"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."
In other words, to beat Merlin in TWR Raptor would have to be less than 1690kg.
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.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.
Which is what incidentally the USAF paid for when entering a contract with SpaceX for this.
Too many coincidences?...XD
That was an excellent article, that even a novice like myself could follow...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.
one question popped up: will the Raptor be more difficult to mass produce than the present Merlin engines?
Thanks...
Gramps...
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.
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 developmentWell, 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.
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.
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.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.
Do we know that the methane channel will indeed use an exchanger?
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
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?
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.
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.
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
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/
The reason is ITAR why SpaceX have to keep details of it's tech. including Raptor secret.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
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)?
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
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/
Mr. Belluscio, a very nice article – thank you.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.
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 (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).
Mr. Belluscio, a very nice article – thank you.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.
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 (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).
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.
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 (http://www.spacex.com/sites/spacex/files/mars_presentation.pdf).
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 (http://www.spacex.com/sites/spacex/files/mars_presentation.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?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)"
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.
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.
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...
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?
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?
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. :)
The 1MN dev. model of Raptor should be mass produced to replace Merlin to do away with the He system on F9 and FH.
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.
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
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.
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: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: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.
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.
382s is for vac-optimized Raptor at vacuum.I agree! Giving anything but the vacuum Isp for a second stage engine makes no sense.
~360s is for sl-optimized Raptor at vacuum.
332s is for sl-optimized Raptor at sea level.
382s is for vac-optimized Raptor at vacuum.I agree! Giving anything but the vacuum Isp for a second stage engine makes no sense.
~360s is for sl-optimized Raptor at vacuum.
332s is for sl-optimized Raptor at sea level.
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.
...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.
...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.
...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.
Am I reading this right? This sounds like it couldn't possibly be more perfect for an enhanced upper stage for Falcon 9.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.
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?
Rescaled the BE-4, Raptor, Merlin picture with latest estimates of size.
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...;)
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.
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.QuoteFrom 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.
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?...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.
Well, the Airforce is paying for 1/3rd of the development cost, so they apparently hope it will be used to carry their payloads.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.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.
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.
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?
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)
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.For comparison here are my models of Raptor (prepared for RPA 2 SE) with all results.
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)
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
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.
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/
Hi Elon,
ITS question:
What SpaceX technology/material still requires the most development for ITS to be a success?
Thank you!
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.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
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.
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 MuskIt 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.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
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.
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.
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.
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.
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.
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.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???
Are there any updates about Raptor development after the September test?
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
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).
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!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.
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 don't know how to force a fixed width font in this forum.
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 |
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 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
Is that what you are after?
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.
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.
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
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?
SpaceX will just use batteries and a solar array. Or just batteries, most likely.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?
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?
SpaceX will just use batteries and a solar array. Or just batteries, most likely.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?
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
SpaceX will just use batteries and a solar array. Or just batteries, most likely.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?
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.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.
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
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.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.
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
Cheers, Martin
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.
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.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.
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.
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.
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.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?
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?
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.
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
Scaled up or down? :P
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.
And, it's largely 3-d printed.
Yup. And 3D printing produces worse strength than forging (for example). And you're limited in your alloy selection.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, 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.
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.
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.
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.
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?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=trueQuoteTensile 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.
https://link.springer.com/article/10.1007%2Fs00170-011-3423-2?LI=trueQuoteTensile 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.
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...
What's the ultimate tensile strength in MPa of this printed sample?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=trueQuoteTensile 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'm distrustful when actual figures are not given in the summary.
Seems to me the summary calling the HIP treated part "hot forged" is confusing things.
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.
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.
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.
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.
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
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
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?
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
I wouldn't. SpaceX learned the lessons and will implement the solutions in their own way.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.
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.
That is more than I thought. I though it was about 550,000 lbs.
Is the Raptor being built with Merlin tooling?Probably not, given that it's a very different engines(different fuels, different cycles, much higher chamber pressure).
I've read the Raptor is about the same size as Merlin. I also read that Raptor will be 3 - 4,000 psi chamber pressure. What is Merlin's chamber pressure?Merlin 1D originally had 1,410psi chamber pressure, but it's been uprated two times since then. Now it's probably around 1,800psi.
If Raptor is going to have a much higher chamber pressure, the turbo pumps will be much stronger right?More powerful engines usually means more powerful turbopumps.
I know my company has ran CNG in vehicles at 2-3000 psi to make a cylinder (welding size), handle the equivelant of 4-5 gallons of gasoline. The compressor for a fleet of 25 vehicles is large. I know the methane is liquid or LNG for the rocket. Is SpaceX going to manufacture these turbo pumps or buy them off shelf?AFAIK turbopumps for rocket engines are usually custom designs and manufactured by the engine manufacturers.
I know the engine also seems to be much smaller than the BE-4. Would it be lighter, thus higher thrust/weight ratio?We don't know, but it seems likely as Raptor will supposedly have a better TWR than M1D, which currently holds the record in that department.
Also, what is the throttle range going to be?Don't think there has been any information on this
Thanks, hopefully some of the others will be answered soon.According to the presentation from last year, Raptor will have a throttle range of 20 to 100% of thrust. That is consistent with 3 engines landing a relatively empty upper stage (barely).
Will be full regen cooled all the way out to the 3 meter (10 ft) nozzle diameter. Heat flux is nuts & radiative view factor is low.
QuoteWill be full regen cooled all the way out to the 3 meter (10 ft) nozzle diameter. Heat flux is nuts & radiative view factor is low.
https://twitter.com/elonmusk/status/877341165808361472 (https://twitter.com/elonmusk/status/877341165808361472)
Elon Musk on twitter. Probably already rumoured but now confirmed
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)
What expansion ratio does this 3m correspond to then? 40?3m Raptor nozzle at ER 40 would produce F-1 class thrust so 3m nozzle dia. for the announced thrust of Raptor would indicate an ER of around 100-150.
What expansion ratio does this 3m correspond to then? 40?3m Raptor nozzle at ER 40 would produce F-1 class thrust so 3m nozzle dia. for the announced thrust of Raptor would indicate an ER of around 100-150.
What expansion ratio does this 3m correspond to then? 40?3m Raptor nozzle at ER 40 would produce F-1 class thrust so 3m nozzle dia. for the announced thrust of Raptor would indicate an ER of around 100-150.
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)
Does it mean it has been shrunk? You could read the statement to mean that the last 2 feet of diameter increase is radiatively cooled.
Interesting! That also means that they have shrunk the Raptor-Vac nozzle, in the original ITS presentation the nozzle diameter was closer to 3.7m / 12 ft. (as measured from the schematic images)
Does it mean it has been shrunk? You could read the statement to mean that the last 2 feet of diameter increase is radiatively cooled.
That would a somewhat contrived reading, IMHO. 'Full regen' implies the whole thing, as opposed to simply 'regen' or 'partial regen'; 'nozzle diameter' implies the whole thing, not a measurement of a point on the nozzle.
It seems slightly more plausible that he simply forgot the .7
This is getting absurd. People in the real world use rounded off numbers when speaking in vernacular context. In those circumstances, you cannot take a number with one significant digit and extrapolate a conversion to another number with four significant digits.And recall that the 12' number was based on measuring screenshots of slides in a presentation video. Arguing over the last foot or two seems pointless.
Haven't picked engine size for Mars vehicle yet, will be 2-3 (probably less than 3) times the size of the sub-scale RaptorEM had selected the Raptor size at 3.05MN SL when ITS was announced at IAC2016. Now SpaceX say they have not selected the Raptor size yet. SpaceX should select a larger not smaller Raptor size for ITS to stop the engine no. of the ITS system spiraling out of control. There are rumors that the final ITS design may end up larger than that announced at IAC2016.
Given what Gwynne Shotwell said yesterday about final thrust of Raptor, the 3m diameter seems to fit quite well.
Plus 3m means the nozzle fits into the Falcon interstage (duck and cover).
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?
I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?
I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.
The BFS will also have to do a TMI burn after a long loiter in LEO. So they either need some sort of onboard propellant cooling capability, or the Raptors will have to be omnivorous and take whatever temperature propellants they are given.
On the Raptor slide of Elon's lecture says, that the engine uses supercooled LOX and methane. Does it mean, that it cannot be used with "normal" density, not supercooled propellants?
I ask this because of the operational complications of supercooling. In case of Falcon 9, supercooled propellants must be loaded immediately before launch - potentially with crew on borad. It is difficult to imagine this with LEO propellant transfer and - even more - at the Martian surface before the existence of a significant launch infrastructure.
The BFS will also have to do a TMI burn after a long loiter in LEO. So they either need some sort of onboard propellant cooling capability, or the Raptors will have to be omnivorous and take whatever temperature propellants they are given.
I think omivorous will be the way they have to go. From reading about supercooling on earth it seems very energy intensive and lots of equipment.
Nitrogen baths. etc.
On the other hand you have vacuum up there and that should provide any temperature you want with lower the pressure enough to get to the boiling point at supercooled temperatures.
Maybe, the simplest solution is to tune the booster stage Raptors for chilled propellants and the spacecarft's ones to the normal propellant density. Obviously, this is even worse performance-wise.Propellant density and starting are both helped by cooling propellants to below their boiling points. Isp doesn't change.
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?According to some, turbines can move more propellant if it's denser. I guess that means turbine speed is more of a limiting factor than turbine power in this case.
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?
Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?
Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?
Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?
Why wouldn't it be? Otherwise there would be no performance gain.
So lox at 66k vapor pressure is .029 bar. So Mars surface at .005 bar and space at 0 bar will cool lox just fine. All you have to do is collect the gaseous oxygen boil off at .029 bar and compress it in a Linde liquefaction cycle and put it back in the tank. Takes energy but not the special stuff you do at 1 bar. Like vacuum pumps or LN2 baths.
So lox at 66k vapor pressure is .029 bar. So Mars surface at .005 bar and space at 0 bar will cool lox just fine. All you have to do is collect the gaseous oxygen boil off at .029 bar and compress it in a Linde liquefaction cycle and put it back in the tank. Takes energy but not the special stuff you do at 1 bar. Like vacuum pumps or LN2 baths.
Good to know. What about LEO?
My understanding was that they use subcooled propellant to eliminate cavitation in the turbopump. Full power would mainly be needed on earth ascent, both in the first and second stage. That can be provided with subcooled propellant on tanking. Can cavitation also be avoided with some throttling? For TMI full power would not be needed, also on Mars ascent it is not as essential.
If subcooled is needed in every phase, can you calculate, how much propellant would be wasted to cool propellant a few degrees below sea level pressure boiling temperature?
Isn't the only reason for super cooling propellants is to provide more propellant in a given fixed space?
Certainly, this is the main point, even if engine operation is affected by propellant density. The question is whenter the TMI, landing and return DeltaV expectations are calculated with assuming the extra propellant, or not?
Why wouldn't it be? Otherwise there would be no performance gain.
Because subcooling prop on-orbit or on the Mars surface is somewhat more challenging than on the launch pad, and the leg from staging to Earth orbit is more challenging than LEO to mars surface or Earth return empty.
I don't think subcooled props are strictly necessary for TMI or Earth return, but if they can solve long term boiling storage than long term subcooled isn't all that much more difficult, so they might do it. It does help with fast transits and next-synod reuse.
Engine pumps are very sensitive to vapor pressure and tank pressure (head or otherwise). In space you don't have head pressure. On Mars, you obviously have less than on earth, but better than in space. In space, BFS will probably need to use either sub-cooled propellants, higher tank pressures or boost pumps or some combination. I was surprised that Raptor doesn't have them (yet). :^)
John
I still don't see the problem. If the propellant is no longer subcooled, it doesn't go away. (conservation of mass and all that) It is still there, just taking up more volume. The engines will certainly be able to handle a bit of temperature range.
I still don't see the problem. If the propellant is no longer subcooled, it doesn't go away. (conservation of mass and all that) It is still there, just taking up more volume. The engines will certainly be able to handle a bit of temperature range.
Assume you have filled the tank in orbit to capacity with subcooled propellant. Then the temperature drifts to the boiling point. Some of the propellant is going to go away unless you keep the vents closed. In that case it will stay until the tanks burst and it all goes away.
It will take 6 tanker launches to fill the tanks. If you can keep it subcooled for that long, what's stopping you from keeping it subcooled until you use it?
Of course.
Didn't Elon say they were using multistage pumps on Raptor? The low pressure pump might be designed to handle lower vapor pressure without cavitation.
It will take 6 tanker launches to fill the tanks. If you can keep it subcooled for that long, what's stopping you from keeping it subcooled until you use it?
Time maybe? They can fill it in a week with daily launches. But if it waits for months in LEO for the Mars window to open it will be hard to keep propellants subcooled without any active measures. Less hard while in interplanetary space away from IR emitting earth.
Edit: I was mostly repying to the "it does not go away".
There are ways to handle it. Fill up to boiling temperature, wait for departure time, with hopefully minimal boiloff. Subcool by opening to vacuum and have a last tanker fill up before departure. One tanker can probably do the topping off for several departing vehicles.
What vehicle will use the Raptor? I know all about the ITS but the Raptor will be done well in advance of ITS and if it's only use is ITS then it seems like the economics of SpaceX won't work. Raptor has got to have more use than that. Is it only me or there is a BIG gap in the SpaceX launch family from F9 to ITS......
The terms "methane" and "liquid natural gas" are sometimes being used interchangeably in discussions of Raptor engines. Is it safe to assume that SpaceX will be removing the other hydrocarbons in LNG to generate pure refined methane? Or is it possible they'll use LNG as is from commercial sources? It seems silly to even ask the question - I presume they will be purifying it - and that LNG is just being used as shorthand for "refined rocket grade liquid methane" in the same way that "kerosene" is sometimes used to refer to RP-1.
The terms "methane" and "liquid natural gas" are sometimes being used interchangeably in discussions of Raptor engines. Is it safe to assume that SpaceX will be removing the other hydrocarbons in LNG to generate pure refined methane? Or is it possible they'll use LNG as is from commercial sources? It seems silly to even ask the question - I presume they will be purifying it - and that LNG is just being used as shorthand for "refined rocket grade liquid methane" in the same way that "kerosene" is sometimes used to refer to RP-1.
LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. It can also have up to 1% nitrogen in it.AFAIK, pure CH4 is planned for Raptor flight versions. This was stated previously in tweets during initial Raptor testing at McGregor Complex.
Yeah. I was thinking more of BO even though the question was specifically about Raptor. It's Jim's fault.LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. It can also have up to 1% nitrogen in it.AFAIK, pure CH4 is planned for Raptor flight versions. This was stated previously in tweets during initial Raptor testing at McGregor Complex.
Somebody reported, that SpaceX did not use the same, refined and expensive LOX as NASA and used lower purity industrial one instead. Because of this I would assume that SpaceX will use the cheapest fuel acceptable.
LNG is kind of a vague term itself and composition varies quite a bit. It can be anywhere from 85% to 95% methane. Even if they did allow some higher anes in it, you'd think they'd want the exact same ratio every time. It can also have up to 1% nitrogen in it.AFAIK, pure CH4 is planned for Raptor flight versions. This was stated previously in tweets during initial Raptor testing at McGregor Complex.
BTW: Is there such a thing as "refined rocket grade liquid methane"? After all, according to astronautix.com there never was a production LOX/LCH4 engine.
Will Raptor be the full 685,000 lb thrust engine or will the initial booster use less thrust?
Some Raptor questions, can't find them anywhere else.
1) How will the existing sub-scale engine be upgraded to a full thrust Raptor? Is the combustion chamber and turbo pumps the same? Just increasing speed of the pumps?
2) Is the sub-scale engine smaller than the full thrust Raptor?
3) How long will it take to go from sub-scale to full thrust?
4) With the possible revelation of a 9m BFR/ITS, will this use 42 sub-scale engines? Or use full scale engines?
5) With the above revelation, can 42 sub-scale engines fit in a 9m BFR?
6) Will there possibly be an in between engine from the 225k lb thrust sub-scale to the 685k thrust full scale to power the 9m BFR/ITS?
Thanks.
...
3) Not quite as long as developing an engine from scratch, but still a while. Probably at least 4 years to first flight.
...
...
3) Not quite as long as developing an engine from scratch, but still a while. Probably at least 4 years to first flight.
...
Since additive manufacturing is used for so much of this engine, making the transition to full scale on this 'scaleable' engine could be much faster than traditional builds.
>
Since additive manufacturing is used for so much of this engine, making the transition to full scale on this 'scaleable' engine could be much faster than traditional builds.
Wow.. Didn't expect that going from a subscale engine to full scale takes that long. What is the rational to do the subscale in the first place? It's practically useless if development of subscale plus the transition time to full scale is longer (and therefore more expensive) than going straight to full scale.
But they tested the raptor engine at McGregor if I recall correctly.
When developing the fulls scale version, the components cant be tested in Stennis either, following your logic.
When developing the fulls scale version, the components cant be tested in Stennis either, following your logic.
Following the capabilities of the Stennis facility.
Building a subscale full engine still makes sense IMO, if they can utilize existing production facilities for Merlin engines to some extent. Like for the combusition chamber and the nozzle.
Establishing it works well on subscale at low cost is a good move before investing in completely new production facilities for the full size. Independently of wether they will ever fly a subscale engine or not.
The only reason I can see, is the Air Force money to develop an upper stage metholox engine. This sub-scale engine can be made for a vacuum engine. Capability is in the range of J2X or two BE-3's. Someone said the way it is designed, that it can be scaled up easily. Hopefully quickly to get the BFR/ITS on the road.
Where does SpaceX test this sub-scale engine? Will they attempt to see what it's maximum thrust can be?SpaceX tests their subscale Raptor at their McGregor test facility in Texas. I think that they will also test the full size Raptor at McGregor as well.
Also, can they test a full scale engine at McGregor? Or do they need Stennis?
The only reason I can see, is the Air Force money to develop an upper stage metholox engine. This sub-scale engine can be made for a vacuum engine. Capability is in the range of J2X or two BE-3's. Someone said the way it is designed, that it can be scaled up easily. Hopefully quickly to get the BFR/ITS on the road.
That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.
John
The only reason I can see, is the Air Force money to develop an upper stage metholox engine. This sub-scale engine can be made for a vacuum engine. Capability is in the range of J2X or two BE-3's. Someone said the way it is designed, that it can be scaled up easily. Hopefully quickly to get the BFR/ITS on the road.
That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.
John
Exactly, No one has done a FFSC Methane engine before. Start small, learn how to run one, then scale up.
Smaller is smaller, faster to build, smaller test stand, less consumables, less damage if it goes boom.
The fact that they may end up with an US engine and that the USAF helped pay for it is a bonus.
The only reason I can see, is the Air Force money to develop an upper stage metholox engine. This sub-scale engine can be made for a vacuum engine. Capability is in the range of J2X or two BE-3's. Someone said the way it is designed, that it can be scaled up easily. Hopefully quickly to get the BFR/ITS on the road.
That is surely part of the reason, but I think the current engine was initially meant to be a sub-scale demonstrator. The full flow staged combustion cycle is completely new to them. A lot of unknowns can be put to rest with a demonstrator. This allows them to better correlate there models to reality. The Air Force contract may have allowed them to take it further.
John
Exactly, No one has done a FFSC Methane engine before. Start small, learn how to run one, then scale up.
Smaller is smaller, faster to build, smaller test stand, less consumables, less damage if it goes boom.
The fact that they may end up with an US engine and that the USAF helped pay for it is a bonus.
