Author Topic: ITS Propulsion – The evolution of the SpaceX Raptor engine  (Read 45526 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.

Online 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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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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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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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.

Offline Lars-J

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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.

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.

Online matthewkantar

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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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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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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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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

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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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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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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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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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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The 1MN dev. model of Raptor should be mass produced to replace Merlin to do away with the He system on F9 and FH.

Online John Alan

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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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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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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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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)?


Online matthewkantar

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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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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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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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Nice article, Baldusi (by the way)
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Offline mheney

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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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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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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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Great work on the article Alejandro, thank you! :)
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Rob

Offline baldusi

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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.
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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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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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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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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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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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.

Online Nilof

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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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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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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 »

Offline Lars-J

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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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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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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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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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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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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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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.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline dglow

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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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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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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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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.  :)

Offline FutureSpaceTourist

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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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...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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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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...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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...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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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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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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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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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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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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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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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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Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.

Offline Dante80

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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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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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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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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.

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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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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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...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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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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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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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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I updated reply #61 on this thread to the correct my estimated size of the engine.

Offline Dante80

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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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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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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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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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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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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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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.

Offline Prettz

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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

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

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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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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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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

I'm super excited. But as you said no info to work with.
bob

Offline baldusi

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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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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
“When it looks more like an alien dreadnought, that’s when you know you’ve won.”

Offline 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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Offline Okie_Steve

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #118 on: 04/01/2017 11:09 PM »
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Offline Apollo100

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #119 on: 04/03/2017 10:54 PM »
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

Offline Stan-1967

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #120 on: 04/03/2017 11:17 PM »
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Off topic, but interesting.  Start a thread to compare this idea to ULA's IVF technology?

Offline kendalla59

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #121 on: 04/06/2017 08:26 PM »
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

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

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #122 on: 04/06/2017 10:31 PM »
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

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

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #123 on: 04/06/2017 11:03 PM »
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

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

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

Though I doubt there is much IPD heritage in Raptor

Offline Stan-1967

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #124 on: 04/07/2017 08:25 PM »
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

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

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

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

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

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

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

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



Offline rsdavis9



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

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


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

If you meant just to power the coolers I guess that works. Still think solar cells are a better choice. Maybe insulation too.
« Last Edit: 04/07/2017 08:46 PM by rsdavis9 »
bob

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #126 on: 04/07/2017 11:53 PM »
Dumping the vehicle's internal heat all into the propellant sounds like a really good way to end up at Mars without any propellant left for landing.

ITS will have radiators. Perhaps body-mounted or something (or even just carefully using the whole side of the vehicle passively, but either way it'll dump heat radiatively
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Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #127 on: 04/08/2017 01:38 AM »
If you wanted to use something like IVF to produce power instead of that 100-200kW of solar for 100 days, it'd consume about 250 tons of propellant. If you want to dump all that produced heat into the propellant, you'd run out of propellant before arriving at Mars, even if the ship somehow was 1950 tons full of propellant after trans-Mars-insertion burn.

People are all "solar is wimpy, use a combustion engine, ha!" but solar actually kicks butt in orbit. For a given mass in orbit (including consumables) you can produce about 200-400 times as much energy with solar as with IVF over 100 days.
« Last Edit: 04/08/2017 02:28 AM by Robotbeat »
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Online meekGee

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #128 on: 04/08/2017 04:05 AM »
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

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

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

... and batteries can recharge.
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Offline MP99

Dumping the vehicle's internal heat all into the propellant sounds like a really good way to end up at Mars without any propellant left for landing.

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

Cheers, Martin

Offline MikeAtkinson

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #130 on: 04/08/2017 08:30 PM »
Dumping the vehicle's internal heat all into the propellant sounds like a really good way to end up at Mars without any propellant left for landing.

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

Cheers, Martin

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

Orbiting Earth the heat load will be higher and cannot be easily controlled by orientation, so if the engine bells were shaded from both the Sun and Earth then your idea would I think be necessary, particularly as the "fleet" might spend months in LEO waiting for the TMI window. 

