).

- Higher Isp (at sea level, not vacuum) is due to higher chamber pressure.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Bananas_on_Mars on 02/16/2019 08:39 pm
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.

Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.

Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.

Demo video (https://youtu.be/UJvjIB0rAUs)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RedLineTrain on 02/16/2019 09:31 pm
What are the approximate dimensions of the Raptor engine manifold?

Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?

Thank you, John.  I was mainly interested in understanding the size of the item that was being casted in Musk's tweet.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/16/2019 09:41 pm
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.

Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.

Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.
...
Not only that, you can also directly print negative molds:
https://www.youtube.com/watch?v=ENKSgeVOHwM (https://www.youtube.com/watch?v=ENKSgeVOHwM)

[edit] Ok, I don't know, would this be compatible for inconel? Also I don't think SpX would use printing service for IP reasons. They probably print/make molds themselves.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/16/2019 09:53 pm
What are the approximate dimensions of the Raptor engine manifold?

Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?

Thank you, John.  I was mainly interested in understanding the size of the item that was being casted in Musk's tweet.

Inconel would mean its either the Lox or CH4 preburner exhaust manifold. We're pretty sure the Lox preburner is symmetrical (at least it use to be), it's probably the CH4 manifold. Probably about 15 inches, give or take.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/16/2019 11:20 pm
What are the approximate dimensions of the Raptor engine manifold?

Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?

I think main injector to nozzle throat contraction is generally far less than 2:1 diameter ratio.
SSME contracts far less than that, looks to be more like about 50% (~2:1 area ratios)
https://upload.wikimedia.org/wikipedia/commons/8/88/SSME_powerhead.jpg
F1 had main injector thurst chamber face diameter of 39inches, and a throat diameter of something like 36inches.
http://heroicrelics.org/info/f-1/f-1-thrust-chamber/f-1-cut-away-thrust-chamber-sm.jpg
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: JohnLloydJones on 02/17/2019 12:32 am
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.

Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.

Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.

Demo video (https://www.youtube.com/watch?v=UJvjIB0rAUs)
A very modern take on the very ancient "lost-wax" process.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rocketguy101 on 02/17/2019 12:47 am
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.

Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.

Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.

Demo video (https://www.youtube.com/watch?v=UJvjIB0rAUs)
A very modern take on the very ancient "lost-wax" process.
It's a shame more people don't take the time to learn about manufacturing processes.  It isn't all blacksmithing anymore!  I guess that is a big positive about SpaceX...they are making manufacturing cool...how many on this site have learned about casting, welding, spin forming, and CNC machining?  BTW a lot job openings in the trades to do this type of work.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/17/2019 02:07 am
What are the approximate dimensions of the Raptor engine manifold?

Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?

I think main injector to nozzle throat contraction is generally far less than 2:1 diameter ratio.
SSME contracts far less than that, looks to be more like about 50% (~2:1 area ratios)
https://upload.wikimedia.org/wikipedia/commons/8/88/SSME_powerhead.jpg
F1 had main injector thurst chamber face diameter of 39inches, and a throat diameter of something like 36inches.
http://heroicrelics.org/info/f-1/f-1-thrust-chamber/f-1-cut-away-thrust-chamber-sm.jpg

- You are right about older NASA engines, they tended towards lower contraction ratios. Vulcain I & II appear to have about 2.5 area ratios.

- SpaceX engines appear to be closer to 4:1. The merlin appears to be about 3.6 from measuring actual photos. Raptor CAD appeared to be about 4. Lets call it somewhere between 3.5 and 4.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/17/2019 02:15 am
Everytime I see one of those fire resistant suits, I think of Homer Simpson. ;^)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: CorvusCorax on 02/18/2019 07:54 am
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?

My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.

I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!


Close, but no cigar. A "full flow" staged combustion engine implies, ALL of the propellant moves through pumps, preburners and turbines, before entering the main combustion chamber.

As you wrote, this turns all the propellant gasseous, but thats not the main thing.

Because all of the propellant moves through the turbines, you essentially get a "low pressure, high flowrate" turbine, as opposed to a " high pressure, low throughput" turbine with the same turbine power - which would be used in a gas generator engine to minimize the amount of propellants wasted.

Even a staged combustion engine like RD180 or SSME - which is not full flow - needs to pump all the propellant with just a fraction of it passing through preburner and turbine. And that means for the same pump horsepower, they need a larger pressure differential in the turbine.

Physics and metallurgy define the maximum pressure in your preburners. Thats the limiting factor. Less pressure drop in the turbines mean more pressure in your main combustion chamber. Which usually gives you a more efficient engine, both in regard to (sea level) ISP and thrust 2 weight metrics.



Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ppb on 02/18/2019 02:44 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?

My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.

I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!


Close, but no cigar. A "full flow" staged combustion engine implies, ALL of the propellant moves through pumps, preburners and turbines, before entering the main combustion chamber.

As you wrote, this turns all the propellant gasseous, but thats not the main thing.

Because all of the propellant moves through the turbines, you essentially get a "low pressure, high flowrate" turbine, as opposed to a " high pressure, low throughput" turbine with the same turbine power - which would be used in a gas generator engine to minimize the amount of propellants wasted.

