Author Topic: ITS Propulsion – The evolution of the SpaceX Raptor engine  (Read 25568 times)


Offline AndyX

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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.

Offline Dante80

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That was a terrific article, many thanks for that!!

Offline Dante80

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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
« Last Edit: 10/03/2016 03:08 PM by Dante80 »

Offline cro-magnon gramps

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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...
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Offline Mongo62

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"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.
« Last Edit: 10/03/2016 03:11 PM by Mongo62 »

Offline Dante80

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"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.
 
« Last Edit: 10/03/2016 03:13 PM by Dante80 »

Offline john smith 19

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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.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline Norm38

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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)
« Last Edit: 10/03/2016 03:36 PM by Norm38 »

Offline F9man

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Very exciting. Can't wait to meet a raptor in person

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #10 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.

Offline clongton

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #11 on: 10/03/2016 03:49 PM »
Awesome write-up. Thank you
Chuck - DIRECT co-founder
I started my career on the Saturn-V F-1A engine

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #12 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.

Online Lars-J

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #13 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.

Offline matthewkantar

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #14 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

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #15 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.

Offline Kansan52

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #16 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!

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #17 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?
« Last Edit: 10/03/2016 05:02 PM by Dante80 »

Offline rsdavis9

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #18 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.
bob

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #19 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.

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #20 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

Offline ellindsey

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #21 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.

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #22 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.

Offline DJPledger

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #23 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.

Offline John Alan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #24 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)

Offline MAC74

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #25 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.

Offline RedLineTrain

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #26 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/
« Last Edit: 10/03/2016 06:13 PM by RedLineTrain »

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #27 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)?


Offline matthewkantar

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #28 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

Offline DJPledger

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #29 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.

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #30 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.

Offline AncientU

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #31 on: 10/03/2016 07:51 PM »
Nice article, Baldusi (by the way)
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Offline mheney

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #32 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.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #33 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

Offline MAC74

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #34 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."

Offline Rocket Science

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #35 on: 10/03/2016 08:50 PM »
Great work on the article Alejandro, thank you! :)
“All engineering experiments generate valuable data, the failures are the ones that yield the most”
Rob

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #36 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.
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.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #37 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.
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.

Offline MATTBLAK

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #38 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?
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Offline MikeAtkinson

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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.

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
« Last Edit: 10/03/2016 09:49 PM by MikeAtkinson »

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #40 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.

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 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.
« Last Edit: 10/03/2016 10:01 PM by dglow »

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #41 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.

Offline SirKeplan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #42 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.

Offline Nilof

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #43 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...
« Last Edit: 10/03/2016 10:20 PM by Nilof »
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #44 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?

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #45 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?
« Last Edit: 10/03/2016 10:26 PM by dglow »

Online Lars-J

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #46 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.  :)

Offline kch

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #47 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.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #48 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.
« Last Edit: 10/03/2016 10:59 PM by dglow »

Offline Toast

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.

Offline wardy89

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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!
« Last Edit: 10/03/2016 11:23 PM by wardy89 »

Offline AS-503

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #51 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).

Offline Elmar Moelzer

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #52 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.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #53 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.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #54 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.
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Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #55 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.

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #56 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.
« Last Edit: 10/04/2016 01:32 AM by Dante80 »

Offline Elmar Moelzer

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #57 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.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #58 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.  :)

Online FutureSpaceTourist

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #59 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!

Offline Elmar Moelzer

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #60 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.

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #61 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
« Last Edit: 10/19/2016 11:22 PM by livingjw »

Online hkultala

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #62 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.

Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #63 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?

Offline ArbitraryConstant

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #64 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.



Offline Nomic

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #65 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. 


Offline Kaputnik

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #66 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.
Waiting for joy and raptor

Online hkultala

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #67 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.