Offline guckyfan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #131 on: 04/08/2017 09:06 PM »
Orbiting Earth the heat load will be higher and cannot be easily controlled by orientation, so if the engine bells were shaded from both the Sun and Earth then your idea would I think be necessary, particularly as the "fleet" might spend months in LEO waiting for the TMI window.

Maybe pointing the heatshield towards earth. It should have reasonable insulation capability with its low weight. But zero boil off is probably not achievable in LEO.

Offline Nomadd

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #132 on: 04/08/2017 09:27 PM »
If you wanted to use something like IVF to produce power instead of that 100-200kW of solar for 100 days, it'd consume about 250 tons of propellant. If you want to dump all that produced heat into the propellant, you'd run out of propellant before arriving at Mars, even if the ship somehow was 1950 tons full of propellant after trans-Mars-insertion burn.

People are all "solar is wimpy, use a combustion engine, ha!" but solar actually kicks butt in orbit. For a given mass in orbit (including consumables) you can produce about 200-400 times as much energy with solar as with IVF over 100 days.
I get about 60 tons a month fuel + LOX for 100kw methane turbine using earthbound generator specs and guessing 3.8 LOX for every 1 fuel. Doesn't really make a case for an engine over solar though.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #133 on: 04/08/2017 10:16 PM »
Stoichiometric is CH4+2*O2, and CH4 is 16 molar mass, and 2*O2 is 64 molar mass. So even using the very optimistic stoichiometric case, you're looking at 4:1. Anyway, I think we're basically in agreement. 100 days is 3 and a third months, so 60 tons per month is 200 tons by your measure. But stoichiometric would likely be way too hot and would burn out the motor. IVF runs very fuel-rich, for instance (while on Earth, 80% nitrogen in air naturally will keep you cool enough). So add at least a bit of methane to keep it cool, and you're at 250 tons for 100 days.

Anyway, whether 200 tons or 250 tons is immaterial. It's way heavier than the ~1-2 tons of solar array that is needed for 100kW at Mars.
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Offline Rocket Surgeon

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #134 on: 04/10/2017 04:43 AM »
Quick question:
Is there any way to find out what kind of ISP the 1/3 Demonstrator would have with a 150:1 expansion ratio and a thrust of 1MN?

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

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

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

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

TL;DR – How much more/less extra payload could you get by putting the 1/3 demonstrator Raptor on a stage the same size/volume as the existing second stage? What ISP would it need to make that meaningful?
« Last Edit: 04/10/2017 04:46 AM by Rocket Surgeon »

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #135 on: 04/10/2017 04:45 AM »
So now it has become 1/3rd scale?

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

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #136 on: 04/10/2017 04:47 AM »
So now it has become 1/3rd scale?

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

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

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

https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/
« Last Edit: 04/10/2017 04:51 AM by Rocket Surgeon »

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #137 on: 04/10/2017 04:50 AM »
So now it has become 1/3rd scale?

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

I thought it was pretty much confirmed it was scaled down, is the stand it was on big enough to take 3MN?
Irrelevant. The THRUST may be scaled down, not necessarily the chamber size. After all, the most challenging part of Raptor is the insane chamber pressures, not the physical size. And even if you had a full-power capability, you'd first run it at lower pressures.
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Online hkultala

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #138 on: 04/10/2017 05:13 AM »
Quick question:
Is there any way to find out what kind of ISP the 1/3 Demonstrator would have with a 150:1 expansion ratio and a thrust of 1MN?

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

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

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

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

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

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

Offline Rocket Surgeon

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #139 on: 04/10/2017 05:49 AM »
Quick question:
Is there any way to find out what kind of ISP the 1/3 Demonstrator would have with a 150:1 expansion ratio and a thrust of 1MN?

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

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

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

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

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

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

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

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

But this is getting off topic, the question is what sort of stats would a sub-scale Raptor (or as I like to think of it, a Raptor 1C :P) have and could that be used to improve GTO payload mass of the Falcon 9 without changing the dimensions of the second stage.