Even a staged combustion engine like RD180 or SSME - which is not full flow - needs to pump all the propellant with just a fraction of it passing through preburner and turbine. And that means for the same pump horsepower, they need a larger pressure differential in the turbine.

Physics and metallurgy define the maximum pressure in your preburners. Thats the limiting factor. Less pressure drop in the turbines mean more pressure in your main combustion chamber. Which usually gives you a more efficient engine, both in regard to (sea level) ISP and thrust 2 weight metrics.
That's the explanation I was looking for. Thanks CorvusCorax.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: WormPicker959 on 02/18/2019 10:36 pm
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?

Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?

My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.

I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!


Close, but no cigar. A "full flow" staged combustion engine implies, ALL of the propellant moves through pumps, preburners and turbines, before entering the main combustion chamber.

As you wrote, this turns all the propellant gasseous, but thats not the main thing.

Because all of the propellant moves through the turbines, you essentially get a "low pressure, high flowrate" turbine, as opposed to a " high pressure, low throughput" turbine with the same turbine power - which would be used in a gas generator engine to minimize the amount of propellants wasted.

Even a staged combustion engine like RD180 or SSME - which is not full flow - needs to pump all the propellant with just a fraction of it passing through preburner and turbine. And that means for the same pump horsepower, they need a larger pressure differential in the turbine.

Physics and metallurgy define the maximum pressure in your preburners. Thats the limiting factor. Less pressure drop in the turbines mean more pressure in your main combustion chamber. Which usually gives you a more efficient engine, both in regard to (sea level) ISP and thrust 2 weight metrics.

Excellent! Thanks for the clarification/explanation. I edited my post so as not to spread future disinformation, and linked to your explanation. This is why I love this forum!
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: StuffOfInterest on 02/19/2019 11:15 am
Here is a fun and irrelevant fact. I actually started (https://en.wikipedia.org/w/index.php?title=Raptor_(rocket_engine_family)&dir=prev&action=history) the Raptor page on Wikipedia (https://en.wikipedia.org/wiki/Raptor_(rocket_engine_family)) in 2009.  Back then it was thought to be a Hydrogen powered upper stage rather than an engine.  Amazing that after 10 years the name has stayed but the meaning behind it is almost 180 degrees different.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Llian Rhydderch on 02/19/2019 12:27 pm
I actually started (https://en.wikipedia.org/w/index.php?title=Raptor_(rocket_engine_family)&dir=prev&action=history) the Raptor page on Wikipedia (https://en.wikipedia.org/wiki/Raptor_(rocket_engine_family)) in 2009.  Back then it was thought to be a Hydrogen powered upper stage rather than an engine.  Amazing that after 10 years the name has stayed but the meaning behind it is almost 180 degrees different.

And if you look at the article today, with the Wikimedia WhoColor tool (https://en.wikipedia.org/wiki/Wikipedia:Tools#Finding_the_responsible_user) that shows editor contribution to the current "as is" article, about 25 of your words are still in that article.  Other NSF posters like Baldusi, Ian The Pineapple, Mmeijeri, and N2e have also improved that article over the years.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Bogeyman on 02/20/2019 06:11 am
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?

(https://upload.wikimedia.org/wikipedia/commons/thumb/e/e2/Raptor_Engine_Unofficial_Combustion_Scheme.svg)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: hkultala on 02/20/2019 06:53 am
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: 2megs on 02/20/2019 05:24 pm

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?


Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure, and thus the cooling loop's inflow is highest of all. There couldn't be an even higher-pressure path going directly from the last pump stage to the preburner, because then you'd have backpressure through the preburner and into the cooling loop's outflow.

(If I'm wrong, I'd really appreciate a gentle education on why.)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/20/2019 06:45 pm
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?

Could be done, but would (probably) increase necessary turbine temperature and necessary 2nd stage pump pressure to compensate, and would be harder to balance all of the various pressure gains and losses for the two flow paths across wide range of throttle and start up conditions (have to ensure no reversed flows occur, ever).  Also for cooling want maximum flow volume to give minimum temperature rise - though this is not too big a deal.

Keeping the turbine temperatures as low as possible is valuable as increases the usable material strength and so reduces the mass of the various manifolds as well as reducing thermally induced strains for longer life.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lar on 02/20/2019 06:49 pm

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?


Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure, and thus the cooling loop's inflow is highest of all. There couldn't be an even higher-pressure path going directly from the last pump stage to the preburner, because then you'd have backpressure through the preburner and into the cooling loop's outflow.

(If I'm wrong, I'd really appreciate a gentle education on why.)
This confuses me a bit because I thought pumps increase pressure as well as move mass. So I would expect pump inlet pressures to be lower than the output.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Slarty1080 on 02/20/2019 07:04 pm

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?


Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure, and thus the cooling loop's inflow is highest of all. There couldn't be an even higher-pressure path going directly from the last pump stage to the preburner, because then you'd have backpressure through the preburner and into the cooling loop's outflow.

(If I'm wrong, I'd really appreciate a gentle education on why.)
This confuses me a bit because I thought pumps increase pressure as well as move mass. So I would expect pump inlet pressures to be lower than the output.

True, the turbopump increases the pressure, but after that its down hill all the way - every step beyond the turbopump is a pressure decrease.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Davidthefat on 02/20/2019 07:12 pm
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?



