SpaceX Raptor engine (Super Heavy/Starship Propulsion) - General Thread 1

[Elmar Moelzer]

...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.

[livingjw]

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

[hkultala]

...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.

[dglow]

...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?

[ArbitraryConstant]

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.

[Nomic]

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. 

[Kaputnik]

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.

[hkultala]

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.

[Dante80]

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?

[Silversheep2011]

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?

[Dante80]

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.

[livingjw]

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

[Dante80]

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...

[baldusi]

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.

[livingjw]

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

[baldusi]

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.

[John Alan]

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... 





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...   

[john smith 19]

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.

[Elmar Moelzer]

...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?

[Nathan2go]

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.