Offline guckyfan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #140 on: 04/10/2017 08:20 AM »
Could it potentially make F9 so much more capable that they can do many flights witout using FH? Could they get away with maybe 50cm more stage diameter without changing the TE? As I expect the new carbon fiber body for a Raptor upper stage they would not be fixed to the same diameter except for TE-restrictions.

Offline cscott

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #141 on: 04/10/2017 02:05 PM »
There have been huge threads on raptor-based upper stages already. Let's not turn this one into a rehash, please.
« Last Edit: 04/10/2017 07:50 PM by cscott »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #142 on: 04/10/2017 09:57 PM »
A scaled Raptor should e able to hit Isp of 375 sec as long as the expansion ratio is near 150 and the pressures are in the neighborhood of the full scale design.

John

Offline Rocket Surgeon

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #143 on: 04/10/2017 11:08 PM »
A scaled Raptor should e able to hit Isp of 375 sec as long as the expansion ratio is near 150 and the pressures are in the neighborhood of the full scale design.

John

Brilliant! Thanks John!

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #144 on: 04/11/2017 01:02 AM »
Scaled up or down? :P
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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #145 on: 04/11/2017 01:26 AM »
Scaled up or down? :P

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

What if they have made a parametric engine design that can be scaled to any size and would just need an extended acceptance test for the first example of a new size? This seems like the sort of thing that Elon "first principles" Musk would figure out how to do.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #146 on: 04/11/2017 02:36 AM »
No.
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Offline TomH

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #147 on: 04/11/2017 06:40 AM »
Performance in rocket engines does not scale linearly in proportion to size. That has been known since liquid engine production first began. Musk said they did extensive modeling and are building at the size which provides optimum performance. Can you scale it? Yes. If you double the mass of the full sized Raptor with all proportions the same, will you get double the performance? No.
« Last Edit: 04/11/2017 06:44 AM by TomH »

Offline TomH

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #148 on: 04/11/2017 07:42 AM »
Let me try to explain with an analogy. I am not a rocket scientist. One of my four teaching credentials is in Physics, though. Let me try to explain this in a manner that a HS physics student would understand.

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

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

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

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

So, again, can you scale the thing up or down proportionally in mass and dimension? Yes. Will you get proportional performance? No. There's nothing simple about any of this stuff. That's why it's called Rocket Science.
« Last Edit: 04/11/2017 08:12 AM by TomH »

Offline MikeAtkinson

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #149 on: 04/11/2017 08:02 AM »
Yes, but a parametric engine design could take those factors into account, it just needs more (and more complex) parameters.

The difficulty for something like Raptor is that simulations would only get them so far, they would actually need to build and test various scales of engines, and then tweek the parameters to agree with reality.

Offline TomH

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #150 on: 04/11/2017 08:21 AM »
And now you have a different engine. You can't just tell a computer controlled 3D printer, Make me a Raptor engine that is scaled proportionally at 1.5 x the size and expect to get 1.5 x the performance. What you are doing is creating a new and different engine.
« Last Edit: 04/11/2017 08:22 AM by TomH »

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #151 on: 04/11/2017 01:03 PM »
Scaling an engine (within reason 1/2 - 2 times) will not have a significant impact on PERFORMANCE. Square cube scaling relations do add to cooling requirements for smaller engines, but as long as there is sufficient cooling capacity in the propellants to handle it it will not effect performance.

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

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

I am a rocket engineer.  ;^)

John
« Last Edit: 04/11/2017 01:15 PM by livingjw »

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #152 on: 04/11/2017 02:28 PM »
OK, so this was precisely the discussion I've been trying to provoke about this for some time.

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

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

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

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

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

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

Offline cscott

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #153 on: 04/11/2017 02:46 PM »
I think the CFD and parametric modeling is the "easy" part.  What is hard (and what takes all the time) is taking the theoretically perfect design, putting it on a test stand, and seeing what breaks.  And this has to be done methodically and slowly because you're talking about real physical objects and test stands that have to be rebuild from scratch if you goof and blow everything up.