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
Yes, I read the full scale sea level will be about 2m in diameter, while the vacuum version will be around 4m. Just wondering the size of the sub-scale and if it could actually be put into production. Maybe not optimal. But a vacuum version could be used as a second stage for F9/FH to improve performance and begin reuse testing.
Yes, I read the full scale sea level will be about 2m in diameter, while the vacuum version will be around 4m. Just wondering the size of the sub-scale and if it could actually be put into production. Maybe not optimal. But a vacuum version could be used as a second stage for F9/FH to improve performance and begin reuse testing.
[bold above mine] I don't think this can be right, at least for the SL version. With the 42 engine arrangement shown in last year's reveal, you have a minimum of 7 engine bells across, plus two sizable gaps. That means the full scale version was planned to be no more than 1.6 meters in diameter.
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.
According to a recent Musk tweet, the Vac Raptor nozzle size is now 3m.
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.
The observed size of the nozzle exit on the sub-scale engine is just over 0.9 meters, so that matches closely. But the nozzle throat appears too large for a 40:1 ER (or for that matter 1,000 kN thrust at 30 MPa).
So, does anyone know what size the sub scale is? Is it Merlin sized since they have the tooling? Also, does anyone know what size the full scale Raptor will be. I've seen some pictures but, they are not scaled, just guessing?
Did this correlation last year with my engine model.
If the sub-scale is the same cycle, the 40:1 nozzle would be the 1.7*sqrt(1/3) = 1.7*.577 = .98 m. Call it a meter.
Length would be around 1.8 m.
The observed size of the nozzle exit on the sub-scale engine is just over 0.9 meters, so that matches closely. But the nozzle throat appears too large for a 40:1 ER (or for that matter 1,000 kN thrust at 30 MPa).
Yes, that fooled me. I think we must be looking at a plug to keep the critters out.
John
Anyone heard anything on the full scale Raptor development? Got to have the engine before we get the 9m ITS going.Not necessarily. It's possible they'd start with some subscale Raptor.
No, because it will be designed to be air start able from the very beginning. Needs to air start for first and second stage both.Thanks for that. Has it been discussed before? I was wondering especially whether a "tap off" might be a part of the solution, and if so, does that encroach on Blue Origin IP?
Ok so I read entire 15pages but didn't find any discussion on this:
If I read it right, the methane/oxygen into main chamber are gasified already, where the gasification happened? preburner? As I find it hard to believe gasification happened during turbine stage.
Or I miss something?
Titus
Ok so I read entire 15pages but didn't find any discussion on this:
If I read it right, the methane/oxygen into main chamber are gasified already, where the gasification happened? preburner? As I find it hard to believe gasification happened during turbine stage.
Or I miss something?
Titus
Here is pretty good video that explains (among other things) Full Flow Staged Combution cycle:
https://www.youtube.com/watch?v=4QXZ2RzN_Oo
Where does SpaceX test this sub-scale engine? Will they attempt to see what it's maximum thrust can be?SpaceX tests their subscale Raptor at their McGregor test facility in Texas. I think that they will also test the full size Raptor at McGregor as well.
Also, can they test a full scale engine at McGregor? Or do they need Stennis?
RESPONSIBILITIES:
• Engineering, design, analysis, material/component selection and procurement, construction, activation, and maintenance of test stands, tooling, and supporting infrastructure
• Provide support and direction to technicians during construction, activation, and maintenance of stands and equipment
• Support testing campaigns by operating ground propellant systems, reviewing data for system health, and modifying equipment or procedures as necessary
• Develop novel ways to streamline site-wide processes and increase the reliability and efficiency of testing operations
• Perform any additional tasks that ensure efficient and effective testing, as required
• It is sometimes necessary to perform hands-on work in all environments (heat, cold, rain), occasionally in tight quarters or at heights
BASIC QUALIFICATIONS:
• Bachelor’s degree in mechanical engineering, aerospace engineering or other engineering discipline
PREFERRED SKILLS AND EXPERIENCE:
• Master’s degree in mechanical or aerospace engineering
• 3+ years of relevant experience in an industrial setting
• Fundamental understanding, intuition, and aptitude of fluid and/or structural design and analysis
• Creative ability to imagine and design from scratch, while retaining low cost, reliability, efficiency, and maintainability
• Experience where quick-thinking and problem solving plays a critical role
• Good response to challenges posed by short deadlines
• Ability to work in a high-concentration, high-stress environment, under possible extended work hours
• Acute attention to detail, ability to see interactions with other systems to avoid problems
• Intermediate skill level using Windows Operating Systems
• Intermediate skill level using Microsoft Office
• Intermediate skill level using CAD (NX a plus)
• Experience with high pressure and cryogenic fluid systems and components
• Experience producing drawings for welders and machine shop fabrication
• Machining, welding, other fabrication techniques, and general hands-on experience
• Experience in FEA or CFD modeling and analysis, with ability to verify by simplified hand calculations
• Piping and pressure vessel design experience per ASME code, work with flanges, gaskets, fasteners
• Instrumentation, testing, data review and analysis, verification against a model
ADDITIONAL REQUIREMENTS:
• General physical fitness is required for some work areas, flight hardware typically is built in tight quarters and physical dexterity is required
• Physical effort including standing, lifting and carrying light weight such as materials or equipment. Must lift up to 30 pounds unassisted
• Occasionally exposed to work in extreme outdoor environments- heat, cold, rain
• Work performed in an environment requiring exposure to fumes, odors, and noise
• Must be available to work extended hours and weekends, which varies depending on site operational needs, flexibility required
Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.So looks like what was termed subscale Raptor that is undergoing testing is now the size of Raptor they will use. Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.
31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar
IAC 2016, it was 300bar, 3050kN
Still thinking they'll go up to 300 eventually.Surely they will, but 250bar->300bar only give thrust change as 1700kN to 2000kN
Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.
That maybe true of the goal is reaching the original ITS(IAC2016) capability.Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.So looks like what was termed subscale Raptor that is undergoing testing is now the size of Raptor they will use. Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.
31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar
IAC 2016, it was 300bar, 3050kN
Based on IAC 2017, we are now looking at much smaller Raptor, roughly 60% of IAC 2016.So looks like what was termed subscale Raptor that is undergoing testing is now the size of Raptor they will use. Saves a lot in dev. costs over IAC2016 size Raptor but I still think they still should have bit the bullet and gone with a larger Raptor to reduce booster engine no.
31 engines with 5400mT thrust. That's 1707kN per engine.
And the chamber pressure now down to 250bar
IAC 2016, it was 300bar, 3050kN
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!
Appears they need more or larger tanks if they can only get a 40 sec run. They’ll need a lot more time running the Raptor.
The blue exhaust and shock diamonds were fantastic.
Still have a hard time imagining 31 of them on a vehicle.
I'm pretty sure this version of the Raptor will eventually produce 3 MN. Initially Merlin 1D had ~70% of block 5 levels; here's the difference is roughly the same.
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?Graph is easier to explain :)
Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
For my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?Graph is easier to explain :)
Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
(using Rocket Propulsion Analysis :))
Titus
Depending on what you define as "oxidizer-fuel" mid-point ;D ;D ;D ;DFor my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?Graph is easier to explain :)
Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
(using Rocket Propulsion Analysis :))
Titus
Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..
@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
Depending on what you define as "oxidizer-fuel" mid-point ;D ;D ;D ;DFor my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?Graph is easier to explain :)
Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
(using Rocket Propulsion Analysis :))
Titus
Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..
@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
"Rich" is always "relatively compared to referencing point" ;D ;D ;D ;D ;D ;D
The test data given was a bit of a surprise to me. I had no idea they had progressed so much already!
Not only progressed, but progressed on the flight engine.
This puts them squarely in the lead on next generation engines and vehicles.
Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.
Depending on what you define as "oxidizer-fuel" mid-point ;D ;D ;D ;DFor my personal understanding, is it correct that the change in mixture ratio makes the combustion colder, hence less pressure and ISP?Graph is easier to explain :)
Also, can someone please explain to me why they are going oxygen rich instead of fuel rich? I always thought a fuel rich environment would create lighter combustion elements, hence higher ISP for the same chamber temperature and pressure. Is that wrong?
(using Rocket Propulsion Analysis :))
Titus
Wow, that is unexpected. The crashing sound you just heard was my intuition thrown out the window and hitting the ground too hard. Thank you for trashing my misconception! :) Now I need to find a new explanation..
@ edit.. wait a second. My mistake, complete nonsense! a mixture ratio of <4 : 1 is actually fuel rich! The 1 is the fuel and the <4 is the oxygen. There need to be 4 oxygen atoms for 1 methane molecule. If there are less, that means fuel rich. So I was correct after all. *reassembles previously assumed misconception*. Sorry buddy. Just a parsing error.
"Rich" is always "relatively compared to referencing point" ;D ;D ;D ;D ;D ;D
Not only progressed, but progressed on the flight engine.
This puts them squarely in the lead on next generation engines and vehicles.
Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.
It is promising.
Do we have any public information about progress on the spark ignition. I couldn't see any initial green tint of TEA/TEB at startup in the most recent test video released with today's (IAC 2017) presentation.
Not only progressed, but progressed on the flight engine.
This puts them squarely in the lead on next generation engines and vehicles.
Next series of tests might be at 250 Bar... then flight qual tests.
Could have a flight qualified engine and production line running as we enter 2019.
It is promising.
Do we have any public information about progress on the spark ignition. I couldn't see any initial green tint of TEA/TEB at startup in the most recent test video released with today's (IAC 2017) presentation.
It does have a beautiful blue/violet flame though.
Night launches for this rocket are going to be a treat.
So during the burn video there are 2 green flame episodes, one near the beginning, one at shutdown. Since these are not actually at startup, and supposedly Raptor uses spark ignition anyway, the only explanation I can think of is a bit of copper chamber or bell vaporizing.
The first frame of the early incident at 5:44 you just see a little streak by the bell, the next frame it's partway down the jet, then it's at the end of the jet. For some reason this frame is doubled. Then the next frame, no more green. However the vapor patterns along the ground don't suggest that any video is missing, they change from frame to frame in a consistent way.
Also, is it possible to tell the exhaust velocity from the spacing of the mach diamonds? They are extremely consistent right until shutdown starts around 6:20. You can see from one frame to the next that they are sliding to the right at that point.
Endless Raptor gif, came out a little small:
(https://i.makeagif.com/media/9-29-2017/-fUtUN.gif)
A quick google suggests that "incomplete combustion" can cool a methane flame from blue to yellow (http://www.elgas.com.au/blog/1585-why-does-a-gas-flame-burn-blue-lpg-gas-natural-propane-methane)Endless Raptor gif, came out a little small:
(https://i.makeagif.com/media/9-29-2017/-fUtUN.gif)
Similarly, here's that green flame episode:
(https://i.makeagif.com/media/9-29-2017/s5Lpag.gif)
And shutdown:
(https://i.makeagif.com/media/9-29-2017/Z1rGm-.gif)
A quick google suggests that "incomplete combustion" can cool a methane flame from blue to yellow (http://www.elgas.com.au/blog/1585-why-does-a-gas-flame-burn-blue-lpg-gas-natural-propane-methane)
If they run fuel rich, the methane may take a few miliseconds longer to turn on and off, resulting in a slightly cooler (green flame) burst as the o2 turns off completely. (or before the O2 turns on)
Still thinking they'll go up to 300 eventually.At around 23 min into the presentation Elon comments on improvements in both ISP (add 10 units or so) and chamber pressure (to 300 bar).
This seems like either copper or boron ions in the flame.
This seems like either copper or boron ions in the flame.
Are you implying engine rich combustion?
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
The question is, will BO follow the road to small engines or not?
Assuming the so called “full scale” Raptor wil never be build.
Blue might not have a choice. I read somewhere that they had to increase the size for Vulcan.
The question is, will BO follow the road to small engines or not?My guess is that BO will stick with fewer, bigger engines. At least I hope they do mostly because I like seeing a wide variety of approaches. :)
Assuming the so called “full scale” Raptor wil never be build.
One aspect that hasn't been talked about much is that Raptor is gong to be able to throttle down to 20%... (Compared to Merlin 1D with ~40%)
This means that the lowest thrust on Raptor will be roughly the same as the lowest thrust on Merlin 1D, since Raptor is going to have ~2x the thrust capability.
One aspect that hasn't been talked about much is that Raptor is gong to be able to throttle down to 20%... (Compared to Merlin 1D with ~40%)
This means that the lowest thrust on Raptor will be roughly the same as the lowest thrust on Merlin 1D, since Raptor is going to have ~2x the thrust capability.
Yes that’s quite important, if Merlin 1D could throttle down to 20%, Falcon 9 could probably hover, and thus also make a softer landing.
The question is, will BO follow the road to small engines or not?I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.
Assuming the so called “full scale” Raptor wil never be build.
Does anybody have a engine layout for 31 engines?Raptor SL nozzle is 1.3m dia. Most likely engine configuration for 31 engines is 1+6+12+12.
I have seen 1+8+16=25 but does the 31 engine layout just have 6 more engines around the outside?
So raptor SL is 2.8m...
Similarly, here's that green flame episode:
And shutdown:
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?Turbomachinery configuration info. is likely covered by ITAR so please don't expect SpX to release any info. on it.
The question is, will BO follow the road to small engines or not?I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.
Assuming the so called “full scale” Raptor wil never be build.
BO will avoid N-1 type architectures like the plague unlike SpaceX.
Does anybody have a engine layout for 31 engines?Raptor SL nozzle is 1.3m dia. Most likely engine configuration for 31 engines is 1+6+12+12.
I have seen 1+8+16=25 but does the 31 engine layout just have 6 more engines around the outside?
So raptor SL is 2.8m...
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?
BO will avoid N-1 type architectures like the plague unlike SpaceX.
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.
Every kg of mass removed from the engine is about 10 kg more to orbit.
has SpaceX released any information on the turbomachinery configuration and if that is involved in the test?
No, but one of the reasons why Raptor is the first production-intent FFSC (full flow staged combustion) engine is that it's difficult (perhaps prohibitively so) to test the fuel pump, oxidizer pump, and main combustion chamber in isolation from each other. Without resorting to extremely elaborate test stand hardware, the engine really has to be tested as a complete unit. This formidable upfront challenge has deterred engine manufacturers for decades.
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.
Every kg of mass removed from the engine is about 10 kg more to orbit.
How much extra payload could be put into orbit for every ton reduction in total engine weight?
The Raptor engine looks really light, with it’s compact turbopumps/combustion chamber.
Every kg of mass removed from the engine is about 10 kg more to orbit.
How does that work? Surely the most it could ever be is a 1:1 trade?
What can we learn about the status of the Raptor development program from the details provided in the IAC talk (1200 seconds total, longest firing 100 seconds, 42 firings)?
A November 2007 SpaceX press release (http://www.businesswire.com/news/home/20071112005019/en/REPLACING-VIDEO-SpaceX-Completes-Development-Merlin-Regeneratively) indicates that the Merlin 1C development program included 3000 seconds total, longest firing 170 seconds, 125 firings. Would the Raptor program be comparable, or would it be at a different scale?
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers. Does anyone have any information to contradict that?
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers. Does anyone have any information to contradict that?
Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers. Does anyone have any information to contradict that?
Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers. Does anyone have any information to contradict that?
Depends on whether the current one is subscale or fullscale. If subscale, then they'll need to scale it up later, in which case you can't claim they're ahead.
BO has the money to dev. an F-1 class engine that SpaceX does not have so SpaceX have downsized the Raptor to fit within their R&D budget.The question is, will BO follow the road to small engines or not?I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.
Assuming the so called “full scale” Raptor wil never be build.
BO will avoid N-1 type architectures like the plague unlike SpaceX.
You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.
Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
BO has the money to dev. an F-1 class engine that SpaceX does not have so SpaceX have downsized the Raptor to fit within their R&D budget.The question is, will BO follow the road to small engines or not?I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.
Assuming the so called “full scale” Raptor wil never be build.
BO will avoid N-1 type architectures like the plague unlike SpaceX.
You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.
Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
As I understand it, none of the flight engines where test fired. They made them in batches and selected a random engine from the batch for testing. If it was good, the batch was declared good.
- The engines where not all test fired beforehand.
As I understand it, none of the flight engines where test fired. They made them in batches and selected a random engine from the batch for testing. If it was good, the batch was declared good.
- The engines where not all test fired beforehand.
It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.
As I understand it, none of the flight engines where test fired. They made them in batches and selected a random engine from the batch for testing. If it was good, the batch was declared good.
- The engines where not all test fired beforehand.
Yes you’re right, once tested it’s probably unusable because of the bad fuel among other things.
Luckily Methalox burns very clean.
It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.
It would be great news if you are right about this. The subscale test Raptor (roughly 0.87mØ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mØ BFR SL engine is projected to be 1.7MN at 250 bar. The test Raptor expansion appears ambient in the recent test footage. The ratio of nozzle areas is 1.3^2 / 0.87^2 = 1:2.23 (a 110% increase). Would the extra 50 bar (a 25% increase) prevent overexpansion with such a large nozzle? Would the increased chamber pressure and larger nozzle be sufficient to boost the test Raptor to 1.7MN, a 70% increase?
In the 2017 IAC presentation, Musk places emphasis on the engine out capability for ITSy during landing. Do people think SpaceX considered that a required design criteria for the ship? If so, did the size of the engine drive the size of the ship, or visa versa?They wanted to be able to man-rate propulsive landing, which requires incredibly high reliability. Having a spare squares your reliability. (technically, squares your failure fraction, making it smaller)
As I play around RPA software, here is my "non mathematics" sense goes ...It looks like to me that the Raptor being tested is sub-scale compared to the one in last year's presentation, but full scale for this years 9 m BFR. They need to increase the pressure and add a bigger nozzle, but not change the linear dimensions.
It would be great news if you are right about this. The subscale test Raptor (roughly 0.87mØ) was quoted as 1MN thrust, and we now know it operates at up to 200 bar. The 'version 1' 1.3mØ BFR SL engine is projected to be 1.7MN at 250 bar. The test Raptor expansion appears ambient in the recent test footage. The ratio of nozzle areas is 1.3^2 / 0.87^2 = 1:2.23 (a 110% increase). Would the extra 50 bar (a 25% increase) prevent overexpansion with such a large nozzle? Would the increased chamber pressure and larger nozzle be sufficient to boost the test Raptor to 1.7MN, a 70% increase?
- I will be cutting numbers Monday for the sub-scale test Raptor and BFR Raptor, but it appears that relatively small (~1.15 linear) scaling up will be needed.
- I believe that engine out during BFS landing probably sized the engine and as a consequence resulted in 31 engines on the BFR. Also, the fact that they already had test fired an engine of about the right size.
John
In the 2017 IAC presentation, Musk places emphasis on the engine out capability for ITSy during landing. Do people think SpaceX considered that a required design criteria for the ship? If so, did the size of the engine drive the size of the ship, or visa versa?Even if BFR or ship has a single benign engine failure and the mission is a total success then there will be a stand down period to find out and fix the issue. 31 engines on booster and 6 on ship will increase the risk of benign engine failures causing stand down periods which SpaceX may not be able to afford.
Is the current test engine "sub-scale" or just under pressure fed? My take from Musk's talk was an increase from the currently tested 20 bar to 25 bar and eventually 30 bar. At no point does it appear any scaling up is called for or necessary. Why are we still calling the as tested Raptor "sub-scale"?Pressures you have quoted should be 200 bar, 250 bar, and 300 bar.
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
Working principle and purpose is much more important than the number of something.
N-1 used differential thrusting for steering. And at least one of the failures were due lack of control.
BRF does not, it uses swiveling engines. It's immune to the "lose control authority due wrong engine failing".
