Poll

When will full-scale hot-fire testing of Raptor begin?

Component tests - 2017
2 (1%)
Component tests - 2018
10 (5%)
Integrated tests -  2017
12 (6%)
Integrated tests -  2018
147 (73.5%)
Integrated tests -  2019
20 (10%)
Raptor is not physically scaled up
8 (4%)
Never
1 (0.5%)

Total Members Voted: 200


Author Topic: ITS Propulsion – The evolution of the SpaceX Raptor engine  (Read 163208 times)

Offline Chris Bergin

https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/ - by Alejandro G. Belluscio.

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

And because this is now updated to what has been revealed, this is the continuation thread.

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.
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Offline Norm38

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In other words, to beat Merlin in TWR Raptor would have to be less than 1690kg.

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

http://forum.nasaspaceflight.com/index.php?action=dlattach;topic=34197.0;attach=1373555;sess=20788
(Tried to quote the image, but can't quote from locked threads)
« Last Edit: 10/03/2016 03:36 PM by Norm38 »

Offline F9man

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

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #10 on: 10/03/2016 03:48 PM »
Also, this makes the thing more intriguing. It might be a big coincidence, but a 1MN dev model with a nozzle area ratio of 150:1 might be very close/exactly what is needed for a Falcon9/FH Mvac methalox replacement.
Which is what incidentally the USAF paid for when entering a contract with SpaceX for this.
Too many coincidences?...XD
I understand that articles are not places to speculate. But yes, now that the size is known, it is, in fact, the perfect size for a Falcon Heavy upper stage. In fact, it might enable SpaceX to make a reusable upper stage for FH. Only issue I see, is that it would seem that the ITS upper stage has 9 engines, and they would only use the inner 3 for landing. At 20% of thrust, that would be 6,67% of thrust. Using a single Raptor would mean 3 times that thrust and thus quite an hoverslam.
But in expendable mode, Dimitry could probably surprise us.

Offline clongton

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

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #12 on: 10/03/2016 03:54 PM »
That was an excellent article, that even a novice like myself could follow...
one question popped up: will the Raptor be more difficult to mass produce than the present Merlin engines?

Thanks...

Gramps...
It will probably cost more to produce, since it will probably need higher tolerances and a lot more material. Which, when 3D printed, means a lot more print time. Also, things like valves, integration, certification and such will also cost more.
But if you look at the previous thread, they appear to have used the 3D printing capabilities in very exiting ways. For example, the LOX TP appears to be integrated straight over the injector. If they can arbitraty passages, they will simplify basically everything because the oxidizer rich gases only need to travel through the preburner/turbine/injector without needed connecting piping.
And the fuel TP case is also integrated to the side, but all the cooling passages also appear to be 3D printed. We will see how the production engines are, but this engine looks a lot like a Tesla, it looks like a conventional car, but the construction and internal layout are completely different.

Offline Lars-J

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #13 on: 10/03/2016 04:07 PM »
An impressive article.  I did not realize the engine Musk showed on the video was a 1/3 full scale unit.

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

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

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

They are clearly aware of past engine development history and some of the pitfalls (SSME, J-2X), which helps a lot.

Offline matthewkantar

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #14 on: 10/03/2016 04:14 PM »
One thing about 3-D printing the innards, I believe it limits what can be coated or left uncoated. Not sure what secret sauce is required, but previous engines of this type relied on some sort of covering to protect engine structures.

Matthew

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #15 on: 10/03/2016 04:27 PM »
The real surprise is the speed with which this engine has been built given the very limited prior art in the West on such designs.  IIRC Aerojet regularly put them into their design proposals but I don't know if many (any?) of them got to development

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

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

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

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

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

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

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

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

But I believe that the actual breakthrough was just daring to the the full flow design. That probably made all the difference.

Offline Kansan52

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #16 on: 10/03/2016 04:41 PM »
Wonderful article and very informative to a lay person (like me). The article presents how difficult this engine is, what they have done to manage the development, and shows the path ahead.

Thanks!

Offline Dante80

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #17 on: 10/03/2016 05:02 PM »
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?
« Last Edit: 10/03/2016 05:02 PM by Dante80 »

Offline rsdavis9

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.
With ELV best efficiency was the paradigm. The new paradigm is reusable, good enough, and commonality of design.
Same engines. Design once. Same vehicle. Design once. Reusable. Build once.

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #19 on: 10/03/2016 05:13 PM »
The only question I have from that article concerns the use of heat exchangers. I always thought that you could tap the methane for pressurization right after it exits the regenerative channels and not need an additional heat exchanger for that.

Do we know that the methane channel will indeed use an exchanger?
We don't know the details, but Elon said theyu usd heat exchangers. Also, expanded methane is not only hot, it is very high pressure, well past its critical point, in fact. So I guess they could use tap off, but I can't see one from the pictures and it would be quite safer to use a heat exchanger.

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