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

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

Although some small amount of testing might be shared among your different parameterized designs (say, you've characterized the actual material characteristics of your 3d-printed parts well enough that you can feed the results back into your parameterization), the vast majority of the work needs to be repeated for each set of parameters.  Practically speaking, it's not worth it: the amount of effort to make a parameterization that yields an actual production-quality engine at two different parameter values is greater than the effort to just design two distinct engines from scratch.  And by forcing the parameterization you're missing out on opportunities for specializing the designs.
« Last Edit: 04/11/2017 02:47 PM by cscott »

Offline rakaydos

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #154 on: 04/11/2017 05:46 PM »
I would have expected there was already a "paramaterized" version of the Raptor, and then they did a "best fit" function on all the theoretical results, to get elon quotes like,

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

(hopefully I didnt mangle that too bad, it's from memory)

Offline Oersted

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #155 on: 04/11/2017 06:11 PM »
You can scale the engine but you cannot scale the molecular and heat-transfer properties of the fuel and oxidiser.

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #156 on: 04/11/2017 09:16 PM »
Are we taking into account SpaceX's advanced CFD system wrt to how fast they can scale a design? Combustion instability is addressed starting about 05:30 in this 2015 video. I assume today they're using NVIDIA's supercomputer since Tesla is using an automotive variant of it.

« Last Edit: 04/11/2017 09:32 PM by docmordrid »
DM

Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #157 on: 04/11/2017 10:22 PM »
OK, so this was precisely the discussion I've been trying to provoke about this for some time.

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

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

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

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

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

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

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

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

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

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

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

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

John

Online gongora

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #158 on: 04/12/2017 01:14 AM »
And, it's largely 3-d printed.

The only information I've seen for the 3D printing is 40% of the demonstrator engine by mass, and the notes from Elon's NRO talks that 3D printing doesn't work nearly as well for Raptor as it does for SuperDraco.  They can't just scale it up a little and hit the Print button.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #159 on: 04/12/2017 03:08 AM »
And, it's largely 3-d printed.

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

And just in general: parametric design works fantastic with simple objects. But as soon as you get to a certain level of complexity, a fully parametric design simply isn't feasible. It gets super complicated, and you get constraints that screw up under certain conditions, and at some point you'll get tired of fighting your model and just redo parts of it from scratch.
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Offline Rei

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #160 on: 04/13/2017 01:39 PM »
And, it's largely 3-d printed.

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

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

http://be-nl.dmgmori.com/blob/334060/67241acc5e196393c59bb68002da7c56/pl1uk15-lasertec-65-3d-pdf-data.pdf
http://en.dmgmori.com/blob/120872/cc1b707f03ee3c2b0bfc81d22c3442ca/pl0uk13-lasertec-series-pdf-data.pdf


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

We're not talking Makerbots here  ;)

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

It doesn't state, but I wonder if you can "resume" a previous build.  If so, you could take your previously-built engine and tweak its geometry without having to print a new one from scratch (since, again, it can both add and subtract).  Now that would be some fast manufacturing.
« Last Edit: 04/13/2017 01:48 PM by Rei »

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #161 on: 04/13/2017 05:26 PM »
Um, yeah, I know all about those types of additive manufacturing. It is, in fact, my job.

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

Again, additively manufactured metal parts are significantly inferior to forged metal parts.
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Offline envy887

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #162 on: 04/13/2017 05:35 PM »
Um, yeah, I know all about those types of additive manufacturing. It is, in fact, my job.

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

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

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

Offline RonM

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #163 on: 04/13/2017 05:44 PM »
As the old saying goes, use the right tool for the job.

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #164 on: 04/13/2017 06:07 PM »
Um, yeah, I know all about those types of additive manufacturing. It is, in fact, my job.