And that "down time after engine failure" just means that after those downtimes, it will be much reliable than any competing rocket.
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture.BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
Any booster with such high engine no. is likely to suffer engine failures.
Is the current test engine "sub-scale" or just under pressure fed? My take from Musk's talk was an increase from the currently tested 20 bar to 25 bar and eventually 30 bar. At no point does it appear any scaling up is called for or necessary. Why are we still calling the as tested Raptor "sub-scale"?
Where's your evidence that they downsized because of budget?Engine dev. costs increase faster than thrust so Raptor has been reduced in size to minimize dev. cost of BFR while keeping engine no. to what they think is acceptable which I don't agree with. BO has almost limitless funding so they can dev. a much bigger engine for their NA.BO has the money to dev. an F-1 class engine that SpaceX does not have so SpaceX have downsized the Raptor to fit within their R&D budget.The question is, will BO follow the road to small engines or not?I see BO going in the opposite direction in engine size than SpaceX and may dev. an F-1 class engine after BE-4 for their NA. SpaceX originally planned SL Raptor to be in the F-1 class but it is now smaller than the SSME.
Assuming the so called “full scale” Raptor wil never be build.
BO will avoid N-1 type architectures like the plague unlike SpaceX.
You see BO going for an even bigger engine than BE-4, despite all the issues they are having with its development? Keep in mind that BE-4 used to be a smaller engine before ULA stepped in and asked for a larger one. So this should tell you something about their love for large engines.
Your posts about this (engines size vs engine count) really sounds like someone from 10 years ago saying “9 engines... CRAZY, it will never be reliable, and it will be so hard to start all 9”. ;)
Engine dev. costs increase faster than thrust so Raptor has been reduced in size to minimize dev. cost of BFR while keeping engine no. to what they think is acceptable which I don't agree with. BO has almost limitless funding so they can dev. a much bigger engine for their NA.
Lets all of us hope that SpaceX does not repeat this footage with their maiden BFR launch.BO will avoid N-1 type architectures like the plague unlike SpaceX.Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
https://www.youtube.com/watch?v=U9fkYIrRwbo
We'll see if SpaceX provides more frequent updates over the next 6-12 months. I certainly hope/expect that they will if they are indeed planning on beginning the first BFR construction before H2 2018. My intuition tells me that there's no way that would happen unless tankage and engine designs were finalized and flight certified.
Something to think about: As far as I can tell from the info we have seen, Raptor is already ahead of Blue Origin and AJ's candidates for Vulcan or Blue's new launchers. Does anyone have any information to contradict that?
N-1 type architecture is any booster having around 30 or more engines so BFR has the N-1 type architecture. Any booster with such high engine no. is likely to suffer engine failures. Even if engine failures are benign and missions are successful there will be down time periods to get the issue found and fixed which SpX may not be able to afford.BO will avoid N-1 type architectures like the plague unlike SpaceX.
Anyone care to elaborate for those of us not as knowledgable? What are "N-1 type architectures?" Why would Blue avoid them and SpaceX not?
If you bet one dollar that the 1st BFR mission is a complete success with no issues then you will win a fortune.
Lets all hope that Raptor works as advertised and that SpX don't lose any BFR's but I think that will be a long shot.
Engine dev. costs increase faster than thrust so Raptor has been reduced in size to minimize dev. cost of BFR while keeping engine no. to what they think is acceptable which I don't agree with. BO has almost limitless funding so they can dev. a much bigger engine for their NA.
Raptor size was fixed by the desire to have two landing engines for reliability. A bigger engine would not allow for engine-out tolerance. Blue Origin doesn't have a reusable upper stage design at the moment, so they don't face this same consideration.
- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.
- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.
John
Also, SpaceX makes extensive use of 3D printers for their engines. The size of Raptor may be limited to the largest 3D printer they could find. As 3D printers get larger, the size of an engine could also increase. 3D printers lower production costs as it reduces labor costs. It seems to also increase reliability.Youre overselling 3D printing here.
My understanding is that additive manufacturing ("3d printing") is really good at the kind of complicated part that otherwise would have needed to be assembled out of multiple components. As I understand, turbopumps especially a re made of lots of that kind of component, and a FFSC engine has twice as many turbopumps as even a regular engine.Also, SpaceX makes extensive use of 3D printers for their engines. The size of Raptor may be limited to the largest 3D printer they could find. As 3D printers get larger, the size of an engine could also increase. 3D printers lower production costs as it reduces labor costs. It seems to also increase reliability.Youre overselling 3D printing here.
- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.
- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.
John
Hi JW,
How did expected cost scale with:
1. Number of parts?
2. Chamber Pressure?
3. Fuel / Oxidizer Choice?
4. Engine Cycle?
5. # of Turbopumps?
6. Any other interesting tid bits I may have missed?
Also, how accurate do you believe these models were? To use your example, how do you believe they verified cost scaled at ~(thrust)^(0.7)?
I've always been curious how well these models work and was just wondering what your opinion of them were?
Thanks,
C
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
As they uprate to 300bar they may even be able to reduce engine count.
"Now", only. I await the progressive uprating.
- Engine development models that I have used (Transcost) show development costs proportional to thrust to the ~.7 power.
- If they were going to develop a booster engine it most probably would be higher thrust, but SpaceX's timeline and funding don't permit it. They are going with what they have, a smaller Raptor similar in size to their demonstrator.
John
Hi JW,
How did expected cost scale with:
1. Number of parts?
2. Chamber Pressure?
3. Fuel / Oxidizer Choice?
4. Engine Cycle?
5. # of Turbopumps?
6. Any other interesting tid bits I may have missed?
Also, how accurate do you believe these models were? To use your example, how do you believe they verified cost scaled at ~(thrust)^(0.7)?
I've always been curious how well these models work and was just wondering what your opinion of them were?
Thanks,
C
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.
- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.
John
Wonderful info, thank you for sharing and I'm sure everyone else here appreciates it as well.
Another question I have for the forum, do we know of any other engines that have been used to date that inject both a gaseous oxidizer and gaseous fuel?
I say gaseous oxidizer and fuel because that's essentially what's coming in as a result of FFSC.
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.
- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.
John
Really great, John. Thanks!
Is that a real picture of the demo engine or just something to use as an illustration? Do you have an inside source that has given you the current throat diameter or is that info generally available and I just missed it?
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.
- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.
John
Really great, John. Thanks!
Is that a real picture of the demo engine or just something to use as an illustration? Do you have an inside source that has given you the current throat diameter or is that info generally available and I just missed it?
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
"Now", only. I await the progressive uprating.
What is this fascination of yours with F-1 class rocket engines?Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
"Now", only. I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
"Now", only. I await the progressive uprating.
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.
John, at what chamber pressure would the 2017 engine produce 1000 kN?
Is it possible that the 1000 kN demo is the same turbopumps and chamber with a lower pressure rating and short nozzle?
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.
- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.
John
Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
"Now", only. I await the progressive uprating.
Why bother? Lots of reliable engines does not increase the risk sufficiently to require big engines, which I think would make the whole system much harder to maintain, and lots more expensive.
I look on it like this. It's fairly easy to work on a small car engine. As they get larger, e.g. a truck engine, they are more difficult to remove and handle, and require more and more specialist handling equipment, and more manpower. Harder to test as well, and also require more expensive production lines.
Not only that, but you start to lose economies of scale. If you are making lots of small engines, your production runs are larger, which usually means cheaper components, which also give more opportunity for manufacturing optimisation.
I can see the Raptor being not much more to build than the Merlin C, once they really get the production line going. Their sizes are fairly similar (from what I understand), there is of course a bit more material and complexity with SC, but I don't think it would be enough to really hike the price up hugely over the Merlin. Happy to be proven wrong on that one though.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.Hopefully SpX will eventually uprate Raptor to F-1 class for future larger ITS systems and/or reducing booster engine no. SpaceX may come up with Plan 3 for the ITS system at next year's IAC conference showing a new design for BFR with fewer larger engines. You know how often EM changes SpX future plans.With a M1D cost range of somewhere between $1-2M the cost factor of ~2.45 for scaling up to a dual turbine 500Klbf Raptor from a 186Klbf M1D gives a cost for Raptor of $2.5 to 5M.
At 31 engines that is a cost in engines for the BFR of $76 to 150M. Putting the cost range of the Booster stage somewhere between $100 to 200M.
Raptor is now only 375 klbf, almost exactly 2x as much thrust as Merlin 1D.
"Now", only. I await the progressive uprating.
Why bother? Lots of reliable engines does not increase the risk sufficiently to require big engines, which I think would make the whole system much harder to maintain, and lots more expensive.
I look on it like this. It's fairly easy to work on a small car engine. As they get larger, e.g. a truck engine, they are more difficult to remove and handle, and require more and more specialist handling equipment, and more manpower. Harder to test as well, and also require more expensive production lines.
Not only that, but you start to lose economies of scale. If you are making lots of small engines, your production runs are larger, which usually means cheaper components, which also give more opportunity for manufacturing optimisation.
I can see the Raptor being not much more to build than the Merlin C, once they really get the production line going. Their sizes are fairly similar (from what I understand), there is of course a bit more material and complexity with SC, but I don't think it would be enough to really hike the price up hugely over the Merlin. Happy to be proven wrong on that one though.
Oxygen rich systems could burn through anything quickly when failed. Engine redundency does not gurantee safty in this case.
Size of Merlin 1C is not big, it could be made to the scale of BE-4 without too much trouble on tooling. This reduce stage 1 engine count to ~15, being much conventional.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.
Approximately one minute and 19 seconds into last night’s launch, the Falcon 9 rocket detected an anomaly on one first stage engine. Initial data suggests that one of the rocket’s nine Merlin engines, Engine 1, lost pressure suddenly and an engine shutdown command was issued immediately. We know the engine did not explode, because we continued to receive data from it. Our review indicates that the fairing that protects the engine from aerodynamic loads ruptured due to the engine pressure release, and that none of Falcon 9’s other eight engines were impacted by this event.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.
This seems like an argument for more smaller Raptors which would fail less energetically.Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
But... there is armor around each engine specifically to avoid such a chain reaction AND the effectiveness of this design has already been demonstrated in flight.Quote from: SpaceXApproximately one minute and 19 seconds into last night’s launch, the Falcon 9 rocket detected an anomaly on one first stage engine. Initial data suggests that one of the rocket’s nine Merlin engines, Engine 1, lost pressure suddenly and an engine shutdown command was issued immediately. We know the engine did not explode, because we continued to receive data from it. Our review indicates that the fairing that protects the engine from aerodynamic loads ruptured due to the engine pressure release, and that none of Falcon 9’s other eight engines were impacted by this event.
The armor has not been proven in an energetic explosion event.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
Raptor Engine Simulation Update. Ran parametric model today on the new Raptor numbers. If you see bad numbers or assumptions please let me know.
- The 2017 Raptor has a throat diameter 16% larger than the Demo engine. And the pump tip speed will need to be about 12% greater.
John
Nice!
Looks like my guess of ~37 and ~120 was pretty close for expansion ratios.
The only thing that may be wrong is I think they switched the mixture ratio from 3.8:1 to ~3.6:1. The spaceship per the presentation holds 860 tonnes of O2 and 240 tonnes of CH4. 860/240 = 3.583
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
There are faaaaar more variables in engine size vs engine count trade-offs... Not just the simplistic "more engines equal more danger" that you operate under.
I'm sure you would have advised SpaceX to not go ahead with with an single F-1 class engine instead of 9 Merlins for the Falcon 9. Thankfully they did not listen to such talk back then.
Smaller engines have higher T/W due to less fluid pressurised volume/shorter flow path lengths - up to some limit set by minimum gauge constraints, and are also likely to have lower thermally induced stresses during start/stop caused by temperature gradients/heat flux through thinner walls - improving long term reusability.
Smaller engines might also have lighter thrust structures, and will lead to shorter interstages and landing legs, and possibly slightly less overall noise as well as greater redundancy, and more rapid throttling (to allow differential throttling steering, with benefits of eliminating flexible joints and actuators + less length), but might also have higher instrumentation and control mass overhead.
Being able to handle smaller engine components by hand will also reduce tooling, manufacturing, assembly and maintenance costs considerably. And of course there are additional learning-curve advantages of making many smaller engines that make Merlin up to an order of magnitude cheaper per unit of thrust than eg RS68.
The ultimate limit to how small is optimal is likely down to the turbomachinery efficiency and turbine inlet temperatures needed to achieve the ~30MPa chamber pressure (limited by reusability issues caused by thermally induced stresses in thrust chamber wall created by through-thickness temperature gradient driven by heat flux). Perhaps there is an additional limit created by Isp hit of making slightly cooler fuel-rich layer near thrust chamber wall.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
I know some here feel like it's a rule of nature, but I think 79 engines is a few too many.
Matthew
Turbomachinery gains in efficiency with size. Cusp losses on the blades decrease along with the square/cube law.
Simple Brayton cycle gas turbines of a few hundred horsepower have thermodynamic efficiencies only in the teens, but large (400MW) gas turbines can reach 40% efficiency on the same cycle.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
Expect this as well they'll either reduce the size of BFS or move to a RD-180 thrust class engine which could be built by making a bigger turbo pump and using two Raptor combustion chambers.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9. More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
Expect this as well they'll either reduce the size of BFS or move to a RD-180 thrust class engine which could be built by making a bigger turbo pump and using two Raptor combustion chambers.
Just keeping Raptor size to that announced in IAC2016 would have allowed the engine no. on BFR to be reduced to 18 which is more sensible than 31 but still more than optimum which is 7-9.
More O2 rich turbopumps on booster = more risk of one burning through causing LOM.
BFS does not have enough sea-level thrust to cope with separation and landing in an abort below some substantial fraction of its normal staging velocity, for good reasons.
In an abort, clearly the engines can be run a little harder, which will push flow separation out a little, and help somewhat, but even that won't help at some point.
Can this cycle of engine in principle be operated dramatically off-mixture, so as to dump either very cold (relatively) gas out of the nozzle (perhaps with damage), or even partially liquid, faster than normal due to the lower pressure drop.
I would assume the answer is no, but was wondering if I was wrong.
[edit] Or ... Well - the rather simpler option of large dump valves.
7-9 is not optimum for the booster as it does not allow engine redundancy during landing.Give Raptor the capability to throttle down to 10% or even less then a 7-9 engined booster can land on 3 engines, 2 engines, or the centre engine only. This would give redundancy on landing without going to some crazy engine no. on 1st stage.
Raptor cant transition to multiple combustion chambers. The turbo machinery for the LOX and its pre-burner is practically integrated into the combustion chamber. Ripping this apart means basically starting from scratch, trashing the thrust to weight ratio and introducing uncountable failure modes. Please stop with this.
7-9 is not optimum for the booster as it does not allow engine redundancy during landing.Give Raptor the capability to throttle down to 10% or even less then a 7-9 engined booster can land on 3 engines, 2 engines, or the centre engine only. This would give redundancy on landing without going to some crazy engine no. on 1st stage.
Raptor cant transition to multiple combustion chambers. The turbo machinery for the LOX and its pre-burner is practically integrated into the combustion chamber. Ripping this apart means basically starting from scratch, trashing the thrust to weight ratio and introducing uncountable failure modes. Please stop with this.
Not necessarily as the RD170,180,and,181 share common heritage and come off the same line though the RD-170 was the first.
The specifics of whether this is practical to do with Raptor is unknown outside of Spacex but it would entitle a larger turbo pump and preburner but such an engine probably can come off the same line as the single chamber ones.
Keep in mind Raptor is still at a fairly early stage there is no full scale engine yet so the specifics can and probably will change.
Here is high-tech mspaint flow chart (as I understand it) of the raptor:
The key to Raptor's high T/W ratio is the minimization of high-pressure piping.
Very nice drawing :) I like it!Very likely Raptor's TWR has not yet been finalized so will not announce it until it is. May also be restricted by ITAR.
How high T/W? why is it not public?
Very nice drawing :) I like it!Very likely Raptor's TWR has not yet been finalized so will not announce it until it is. May also be restricted by ITAR.
How high T/W? why is it not public?
Also notice that SpX have only shown videos of Raptor firings in the dark to deliberately hide the engine's turbomachinery. Again could be down to ITAR.
Here is high-tech mspaint flow chart (as I understand it) of the raptor:
The key to Raptor's high T/W ratio is the minimization of high-pressure piping.
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.Raptor TWR of 600 is likely impossible but it may exceed 250.
Here is high-tech mspaint flow chart (as I understand it) of the raptor:
The key to Raptor's high T/W ratio is the minimization of high-pressure piping.
Doesn't the fuel flow go: first compressor stage -> regen nozzle cooling -> second compressor stage -> fuel preburner -> injectors?
You have it going through the regen channels after all compressor stage, meaning that the regen nozzle has to handle fuel at greater than chamber pressure. That's a LOT of high-pressure piping, which is what you say it's trying to minimize.
If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.Raptor TWR of 600 is likely impossible but it may exceed 250.
If Elon says at IAC 2016 on why such small engines? “similar engine size but 3 times the thrust” I asume T/W is around 600 because Merlin 1D is 200.
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.Raptor TWR of 600 is likely impossible but it may exceed 250.
Why imposible?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.Raptor TWR of 600 is likely impossible but it may exceed 250.
Why imposible?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.If Elon says similar engine size but 3 times the trust I asume T/W is around 600 because Merlin 1D is 200.Raptor TWR of 600 is likely impossible but it may exceed 250.
Why imposible?
And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
Here is what I mean by minimizing high pressure piping;
The weight of pressure vessels scales with the the mass in the the pressure vessel and the pressure.
Here is the RD-180:
http://www.markelwood.com/images/spaceart/RD-180.jpg
Those two big pipes coming out of the top of the turbine are the O2 rich gas pipes. ~73% of the mass flow of the fuel is going through those two big pipes, it looks like it has to travel 2-3 meters before getting to the combustion chamber. In the Raptor, which is even more oxidizer-rich (oxidizer is ~78% of propellant), it only travels inches.
...
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.
Raptor and M1D are completely different engines being FFSC LOx/LCH4 and GG LOx/RP-1 respectively. Raptor no longer has 3x M1D thrust more like 2x now. So your assumption of Raptor TWR of 600 based on Raptor having same size as M1D with 3x the thrust is not credible.Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.
It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could say “same size engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust
They are all bigger or have the same thrust. So why does he say that? Probably he meant by size, wheight, then it makes sense. Otherwise not
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.
It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could easily say “same size engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust
They are all bigger or have the same thrust. So why does he say that? Probably he meant by size, wheight, then it makes sense. Otherwise not
Really... we need to wait if or until SpaceX releases some specs at a later date TBD...
That said...
My guess of a mass of ~980kg (almost 1 metric ton) and a thrust stated at 170 to 190 metric tons
Puts Raptor in about the same thrust to weight ratio as Merlin 1D full thrust... 180 to 1
I will add my thought of "Good Enough" to this... No real need to try and beat that... ;)
200:1 is the planned 300 bar later version... in my opinion...
If Elon says at IAC 2016 on why such small engines? “similar engine size but 3 times the thrust” I asume T/W is around 600 because Merlin 1D is 200.You're taking things way too literally. In this context "similar size" does not mean "equal mass". WTF?
Due to the mass of Raptor's turbomachinery and plumbing. Pipes need to be a minimum thickness to withstand the operating pressures expected in Raptor. If Raptor could be made out of CNT then it's TWR could easily reach or exceed 600.
And have you accounted for the fact that the plumbing is much shorter than Merlin 1D?
