Poll

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

Component tests - 2017
3 (0.6%)
Component tests - 2018
21 (4.2%)
Integrated tests -  2017
19 (3.8%)
Integrated tests -  2018
237 (47%)
Integrated tests -  2019
181 (35.9%)
Raptor is not physically scaled up
33 (6.5%)
Never
10 (2%)

Total Members Voted: 504


Author Topic: SpaceX Raptor engine (Super Heavy/Starship Propulsion) - General Thread 1  (Read 869746 times)

Offline 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

Offline 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 »
Wildly optimistic prediction, Superheavy recovery on IFT-4 or IFT-5

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
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 Manabu

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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."

Offline 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.
Wildly optimistic prediction, Superheavy recovery on IFT-4 or IFT-5

Offline rsdavis9

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

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