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SpaceX Vehicles and Missions => SpaceX BFR - Earth to Deep Space => Topic started by: Chris Bergin on 10/03/2016 02:28 PM

Title: ITS Propulsion – The evolution of the SpaceX Raptor engine
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 ฐC. 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 Hastelloyฎ 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 Hastelloyฎ 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 Hastelloyฎ 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 Hastelloyฎ 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 Hastelloyฎ 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 Hastelloyฎ 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 Hastelloyฎ 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 Hastelloyฎ 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: TheSpaceRod 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://www.youtube.com/watch?v=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://www.youtube.com/watch?v=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 0.87mุ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mุ 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 0.87mุ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mุ 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 0.87mุ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mุ 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 200-250ฐC 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 200-250ฐC 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://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s (https://www.youtube.com/watch?v=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