How do *you* account for your T/W number of 600? It surely sounds like a number pulled out of thin air. M1D has record breaking T/W, so what - based on existing engines - makes it likely that somehow Raptor has bested that number by over 3 times? It just isn't credible.
It’s mainly based on what was said by Elon, suppose he knows the number is 600, then he could easily say “similar sized engine as Merlin 1D, 3 times the thrust”. I don’t know any Raptor engine which is the same size and has 3 times the thrust?
They are all bigger or have the same thrust. So why does he say that? Probably he meant wheight instead of size, then it makes sense, otherwise not.
But I dont know either, only that its “the highest TWR of any engine, of any kind”, so above 200.
I wonder if any of the Raptor plumbing is carbon fiber overwrapped. It sure saves a lot of mass in COPVs, and not all the fluid flows are high temp.
Carbon-carbon, on the other hand...I wonder if any of the Raptor plumbing is carbon fiber overwrapped. It sure saves a lot of mass in COPVs, and not all the fluid flows are high temp.
No chance, too much mismatch in thermal expansion coefficients, near impossible to do wrapping due to poor accessibility, potential fire danger around oxygen, and carbon fibre doesn't have the ability to handle more than about 200-250°C during reentry.
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
Raptor is on the path of needing to exceed LM ASC engine reliability. That's a tall order to fill. (If those point to point transport on Earth graphics are "real", likely engine reliability would have to approach commercial transport turbofan reliability, which is three orders plus of magnitude higher.) To prove this would require extreme testing/use/reuse.
One could "concern troll" that if AJR/BO can't test to such, then SX couldn't ever do such, omitting the fact that they seem to be able to meet reliability margins above industry norms.
What are the industry norms on development testing? Was the SSME Block III proposal especially gold-plated? Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?
What are the industry norms on development testing? Was the SSME Block III proposal especially gold-plated? Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?
Some things to educate yourself with:
Test and Evaluation Guideline for Liquid Rocket Engines (http://www.dtic.mil/dtic/tr/fulltext/u2/a554916.pdf)
Liquid Rocket Engine Flight Certification (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910018936.pdf)
In general, look at the acceptance criteria of contracts for vehicles engines.
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
What’s the Block 5 Merlin ?
Is it Merlin 1E ?
Is it 145% more heavy?
What are the industry norms on development testing? Was the SSME Block III proposal especially gold-plated? Or do some substitute testing of production engines for development testing -- i.e., like I assume SpaceX did with Merlin 1D?
Some things to educate yourself with:
Test and Evaluation Guideline for Liquid Rocket Engines (http://www.dtic.mil/dtic/tr/fulltext/u2/a554916.pdf)
Liquid Rocket Engine Flight Certification (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910018936.pdf)
In general, look at the acceptance criteria of contracts for vehicles engines.
Very interesting, thanks. One thing I picked up is "Testing should demonstrate margin on maximum specified operating life." If you read that literally and simplistically, then all the claims for BFR booster design life imply an extremely long test program.
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
What’s the Block 5 Merlin ?
Is it Merlin 1E ?
Is it 145% more heavy?
The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9. It is basically the same mass engine AFAIK. The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
What’s the Block 5 Merlin ?
Is it Merlin 1E ?
Is it 145% more heavy?
The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9. It is basically the same mass engine AFAIK. The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'
A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.
John
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
What’s the Block 5 Merlin ?
Is it Merlin 1E ?
Is it 145% more heavy?
The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9. It is basically the same mass engine AFAIK. The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'
A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.
John
Understood, but the original question was asking for the BFS Abort system thread. While running engines at higher than rated values greatly reduces the life of the engine, I am asking whether it may, in a dramatic situation, contribute to an increase in the life of the payload, given what we know about expansion ratios and TWR. :p
Very nice drawing :) I like it!Musk said it was the best ever, so better than M1D.
How high T/W? why is it not made public?
Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?
What’s the Block 5 Merlin ?
Is it Merlin 1E ?
Is it 145% more heavy?
The latest version of Merlin (M-1D) is a standard engine with upgrades for human rating the Falcon 9. It is basically the same mass engine AFAIK. The test to 145% power was a validation of margin, I believe, not a new rating for the engine or even a planned operating 'option.'
A long life engine will experience "plastic creep" of its hot and highly stressed parts. (turbines, combustion chamber coolant passages). Running engines at higher than rated temperatures, pressures and speeds greatly reduces life. Running at these conditions during development testing shows margin and provides information about plastic creep and other failure modes.
John
Understood, but the original question was asking for the BFS Abort system thread. While running engines at higher than rated values greatly reduces the life of the engine, I am asking whether it may, in a dramatic situation, contribute to an increase in the life of the payload, given what we know about expansion ratios and TWR. :p
Falcon 9 Block 5 will have higher thrust than Falcon 9 Full Thrust. We don’t know if the 145%, is higher than planned operation. Since we don’t know how much more thrust the Fuller than Full thrust will be...
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
But it wouldn't contribute to thrust, it would be ejected from the severed regen channels and burn somewhere behind the vehicle.
It would also cut off cooling to the remaining part of the nozzle and the chamber, leading to very rapid engine failure.
So any sort of jettisonable nozzle is going to have to address this anyway by redirecting the coolant pathway.
Edited to add: simply chopping off the nozzle would actually lower T:W because all that fuel is lost rather than going to the combustion chamber. So I would assert that even if it is only for use in dire emergencies, it is essential that any sort of nozzle jettison capability must be accompanied by a redirect of the regen pathway. Not impossible, I'm sure, just an added complication.
Keep in mind spewing unburnt fuel isnt exactly a drawback when you're struggling to raise your TWR as quickly as possible.Question for the thread: I know the Block 5 merlin has been tested to 145% thrust without issue- assuming the raptor had an equivilant level of engineered reserve, Given the expansion ratio of the Raptor vac, could you use the RaptorVacs to get useful emergency thrust during a near-sea level Abort scenerio?Raptor vac. could be designed with a detachable nozzle extension which can be jettisoned in an emergency abort situation. This would allow it to operate at SL along with an emergency thrust reserve to push the ship safely away in an emergency.
Given it is a full regen nozzle, you'd need to set up the plumbing to allow this (upper and lower circuits?) and have valves that shut off in an emergency- otherwise you'll be spewing unburned fuel into the exhaust, within the relatively enclosed interstage area. I'd imagine the base heating would ramp up extremely quickly in this scenario, limiting how long you could burn the RapVacs.
But it wouldn't contribute to thrust, it would be ejected from the severed regen channels and burn somewhere behind the vehicle.
It would also cut off cooling to the remaining part of the nozzle and the chamber, leading to very rapid engine failure.
So any sort of jettisonable nozzle is going to have to address this anyway by redirecting the coolant pathway.
Edited to add: simply chopping off the nozzle would actually lower T:W because all that fuel is lost rather than going to the combustion chamber. So I would assert that even if it is only for use in dire emergencies, it is essential that any sort of nozzle jettison capability must be accompanied by a redirect of the regen pathway. Not impossible, I'm sure, just an added complication.
Does the extension skirt actually need active cooling, or is it radiatively cooled? I'd expect the exhaust to be rather cold when it has expanded 30 times or so...
Will be full regen cooled all the way out to the 3 meter (10 ft) nozzle diameter. Heat flux is nuts & radiative view factor is low.
Historically rocket builders have frequently launched extremely expensive engine development programs in search of larger engines. Were they wrong to do so?Probably not.
Historically rocket builders have frequently launched extremely expensive engine development programs in search of larger engines. Were they wrong to do so?Probably not.
Historically, rockets used analog computers to control the engines. More engines increase complexity more than linearly, which means both heavier and more difficult to design avionics. Think N1 KORD as an extreme and failed example.
Miniaturized digital computers you can program, optical cables, etc, diminish the mass and complexity of such systems significantly, allowing more engines to be used economically. So, probably since the 90s, multiple engine rockets became more viable.
Future BFR’s with 200 or much more engines really wouldn’t surprise me at all.Raptor no. on BFR system has been determined by the need for the ship to have engine out capability for landing and using single engine design throughout the system while keeping complexity to the minimum required level. For future larger BFR's just scale up Raptor thrust with BFR system mass to keep engine no. same as current BFR.
The flight engine design is much lighter and tighter, and is extremely focused on reliability. The objective is to meet or exceed passenger airline levels of safety. If our engine is even close to a jet engine in reliability, has a flak shield to protect against a rapid unscheduled disassembly and we have more engines than the typical two of most airliners, then exceeding airline safety should be possible.
Q: Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?I'm curious what kind of ignition these will use. I would think even sparker ignition might be too slow. Could they have a hot ignition coil in the combustion chamber that stays heated when it is anticipated that they might need to be fired?
A: The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).
Long haul intercontinental jet flights run the engines 100 to 1000 times longer per flight then a P2P rocket would.Quote from: EM AMAThe flight engine design is much lighter and tighter, and is extremely focused on reliability. The objective is to meet or exceed passenger airline levels of safety. If our engine is even close to a jet engine in reliability, has a flak shield to protect against a rapid unscheduled disassembly and we have more engines than the typical two of most airliners, then exceeding airline safety should be possible.
The NASA/Richards document continues to be golden. Thanks again, Space Ghost. It shows a jet fighter engine qualification requirement to be 150 hours (540,000 seconds), or roughly two orders of magnitude more than the original SSME qualification requirement.
The 150 hour requirement also appears to be replicated in the FAA type certification requirements for endurance testing (https://www.ecfr.gov/cgi-bin/text-idx?SID=eed43786296c5051130faf9170d05790&mc=true&node=pt14.1.33&rgn=div5#se14.1.33_149). Perhaps because Raptor only fires for a short time compared to jet engines, the qualification requirements arrived at for Raptor may be less, at least in duration.
Edit: Reliability for modern jet engines seems pretty extreme. GE's G90 powerplant (used on the Boeing 777) is said to have an in-flight shutdown rate of one per million engine flight hours (https://www.geaviation.com/press-release/ge90-engine-family/record-year-worlds-largest-most-powerful-jet-engine).
I don't have any inside information, but I would bet they will use normal spark torch igniters. I have worked with them before, and they have a pretty fast response time that should be more than adequate for an RCS system.QuoteQ: Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?I'm curious what kind of ignition these will use. I would think even sparker ignition might be too slow. Could they have a hot ignition coil in the combustion chamber that stays heated when it is anticipated that they might need to be fired?
A: The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).
I don't have any inside information, but I would bet they will use normal spark torch igniters. I have worked with them before, and they have a pretty fast response time that should be more than adequate for an RCS system.QuoteQ: Will the BFS methalox control thrusters be derived from Raptor or from SuperDraco engines?I'm curious what kind of ignition these will use. I would think even sparker ignition might be too slow. Could they have a hot ignition coil in the combustion chamber that stays heated when it is anticipated that they might need to be fired?
A: The control thrusters will be closer in design to the Raptor main chamber than SuperDraco and will be pressure-fed to enable lowest possible impulse bit (no turbopump spin delay).
How will they get the required tank pressure if they only need a short burst? Seeing as they need to build up some energy in the heat exchanger.
(Unless he was serious about the Harry Potter thing)...
How will they get the required tank pressure if they only need a short burst? Seeing as they need to build up some energy in the heat exchanger.
(Unless he was serious about the Harry Potter thing)...
Are RCS motors fed with gaseous propellants or liquid? I seem to remember someone stating that they would be gaseous, but now I am not so sure.
John
From the Reddit AMA:
Redditor comment:
You can't land on moon using 3MN engine
Elon Musk reply:
Yes, you can. - Bob, the Builder
In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you've lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.
I'm interested in the phrase "a third medium area ratio Raptor engine". Everyone's taken this to mean a third SL engine; but, if so, why didn't he just say so? (Are SL engines referred to as medium area ratio engines?) But, I wonder if he actually meant a third type of engine, with an area ratio between that of the SL and Vac engines. One that will work at sea-level without the usual adverse consequences of attempting to run a Vac engine at those atmospheric pressure, but one that is more efficient than the SL engines when used in regimes of lower atmospheric pressure (though not as efficient as a Vac engine).
In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you've lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.
I'm interested in the phrase "a third medium area ratio Raptor engine". Everyone's taken this to mean a third SL engine; but, if so, why didn't he just say so? (Are SL engines referred to as medium area ratio engines?) But, I wonder if he actually meant a third type of engine, with an area ratio between that of the SL and Vac engines. One that will work at sea-level without the usual adverse consequences of attempting to run a Vac engine at those atmospheric pressure, but one that is more efficient than the SL engines when used in regimes of lower atmospheric pressure (though not as efficient as a Vac engine).
Are they proposing to land on the Moon and Mars using the SL engines? If so, such an engine would be more efficient. How high will the BSF be when it separates from the booster and could such an engine be useful at such an altitude? I'm not a rocket engineer (does it show! :) ), but could there be benefits from having such an intermediate engine?
The obvious argument against is having a third type of engine, with the design and manufacturing complexity etc. I suppose this depends on the level of commonality with the SL and/or Vac engines, and whether any benefits are worth the cost.
I'm interested in the phrase "a third medium area ratio Raptor engine".
I'm interested in the phrase "a third medium area ratio Raptor engine".
Booster might have a lower ER Raptor then the landing Raptors on BFS? ISP doesn't matter as much for the booster, can squeeze more engines on to it with a smaller nozzle and save a bit of weight.
So BFR with low ER raptor, BFS has mid size and vac?
I'm interested in the phrase "a third medium area ratio Raptor engine".
Booster might have a lower ER Raptor then the landing Raptors on BFS? ISP doesn't matter as much for the booster, can squeeze more engines on to it with a smaller nozzle and save a bit of weight.
So BFR with low ER raptor, BFS has mid size and vac?
Blue Origin recently reported a successful test of the B-4. Does this put the B-4 ahead or behind raptor in terms of development? or is it hard to compare given SpaceX's decision to test a subscale engine first?
They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.
They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.
Are you seriously claiming that BE-4 is only running at about 40% thrust currently?
They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.
Are you seriously claiming that BE-4 is only running at about 40% thrust currently?
Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/
They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.
Are you seriously claiming that BE-4 is only running at about 40% thrust currently?
Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/
Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.
They're both at about the same thrust currently, though this will change as duration and power level increases with next steps.
Are you seriously claiming that BE-4 is only running at about 40% thrust currently?
Eric Berger's article at ARSTechnica says the test was at 50% thrust: https://arstechnica.com/science/2017/10/blue-origin-has-successfully-tested-its-powerful-be-4-rocket-engine/
Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.
Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.
Yes, there is a big difference. Running subscale at 3000 psi is way harder than full scale at 1000 psi.
SpaceX gets another $40.8 million in Pentagon funding for Raptor engine
Does any one have more on this tweet earlier today:QuoteSpaceX gets another $40.8 million in Pentagon funding for Raptor engine
https://twitter.com/R_Wall/status/921257396797870080 (https://twitter.com/R_Wall/status/921257396797870080)
The tweet doesn't appear to be a reply to anything else, sounds like new money?
Edit: forgot to say that Robert Wall is aerospace reporter for WSJ
Does any one have more on this tweet earlier today:QuoteSpaceX gets another $40.8 million in Pentagon funding for Raptor engine
https://twitter.com/R_Wall/status/921257396797870080 (https://twitter.com/R_Wall/status/921257396797870080)
The tweet doesn't appear to be a reply to anything else, sounds like new money?
Edit: forgot to say that Robert Wall is aerospace reporter for WSJ
ULA has agreed to initially add $40.8 million under the terms of the government award.I wonder if some wires have been crossed.
I really hope this 40 million doesn't reignite Raptor upper stage fever.
Related to this?
https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/ (https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/)
Related to this?
https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/ (https://www.dodbuzz.com/2017/10/06/air-force-seeks-next-gen-launch-vehicles-for-space-missions/)
Space Exploration Technologies, Corp. (SpaceX), Hawthorne, California, has been awarded a $33,660,254 other transaction agreement for the development of the Raptor rocket propulsion system prototype for the Evolved Expendable Launch Vehicle (EELV) program. This agreement implements Section 1604 of the Fiscal Year 2015 National Defense Authorization Act, which requires the development of a next-generation rocket propulsion system that will transition away from the use of the Russian-supplied RD-180 engine to a domestic alternative for National Security Space launches. An other transaction agreement was used in lieu of a standard procurement contract in order to leverage on-going investment by industry in rocket propulsion systems. This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles. The locations of performance are NASA Stennis Space Center, Mississippi; Hawthorne, California; and Los Angeles Air Force Base, California. The work is expected to be completed no later than Dec. 31, 2018. Air Force fiscal 2015 research, development, test and evaluation funds in the amount of $33,660,254 are being obligated at the time of award. SpaceX is contributing $67,320,506 at the time of award. The total potential government investment, including all options, is $61,392,710. The total potential investment by SpaceX, including all options, is $122,785,419. This award is the result of a competitive acquisition with multiple offers received. The Launch Systems Enterprise Directorate, Space and Missile Systems Center, Los Angeles Air Force Base, California is the contracting activity (FA8811-16-9-0001)..
...QuoteAir Force fiscal 2015 research, development, test and evaluation funds in the amount of $33,660,254 are being obligated at the time of award. SpaceX is contributing $67,320,506 at the time of award The total potential government investment, including all options, is $61,392,710. ..
Still... there is a big difference between a fullscale engine being tested at a lower thrust level and a subscale development engine.
Yes, there is a big difference. Running subscale at 3000 psi is way harder than full scale at 1000 psi.
This. They have not even reached the chamber pressures that M1D runs at, much less Raptor or RD-180. Also livingjw's calculations show that it's not a very big scale-up that SpaceX needs, only 15%.
$61M-$33M does not equal $40M. I have not looked if more money has been allocated.
Air Force adds more than $40 million to SpaceX engine contract
by Jeff Foust — October 21, 2017
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.
What I find typical is that: 33.6mln + 67.3 mln = 100.9mln development cost for 1MN raptor.
?what was the prometheus engine going to cost?
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.I fail to see any evidence of which to draw such an extreme conclusion from.
What I find typical is that: 33.6mln + 67.3 mln = 100.9mln development cost for 1MN raptor.
?what was the prometheus engine development going to cost?
The locations of performance are NASA Stennis Space Center, Mississippi; Hawthorne, California; and Los Angeles Air Force Base, California.
Work will be performed at NASA Stennis Space Center, Mississippi; Hawthorne, California; McGregor, Texas; and Los Angeles Air Force Base, California[.]
This is proof that Raptor is way behind even ULA's AR-1 engine. They have to start power-pack tests again for the full scale engine at Stennis. Full Raptor development is going to take at least 2 years.
How did you get to 105 M$ investment from US Air Force ? They only communicated on 33+40 M$ contracts to SpaceX. And only 66 M$ investment from SpaceX has been confirmed so far in the frame of this OTA with USAF. Do you have complementary information ?The original AF commitment of the contract was $95M over a execution period covering 3 years. This option execution just made some modifications by adding McGregor and increasing the total by ~$8M by upping the amount on this option from $33M to $40.7M. Making the new total of AF funding increasing from the $95M to ~$105M I think the new value may really be closer to $103M.
Wow, I poked a hornet's nest of dickishness on Twitter. Really disappointed at how consistently arrogant Blue's own purported engineers are, especially publicly so on Twitter.
In response to Jeff Foust's simple Raptor funding article (http://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/), a BO propulsion engineer commented, "Or they can pay $0 for a more reliable engine that produces more thrust and already tested at full scale 🙃". An extraordinary statement that is really hard to rationally parse for an engine that has fired for no more than 3 seconds at half thrust and experienced at least one serious failure during testing.
Another BO employee chimed in, "How exactly is a subscale version of an engine "ages closer" to flight readiness than a full scale version of an engine?"