It is possible to tap the coolant off the first stage, run it through the regen and dump to the low pressure side of the first stage.

edit: the reason to do that is to reduce the delta P between the cooling channels and thrust chamber. There is an incentive to make the liner as thin as possible for weight and heat transfer reasons. Increasing that delta P drives up the thickness of the liner for structural reasons, driving up coolant flow requirements (if they are putting the full flow of CH4, they are set) and weight. If that weight offsets the extra plumbing? That's the SpX engineers' jobs to figure out.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tulse on 02/20/2019 08:00 pm
Does the cooling loop have to be roughly at that pressure to stay liquid while cooling the nozzle?  Would the heat otherwise vaporize the methane?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Bananas_on_Mars on 02/20/2019 08:08 pm
Does the cooling loop have to be roughly at that pressure to stay liquid while cooling the nozzle?  Would the heat otherwise vaporize the methane?
Methane should turn supercritical in the cooling loop, at the high pressures they have there's no phase change between liquid and supercritical fluid IMO. So no explicit boiling, the fluid just gets hotter and more compressible.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: hkultala on 02/20/2019 08:30 pm

The methane pump has multiple stages.  Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?


Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure

It seems you totally misunderstood what I was proposing.

No, if there is another pump stage after the cooling loop, increasing the pressure.

First pump would stage increases the pressure to something like 300 bars, then there would be the cooling loop at about 300 bars, then the second pump stage would increase the pressure to about 600 bars for the preburners.

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/20/2019 09:50 pm
I think you can not push the supercritical hot fuel in the pump; cool fuel flow is there for keeping whole prebuner/turbopump assembly cool. Both SSME (410 bar) and RD-180 have similar setup, where fuel goes right from the last stage of pump to the cooling jacket*.

*actually it's bit more complicated:
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Hominans Kosmos on 02/20/2019 10:51 pm
Doesn't the regenerative cooling loop have to significantly exceed main combustion channel pressure for purposes of film-cooling to begin with? 

What of the pressure-stiffening potential for having the walls inflated, strengthening against acoustic stresses? 

Add on top of it lower pump temperatures, is it really that bad an idea to have the coolant loop after all main pump stages?

Edit: counter-point: I have no idea how the layers of the engine wall are supposed to remain joined in the face of such pressures, but the middle point above assumes such powerful welding is possible and practical.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Hitech on 02/20/2019 11:12 pm
There might be some confusion with the term "high pressure" in the turbine area. A better term would be "high pressure ratio low flow" turbine vs "low pressure ratio high flow".
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: kendalla59 on 02/21/2019 12:48 am
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.

Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.

Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.

You can also print positives in wax, compatible with traditional IC methods.
3DS video (https://www.youtube.com/watch?v=MhpU8FtmpAs)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: hkultala on 02/21/2019 09:29 am
Doesn't the regenerative cooling loop have to significantly exceed main combustion channel pressure for purposes of film-cooling to begin with? 

The chamber and nozzle are cooled regeneratively, NOT by film cooling. These are two totally different things.

For regenerative cooling, there should be no need for high pressure for the cooling loop.

Quote
Add on top of it lower pump temperatures, is it really that bad an idea to have the coolant loop after all main pump stages?

The pump temperature is a good point.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: jwade on 02/21/2019 10:24 am
Upthread, people were speculating the the raptor used regen plus film cooling for the nozzle.  This was based on the appearance of the exhaust jet from the picture of the second raptor test ( as tweeted by Elon Musk).   The assumption was that they had increased the film cooling flow to address the copper chamber erosion seen in the video of the first test.    As far as I know there is no concrete information that raptor uses film cooling, but it is plausible
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: tyrred on 02/21/2019 10:27 am
Thought exercise... What are the ramifications of using regenerative cooling + film cooling?

*Not necessarily advocating that film cooling is actually being used.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: rakaydos on 02/21/2019 10:41 am
Thought exercise... What are the ramifications of using regenerative cooling + film cooling?

*Not necessarily advocating that film cooling is actually being used.
Basically the same as the heat shield- the film is too thin to stop radiative heating, but it can minimize convective heating, while the regenerative cooling handles the radiative heating from the plasma fireball it's containing.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/21/2019 10:56 am
Like was said before in this thread, all the modern staged combustion engines use both regenerative and film cooling, but they are separated systems. Coolant (fuel) for the film cooling is tapped off before the regenerative cooling loop in RD-180.

[1] Space Transportation System, Training Data, Space Shuttle Main Engine Orientation
[2] Atlas V Launch Services User’s Guide
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/21/2019 11:28 am
Elon has tweeted a response to the Roscosmos criticism of his claims about Raptor pressure record:

Quote from: @macodieas
The point RosKosmos is trying to make is that in your tweet you compared a liquid/gas engine to a gas/gas engine, and that the operating parameters of those two are not comparable. (Just translating the article, I definitely don't know enough to make a judgement.)

https://twitter.com/elonmusk/status/1098536496494108673

Quote
Not true. Limiting factor in any staged combustion rocket engine, liquid/gas or gas/gas, is pressure & temperature in oxygen preburner
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 12:31 pm
Like was said before in this thread, all the modern staged combustion engines use both regenerative and film cooling, but they are separated systems. Coolant (fuel) for the film cooling is tapped off before the regenerative cooling loop in RD-180.