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

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

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

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

Matthew

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #165 on: 04/13/2017 06:14 PM »
Yeah, for small cast parts, additive is a big threat. Strength can be greater for additive. Really expensive, but not a problem for low part count runs.
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Offline Rei

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #166 on: 04/17/2017 09:12 AM »
Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

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

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

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


Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #167 on: 04/17/2017 03:41 PM »
Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

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

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

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

I'm distrustful when actual figures are not given in the summary.
« Last Edit: 04/17/2017 06:17 PM by Robotbeat »
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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #168 on: 04/17/2017 03:49 PM »
https://link.springer.com/article/10.1007%2Fs00170-011-3423-2?LI=true

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

I think this still does not address at all the comparison to a conventionally manufactured part that starts with a forged blank. Seems to me the summary calling the HIP treated part "hot forged" is confusing things.

Offline JamesH65

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #169 on: 04/18/2017 12:00 PM »
Surely comparing the different process strengths is mostly irrelevant - just use the one that is strong enough? And if its AM, then fill yer boots?


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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #170 on: 04/18/2017 01:31 PM »
Surely comparing the different process strengths is mostly irrelevant - just use the one that is strong enough? And if its AM, then fill yer boots?
Anything can be strong enough if you make it beefier...

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

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #171 on: 04/18/2017 03:17 PM »
Additive can compete with and exceed /cast/ properties, but forgings are much stronger due to an aligned grain structure. And this is something that simply cannot be done to the same degree with additive approaches.

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

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

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

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

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

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

There's actually an additional category in there: SLM, SLM + HIP, and hot forged / no SLM.  SLM has the strongest UTS, followed by SLM + HIP (838-845 MPa), followed by hot forged (767 MPa). The images of the microstructure in figure 5 are telling; it makes very fine, very regular dendrites surrounded by precipitates, with the dendrites oriented in the building direction. After HIP the dendrites coarsen and become more irregular, while in the purely hot forged version, the microstructure is coarse grains.
 
HIP did however improve the fatigue life by removing cracks, decreasing porosity, eliminating embedded unmelted powder, etc. But it comes at a cost of tensile strength.
« Last Edit: 04/18/2017 03:24 PM by Rei »

Online acsawdey

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #172 on: 04/18/2017 03:42 PM »
You could read more than the summary if you wanted, it's not an incredibly long paper  ;)  923-937 MPa, depending on where on the sample they tested.

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

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #173 on: 04/18/2017 04:01 PM »
Here's another paper you can grab a pdf of, it references the one Rei linked.

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

Shows the properties are strongly anisotropic with respect to the build direction.

Offline DanielW

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #174 on: 04/18/2017 05:09 PM »
Surely comparing the different process strengths is mostly irrelevant - just use the one that is strong enough? And if its AM, then fill yer boots?
Anything can be strong enough if you make it beefier...

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

This is not true. There will always be important properties involved in "strength" that don't scale with "beefy" This is especially true for anythings that requires cooling.

Offline Lar

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #175 on: 04/19/2017 01:49 AM »
This is a fascinating discussion. Can we draw any conclusions? How likely is it that AM state of the art will advance fast enough to rival forging by the time Raptor goes into serial production?  And even if not, SpaceX optimizes for cost. In this case, weight has a big leverage, presumably, but does that change the answer at all? 

Not sure if there's a better thread but maybe?
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Offline Rei

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #176 on: 04/19/2017 12:30 PM »
You could read more than the summary if you wanted, it's not an incredibly long paper  ;)  923-937 MPa, depending on where on the sample they tested.

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

If you don't have a subscription and can't get to a place that does, there's always sci-hub  ;)

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #177 on: 04/19/2017 01:18 PM »
I read the article. AM parts get higher strength than regular parts, but if you cold forge (cold draw) the metal, you get 1100MPa ultimate strength, which is a good 20% stronger than the figure they use in the paper (780MPa, I think?). Heat aging the metal also helps a lot.