Me: "That 1MN Raptor has been tested for 100s nonstop and > 1200s total should be self-explanatory"
Me: "I would also be a fool to totally discount a company's CTO saying that it is "simple to scale the dev Raptor to 170 tons""
BO guy: "You'd also be a fool to just blindly believe everything that person says when they've proven to not do things when they say they will."
BO guy: "But since you admittedly have no tech expertise, sure just believe what others say. I have technical expertise and know it's not that simple"
It doesn't exactly take a genius to understand that Musk has a habit of understating the difficulty of doing relatively hard things, but both of these BO engineers were dead-set on a single 3s 50% thrust firing of a full-scale engine indicating that BE-4 was somehow closer to flight-readiness than subscale Raptor, with (probably multiple) successful ~100s hot-fires and more than 1200s total. It boggles the mind.
I really want to cheer on Blue Origin but s*** like this makes it rather difficult to support a company with such a seemingly arrogant culture. These are anecdotes, of course, I can only hope that they are representative of a tiny minority. It's almost as if Jeff "Welcome to the club" Bezos managed to only hire clones of himself...
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.
and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.https://twitter.com/SciGuySpace/status/921106486272675840
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
That's okay. He doesn't actually follow SpaceX very closely.I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.https://twitter.com/SciGuySpace/status/921106486272675840
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.https://twitter.com/SciGuySpace/status/921106486272675840
I think Blue will be able to compete eventually. I just think SpaceX is ahead with BFR.Wandering somewhat OT. But anyway, if BFR was just a much much larger version of F9 then I would agree that BFR would be ahead and would get top flight first. But BFR is much more complex with more testing gates to successfully pass than what NG has to.
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.
and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.
I don't know if you really understand what he's saying:
What he's saying is that the Blue Origin engine is full size ... Combustion Chamber, injectors, preburner injectors, pumps, everything is full size, that requires no drawing changes (in theory) when they go to flight, full thrust SHOULD be as easy as opening up the fuel valve to the preburner to let the turbopump spin faster... i don't know if this is what they're using to alter pump speed (inevitably engine power level) or not... whether this happens or not is yet to be seen
SpaceX's Raptor is: what, help me out here ... 80% geometrically the size of the flight engine size, this means new part numbers for the combustion chamber, injector, preburner(s) injectors, pumps, if i'm understanding what Elon has said correctly, everything has to be geometrically scaled up to reach flight engine size, that is not a small task, also, dynamic similitude in fluid mechanics doesn't mean you multiply or divide everything by 0.80 ...
^ my $0.02
C
Between SX and BO, the former has launched multi-engine, multi-stage rockets to LEO. And while SX has developed almost all the key elements for BFR/BFS into LEO [excluding the refuel maneuver for other uses] in some form or other, BO has done little. Maybe BO can compete with SLS ME2, maybe they will all be close when the combusted fuel hit the launch pad, maybe they will spread over many years, maybe some will fail. But Berger is out on a fishing expediting for bad analysis.
I think Blue will be able to compete eventually. I just think SpaceX is ahead with BFR.
I tell ya, if ULA start flying a vehicle with a Raptor engine I'm going to have to go buy some new ice skates.
Does the new money the Air Force possibly invest in Raptor mean SX can progress faster?Only if they were ressource constrained before. Otherwise I would go with the old adage: Throwing manpower at a late project makes it later.
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
Does the new money the Air Force possibly invest in Raptor mean SX can progress faster?Only if they were ressource constrained before. Otherwise I would go with the old adage: Throwing manpower at a late project makes it later.
FWIW, I expect we'll see full scale testing begin before the end of 2017.
FWIW, I expect we'll see full scale testing begin before the end of 2017.
If they had a full-scale engine almost ready for testing I expect Musk would have shown off some pictures of full-scale hardware at IAC. The fact that he didn't might suggest that they are still a ways off.
FWIW, I wouldn't be surprised if we didn't see a full scale Raptor test fire until late 2018 or even 2019. These things take time.
Although Raptor will help revolutionize access to space it will not cause the death of all other (new) launch systems.I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
Unmanned SLS will get there first. Then you'll see manned BFS on Luna and SLS will never fly again.
Most likely none of those others will ever fly again. Why sail galleons when container ships suddenly show up?
Although Raptor will help revolutionize access to space it will not cause the death of all other (new) launch systems.Off topic, but I think very true, other nations will not cede space to SpaceX, they will invest what is needed to catch up. That is the really exciting time. Landing on mars is not as fundamental a milestone to me as the moment we see China test their first grasshopper, and the age of reusable rockets is here no matter how badly SpaceX may stuff up in any future endeavour.
For example: Ariane 6 will fly, and after it a next generation Ariane vehicle will as well. The reason is simple: Europe wants it's own independently assured access to space.
I bet BFS will get to space before New Glenn or Vulcan or Ariane 6 or SLS.
Unmanned SLS will get there first. Then you'll see manned BFS on Luna and SLS will never fly again.
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.
Although Raptor will help revolutionize access to space it will not cause the death of all other (new) launch systems.
For example: Ariane 6 will fly, and after it a next generation Ariane vehicle will as well. The reason is simple: Europe wants it's own independently assured access to space.
When the original Ariane vehicle was being developed there was a lot of pressure from the United States to stop that development. The thinking was that Europe could get all the launch services they ever needed by buying them from the United States.
Europe developed Ariane regardless, despite that seeming to be the more expensive option.
The same applies to China.
So, once BFR/BFS is flying, there will still be vehicles such as Ariane 6 and Long March (insert a number here).
...I think very true, other nations will not cede space to SpaceX, they will invest what is needed to catch up. That is the really exciting time. Landing on mars is not as fundamental a milestone to me as the moment we see China test their first grasshopper, and the age of reusable rockets is here no matter how badly SpaceX may stuff up in any future endeavour.
You say it like it's indicative of incompetence... no private company has ever put a manned spacecraft into orbit ever, and in the case of BFS/BFR SpaceX can set their own requirements and distribute their milestones in developing a manned ship as they like.Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.
I think that's optimistic. I am confident BFR/RFS are going to succeed, but let's remember, SX hasn't even put a manned Dragon into orbit yet.
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.
Robotbeat said BFS going to space. Which would be an early test flight of the SSTO kind. I fully expect that to happen in 2019 or early 2020. It may very well be BFS first. Though for the full stack I agree. It will likely be SLS.
I suspect Elon means sub-orbital flight testing, with the first orbital flights not happening until they've ironed out the kinks.
To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.
and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.
Wrong take away.To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.
and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.
Falling in to the same Dunning-Kruger hole. You are calling them arrogant and irrational, and yet, by your own admission, THEY are the experts. Which probably means they are not irrational or arrogant, but in fact know more about it than you/other commentators do. I KNOW they know more about it than I do, which is why I let them design rocket engines without me putting my oar in.
ISTM the test Raptor falls squarely within the definition of a prototype. This from wiki:To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.
and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.
I don't know if you really understand what he's saying:
What he's saying is that the Blue Origin engine is full size ... Combustion Chamber, injectors, preburner injectors, pumps, everything is full size, that requires no drawing changes (in theory) when they go to flight, full thrust SHOULD be as easy as opening up the fuel valve to the preburner to let the turbopump spin faster... i don't know if this is what they're using to alter pump speed (inevitably engine power level) or not... whether this happens or not is yet to be seen
SpaceX's Raptor is: what, help me out here ... 80% geometrically the size of the flight engine size, this means new part numbers for the combustion chamber, injector, preburner(s) injectors, pumps, if i'm understanding what Elon has said correctly, everything has to be geometrically scaled up to reach flight engine size, that is not a small task, also, dynamic similitude in fluid mechanics doesn't mean you multiply or divide everything by 0.80 ...
^ my $0.02
C
Thanks, your thoughts are appreciated.
I completely agree, and that's largely how I understood the situation. Scaling up both physical dimensions and chamber pressure by 15-25% is not said and done by any means, and the complexity of RPS and plumbing necessitate that it will be more difficult than "enlarging the CAD model by 15%", as one of the BO employees condescendingly suggested.
Howeverrrrr, I also have little doubt that SpaceX has been iterating and exploring full scale Raptor hardware during the 12+ months they've been testing its scaled prerequisites, thus learning many lessons about running an integrated 1MN methalox FFSC engine. Dozens of times and at considerable duration, as well. (Also some L2 info that strengthens this feeling, but can't say more)
Given how little Raptor will have to grow to reach its current operational performance specs, as well as SpaceX's vast (compared to BO) experience producing rocket propulsion systems, it seems implausible to say that BE-4 is closer to flight readiness because they successfully fired a full sized engine for 3 seconds, after suffering at least one major hardware failure.
Another main difference I perceive simply lies in SpaceX's decision to begin with subscale testing. They've developed some level of expertise with Raptor, even if it may not all remain applicable after scaling thrust by an additional 70%. BO has a sum total of 3 seconds of experience testing an integrated engine, even if it's full scale. Their test program could proceed utterly flawlessly, but that seems improbable. I'm sure SpaceX has had to deal with many issues with scale Raptors over 40+ tests, and I would bet money that a lot of the lessons learned with scale Raptor will transfer to full scale testing.
Again, I am self-admittedly not a technical expert. I don't currently have time to do so, due to school, but my hope is to build a decent foundation of the basics of rocketry and RPS when I have the free time. What minimal reading I've done has informed the above opinions, and I welcome any and all criticisms and corrections, as well as complete refutation. Just trying to better understand things and tweak my intuition along the way.
(trimmed)
I'd bet money that full scale hardware ready for testing already exists and full scale preburner testing has already begun.
But just pure speculation at this point. That's it from me!
I don't know, the jump from a subscale, lower chamber pressure, development Raptor to a full scale Raptor with ~2x the thrust seems a lot more like the jump from M1C to M1D, rather than the jump from a prototype M1D to a production M1D.
(trimmed)
I'd bet money that full scale hardware ready for testing already exists and full scale preburner testing has already begun.
But just pure speculation at this point. That's it from me!
Do we expect that the larger pre-burner unit will require a new round of testing? Would that likely occur at Stennis again? Is that something that the public could gain insight into if it is underway or has already completed?
Stennis has unique capabilities. The test stand delivers hot gases and allows for tests of preburners or fuel injection by itself. A normal teststand like in McGregor can only test full engines or at least full powerheads.
My understanding was that the components tested in Stennis used the full capacity and larger preburners or fuel injectors could not be tested there. I may be wrong on this one.
Since the final thrust level of the Raptor had not been settled, it was decided that the first integrated test engine would be a 1MN sub-scale engine.
It enabled the full testing at Stennis E2 and allowed for the development of robust startup and shutdown sequences, characterize hardware durability and anchor analytical models that would be used for future designs.
Once the final engine thrust was defined, the engine could be scaled up with relative ease. The full flow cycle is very helpful in that sense and the 1MN thrust level would already be considered a big engine.
With the production engines – as currently envisioned – it would need to triple its thrust. Not trivial, but still within what could be considered highly representative as a demonstrator.
Wrong take away.To me it just sounds like actual rocket engineers being cheesed off that they've been told what's what by armchair rocket engineers.
I fully agree that I am a complete non-expert in comparison to actual propulsion engineers, but that doesn't excuse the highly irrational and arrogant attitude towards Raptor. Even if scaling thrust by ~70% is far more difficult than SpaceX's RPS engineers believe it to be, almost completely discounting 1200 seconds of hot-fires, half a decade of Merlin 1D mass production, and orbital and vacuum rocketry experience fly directly counter to the ideals a functional and rational engineer/scientist ought to hold.
and FWIW, I did not start the discourse. The trash talking was begun unsolicited by a BO propulsion engineer.
Falling in to the same Dunning-Kruger hole. You are calling them arrogant and irrational, and yet, by your own admission, THEY are the experts. Which probably means they are not irrational or arrogant, but in fact know more about it than you/other commentators do. I KNOW they know more about it than I do, which is why I let them design rocket engines without me putting my oar in.
The BO guys know all about the BO engine (BE-4)
The SpaceX guys know all about the SpaceX engine (Raptor)
The BO guys do NOT know all about the SpaceX engine (Raptor)
The SpaceX guys do NOT know all about the BO engine (BE-4)
Any time that a BO guy says that their engine is better (or further ahead in development) than the SpaceX engine he is stating an ASSUMPTION.
Any time a SpaceX guy says their engine is better (or further ahead in development) than the BO engine he is stating an ASSUMPTION.
Anyone else, armchair engineers included, best not comment on aspect like "better" or "further along in development" at all.
And yet they are all rocket engine engineers, which gives them a greater insight in to the issues involved, whether at BO or SpaceX. Of course there are assumptions, but they are very educated assumptions from people who are experts in the field, which should give them greater validity than most of the people who comment on it, including myself.They are also extremely biased (as they should be!) and posting off-the-cuff (as you do) on twitter. Frankly, it's juvenile and embarrassing, and not a good look. Fortunately the engine looks great and I look forward to SpaceX and BO trying to one-up each other, hopefully for some time to come.
Now, can we all get back to Raptor, pretty please? (Not intended to call out anyone specifically, and I include myself in that request).
Now, can we all get back to Raptor, pretty please? (Not intended to call out anyone specifically, and I include myself in that request).
Yes please.
Does anyone tellif the Raptor test bay in McGregor that is being outfitted is maybe being out fitted for pre-burner testing of the full scale raptor?
That would seem like a logical place to conduct that testing.
The real question here, is which engine will create the best rocket?
A few heavy 2400 kN BE4’s or many light 1700 KN Raptors?
We miss some data to answer that question, for instance the mass of both engines.
Thats probably question number one, the airforce wanted to know...
Now, can we all get back to Raptor, pretty please? (Not intended to call out anyone specifically, and I include myself in that request).
Yes please.
Does anyone tellif the Raptor test bay in McGregor that is being outfitted is maybe being out fitted for pre-burner testing of the full scale raptor?
That would seem like a logical place to conduct that testing.
There's some tangential L2 info related to your question.
Unrelated to the L2 info, given how successful the subscale test program has been, I wouldn't be surprised if SpaceX jumped headfirst into full scale integrated testing. Could start with a burp test this time around, if they want to be extra cautious.
Good point, Build, test, iterate is how EM likes to work.
If the testing results have closely matched the design and expectation then jumping to a more complete next step may work well.
The Raptor is being developed in similar manor that the M1D was done.
First there was a prototype that was used to test assumptions and update the design models. Then the updated design models were used to create a production design with a resulting fairly accurate physical set of specifications in thrust ISP and weight. What happened was the prototype was a design with a probability of a higher level of successful operation. But then the data gathered is used to "tighten" the engine design models such that when entered a set of constraints like thrust, bell size, ISP, TC pressure, and a few others the design software developed will generate parts drawings that has a high probability to create an engine with almost those exact specifications that will work!
Thus at the end of the prototype testing a specification for a production engine which was delivered 4 months later actually met the design specifications. It also had a near trouble free testing.
So to is the Raptor strategy. The Raptor is at a similar point that the M1D was at in April 2012. And that less than 6 months later production engines were proceeding successfully thru qualification testing and flight units were being delivered and successfully acceptance tested.
Then just a little more than 2 years from the April 2012 point a F9v1.1 flew.
So it is possible that a tanker/cargo version of the SC could start into tests late 2019 or early 2020.
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
Is there a poll as to when we expect to see a full scale production Raptor engine?
Yes, I meant 135 bar, 67 is the test pressure. Fixed my post.4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.
Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.
Sorry if this is a wrong thread or this has been discussed already...
I was wondering about using BFS as it's own escape system and I understand one of the problems is that complex engines with turbopumps need more time to start running. But what order of magnitude are we talking about? What could be the possible minimum time they might achieve for Raptor startup sequence? Miliseconds, seconds, a minute?
Raptor measurements and specs are earlier in this thread as well as on L2. It is not much bigger than Merlin, so it would seem to be much less mass than BE-4. A BFR with 31 of these subscale engines would get it in Saturn V range. Probably 75 to 100 tons to LEO with a full up BFR/BFS with subscale engines.
I added a poll! Feel free to let me know if you think I missed important options.
I added a poll! Feel free to let me know if you think I missed important options.
I don't see a poll in the polls section yet.
4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.
Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.
If you are going for a "wisdom of the crowds" estimate it works better if it's done blind. Otherwise you get herding.4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.
Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.
Can anyone with engine design background estimate the weight of both engines?
My layman estimation is:
400kg for Raptor 1000kN (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)
Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
If you are going for a "wisdom of the crowds" estimate it works better if it's done blind. Otherwise you get herding.4x demo Raptors, if operating at 200 bar and 1,000 kN as claimed, would almost certainly outperform 2x BE-4 at full design specs (67 bar and 2,450 kN) if both were pushing the same booster and upper stage.
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.
Based on what data? As far as I know the specific impulse and weight of the dev Raptor have not been published, likewise with BE-4.
Also, BE-4 chamber pressure is 135 bar, not 67. Although either way Raptor is still a much more aggressive design.
Can anyone with engine design background estimate the weight of both engines?
My layman estimation is:
400kg for Raptor 1000kN (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)
Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
I could have answered I’m part of the crowd, but I don’t consider the members of this particular discussion forum having a herd mentality. Quite the (sometimes annoyingly) opposite actually :)
Suggest you look up the published dry weight of the Merlin 1D SL and maybe rethink your numbers...
That said... I have already posted (somewhere here) a dry weight guess of ~980kg for a single SL Raptor (1700kN) as fitted to BFR...
On edit...
Source of said guess I made...
https://forum.nasaspaceflight.com/index.php?topic=41363.msg1732978#msg1732978 (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1732978#msg1732978)
Elon said he believes the Raptor will have the best T/W ratio of any engine ever made, but said it hesitantly as if it is only just true. Their M1D holds the title currently at 180:1 (well maybe as high as 199:1 if the weight hasn't grown with the block 5 thrust upgrade). That implies to me that Raptor will be close to that but slightly better. Considering Raptor is as complex an engine as it is, it should have a lower thrust to weight (see the SSME at ~54:1), so to be better than the comparatively dead simple M1D is incredibly impressive already.
If you assume a minimum of 200:1 T/W and the a thrust of 1,900kN, that puts Raptor's max weight around 975kg.
Elon said he believes the Raptor will have the best T/W ratio of any engine ever made, but said it hesitantly as if it is only just true. Their M1D holds the title currently at 180:1 (well maybe as high as 199:1 if the weight hasn't grown with the block 5 thrust upgrade). That implies to me that Raptor will be close to that but slightly better. Considering Raptor is as complex an engine as it is, it should have a lower thrust to weight (see the SSME at ~54:1), so to be better than the comparatively dead simple M1D is incredibly impressive already.
If you assume a minimum of 200:1 T/W and the a thrust of 1,900kN, that puts Raptor's max weight around 975kg.
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.
I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.
140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.
What do you think is the weight of the 1000kN sub scale engine?
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.
I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.
140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.
I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.
140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.
https://en.wikipedia.org/wiki/RD-270
N2O4/UDMH 190:1 thrust to weight ratio full flow stages combustion at 26MPa chamber pressure in 1969 Russian design. 2018 SpaceX can probably do better than that given advances in analytical tools with their thrust-chamber top integrated LOX turbopump even if LOX/CH4 bulk fuel density is only around 0.82 vs 1.12 for UDMH/N204
Yeah... just look at the RD-180 with 80:1 T/W. if Raptor ends up anywhere close to the current Merlin T/W, it will be an *incredible* achievement.
I suspect their end goal for Raptor is 190-200 T/W but I'm betting that the dev Raptor is no where close to that, and the first iteration of flight Raptor won't achieve it either.