[1] Space Transportation System, Training Data, Space Shuttle Main Engine Orientation
[2] Atlas V Launch Services User’s Guide

You can see the effect of this film cooling in the dark "curtains" coming from below the engine bell on the RD-180. It's the same effect that the turbine exhaust cooling had for the F1, but on a much, much smaller scale.

RD-180
(https://i.pinimg.com/originals/7f/2c/47/7f2c47101a27d89af464feb4d394ef76.jpg)

F-1
(http://www.collectspace.com/images/news-032812a.jpg)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/21/2019 03:33 pm
Interesting Merlin vs Raptor comparison by Elon:

https://twitter.com/elonmusk/status/1098613993176850432

Quote
SpaceX Merlin architecture is simpler than staged combustion (eg SSME or RD), but it has world record for thrust/weight & thrust/cost engine. Raptor has better Isp, but I’m worried it may fall short on those two critical metrics.

In May last year Tom Mueller seemed to be expecting a different outcome:

https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/21/2019 04:01 pm
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?


Don't think so. If cooling passage pressure is 640 bar, slots are 3mm wide, and allowable stress is 70 MPa, the needed wall thickness is about 1.42 mm.

If the main fuel line coming out of the pump is 76 mm diam, pressure 640 bar, allowable stress 700 MPa, the needed wall thickness is about  3.6 mm.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Yellowstone10 on 02/21/2019 05:45 pm
https://twitter.com/elonmusk/status/1098653939141009408

Godspeed, Raptor SN1...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 05:54 pm
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.

- Added pressures, temperatures and pump powers to engine schematic.

- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.

- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power

- At 320 bar I get  2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.

Added some more flow detail.

John

Something I've been thinking about; perhaps the "full thrust" raptor drops the ER slightly while increasing chamber pressure?

how much would ISP change going from 270-280 bar with and ER of 35 (~0.22m throat) to ~320 bar with an ER of 32 (~0.23m throat)? Those numbers should get you 250 tonnes of thrust too.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: FutureSpaceTourist on 02/21/2019 06:07 pm
Quote from: @gavbrowne
Are there component differences between SN1 and SN2 that would prevent damage?

https://twitter.com/elonmusk/status/1098656849740451840

Quote
SN2 has changes that should help
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/21/2019 06:43 pm
https://twitter.com/elonmusk/status/1098653939141009408

Godspeed, Raptor SN1...

Good way of running final part of a complex mechanical design optimisation, run near limits and see what gets damaged and needs tweaking.  Simulation only gets you so far (though it is improving all the time)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 06:47 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: TorenAltair on 02/21/2019 07:26 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

Insane :D In case you're bored, how about adding an F1-engine?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/21/2019 07:32 pm
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.

- Added pressures, temperatures and pump powers to engine schematic.

- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.

- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power

- At 320 bar I get  2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.

Added some more flow detail.

John

Something I've been thinking about; perhaps the "full thrust" raptor drops the ER slightly while increasing chamber pressure?

how much would ISP change going from 270-280 bar with and ER of 35 (~0.22m throat) to ~320 bar with an ER of 32 (~0.23m throat)? Those numbers should get you 250 tonnes of thrust too.

Note, ER limited by 1.3 m exit diameter constraint.

pc (bar)    Dthroat      ER       IspSL (sec)       IspVac (sec)     Thrust SL (MN)      Thrust Vac (MN)
270           .22            35        332                   356                   1.84                    1.97
270           .22          120        ----                   375                    ----                     2.08
320           .23            32        335                   354                   2.41                    2.54 
320           .23          120        ----                   375                    ----                     2.70

Would have to reduce ER to 32.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nilof on 02/21/2019 07:33 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/21/2019 07:39 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.

Big engines? He hasn't drawn the F1 yet.  D=3.7m ;^)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 07:42 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.

Big engines? He hasn't drawn the F1 yet.  ;^)

New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

Insane :D In case you're bored, how about adding an F1-engine?

Not enough room left to fit it on the paper...  ;D
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 07:45 pm
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.

- Added pressures, temperatures and pump powers to engine schematic.

- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.

- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power

- At 320 bar I get  2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.

Added some more flow detail.

John

Something I've been thinking about; perhaps the "full thrust" raptor drops the ER slightly while increasing chamber pressure?

how much would ISP change going from 270-280 bar with and ER of 35 (~0.22m throat) to ~320 bar with an ER of 32 (~0.23m throat)? Those numbers should get you 250 tonnes of thrust too.

Note, ER limited by 1.3 m exit diameter constraint.

pc (bar)    Dthroat      ER       IspSL (sec)       IspVac (sec)     Thrust SL (MN)      Thrust Vac (MN)
270           .22            35        332                   356                   1.84                    1.97
270           .22          120        ----                   375                    ----                     2.08
320           .23            32        335                   354                   2.41                    2.54 
320           .23          120        ----                   375                    ----                     2.70

Would have to reduce ER to 32.