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

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #178 on: 04/19/2017 07:58 PM »
I read the article. AM parts get higher strength than regular parts, but if you cold forge (cold draw) the metal, you get 1100MPa ultimate strength, which is a good 20% stronger than the figure they use in the paper (780MPa, I think?). Heat aging the metal also helps a lot.

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

Reference to that 1100 MPa figure just from cold rolling, if you would. I've been checking a variety of references for Hastelloy X and the only ones that show figures that high are from tempering.  And you can temper 3d prints, just like you can temper forged products.
« Last Edit: 04/19/2017 09:41 PM by Rei »

Offline Navier–Stokes

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #179 on: 04/21/2017 11:44 PM »
New job posting for a Raptor Test Specialist at McGregor:
Quote
Responsibilities:
*    Work with design engineers to develop and document test procedures
*    50% hands on working with hardware, 50% control systems/operation work
*    Perform tests according to procedure
*    Design fixtures and adaptors needed to perform tests
« Last Edit: 04/21/2017 11:47 PM by Navier–Stokes »

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #180 on: 04/22/2017 12:29 AM »
Any significance to specifying flight assemblies & hardware?

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

General physical fitness is required for some work areas, flight hardware is typically built in tight quarters and physical dexterity is required
DM

Offline philw1776

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #181 on: 04/22/2017 01:35 PM »
I wouldn't get too excited about something likely massaged by HR
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Offline robert_d

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #182 on: 05/04/2017 02:04 AM »
My question is what conditions/factors must be accounted for if this new engine is to be restartable?
Will there be a separate restartable version? Does performance suffer overall? Is there extra weight involved for other equipment/fluids? What about power required before the engine can produce any of its own?


Offline livingjw

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #183 on: 05/04/2017 11:18 PM »
My question is what conditions/factors must be accounted for if this new engine is to be restartable?
Will there be a separate restartable version? Does performance suffer overall? Is there extra weight involved for other equipment/fluids? What about power required before the engine can produce any of its own?

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

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

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

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

John



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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #184 on: 05/08/2017 05:52 PM »
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

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

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

Though I doubt there is much IPD heritage in Raptor

Thanks for the replies and apologies for the delayed response.... Given that SX acquired the IPD Final report, all of the drawings, and all of the hardware, I would imagine that there is quite a bit of IPD heritage in the Raptor engine.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #185 on: 05/09/2017 12:42 PM »
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

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

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

Though I doubt there is much IPD heritage in Raptor

Thanks for the replies and apologies for the delayed response.... Given that SX acquired the IPD Final report, all of the drawings, and all of the hardware, I would imagine that there is quite a bit of IPD heritage in the Raptor engine.
I wouldn't. SpaceX learned the lessons and will implement the solutions in their own way.
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Offline spacenut

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #186 on: 05/09/2017 12:54 PM »
So, what is the proposed thrust SL and Vacuum?  I've seen it all over the map.  In pounds thrust, please.  I'm retired and grew up and used the English system all my life.  I compare it to old engines from the 1960's like the F-1 and H-1, etc. 

Offline philw1776

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #187 on: 05/09/2017 02:13 PM »
So, what is the proposed thrust SL and Vacuum?  I've seen it all over the map.  In pounds thrust, please.  I'm retired and grew up and used the English system all my life.  I compare it to old engines from the 1960's like the F-1 and H-1, etc.

R SL  685,000 LBS   3050 KN

Rvac  787,000 LBS   3500 KN

Source ITS presentation Sept 2016
« Last Edit: 05/09/2017 02:42 PM by philw1776 »
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Offline spacenut

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #188 on: 05/10/2017 07:05 PM »
That is more than I thought.  I though it was about 550,000 lbs. 

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #189 on: 05/13/2017 05:33 PM »
That is more than I thought.  I though it was about 550,000 lbs.

That was the # announced years before the September reveal.  Even before that it was up to F-1 levels.
In the BFR threads here I predicted the thrust upgrade in Elon's reveal and made the obvious (sun to rise in East tomorrow) prediction that all BFR #s would continue to evolve long after those on the September Tablets From The Mount.
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