140-160 seems a lot more realistic, and still *very* impressive for a staged combustion engine. I suspect the mass of flight Raptor will start at around 1300kg.
https://en.wikipedia.org/wiki/RD-270
N2O4/UDMH 190:1 thrust to weight ratio full flow stages combustion at 26MPa chamber pressure in 1969 Russian design. 2018 SpaceX can probably do better than that given advances in analytical tools with their thrust-chamber top integrated LOX turbopump even if LOX/CH4 bulk fuel density is only around 0.82 vs 1.12 for UDMH/N204
I'm not sure if they actually achieved that, though.
There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.
http://www.russianspaceweb.com/ur700.html
Всего за период с октября 1967 г. по июль 1969 г. было проведено 27 огневых испытаний двигателя РД-270. Все они имели аварийный исход.
In total, during the period from October 1967 to July 1969, 27 fire tests were performed on the RD-270 engine. All of them had an emergency outcome.
There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.
http://www.russianspaceweb.com/ur700.html
there's a thread dedicated to this engine on Novosti Kosmonavtiki forum:
http://novosti-kosmonavtiki.ru/forum/messages/forum13/topic4559/message176620/#message176620
you may try google translate it; it states that all test fires ended in failures, sometimes serious. All they achieved is a startup at reduced pressure and smaller bell.
Author of the message on forum claims to quote it from official Energomash history book
Exact russian quote:QuoteВсего за период с октября 1967 г. по июль 1969 г. было проведено 27 огневых испытаний двигателя РД-270. Все они имели аварийный исход.
Google translate:QuoteIn total, during the period from October 1967 to July 1969, 27 fire tests were performed on the RD-270 engine. All of them had an emergency outcome.
There is depressingly little info available that isn't written in Russian. What little there is can be found here, suggesting that dozens of successful test fires were done with a bunch of test articles, 9 tests of which were apparently flawless. So they likely hit their goals if those test articles were full scale.
http://www.russianspaceweb.com/ur700.html
there's a thread dedicated to this engine on Novosti Kosmonavtiki forum:
http://novosti-kosmonavtiki.ru/forum/messages/forum13/topic4559/message176620/#message176620
you may try google translate it; it states that all test fires ended in failures, sometimes serious. All they achieved is a startup at reduced pressure and smaller bell.
Author of the message on forum claims to quote it from official Energomash history book
Exact russian quote:QuoteВсего за период с октября 1967 г. по июль 1969 г. было проведено 27 огневых испытаний двигателя РД-270. Все они имели аварийный исход.
Google translate:QuoteIn total, during the period from October 1967 to July 1969, 27 fire tests were performed on the RD-270 engine. All of them had an emergency outcome.
Can anyone with engine design background estimate the weight of both engines?
My layman estimation is:
400kg for Raptor 1000kN (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)
Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
Can anyone with engine design background estimate the weight of both engines?
My layman estimation is:
400kg for Raptor 1000kN (Isp lower than 330)
450kg for Raptor 1700kN (Isp= 330 s and 356 s in vacuum)
580kg for Raptor 1900kN (Isp= 375 s)
1800kg for BE4 2400kN (Isp = ?)
Individually it’s quite hard to have a good estimation.
But collectively we could get close to the real values
No, you don't get closer to the truth based on just opinions and data pulled from thin air. (Like your numbers) More people doing the same doesn't make the conclusion more valid.
Rather unintuitive , but there is actually some veracity to the "wisdom of the crowd" concept (https://www.ncbi.nlm.nih.gov/pmc/articles/PMC3107299/). The effect is far more pronounced when in highly selective groups like these forums, too :)
Nevertheless, this thread 100% should lean more towards discussion based on available data (like mass estimates from known thrust and TWR figures above) than base speculation.
The only reason that Raptor isn't MUCH closer to flying than BE-4 is because Raptor's design goals are so much higher. It's already a better RD-180 replacement.You can't say this yet.
...
BE-4 will likely eclipse RD-180 (a very hard act to follow) in all ways, including in vacuum thrust.
...
That's the current goal....
BE-4 will likely eclipse RD-180 (a very hard act to follow) in all ways, including in vacuum thrust.
...
Not sure where this comes from... 2x BE-4 about equals RD-180, and BE-4 is shooting for less extreme operating parameters across the board with respect to RD-180.
What do you mean by eclipse -- replace, or exceed RD-180 performance cross the board? If the latter, I believe Blue has stated otherwise.
...
But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.
...
But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.
Understand where you're going with this, but don't think Blue has the push-the-tech-to-the-limit DNA that is driving Raptor.
Started with much lower goals and will end with much lower performance, IMO.
Whatever you mean by 'effectiveness' (T/W, ISP, ?) will be lower, too -- lower than RD-180 and Raptor.
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.Why is a FFSC higher chamber pressure than a OSRC?
I think I know the answer.Full flow means you are extracting more energy for pumping.QuoteThese guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
Advantages of a full-flow staged combustion cycle engine system (https://arc.aiaa.org/doi/abs/10.2514/6.1997-3318)I think I know the answer.Full flow means you are extracting more energy for pumping.QuoteThese guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
I think I know the answer.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
The aforementioned test stand pic.
The aforementioned test stand pic.
Full-scale engine?
The aforementioned test stand pic.
Full-scale engine?
Re: cold turbine exhaust temperatures.
It appears to me you want the turbine exhaust at all thrust regimes to be above the condensation temperature of steam. At 20..30 bar steam should condense below 230..240 celsius.
After the preburner/heater the propellant is a mix ture including combustion products.
No?
From my understanding water droplets in a gas turbine is a bad thing, abrasive blasting or some such effect.
Again I must say... the secret sauce has got to be the light off sequence... 8)
Near the end of this video, 2 new clips of raptor including startup (18 video frames between the start of what is shown and main ignition)
https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s (https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s)
Near the end of this video, 2 new clips of raptor including startup (18 video frames between the start of what is shown and main ignition)
https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s (https://www.youtube.com/watch?v=TXYh4re0j8M?t=2m45s)
Check the comment tree:
Elon Musk: Can confirm, pretty cool place.
Jeff Bezos: Want me to show you a real rocket?
Check the comment tree:
Elon Musk: Can confirm, pretty cool place.
Jeff Bezos: Want me to show you a real rocket?
Is that a green flash when the Raptor starts up? TEA/TEB shot?Isn't it spark ignited?
I think it's TEA-TEB. Spark igniters are external
I think it's TEA-TEB. Spark igniters are external (?right?) and there's no evidence of one present. There's also no illumination that would suggest sparks prior to ignition. SpaceX is very comfortable with and experienced at using TEA-TEB. The idea they would use it for the prototype engine seems very unsurprising to me as a result.
As far as the green flash in the middle of the previous fire, maybe some residue somehow? That seems unlikely. Or a leak. That would be troubling, not really for the engine itself, but more for the test program. So that's definitely a data point going the other direction.
Could the green be a very slight engine rich combustion?
Could the green be a very slight engine rich combustion?
Most likely it's just a camera artifact from the sudden increase in brightness.
Could the green be a very slight engine rich combustion?
Most likely it's just a camera artifact from the sudden increase in brightness.
I think so too. I remember distinctly that Elon said in the 2016 presentation that Raptor has a little torch inside that is spark ignited which in turn ignites the main combustion cycle. Creating a prototype with hypergolics makes no sense since this is one of the core problems in engine development.
[...]
The LOX and RP-1 tanks are pre pressurized with helium.
High pressure helium spins up the turbo pump. LOX and RP-1 are ignited by TEA-TEB in the gas generator and takes over from the helium. The propellants meet in the combustion chamber and are also ignited by TEA-TEB.
[...]
They said it uses autogenous pressurization, so use some of those gases.
Spinning the wheel a bit further (pun intended), how does the Raptor actually start? I mean, spark ignition or not, it needs to spin up its turbines. Following the ongoing discussion on the Merlin:[...]
The LOX and RP-1 tanks are pre pressurized with helium.
High pressure helium spins up the turbo pump. LOX and RP-1 are ignited by TEA-TEB in the gas generator and takes over from the helium. The propellants meet in the combustion chamber and are also ignited by TEA-TEB.
[...]
But the Raptor doesnt have high pressure helium available. Its tanks are autogenous pressurization. So how do the turbine wheels of Raptor start? I do have ideas how it could be done but I dont want to wildly speculate. Does anyone has info on that?
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
I would expect the combustion chamber and turbine would be wider than the pipe into the combustion chamber and pump connected to the turbine, so the shock wave only applies a few square CM of the pressure wave back toward the tanks, but many more times the pressure foreward toward the combustion chamber.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
Thats exactly what I am interested in. So initially, the propellant flows through the not-jet-rotating pumps until it reaches the preburner, is than ignited. It therefore puts pressure onto the turbine which starts to turn. But at the same time, the preburner also puts pressure back up the pumps and into the tanks. Because the pumps are not yet rotating. They are about to start rotating but they dont do it yet. It looks to me like a hen and a egg problem. How can you start the turbines/pumps under these conditions? Are there valves in front of the preburner that quickly close once some propellant is in the preburners and push it out the turbine only to open a fraction of a second later to allow new fuel to reach the preburner and further turn the turbine? And now my thought process looks like a moebius strip...
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.
It doesn't increase the flow rate (mass-per-second). It does increase the velocity of the propellants.Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.
It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure?
Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.Tank head start is possible, but it's a slow process that could easily give rough starts/problems with startup sequences. Venting Helium (or other start gases) through the turbine is something that can be precisely controlled, is highly predictable and spins up the turbine very quickly.
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.
It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.
It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.
- My mistake. I should have said: Everything downstream of the PRE-BURNER has a lower total pressure. The static pressure will rise in the pre-burner as combustion products back pressure the turbine. But, it cannot increase above the pressure upstream of the pre-burner injectors.
- The start mode I outlined is what NASA SP-125 (pg 68) calls as "main tank head start". If this type of start takes too long (> 3 seconds or so) a "turbine spin start" may be added to the system to decrease the start time. I do not know which method the Merlin uses. I would guess a "main tank head start". Does anyone know?
Also see SP-125 pg 181 for different types of starts.
I need to digest all this ...
Everything downstream of the pumps has a lower total pressure. That's why the propellants flow. This is true as long as there is pressure in the tanks even if the pumps are not rotating. Preburner combustion greatly increased the volume of the propellants and hence their velocity (dynamic pressure) not their static pressure. Dynamic pressure spins the turbine.
It seems I lack some basic physics here. How can the preburner ignite the propellant, increase its volume and flow rate and not increase the static pressure? Rakaydos gave a good reason why there is more pressure on the turbine than on the tank. But if it's all that simple, why would the merlin need to prestart the turbopumps with high pressure helium? They could do the same/similar procedure with the gas generator.
- My mistake. I should have said: Everything downstream of the PRE-BURNER has a lower total pressure. The static pressure will rise in the pre-burner as combustion products back pressure the turbine. But, it cannot increase above the pressure upstream of the pre-burner injectors.
- The start mode I outlined is what NASA SP-125 (pg 68) calls as "main tank head start". If this type of start takes too long (> 3 seconds or so) a "turbine spin start" may be added to the system to decrease the start time. I do not know which method the Merlin uses. I would guess a "main tank head start". Does anyone know?
Also see SP-125 pg 181 for different types of starts.
I need to digest all this but I think the Merlin has a face start sequence (I have no idea what that means). I remember Mueller in an interview reporting that they blew up 100 engines before they go it right.
I need to digest all this ...
Semmel - I think the key question here is, ignoring the actual combustion chamber, does the powerpack run a thermodynamic cycle.
If yes, then startup would be like a jet engine's, which cannot be done by simply "lighting it up".
But I think the power pack is different. The power extracted from the exhaust is not used to pump fuel into the pack, but to pump it into the combustion chamber. I think that's why it is possible.
However, with all the phase changes that are going on, it is far from trivial, and the explanation upthread is too simplistic - I wouldn't take it literally.
I need to digest all this ...
Semmel - I think the key question here is, ignoring the actual combustion chamber, does the powerpack run a thermodynamic cycle.
If yes, then startup would be like a jet engine's, which cannot be done by simply "lighting it up".
But I think the power pack is different. The power extracted from the exhaust is not used to pump fuel into the pack, but to pump it into the combustion chamber. I think that's why it is possible.
However, with all the phase changes that are going on, it is far from trivial, and the explanation upthread is too simplistic - I wouldn't take it literally.
- By powerpack I assume you mean a turbo-pump assembly with its associated gas generator. Yes, it runs a thermodynamic cycle.
- This "powerpack" can be started with only main tank head pressure and an igniters, but may be slow to spool up. If this is the case a "spin turbine" may be added. The tank pressure is the initial motive force.
- The power extracted from the exhaust (along with the tank pressure ~3 atms) is used to pump fuel into the powerpack as well as the main chamber.
- The gas generator, or pre-burner, gasifies the propellants either fuel rich or oxidizer rich. This is well understood. I fail to see the problem in my start sequence?
- According to Sutton, the F1, MA-3 and SSME are all started using "tank head" starting.
John
For all that I know, they might put an electrical motor on the shaft...
Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.No, the the powerpack fuel is also pumped by the powerpack. This is certainly a requirement for raptor, since in FFSC the pressure in the preburner must be greater than chamber, and since all fuel goes through the preburners, there would be nowhere else to pump to.
I'm not sure about the cycle.
In a jet engine, you have a clear "cycle", since the far field inlet and outlet conditions are sinked into the same atmosphere, and mechanical power extracted from the exhaust goes into compressing the inflow.
Here, the conditions in the far field inlet are simple the tanks (with head pressure), and the outlet goes into the combustion chamber. Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.
If you include the combustion chamber, then far field outlet conditions are the cold hard cynical vacuum of space.
If there's no cycle, then in theory you could just "light it up", but as you say, practicalities may dictate that the spin up will be impractically slow.
Whichever way, I don't think it's an intractable problem. For all that I know, they might put an electrical motor on the shaft...
Saying something is possible or even a good idea is not speculation...For all that I know, they might put an electrical motor on the shaft...
That was the idea I had and didn't voice because of the danger of baseless speculation. Once running, the motor would generate electricity to heat up and gasify some of the propellant in the tanks to create the autogenous pressure. Safes the running of hot fuel pipes in favor of electrical cables. No idea what is lighter but it probably would safe a lot of headaches with the hot pure oxygen.
Again, total speculation on my part and probably wrong.
Depends where you draw the boundary of the control space.
I'm not sure about the cycle.
In a jet engine, you have a clear "cycle", since the far field inlet and outlet conditions are sinked into the same atmosphere, and mechanical power extracted from the exhaust goes into compressing the inflow.
Here, the conditions in the far field inlet are simple the tanks (with head pressure), and the outlet goes into the combustion chamber. Mechanical power extracted from the exhaust goes into pumping the combustion chamber - not into the powerpack.
If you include the combustion chamber, then far field outlet conditions are the cold hard cynical vacuum of space.
If there's no cycle, then in theory you could just "light it up", but as you say, practicalities may dictate that the spin up will be impractically slow.
Whichever way, I don't think it's an intractable problem. For all that I know, they might put an electrical motor on the shaft...
- I'm sure of the cycles both gas generator and pre-burners.
- Before starting, the main chamber is at what ever the outside pressure is (which could be vacuum).
- Mechanical power extracted from a gas generator's or pre-burner's exhaust all goes into pumping the propellants!
- The propellants then either go to the main chamber or gas generator / pre-burner for combustion.
In a gas generator cycle only a small portion of the propellants is burnt and it is exhausted separately from the main chamber. In a pre-burner a larger portion of the propellant is burnt and it is exhausted into the main chamber. The pre-burner obviously needs to be at pressure higher than the main chamber.
- No electric motors. Pumps for large rocket engines require 10s of thousands of horse power.
John
Depends where you draw the boundary of the control space.
If around the powerpack, then no, power doesn't go to pump propellant into it. It is fed by head pressure. Power goes into pumping into the main chamber.
If around the whole motor, then yes, but then a rocket engine as a whole - does it run a thermo cycle? I'm not sure. It's very different from a jet engine...
Depends where you draw the boundary of the control space.
If around the powerpack, then no, power doesn't go to pump propellant into it. It is fed by head pressure. Power goes into pumping into the main chamber.
If around the whole motor, then yes, but then a rocket engine as a whole - does it run a thermo cycle? I'm not sure. It's very different from a jet engine...
Not that much different.
John
- Yes, you could use an electric motor to spin up the turbine for faster starting, or you could use another high pressure gas source, or you could just use tank head pressure like the F1, MA-3 and SSME. I'm just saying that I have seen no evidence that the Raptor or Merlin use any type of spin up system. Does anyone know different?
They are similar, and turbojets use electric motors to spin up the turbines to get the compressors feeding air pressure. Why couldn't a FFSC engine spin the turboshaft with a motor, just to get greater than tank head pressure in the preburner?
A key difference is that a turbojet has zero pressure differential between the inlet and the burner before spinning the compressor up, while the rocket has several atmospheres (~50 psi) of pressure pushing oxidizer into the burner. So a turbojet can't do a head start, while a SC rocket engine can.
They said it uses autogenous pressurization, so use some of those gases.
I am pretty sure the tank pressure provided by autogenous pressurization system is not enough to start the spin of the turbines. If that was the case, F9 would be able to do the same with LOX and RP1 but they use high pressure helium instead. Probably a lot of it. But I am not an expert and happy to be proven wrong.
I'm just saying that I have seen no evidence that the Raptor or Merlin use any type of spin up system. Does anyone know different?
Thank you.I'm just saying that I have seen no evidence that the Raptor or Merlin use any type of spin up system. Does anyone know different?
Merlin uses a spin up system driven by high-pressure helium. The actual valves used to control the flow of fuel and LOX are spring-actuated (built into the pintle in the case of the combustion chamber, not sure about the preburner but presumably it has something similar?), so in order for fuel to even be injected the pressure must be high enough, which requires first spinning up the turbopump somehow.
Any new information from SpaceX on Raptor development?Not heard anything. Perhaps you should not expect any more new info. on Raptor dev. until IAC2018 knowing SpaceX.
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology? Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future? It seems Blue is going for a slightly more proven ORSC methalox design, and is slowly making some progress, but even that seems years ahead of any new motor on the horizon.
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology?
Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future?
And methalox is not an optimal booster propellant. Kerolox allows lighter tanks, and for similar engine, allows better T/W.
>
BUT... can't make propane on mars... ;)
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage... :o
BUT... can't make propane on mars... ;)
>
BUT... can't make propane on mars... ;)
Sez here you can, using Fischer-Tropsch Synthesis.
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140002709.pdf
Subcooled propane with Falcon 9 recovery IS a good architecture. They should cancel Ariane 6 and go straight to it, over-sized first stage that would allow upper stage reuse down the road.Won't work.
@John Alan
Do you mean those studies of the German Aerospace Center? http://elib.dlr.de/114430/2/PresentationIAC-17%20-%20D2.4.3_f.pdf
That's not it directly...
HOWEVER... it reads like that was written after the prior study (I can't find) was published...
And it seems to be a follow on presentation based on that prior dry wordy study...
Nice Find... :)
[1] getting back on topic Lar... I swear... ;)long way round the barn but ok :)
That said I can't see them licensing F9 tech and tooling.OTOH, the Russians did license their 50 year old Soyuz to the Europeans, so you never know. In fact, why not? If you have already reduced launch cost by an order of magnitude with your methalox FFSC powered RLV?
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage... :o
BUT... can't make propane on mars... ;)
Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577
But methane is perfectly acceptable, especially for higher delta-v stages.