That would be a monster engine....I'm thinking more and more then that the "Full Thrust" version is going to drop the ER slightly, and have 310-320 bar of pressure.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/21/2019 07:47 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)
Vector graphics:
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 07:51 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)
Vector graphics:

She cleans up nice  ;)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Oersted on 02/21/2019 08:23 pm
Interesting Merlin vs Raptor comparison by Elon:

https://twitter.com/elonmusk/status/1098613993176850432

Quote
SpaceX Merlin architecture is simpler than staged combustion (eg SSME or RD), but it has world record for thrust/weight & thrust/cost engine. Raptor has better Isp, but I’m worried it may fall short on those two critical metrics.

In May last year Tom Mueller seemed to be expecting a different outcome:

https://twitter.com/spacecom/status/999691403172036608

Quote
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018

It is worrying to read that Elon is worried!  :-O
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/21/2019 08:33 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.

Big engines? He hasn't drawn the F1 yet.  ;^)

New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

Insane :D In case you're bored, how about adding an F1-engine?

Not enough room left to fit it on the paper...  ;D
I removed copyrighted parts and added NASA Public Domain F-1...
(I hope I get scale right; height 5.8 m )
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: cuddihy on 02/21/2019 08:33 pm
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”

It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.

Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: envy887 on 02/21/2019 08:38 pm
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”

It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.

Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.

Maybe. But there is a reason to find out what it can't do early on, rather than taking a lot of time characterizing what it can do: if it ultimately falls short, you've wasted that time.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/21/2019 08:49 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.

It's also a sustainer stage, hence the big nozzle. The thrust chamber itself is only about a meter in diameter (don't quote me on that, just eye balling the hardware)

It's actually much less than that, there is just a lot of equipment surrounding it. The throat is about 0.29m, the entire combustion chamber is probably ~0.5m wide.

(https://external-preview.redd.it/Hzz7Ej0gg_esYGhY97XmqTjx-miy4LBfxfk51gMOycM.jpg?width=1200&height=628.272251309&auto=webp&s=9e92e26e21772c0c369d34746a344aac804be7da)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Davidthefat on 02/21/2019 08:50 pm
New Drawing comparison:

BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure  8)

That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.

RS-25 is a sustainer stage, meaning the nozzle is expanded out much higher than the boost stage engines.

Having stood next to the thrust chamber it's not all that much bigger than a merlin's. F1 on the other hand...


Edit: the RS 68 is a big boy too.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: abaddon on 02/21/2019 10:58 pm
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”

It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.
From a previous tweet:

https://twitter.com/elonmusk/status/1094790663646760961

Quote
Propellant was not deep cryo. CH4 & O2 were just barely below liquid temp at 1 bar. In theory, Raptor should do ~300 bar at deep cryo, provided everything holds together, which is far from certain. However, only 250 bar is needed for nominal operation of Starship/Super Heavy.

We can imagine that the pressure was higher than the previously announced level, although how close it was to 300 we do not know.  It is plausible that they did in fact push it to see how far it could go, knowing that there was a good chance it would break.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: philw1776 on 02/21/2019 11:13 pm
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”

It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.

Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.

Huh?  Thrust is already higher than needed to toss 100T to LEO running the rocket equation and assuming a heavy Starship even though Elon said it's lighter.  And if something yet unexpected falls short, it's 80 or 90 tonnes worst case each launch.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: TomH on 02/21/2019 11:40 pm
I removed copyrighted parts and added NASA Public Domain F-1...

Who remembers the M-1? (Thanks to Blackstar for posting this awhile back). And no, that M does not stand for Merlin.

(https://forum.nasaspaceflight.com/index.php?action=dlattach;topic=31401.0;attach=503435;image)

(http://www.astronautix.com/graphics/s/satupeng.gif)
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ThePonjaX on 02/22/2019 02:09 am
I'm sorry to say that. You know a lot about rocket engines guys but it's really hard to follow a thread about the Raptor and the news about it when it's full of diagrams which a lot of cases aren't related to the Raptor.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: HVM on 02/22/2019 02:42 am
I'm sorry to say that. You know a lot about rocket engines guys but it's really hard to follow a thread about the Raptor and the news about it when it's full of diagrams which a lot of cases aren't related to the Raptor.

Tom Müller and other Elven-smiths are forging Twi~Essɛn right now, but it takes some time before that bird of prey expels its fire from its gullet and out of its mouth... So we need to entertain ourselves some how.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/22/2019 03:40 am
I'm sorry to say that. You know a lot about rocket engines guys but it's really hard to follow a thread about the Raptor and the news about it when it's full of diagrams which a lot of cases aren't related to the Raptor.

Tom Müller and other Elven-smiths are forging Twi~Essɛn right now, but it takes some time before that bird of prey expels its fire from its gullet and out of its mouth... So we need to entertain ourselves some how.

Yes, but the valid point is that general engine discussion does not need to happen in THIS thread. Show some consideration. When people subscribe to to threads they get emails for every reply.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Okie_Steve on 02/22/2019 05:22 am
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”

It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.

Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.

Maybe. But there is a reason to find out what it can't do early on, rather than taking a lot of time characterizing what it can do: if it ultimately falls short, you've wasted that time.
My guess would be they learned something important on the first runs that immediately got incorparated into sn2 and would make further testing of that part of the design less useful so they decided to stress some other parts hard since sn2 is coming soon
Title: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Lars-J on 02/22/2019 06:27 am
With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.