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #68 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?
« Last Edit: 10/04/2016 11:55 AM by Dante80 »

Offline Silversheep2011

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #69 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?


see 1:37 to 2:31 that makes the S.L. raptors that little bit more efficient in the vacuum of outer space?

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #70 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.
« Last Edit: 10/04/2016 12:28 PM by Dante80 »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #71 on: 10/04/2016 02:08 PM »
Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #72 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...;)
« Last Edit: 10/04/2016 02:55 PM by Dante80 »

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #73 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.
« Last Edit: 10/04/2016 06:01 PM by baldusi »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #74 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
« Last Edit: 10/04/2016 06:00 PM by livingjw »

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #75 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.

Offline John Alan

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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)





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)

« Last Edit: 10/04/2016 08:55 PM by John Alan »

Offline john smith 19

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #77 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.
« Last Edit: 10/05/2016 08:46 AM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline Elmar Moelzer

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #78 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?

Offline Nathan2go

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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.
« Last Edit: 10/05/2016 02:40 AM by Nathan2go »

Online TrueBlueWitt

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #80 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.

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #81 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. 

« Last Edit: 10/05/2016 06:50 AM by Dante80 »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #82 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)
« Last Edit: 10/16/2016 01:22 PM by livingjw »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #83 on: 10/06/2016 04:11 PM »
I updated reply #61 on this thread to the correct my estimated size of the engine.

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #84 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).
« Last Edit: 10/07/2016 06:40 AM by Dante80 »

Offline AP3

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #85 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)

Offline SirKeplan

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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?
« Last Edit: 10/07/2016 09:54 PM by SirKeplan »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #87 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

Online wannamoonbase

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #88 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.
« Last Edit: 10/18/2016 07:08 PM by wannamoonbase »
I know they don't need it, but Crossfeed would be super cool.

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #89 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.

Offline RedLineTrain

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #90 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
« Last Edit: 10/24/2016 04:08 PM by RedLineTrain »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #91 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.

Online Prettz

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #92 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.

Offline rsdavis9

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #93 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
bob

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #94 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?

Offline Llian Rhydderch

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #95 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.

Re arguments from authority on NSF:  "no one is exempt from error, and errors of authority are usually the worst kind.  Taking your word for things without question is no different than a bracket design not being tested because the designer was an old hand."
"You would actually save yourself time and effort if you were to use evidence and logic to make your points instead of wrapping yourself in the royal mantle of authority.  The approach only works on sheep, not inquisitive, intelligent people."

Online wannamoonbase

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #96 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.
I know they don't need it, but Crossfeed would be super cool.

Offline rsdavis9

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #97 on: 11/03/2016 06:42 PM »
I'm super excited. But as you said no info to work with.
bob

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #98 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.

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #99 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???

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #100 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.

Offline jpo234

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #101 on: 01/09/2017 02:10 PM »
Are there any updates about Raptor development after the September test?

Offline philw1776

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #102 on: 01/10/2017 07:30 PM »
Are there any updates about Raptor development after the September test?

None about development or test
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Online FutureSpaceTourist

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #103 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
« Last Edit: 01/13/2017 10:40 AM by FutureSpaceTourist »

Offline rockets4life97

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #104 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?

Offline envy887

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #105 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.

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #106 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 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 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 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, 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, 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.
« Last Edit: 03/16/2017 12:09 AM by Manabu »

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #107 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.

Offline envy887

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #108 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.

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #109 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.
« Last Edit: 03/16/2017 12:12 AM by Manabu »

Offline OneSpeed

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #110 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?

Offline nacnud

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #111 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

Offline AnalogMan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #112 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 |
« Last Edit: 03/15/2017 08:43 PM by AnalogMan »

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #113 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.
« Last Edit: 03/16/2017 12:37 AM by Manabu »

Offline spacenut

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #114 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?

Online macpacheco

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #115 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.
« Last Edit: 03/21/2017 10:20 PM by macpacheco »
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Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #116 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.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #117 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).
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