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage... :o
BUT... can't make propane on mars... ;)
Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577
But methane is perfectly acceptable, especially for higher delta-v stages.
Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.
That said I can't see them licensing F9 tech and tooling.OTOH, the Russians did license their 50 year old Soyuz to the Europeans, so you never know. In fact, why not? If you have already reduced launch cost by an order of magnitude with your methalox FFSC powered RLV?
A better question is why is the ESA bothering with Soyuz or potentially Falcon 9 in the future? What the heck is a "Europeanized" Soyuz anyways? The answer lies somewhere within providing launch services without the need for any R & D fixed costs and timely availability.
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"
Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?
It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"Reminds me of Burt Rutan not wanting a liquid cooled engine for Voyager because it weighed an extra 15 pounds on a plane where they were scrounging for ounces of weight savings. Then someone figured the extra efficiency of the engine would mean 1,000 pounds less fuel used on the trip.
Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?
It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?
Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...
Purified Propane could be brought in at outside temp in pressurized tanks... Infinite shelf life...
(I'm picturing 40ft ISO container/tanks floated into port on cargo ships and trucked to the launch site)
Then run the propane thru a sub cooler system and load on the rocket at atmospheric pressure before launch...
Point was, both NASA and the Germans published papers (links up above) that indicate a scPropLox Rocket has the best mass fraction and ISP trade off of any system they have studied... IF your starting over clean sheet, you may want to look at this, was my take away from both of them...
My 2nd point was... there is no reason I can see that the F9 basic design couldn't be reconfigured to fly with scPropLox and likely would gain some additional performance for the trouble of doing so...
SpaceX has already said BFS/BFR is our future... F9/FH will die someday as obsolete...
3rd and last point was... SpaceX could stand to make some big bucks by selling either the designs OR just sell prebuilt rocket assemblies to the EU or Japan as an additional income stream down the road...
Have to remember the big picture in the world...
Many countries maintain a space and rocket launching ability for national pride and me too standings...
They will not give that up... even when SpaceX gains a monopoly with BFS/BFR...
Meanwhile... No new news on Raptor [1]
[1] on topic... ;)
Ugh, don't call it scPropLox. It's ambiguous whether you mean propane or propylene.Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage... :o
BUT... can't make propane on mars... ;)
Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577
But methane is perfectly acceptable, especially for higher delta-v stages.
Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.
Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...
Purified Propane could be brought in at outside temp in pressurized tanks... Infinite shelf life...
(I'm picturing 40ft ISO container/tanks floated into port on cargo ships and trucked to the launch site)
Then run the propane thru a sub cooler system and load on the rocket at atmospheric pressure before launch...
Point was, both NASA and the Germans published papers (links up above) that indicate a scPropLox Rocket has the best mass fraction and ISP trade off of any system they have studied... IF your starting over clean sheet, you may want to look at this, was my take away from both of them...
My 2nd point was... there is no reason I can see that the F9 basic design couldn't be reconfigured to fly with scPropLox and likely would gain some additional performance for the trouble of doing so...
SpaceX has already said BFS/BFR is our future... F9/FH will die someday as obsolete...
3rd and last point was... SpaceX could stand to make some big bucks by selling either the designs OR just sell prebuilt rocket assemblies to the EU or Japan as an additional income stream down the road...
Have to remember the big picture in the world...
Many countries maintain a space and rocket launching ability for national pride and me too standings...
They will not give that up... even when SpaceX gains a monopoly with BFS/BFR...
Meanwhile... No new news on Raptor [1]
[1] on topic... ;)
Not to change the subject, but does anyone know? "Has SpaceX considered variable nozzles for the Raptor engine?"
Variable nozzles are heavy, but how does that compare to the "extra" landing engines on the BFS?
Might variable nozzles allow SSTO performance due to the higher nozzle thrust coefficient?
Would all of the engines need variable nozzles or perhaps only the landing engines?
It boils down to, "What is the mass of the variable nozzle hardware?"
"How does the improvement in thrust coefficient from the variable nozzle / extra nozzle mass translate to the payload to orbit?
Variable exhaust nozzle for a FFSC engine? I think that's too much cutting edge tech to implement with the Raptor. It is impressive enough that SX got a working large FFSC engine. Lets not tempt fate.
They would launch these from where?Flordia of course. If it is thought that untested hardware is high risk, then recall that three of the engines are restartable and they are not needed for lift-off without the mass of the fully fueled second stage atop the booster. Launch with the DUT quiescent, then start engines with modified nozzles after the booster clears the beach. After they are somewhat proven then launch normally just as the FH is doing.
Ok, does anyone have any updates on Raptor? Have they tested it to a higher pressure yet?No more news on Raptor as far as I know. We will likely have to wait until IAC2018 for the next Raptor update.
Or a BFS update...Ok, does anyone have any updates on Raptor? Have they tested it to a higher pressure yet?No more news on Raptor as far as I know. We will likely have to wait until IAC2018 for the next Raptor update.
On the Methane side of Raptor...
I've always thought of it as a modified expander cycle with the preburner there to kick start it from cold and add some heat to vaporize the LNG full flow pre turbine once running...
In short... BOTH turbines will run at near room temps once going... (my opinion) ;)
That said... the hard part of Raptor is starting it... (I think)
I'm thinking a supply of very high pressure gaseous oxygen and gaseous methane is needed to bring Raptor to life from a cold start...
700 bar room temp COPV's anyone?... :o
All preburners (and the RCS system) share this common supply (maybe with some redundancies)
Once a Raptor is running... It can be tapped to refill such a bottle supply and keep it topped up...
(tap high pressure liquid into a small "boiler" to batch flash it into the higher pressure of the storage system)
Batch boilers may be electric heated... I'm not sure on that...
It's all a system... thinking system and not just a rocket engine here...
On the Methane side of Raptor...
I've always thought of it as a modified expander cycle with the preburner there to kick start it from cold and add some heat to vaporize the LNG full flow pre turbine once running...
In short... BOTH turbines will run at near room temps once going... (my opinion) ;)
That said... the hard part of Raptor is starting it... (I think)
I'm thinking a supply of very high pressure gaseous oxygen and gaseous methane is needed to bring Raptor to life from a cold start...
700 bar room temp COPV's anyone?... :o
All preburners (and the RCS system) share this common supply (maybe with some redundancies)
Once a Raptor is running... It can be tapped to refill such a bottle supply and keep it topped up...
(tap high pressure liquid into a small "boiler" to batch flash it into the higher pressure of the storage system)
Batch boilers may be electric heated... I'm not sure on that...
It's all a system... thinking system and not just a rocket engine here...
Is there a commonly recognized name for this startup sequence? "Expander pressure vessel-boosted bootstrap startup"? :)
The RS-25 uses "pure" bootstrap startup, but the start sequence is slow, three to six seconds according to blogs.nasa.gov/J2X/2014/01/24/inside-the-leo-doghouse-light-my-fire/ (http://blogs.nasa.gov/J2X/2014/01/24/inside-the-leo-doghouse-light-my-fire/), which seems a bit slow for inflight and landings. Boosting the spin-up with pressurized gaseous propellant seems quite ingenious to me.
Quote from: John Alan
That's not it directly...
HOWEVER... it reads like that was written after the prior study (I can't find) was published...
And it seems to be a follow on presentation based on that prior dry wordy study...
Nice Find... :)
You'll find the whole study at http://elib.dlr.de/114430/
Actually you can find a lot of similar studies at the DLR servers. Their ftp servers never forget anything (Experiencing a Vulcain full run engine test from 200 meters away is something you won't forget as well ;) )
Somewhere (can't find it now) I saw a paper by someone in the EU rocket group, that sub cooled Propane and sub cooled LOX actually works out as the best mass fraction (tank sizes, weights etc) propellant to use in a booster stage... :o
BUT... can't make propane on mars... ;)
Propylene is even better, see this post (and several other linked to it)
https://forum.nasaspaceflight.com/index.php?topic=42302.msg1642577#msg1642577
But methane is perfectly acceptable, especially for higher delta-v stages.
Methane/NG is much cheaper though, not really as important now, but could be the difference of millions of dollars in a BFR-sized rocket.
Don't forget the EU has their launch site at Guiana in South America...
I'm not 100% sure LNG will be the cheapest hydrocarbon to get on site... ready to load on the rocket...
Purified Propane could
It appears the Raptor is still the only FFSC engine currently under development. It has been shown to work at full thrust for many minutes in sub-scale form. Why are other space agencies not pursuing this very efficient technology? Why aren't Russia, China, India and the ESA pursuing Methane as the fuel of the future? It seems Blue is going for a slightly more proven ORSC methalox design, and is slowly making some progress, but even that seems years ahead of any new motor on the horizon.
BTW, "subscale" is in the eye of the beholder. Current Raptor would work fine in a prototype BFS doing Grasshopper-like hoops.
BTW, "subscale" is in the eye of the beholder. Current Raptor would work fine in a prototype BFS doing Grasshopper-like hoops.
Agreed, EM has shown with rockets and cars that he prefers to get something done and iterate. He said they'd have a BFS flying. He didn't say it would have the final Raptor or be the final design (edit: of the BFS).
I could see them building something that flys and has the equivalent of a Merlin 1A engine. Get it up, learn and iterate.
Compared to how NASA has spent $10 of billions in the last 30 years on vehicles that have never left the ground I prefer the SpaceX development method.
Let's be fair, Ares IX flew a sum total of one time ;)And that was a subscale booster and lacked a second stage...
Let's be fair, Ares IX flew a sum total of one time ;)And that was a subscale booster and lacked a second stage...
A smaller raptor would mean an even smaller initial vehicle.
Not going to happen.
A smaller raptor would mean an even smaller initial vehicle.
Not going to happen.
They could drop in 1000 kN Raptors on the 2017 BFS and still launch with up to half a fuel load.
Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.There were interim versions of Merlin used on actual Falcon 9 missions, and I don't think I need to tell you that falcon 9 changed substantially during its service life.
They could use the subscale Raptor for BFR initially, but the vehicle (like Falcon 9) would have to be smaller.
Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.There were interim versions of Merlin used on actual Falcon 9 missions, and I don't think I need to tell you that falcon 9 changed substantially during its service life.
They could use the subscale Raptor for BFR initially, but the vehicle (like Falcon 9) would have to be smaller.
no, it would not have to be smaller. It could launch with only partially fueled tanks.
When they were developing Merlin 1A, falcon 1 and falcon 9 1.0, they did not know much more thrust they will eventually get from the updated later merlin engine variants, and they really, really had to get SOMETHING flying.
So they made the craft for those engines they had at that point.
Falcon 9 never flies with partially-fueled tanks.
Dec. 5, 2010 Proton-M/ Blok-DM-3 Uragan-M #739 Uragan-M #740
Uragan-M #741 Failure Rocket failed to reach orbital velocity after upper stage overfilled with propellant.
Moving the goalposts.
Launching with half-size tanks is not the same as launching with full-size tanks that are half-empty.
You also seem to think that SpaceX will build a BFR core with undersized engines, and then swap out those engines later on with the full-size version instead of just building a larger core for the larger engines.
It's not quite insane to imagine that at some point there might be a reasonable trade where it was easier to leave the tanks a bit long, and later upgrade the engines.is very far from a ringing endorsement of the concept.
Falcon 9 never flies with partially-fueled tanks.Considering that the Merlin 1A (340kN thrust) is basically a subscale version of the Merlin 1D (845kN thrust), and SpaceX developed Falcon 1 and the three-engine Grasshopper and F9dev as "subscale" test vehicles, I don't see why a subscale raptor wouldn't be used on initial test vehicles.There were interim versions of Merlin used on actual Falcon 9 missions, and I don't think I need to tell you that falcon 9 changed substantially during its service life.
They could use the subscale Raptor for BFR initially, but the vehicle (like Falcon 9) would have to be smaller.
no, it would not have to be smaller. It could launch with only partially fueled tanks.
When they were developing Merlin 1A, falcon 1 and falcon 9 1.0, they did not know much more thrust they will eventually get from the updated later merlin engine variants, and they really, really had to get SOMETHING flying.
So they made the craft for those engines they had at that point.
The Grasshopper test vehicle used a significantly lower thrust (development) variant of Merlin 1D than is in use today, and even the current Raptor is only somewhat lower thrust than the planned operational Raptor. This isn’t “weaseling,” it’s factual.SpaceX has supposedly done a fair amount of engine firings on the current raptor already, so there is probably a decent data set showing what the current engine is capable of. I do not believe this is publicly available at this time, however it would be interesting to know where we are in thrust and specific impulse right now vs the current plan for 'operational' raptor.
SpaceX has supposedly done a fair amount of engine firings on the current raptor already, so there is probably a decent data set showing what the current engine is capable of. I do not believe this is publicly available at this time, however it would be interesting to know where we are in thrust and specific impulse right now vs the current plan for 'operational' raptor.
We do know the current chamber pressure is lower, Elon mentioned what they achieved so far but I forget where he talked about this.
The test engine currently operates at 200 atmospheres, 200 bar, the flight engine will be at 250 bar, and then we believe over time we could probably get that to a little over 300 bar.
I think we might, if we get lucky, be able to do short hop flights with the spaceship part of BFR maybe next year.
SpaceX has supposedly done a fair amount of engine firings on the current raptor already, so there is probably a decent data set showing what the current engine is capable of. I do not believe this is publicly available at this time, however it would be interesting to know where we are in thrust and specific impulse right now vs the current plan for 'operational' raptor.
We do know the current chamber pressure is lower, Elon mentioned what they achieved so far but I forget where he talked about this.
IAC.QuoteThe test engine currently operates at 200 atmospheres, 200 bar, the flight engine will be at 250 bar, and then we believe over time we could probably get that to a little over 300 bar.
It would not surprise me to learn they've hit 250 already, but who knows.
There is a big gap between 'it works on the test stand' and 'we are leaping in with both feet at the highest pressure we can justify'.
The margin, if other things don't go wrong is very large (in space terms) for BFS/R two stage.
Post FH conference -QuoteI think we might, if we get lucky, be able to do short hop flights with the spaceship part of BFR maybe next year.
Maybe we'll get more on Raptor next year.
If they used the sub scale raptor in an upper stage now, does anyone have a good idea of the additional mass enabled for geostationary orbit? I assume low earth orbit would have less gain, but could be substantial if used as US for the falcon heavy.
Would a Raptor US also be easier to start testing for reuse?
Any news on Raptor testing above sub-scale? Seems like Raptor has to be finished first before BFR/BFS can really get started.No news as far as I know. SpaceX are being very secretive regarding Raptor dev. with updates only during the annual IAC events.
Any news on Raptor testing above sub-scale? Seems like Raptor has to be finished first before BFR/BFS can really get started.I think they can do BFS tests with subscale Raptor just fine (while full scale is being finished up). Don't know if they would bother to do so, though.
We are only talking about roughly a 15% linear scale up. I think they will be full size. Why invest in creating multiple engines of a size you don't plan on developing fully? Ultimately, a waste of time and manpower.My theory is that the BFR/BFS specs were in flux, so to avoid blocking engine dev on BFR dev they told the propulsion guys to just pick a reasonable number and go with it. While they were doing engine dev they settled on a 15% larger number in the BFR dev process, but to avoid the distraction of a continually moving target the raptor guys are going to keep going with the subscale engine to the end of their dev cycle (whatever that involves), instead of pivoting instantly to match whatever # the BFR guys prefer this week.
John
That makes no sense.
That makes no sense.CScott's? Makes perfect sense, exactly as he described it.
They are exploring the unknown. During engine development there's some uncertainty in exactly how much thrust a given physical scale will be able to provide, and during BFR development there's uncertainty in how much the rocket will need to weigh. You have to untie the knot somehow, and you can't throw away your work and start from scratch every time any component diverges from its initial estimates. We only get snapshots of progress once a year, and at that point the Raptor on the test stand was underperforming the BFR design specs by 15%. (Or the BFR was 15% overweight.) Who knows, Tom might squeeze out more performance from the subscale Raptor (or the BFR might slim down) and the final version will only be (say) a 10% scale up.That makes no sense.
It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
That makes no sense.
It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
Didn't they scale down the size of the production engines to be very close to the engine they already had under test? Or did I just dream that 2017 IAC presentation?
Didn't they scale down the size of the production engines to be very close to the engine they already had under test? Or did I just dream that 2017 IAC presentation?
I don't think 'scale down' is the right word.
BFS/R was scaled down to fit market needs.
This meant that for engine out and packing reasons, you pretty much need that size of engine.
That makes no sense.
It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
Didn't they scale down the size of the production engines to be very close to the engine they already had under test? Or did I just dream that 2017 IAC presentation?
While incapable of handling the full size of the expected Raptor engine unit, the Stennis test stand enabled the individual testing of each subcomponent of the 1MN scaled prototype that SpaceX currently has at its test facility in McGregor, Texas.
Since the final thrust level of the Raptor had not been settled, it was decided that the first integrated test engine would be a 1MN sub-scale engine.
It enabled the full testing at Stennis E2 and allowed for the development of robust startup and shutdown sequences, characterize hardware durability and anchor analytical models that would be used for future designs.
That makes no sense.
It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
Didn't they scale down the size of the production engines to be very close to the engine they already had under test? Or did I just dream that 2017 IAC presentation?
I think they did. See the size comparison on my chart.
John
Didn't they scale down the size of the production engines to be very close to the engine they already had under test? Or did I just dream that 2017 IAC presentation?
I think they did. See the size comparison on my chart.
John
Correlation is not causation. The vehicle architecture is not being designed to match the engine, but the expected economic of the entire operation. Looking it from the opposite direction is the path to being mislead. Size of vehicle is not determined by one parameter (difficulty of engine scaling) but by a hundred.
Dimension | Ratio |
Linear | 0.75 |
Area | 0.5625 |
Mass | 0.421875 |
That makes no sense.
It makes a little sense. BFR's engines seem to be under design for the mission they have in mind, rather than the vehicle being specifically scaled to the engines they have available. This hypothesis does require the ability to scale the engine almost arbitrarily once they have the basic design worked out, though.
Didn't they scale down the size of the production engines to be very close to the engine they already had under test? Or did I just dream that 2017 IAC presentation?
I think they did. See the size comparison on my chart.
John
Correlation is not causation. The vehicle architecture is not being designed to match the engine, but the expected economic of the entire operation. Looking it from the opposite direction is the path to being mislead. Size of vehicle is not determined by one parameter (difficulty of engine scaling) but by a hundred.
Is it possible that some of the pre-burner( or main injectors) structure is copper like many other engines. and some local impingement occurs during start and shutdown?
Jeff Foust @jeff_foust
If you squint at this chart, you can see ongoing and planned test activity at Stennis by Aerojet Rocketdyne, Relativity, Stratolaunch and SpaceX, among others.
8:41 AM - 26 Mar 2018
What is the "SpaceX combustion device"?Quote from: https://twitter.com/jeff_foust/status/978295808679464968Jeff Foust @jeff_foust
If you squint at this chart, you can see ongoing and planned test activity at Stennis by Aerojet Rocketdyne, Relativity, Stratolaunch and SpaceX, among others.
8:41 AM - 26 Mar 2018
What is the "SpaceX combustion device"?Quote from: https://twitter.com/jeff_foust/status/978295808679464968Jeff Foust @jeff_foust
If you squint at this chart, you can see ongoing and planned test activity at Stennis by Aerojet Rocketdyne, Relativity, Stratolaunch and SpaceX, among others.
8:41 AM - 26 Mar 2018
I thought this was more interesting. What's "Mars Lander"?
Why the heck not?I thought this was more interesting. What's "Mars Lander"?