There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.

I'm sorry, but what do you base that on? What have we been told to indicate a significant ISP shortfall? What observations support this?

Because if you can't source that, the rest of your lengthy post is even less based in reality.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Stan-1967 on 02/22/2019 07:15 am

With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.

There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.

There is zero chance that SpaceX was expecting an ISP of 380 for this engine & expansion ratio.  The target would be more 340 ish.  Chamber pressure is more key in getting the T/W they want, they are already pretty far up the ISP vs. chamber pressure curve above 200 bar so that ISP is more incremental.  Others correct me if I am wrong please, but the ISP is not linear to chamber pressure, whereas thrust is more proportional to chamber pressure in that is is the first order determining factor in the mass flow rate.

Way too much concern for such limited information.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: niwax on 02/22/2019 07:50 am

With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.

There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.

There is zero chance that SpaceX was expecting an ISP of 380 for this engine & expansion ratio.  The target would be more 340 ish.  Chamber pressure is more key in getting the T/W they want, they are already pretty far up the ISP vs. chamber pressure curve above 200 bar so that ISP is more incremental.  Others correct me if I am wrong please, but the ISP is not linear to chamber pressure, whereas thrust is more proportional to chamber pressure in that is is the first order determining factor in the mass flow rate.

Way too much concern for such limited information.

Their stated targets for a future production iteration of Raptor were 330s SL and 380s vac. But they have made it clear that the vacuum engine is some ways off, and looking at Merlin Vac with its 140+ ER it will be significantly different from the thing currently on the test stand. The current first version of Starship has also explicitly been designed to have SL engines on the first stage for now, so I don't really see the issue.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: uhuznaa on 02/22/2019 09:46 am

With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.

There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.

There is zero chance that SpaceX was expecting an ISP of 380 for this engine & expansion ratio.  The target would be more 340 ish.  Chamber pressure is more key in getting the T/W they want, they are already pretty far up the ISP vs. chamber pressure curve above 200 bar so that ISP is more incremental.  Others correct me if I am wrong please, but the ISP is not linear to chamber pressure, whereas thrust is more proportional to chamber pressure in that is is the first order determining factor in the mass flow rate.

Way too much concern for such limited information.

I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: 50_Caliber on 02/22/2019 11:26 am
I wonder how fast can SpaceX iterate a design? Could they have tested version 1.0 then incorporate design changes into the next one and be testing version 2.0 the next week or two?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Chris Bergin on 02/22/2019 12:08 pm
Going to let this go to 100 pages - just because it's a great thread - and then move to the "Thread 2" on this, so that'll be this weekend.

Remember, massive threads can go a bit off topic at times, so let's keep it focused on Raptor. (A few report to mods - letting it slide per the next thread incoming).
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: sevenperforce on 02/22/2019 05:18 pm
I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.
No unburned fuel is dumped.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: OxCartMark on 02/22/2019 05:39 pm
I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.
No unburned fuel is dumped.

?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Davidthefat on 02/22/2019 05:42 pm
I wouldn't be surprised if they film cooled in the chamber near the injector.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/22/2019 07:20 pm
I wonder how fast can SpaceX iterate a design? Could they have tested version 1.0 then incorporate design changes into the next one and be testing version 2.0 the next week or two?
Super allow foundry allows for rapid iteration of Raptor. So it is entirely possible that SpaceX have made changes to Raptor SN2 based on data from SN1 firings. Modern CAD/CAM + 3D printing of mold patterns for cast parts can allow for very rapid iteration of designs.

EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Tuna-Fish on 02/22/2019 07:39 pm
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.

They can't do that unless they also make the upper stage a lot heavier. The upper limit on Raptor size is set by the requirements to be able to hover and to be able to safely land even if they lose 2 engines during landing. This means that the thrust of the raptor at min safe throttle at sea level cannot be more than 1/3rd of the mass of the upper stage when near-empty.

They might relax those requirements a bit for the initial, unmanned versions, but I really don't see them making raptors bigger than the current design without also going to a larger diameter craft. Besides, Merlins are the champions of thrust/cost already, even with Raptors being a little worse than that, so long as they are reusable enough times, they should be fine.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/22/2019 07:52 pm
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.

They can't do that unless they also make the upper stage a lot heavier. The upper limit on Raptor size is set by the requirements to be able to hover and to be able to safely land even if they lose 2 engines during landing. This means that the thrust of the raptor at min safe throttle at sea level cannot be more than 1/3rd of the mass of the upper stage when near-empty.

They might relax those requirements a bit for the initial, unmanned versions, but I really don't see them making raptors bigger than the current design without also going to a larger diameter craft. Besides, Merlins are the champions of thrust/cost already, even with Raptors being a little worse than that, so long as they are reusable enough times, they should be fine.

Elons wording also seems to suggest that it will still be close though.

A 300 bar FFSC engine with 330+ sl ISP and and 'only' the second best TWR and cost/thrust ratio is still an engine that is so much better than all of the competition it's almost crazy... Merlins are insanely cheap by industry standards (under $1m).

If I had to guess Raptor is probably ~1200kg and costs a little over $3m each... I bet they still have a good chance of beating the TWR figures with the "full thrust" Raptor. Other LV providers would probably kill to have such an engine at a little over $3m.

Cost/thrust is probably out of reach just for the FFSC architecture.

The Raptor engine really is almost a work of art...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: punder on 02/22/2019 07:54 pm
A few days ago it was all "OMG look how fast they're going!!" and now it's all doom-n-gloom.

This is the very first full-scale, flight-worthy Raptor, in its very first series of firings, correct?

A quick review of previous engine development efforts might provide some perspective.

Okay off my high horse. Just sayin.   :D

Mr. Obvious

WOW my 500th post! An extra Glenmorangie tonight.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: DJPledger on 02/22/2019 08:01 pm
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.

They can't do that unless they also make the upper stage a lot heavier. The upper limit on Raptor size is set by the requirements to be able to hover and to be able to safely land even if they lose 2 engines during landing. This means that the thrust of the raptor at min safe throttle at sea level cannot be more than 1/3rd of the mass of the upper stage when near-empty.

They might relax those requirements a bit for the initial, unmanned versions, but I really don't see them making raptors bigger than the current design without also going to a larger diameter craft. Besides, Merlins are the champions of thrust/cost already, even with Raptors being a little worse than that, so long as they are reusable enough times, they should be fine.
Future iterations of Raptor could have even deeper throttle capability which could allow the no. of Raptors on SS to be reduced from 7 to 5 which would allow Raptor to be sized for 19 on SH. Also 19 is one of the nos. that gives max. packing density on SH while 31 is not. OTOH, SpaceX could make Raptor smaller for 9 on SS and 37 on SH which would also give max. packing density on SH. Still think that smaller Raptor may cost more than larger one for entire SS/SH propulsion system.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RoboGoofers on 02/22/2019 08:33 pm
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.
he is saying here that Raptor might not be a record-breaker, not that it won't meet its design goals.
https://twitter.com/elonmusk/status/1098613993176850432
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: livingjw on 02/22/2019 09:43 pm
I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.
No unburned fuel is dumped.

Fuel film cooling is typically very small fraction  of total flow and is introduced around the main chamber at the injector plane. It is also used at the throat. The cycle is fuel rich, so of course unburnt fuel is expelled, but mostly in the form of H2 and CO.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: spacenut on 02/22/2019 10:02 pm
Even if it didn't meet it's thrust/weight and thrust/cost goals, being reusable, will save money in the long run.  Raptor also will not have to be cleaned or rebuilt as often as Merlin. 
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: RobLynn on 02/23/2019 03:01 am
Superheavy would work fine with Merlin, or a modified Merlin burning Methalox if that turned out to be more cost effective (and at this stage who knows, it might, as per-flight engine depreciation is quite possibly the dominant cost for the whole SS/SH system).

Regarding film cooling: methane has low molecular mass and high specific heat.  At point it enters the chamber it is already at 700-800K and even it it only stays at that temperature without combusting with LOX will expand and be ejected with a velocity of nearly 2500m/s (Isp 250s).  In reality it will be heated further by radiation, so that no-oxygen thin film region next to the wall costs very little in reduced average Isp.

Hydrogen film cooling is even better- it can potentially have higher Isp than the main combusted flow.
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Coastal Ron on 02/23/2019 03:35 am
Superheavy would work fine with Merlin...

Something LIKE the CH4/LOX powered could be built that uses the RP-1/LOX Merlin engines, but I don't know why SpaceX would want to make such a radical change at this point.

Quote
...or a modified Merlin burning Methalox if that turned out to be more cost effective...

I'm not a rocket engineer, but from what I do understand you can't simply swap fuels on rocket engines because they are optimized for a specific fuel.

I don't see any need for SpaceX to change away from Raptor and methane at this point...
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: Nilof on 02/23/2019 08:21 am
I'm mostly curious to see whether the Raptor will see continued improvements over the year like the Merlin did.

Let's imagine the case where they managed to uprate the chamber pressure by some absolutely crazy factor to say, 35-40 MPa. What would the most likely change to the design be? A tank stretch like on the Falcon to increase payload, or maybe reducing the number of engines to match the increased thrust, and using the extra space to increase the expansion ratio to squeeze out more Isp?
Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: CorvusCorax on 02/23/2019 04:05 pm
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”

It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.

Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.

Maybe. But there is a reason to find out what it can't do early on, rather than taking a lot of time characterizing what it can do: if it ultimately falls short, you've wasted that time.
My guess would be they learned something important on the first runs that immediately got incorparated into sn2 and would make further testing of that part of the design less useful so they decided to stress some other parts hard since sn2 is coming soon

That!  I think, as seen in the first videos posted by Elon, the engine had a flaw. Maybe a hotspot at the throat or in the chamber with insufficient regen cooling. Sonething they could fix for SN2, like a a minor cooling channel redesign or a slight change in geometry.

In the interim, they worked around the issue with excessive "film" cooling. To the point you could see the flaming flow of burning Methane in the exhaust. This amount likely eats into their ISP and wouldnt be practical for a flight engine, but it allowed them to continue the test campaign and rattle the engine to see what else needs fixing.

Apoarently nothing else came loose, so they rattled harder. Got a neat chamber pressure miledtone, rattled even harder ( see L2 ) until they found something that breaks, which then gets incorporated into the next design change.

You wouldn't stop testing when you find the first small issue with the first full size prototype. If you did, iteration would take far too long. You rather work around it and try to learn as much as you possibly can from the thing.

To quote Elon Musk ( in the context of grasshopper, but it applies here )  " If it didn't blow up, you did not test hard enoigh"

Title: Re: The evolution of the SpaceX Raptor engine (Super Heavy/Starship Propulsion)
Post by: ZachF on 02/23/2019 04:14 pm
Superheavy would work fine with Merlin, or a modified Merlin burning Methalox if that turned out to be more cost effective (and at this stage who knows, it might, as per-flight engine depreciation is quite possibly the dominant cost for the whole SS/SH system).

Regarding film cooling: methane has low molecular mass and high specific heat.  At point it enters the chamber it is already at 700-800K and even it it only stays at that temperature without combusting with LOX will expand and be ejected with a velocity of nearly 2500m/s (Isp 250s).  In reality it will be heated further by radiation, so that no-oxygen thin film region next to the wall costs very little in reduced average Isp.

Hydrogen film cooling is even better- it can potentially have higher Isp than the main combusted flow.

It would "work" but there is much more than just Thrust per $.

Raptor's potent combo of TWR, T/$, ISP, Impulse density, and cheap fuel cost is what makes it the best engine out there, IMHO.

Let's do a little thought exercise:
------------------------------------------------------------------------------

Suppose you want to switch out Raptors for GG "Methamerlin" engines, How would that impact the system as a whole?

A gas generator Merlin-style engine could probably get an ISP of 300/320 on a sea-level version, and ~358 in a vacuum version. So lets plug these in to see what we'd need to match the current design, but first a few points on how these numbers were derived:

(See attachment)

-The landing dVs were derived from the original 2016 ITS slides. The original required 7% of it's fuel, so I calculated the dV given from that and set it as the landing load. Landing fuel for the upper stage was taken from measurements of the landing tanks from the 2016 iteration an dV calculated.

-Dry masses are taken from the 2016 and 2017 proposals. On a percentage basis. I included a little mass increase on the upper stage as well.

-Lower stage was stretched some in latest proposal. Estimated here.

These numbers are not perfect but a good enough for the purposes of showing the comparison

The top three are, left to right:
-First iteration with all SL Raptors
-GG "Methamerlin" lower stage with SL Raptor upper stage... doesn't make much sense, but I put it there for comparison's sake.
-GG Methamerlins on both the upper and lower stage

Bottom four have vacuum engines and a larger payload, left to right:
-Raptors with Raptor vacs
-GG Methamerlins with Raptor vacs
-GG Methamerlins with Methamerlin vac upper stage
-Kerolox Merlin lower stage with vacuum raptors on upper stage


You'll notice in the non-Raptor iterations mass increases massively. Lower ISP also means a larger fraction of the fuel must be held as a landing reserve. 

Some findings:
------------------------------------

-Switching out Raptors with GG Methamerlins on the lower stage increases the mass of the lower stage by 37%. It's likely that switching to GG engines would require the mixture ratio to be dropped from ~3.7:1 to ~3:1, reducing fuel density by ~5% and increasing volume by a larger factor than mass. Mass of the stack as a whole increases by 28%, thus you need 1.28x thrust if you wanted to switch out Raptors for GG Methalox engines. If you kept Raptor Vacs for the upper stage but used Methamerlins on the lower stage you'd no longer have common engines and the cost for both would rise, probably by at least 20%. So Raptor could cost >50% more in Thrust/$ and still come out on top in this scenario... But that does not factor in the >37% increase in the lower stage mass which would increase the fabrication costs, with those Raptor could cost much much more and still come out on top.

-Switching out Raptors for GG engines on both the upper and lower stages increases the mass of the stack as a whole by 60% when comparing the vacuum-engined versions. Raptor could thus cost >60% more than the GG Methamerlin and come out on top. Again, this does not factor in the increased cost from dry mass increasing by over 60%. So Raptor could again cost much more in thrust/$ and still win out big time on a systemic basis. Also, in this version, because of the poorer ISP in the vacuum Methamerlins, it now takes 30% more tanker flights to refuel in orbit.

-Switching out a Raptor powered lower stage with a Kerolox Merlin-powered lower stage increases the size of the lower stage by over 50%. Dry mass would increase by a similar amount. So again, Raptor could again cost much more in thrust/$ and still win out big time on a systemic basis. Also, on this proposal, fuel costs per launch go from about $500k for the Methane, to $2+ million for 5,140 tonnes of kerolox fuel.


Long story short, Raptor can cost much more on a thrust/$ basis but it will still win on a systemic basis. If it is even close to Merlin it is a no-brainer to go with Raptor.

(EDIT- noticed I used vacuum figures for the SL Methamerlins, so the numbers are actually worse...)
Title: Re: SpaceX Raptor engine (Super Heavy/Starship Propulsion) - General Thread 1
Post by: Chris Bergin on 02/23/2019 04:28 pm
As promised, thread 2 now live:

https://forum.nasaspaceflight.com/index.php?topic=47506.0