Whatever it is, it's not likely to have anything to do with Raptor.
The first flight of Falcon 9 had a failed restart of the Vacuum Merlin because roll control froze up in vacuum, a problem they probably could've caught in a test like this. At the time (correct me if I'm wrong), this facility had been basically mothballed, thus they would've had to spend like hundreds of millions to restart it or something vs the ~$40 million price of another Falcon 9 launch which they may have had to do anyway. So at the time, it was a good trade, even especially in retrospect.
Apparently now the facility is being used by upper stages again, which means SpaceX won't have to pay the full cost of un-mothballing. Also, a BFR and BFS cost about an order of magnitude more than a Falcon 9 v1.0, and SpaceX would have to wait another 26 months to try again, thus pushing back their crewed flight another synod.
Pretty sure it makes sense to test for integrated-Raptor landing failure modes in Mars-like conditions.
Based on what, exactly?The first flight of Falcon 9 had a failed restart of the Vacuum Merlin because roll control froze up in vacuum, a problem they probably could've caught in a test like this. At the time (correct me if I'm wrong), this facility had been basically mothballed, thus they would've had to spend like hundreds of millions to restart it or something vs the ~$40 million price of another Falcon 9 launch which they may have had to do anyway. So at the time, it was a good trade, even especially in retrospect.
Apparently now the facility is being used by upper stages again, which means SpaceX won't have to pay the full cost of un-mothballing. Also, a BFR and BFS cost about an order of magnitude more than a Falcon 9 v1.0, and SpaceX would have to wait another 26 months to try again, thus pushing back their crewed flight another synod.
Pretty sure it makes sense to test for integrated-Raptor landing failure modes in Mars-like conditions.
That facility won't support a single Raptor.
Adjusting the fiscal year quarters to actual calendar dates, doesn't that test window close at the end of March 2020? Additionally, doesn't the Mars 2020 rover have a NET of July 2020?That would be the Mars rover, not the "Mars Lander." The 2020 rover uses the Skycrane stage, just like MSL, therefore there is no "lander" part.
3-4 months to button up and ship it?
To reach Mars, a spacecraft must travel through the cold vacuum of space for nine months. These extreme conditions are recreated at NASA’s Glenn Research Centre in Cleveland. In Vacuum Chamber 5, powerful pumps suck air out of the 4.5 metre tall space. Even a spacecraft’s thrusters can be tested here – it’s designed to handle the heat and gas they generate. Panels cooled to -262°C line the walls, chilling the thruster’s exhaust. Pumps lining the bottom of the tank concentrate and recycle the precious xenon gas that the thrusters blast out as propellant.
https://cosmosmagazine.com/space/nasa-prepares-mars-missionsQuoteTo reach Mars, a spacecraft must travel through the cold vacuum of space for nine months. These extreme conditions are recreated at NASA’s Glenn Research Centre in Cleveland. In Vacuum Chamber 5, powerful pumps suck air out of the 4.5 metre tall space. Even a spacecraft’s thrusters can be tested here – it’s designed to handle the heat and gas they generate. Panels cooled to -262°C line the walls, chilling the thruster’s exhaust. Pumps lining the bottom of the tank concentrate and recycle the precious xenon gas that the thrusters blast out as propellant.
The timeline is odd for Mars 2020 though.
I find references to wheels and nuclear work being done for rovers at Glenn.
Perhaps final installation or preparation for the RTG would make more sense than thrusters or BFS/raptor tests.
With planned schedule, there will be very little need for Glenns vacuum test capability, as actual (possibly suborbital) space works just fine.
https://cosmosmagazine.com/space/nasa-prepares-mars-missionsQuoteTo reach Mars, a spacecraft must travel through the cold vacuum of space for nine months. These extreme conditions are recreated at NASA’s Glenn Research Centre in Cleveland. In Vacuum Chamber 5, powerful pumps suck air out of the 4.5 metre tall space. Even a spacecraft’s thrusters can be tested here – it’s designed to handle the heat and gas they generate. Panels cooled to -262°C line the walls, chilling the thruster’s exhaust. Pumps lining the bottom of the tank concentrate and recycle the precious xenon gas that the thrusters blast out as propellant.
The timeline is odd for Mars 2020 though.
I find references to wheels and nuclear work being done for rovers at Glenn.
Perhaps final installation or preparation for the RTG would make more sense than thrusters or BFS/raptor tests.
With planned schedule, there will be very little need for Glenns vacuum test capability, as actual (possibly suborbital) space works just fine.
The "recycle the precious xenon gas" part indicates the chamber is dedicated for ion engine test. RTGs would be one way of providing the electric power needed.
So would Kilopower, potentially at higher power levels.The RTG I mentioned was in the context of wondering if the Mars 2020 lander RTG was going to be installed there.
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?
I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.
Are the turbopumps and combustion chamber different between the sea level and vac versions of the Merlin?
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?
I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.
Are the turbopumps and combustion chamber different between the sea level and vac versions of the Merlin?
I'm not sure this belongs here or in the BFR thread:No. I think the timing is due to finally getting the permission to use that piece of port land. Raptor has been pretty far along in testing for a while, now. There's not a ton of uncertainty there.
Can we infer by the arrival of parts of BFS tooling that Raptor passed a SpaceX internal milestone a few months ago?
I have a question about the Raptor engine. The Merlin vac is quite different to the SL-Merlin. Can we expect that Raptor vac will be very similar to the SL-Raptor? They both have regeneratively cooled nozzles. Would the set of turbopumps and combustion chamber be identical and differentiate only by mounting different nozzles or will they too be different?
I am asking from general curiosity but also thinking of testing the engines. Could an engine assembly be tested with a SL or intermediate nozzle then become a vac engine by changing the nozzle? I know that Elon Musk said the vac engines can be fired at SL with full thrust but it is not advisable, so maybe not feasible for testing.
Speaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.”https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/
Was it previously announced that a larger Raptor was under development?
I could easily have missed it...QuoteSpeaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.”https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/
Since this is six-month old news, we should be seeing something on the stand yet this year... The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.
Was it previously announced that a larger Raptor was under development?
I could easily have missed it...QuoteSpeaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.”https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/
Since this is six-month old news, we should be seeing something on the stand yet this year... The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.
Yup. Makes sense to scale up before you have production tooling made, but honestly, they probably could get away with the dev engine just fine if they had to. It's barely any smaller.Was it previously announced that a larger Raptor was under development?
I could easily have missed it...QuoteSpeaking at a private talk given to MIT campus members in October 2017, attendees reported that Shotwell stated that although “[BFR’s] composite tanks [would] be a challenge [for SpaceX],” the company was already working on maturing the technologies required, and also noted that SpaceX was “building a larger [version of] Raptor right now.”https://www.teslarati.com/spacex-shotwell-bfr-mars-rocket-texas/
Since this is six-month old news, we should be seeing something on the stand yet this year... The BFS prototype, though, may fly the sub-scale version to move forward, but the BFBooster and orbital versions of BFS may have the bigger engines.
It would also be pretty damn hard to tell that the "larger" Raptor had been swapped in, too. If livingjw is more or less correct, the final-ish 1700kN would require a physical scale-up of maybe 10-20%. That's deeeeeep in the weeds for the photos we have at hand (both L2 and now public).
Attached my favorite of the Raptor bays.
Making the Raptor engine 10-20% larger is not to get it to produce the 1700kN of thrust, so it meets the specs of IAC 2017.
The “non-scalled” engine which will be revealed at IAC2018, probably has more thrust than 1700 kN since TED talk animations show the BFR has become higher since IAC2017.
The exact specs don’t dictate the design. The specs follow out of the optimal design.
Attached are two pictures of the Raptor test stand taken 1 year apart. The first one taken in early 2017 shows some evidence of operation can be seen in the gravel apron and close in in the "grass". The other picture taken a year later shows evidence of much heaver erosion on the Gravel apron and "grass damage out about 1000 feet.
Attached are two pictures of the Raptor test stand taken 1 year apart. The first one taken in early 2017 shows some evidence of operation can be seen in the gravel apron and close in in the "grass". The other picture taken a year later shows evidence of much heaver erosion on the Gravel apron and "grass damage out about 1000 feet.
First photo may just show that the only testing up to that point was only cold flow tests. BTW great photos.
The last rocket with this amount of innovation was Saturn V; there was much iteration leading up to that 'optimal design.'
The last rocket with this amount of innovation was Saturn V; there was much iteration leading up to that 'optimal design.'
Keep in mind that engineers had already optimized far beyond that point when the program was cancelled. F-1A was already done and certified @ 1.8M lb thrust. Designs for new blocks of Saturn V with extended tanks had been completed. Solid and liquid boosters for the S1 had been designed. A perusal of Encyclopedia Astronautica (http://www.astronautix.com) in this area is fascinating.
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
"I don’t want to say too much. We’re building up the test stand right now. We’ve got the first flight version of that engine in work. We’ve been running the development engine quite a bit. It’s running great," Mueller told the audience.
I guess this is new...Would the current test stand be destroyed by a full thrust raptor?Quote"I don’t want to say too much. We’re building up the test stand right now. We’ve got the first flight version of that engine in work. We’ve been running the development engine quite a bit. It’s running great," Mueller told the audience.
From GeekWire (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/)
https://twitter.com/spacecom/status/999691403172036608Quote"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
I guess this is new...Quote"I don’t want to say too much. We’re building up the test stand right now. We’ve got the first flight version of that engine in work. We’ve been running the development engine quite a bit. It’s running great," Mueller told the audience.
From GeekWire (https://www.geekwire.com/2018/spacex-propulsion-guru-tom-mueller-looks-ahead-rocket-engines-mars/)
It almost doesn't matter.https://twitter.com/spacecom/status/999691403172036608Quote"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Ive been waiting for this number, for a while now, since I know it will be insanely high.
It almost doesn't matter.SL Merlins are now at 185:1 T/W at SL and close to 200:1 in a vacuum.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)
It almost doesn't matter.https://twitter.com/spacecom/status/999691403172036608Quote"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Ive been waiting for this number, for a while now, since I know it will be insanely high.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)
1800 tons of thrust for BFS would mean eleven tons of engines at this figure. Halving it to 300:1 only drops five tons.
5 seconds more ISP would make up that limit.
Not that more margin isn't better of course.
But if doubling weight lets you make a simple trade for reliability - it would make sense.
(it is regrettably unlikely to be this simple)
Yes, it is.
Its the Final Mass that determines Delta V in the Tsiolkovsky Rocket equation
https://en.m.wikipedia.org/wiki/Tsiolkovsky_rocket_equation
It almost doesn't matter.https://twitter.com/spacecom/status/999691403172036608Quote"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Ive been waiting for this number, for a while now, since I know it will be insanely high.
Merlin's at 158:1 or so (perhaps more, I diddn't carefully check if these were latest numbers.)
1800 tons of thrust for BFS would mean eleven tons of engines at this figure. Halving it to 300:1 only drops five tons.
5 seconds more ISP would make up that limit.
Not that more margin isn't better of course.
But if doubling weight lets you make a simple trade for reliability - it would make sense.
(it is regrettably unlikely to be this simple)
I see your logic, but It’s not only the 5 tons of the BFS engines than needs to be substracted from the 150 tons payload to LEO. Its also the extra weight of the 31 booster engines, and the extra weight ofIt is actually not.
extra landing fuel for the heavier booster that (partly) needs to be subtracted.
It’s quite a difficult calculation actually.
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.
To be sure. However, that doesn't necessarily make a meaningful difference in whether the vehicle can perform the Mars mission.
On the contrary, I'm pretty sure TWR for engines on the booster significantly affects landing dry mass, a savings that propagates back through reserved fuel mass for landing to staging velocity.To reiterate - I did a more-or-less reasonable calculation assuming mars-entry like delta-v for earth return of BFS (700m/s, which ties in neatly with the capacity of the shown tanks), return velocity of BFR being akin to F9S1, ... counting the structure, published ISP, ...
Anner J. Bonilla
@annerajb
How's raptor testing going?
Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.
https://twitter.com/elonmusk/status/1001565360783474688
Anner J. Bonilla
@annerajb
How's raptor testing going?
Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.
https://twitter.com/elonmusk/status/1001565360783474688
Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor? I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?
Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor? I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?
Anner J. Bonilla
@annerajb
How's raptor testing going?
Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.
https://twitter.com/elonmusk/status/1001565360783474688
Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor? I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?
Anner J. Bonilla
@annerajb
How's raptor testing going?
Elon Musk
Verified account
@elonmusk
Good progress. Really proud of this design & SpaceX propulsion team. This engine is something special.
https://twitter.com/elonmusk/status/1001565360783474688
Since Raptor is supposed to easily take the thrust-to-weight crown from Merlin (which is somewhere around 200 at launch and 220 in vacuum), can we estimate what the theoretical T/W could be for Raptor? I heard numbers over 300 (I think 350 was quoted) a while back... is this even possible?
I was the one who predicted at least 350 based on the cad drawings of the design and chamber pressure.
Made a bet for $50 dollars on it, in this forum.
Another way to approximate T/W of the Raptor engines is to calculate the % mass to orbit of the whole rocket.
Theoretical T/W maximum of a rocket engine can become very high, the smaller you make your engines while retaining high Isp. Practical limit of T/W of methane burning engine would be around 10.000.
I saw these also, including valves turbo pumps and sensors
With t/w of 1000. (See pic)
Interesting Quote from Elon Musk:
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"
http://spacenews.com/elon-musks-ask-me-anything-qa-just-the-space-parts/
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.
Both Musk and Mueller have made it pretty clear that they expect Raptor to have higher TWR than Merlin 1D.
There are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Interesting Quote from Elon Musk:
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"
The T/W ratio of the raptor is going to be lower than the Merlin strictly due to the fact that its components are operating 2-5 times the pressure that the Merlin's operates at the various points in the cycle (pump discharges at kick pumps is on the order of 4-5x the discharge of the Merlin's pumps, main pumps are at like 2-3x, ect). Flanges are going to be bigger, pipes are thicker, there are more pumps (boost, main, kick per propellant), bigger power head, separate loop for TVC hydraulics, ect.
Both Musk and Mueller have made it pretty clear that they expect Raptor to have higher TWR than Merlin 1D.
There are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.
QuoteThere are a number of significant design changes that enable this, e.g. the oxidizer pump discharges directly into the injector, so there is no high-pressure full flow oxidizer plumbing anywhere on the engine.
... but instead of liquid medium-pressure fuel pipes there are fuel-rich higher-pressure gas pipes.
snip...
Interesting Quote from Elon Musk:
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"
http://spacenews.com/elon-musks-ask-me-anything-qa-just-the-space-parts/
Interesting Quote from Elon Musk:
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"
I was talking to Baldusi about that.
There are lots of factors to how heavy you make various engine parts, but the simplest rule to start with is the square/cube thing. If you make an engine twice as big in all dimensions, it weighs 8 times as much but only has 4 times the thrust.
Of course, it's way more complicated since the thickness of various components won't always exactly double, but it's a good basis for understanding why lots of smaller engines can be better than a few big ones.
...Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines. All the physics points to less performance and less thrust to weight for smaller engines......this isn't quite true. Particularly if you look at the nozzle. The nozzle mass scales as dimension cubed, but thrust only scales as dimension squared. There are other, more subtle examples. But if you look at nozzle mass, clusters of small engines win handily over one big nozzle (past a certain size). They're much shorter, too, for the same expansion ratio.
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
.... Some things make me wonder..
* Where is the fuel turbine? I can sort of see a connection of the preburner to the injector face but there should be a turbine somewhere in between. Cant see it.
* The fuel regen out splits to the oxygen preburner and into the fuel preburner. But at the bottom. Looks strange to me, how can they inject the fuel into the preburner at a different position than the main fuel flow?
* I assume the fuel regen in has to be connected to the exit of the fuel pumps. I cant find that connection.
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.
...Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines. All the physics points to less performance and less thrust to weight for smaller engines......this isn't quite true. Particularly if you look at the nozzle. The nozzle mass scales as dimension cubed, but thrust only scales as dimension squared. There are other, more subtle examples. But if you look at nozzle mass, clusters of small engines win handily over one big nozzle (past a certain size). They're much shorter, too, for the same expansion ratio.
Aware of no “published” study, although surely some exists. This comes from scaling laws (I’ve analyzed this in detail) and is borne out in historical example. Compare the bottom of N1 with the bottom of Saturn V....Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines. All the physics points to less performance and less thrust to weight for smaller engines......this isn't quite true. Particularly if you look at the nozzle. The nozzle mass scales as dimension cubed, but thrust only scales as dimension squared. There are other, more subtle examples. But if you look at nozzle mass, clusters of small engines win handily over one big nozzle (past a certain size). They're much shorter, too, for the same expansion ratio.
- The mass of most of the expansion portion of the nozzle is relatively low pressure. Its mass, to a large degree is determined by regenerative cooling, which is proportional to dimension squared.
- Clusters of small engines win handily? What is the physics of this? Do you know of a paper or tests comparing one large regenerativly cooled LRE MCC versus many smaller one?
John
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
.... Some things make me wonder..
* Where is the fuel turbine? I can sort of see a connection of the preburner to the injector face but there should be a turbine somewhere in between. Cant see it.
* The fuel regen out splits to the oxygen preburner and into the fuel preburner. But at the bottom. Looks strange to me, how can they inject the fuel into the preburner at a different position than the main fuel flow?
* I assume the fuel regen in has to be connected to the exit of the fuel pumps. I cant find that connection.
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.
- The fuel turbine is in the upper part of the fuel preburner. The flow is upward towards the fuel pump. The exhaust does a 180 degree turn downward and exits the horizontal pipe to the MCC.
- The regen. fuel exits the MCC near the injector face and part way down the expansion nozzle. They both appear to be routed to the bottom of the fuel pre-burner. A smaller pipe branches off just before one of the regen. fuel pipe enters the pre-burner and heads to a valve on its way to the Lox pre-burner.
- Fuel out of the pump must be hidden behind the pump. I am assuming it goes straight to the manifold at the throat. From the throat, fuel is directed upward towards the main injectors and downward to the expansion nozzle.
- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.
John
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
I love the Raptor design. The fact that they mounted the oxygen powerpack directly ontop of the injector seems revolutionary in oxygen rich staged combustion. I havent seen that trick before.
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.
My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.
Ross.
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.
My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.
Ross.
Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F. Also, it doesn't look like the MCC is using a pintle injector. Wrong shape. Looks more like a planar array of coaxial injectors, of course its hard to tell.
John
I think the RD-170 injector design gives a good idea on what the injectors will look like. The differences of course will be due to FF and liquid methane vs RP1 and full "gas" phase injection in the MCC. and the use of partitions in the MCC may be unnecessary.
When Spacex was testing their pintle injector at Stennis, that made sense since both the Kestrel and all versions of the Merlin used the face shutoff pintle for the MCC.
While we have no info from the Raptor team, on their choice of injector. I would be surprised if they went in a different direction than what they had built with the Merlin engines.
I think the RD-170 injector design gives a good idea on what the injectors will look like. The differences of course will be due to FF and liquid methane vs RP1 and full "gas" phase injection in the MCC. and the use of partitions in the MCC may be unnecessary.
With gas phase injection most of the cyclic excitation caused by the liquid interaction in the burn area will be absent.
Could you care to explain your thought process on the injector geometry choice and why you think the "partitions" wouldn't need it?
edit: I ask because your statements are contrary to what I believe. For the preburners, the RD 170 preburner injector style geometry might well be used, but I don't agree with the main injector.
Goes to show that SpaceX sized Raptor to minimize dev. costs of BFR/BFS system by having just one engine design throughout the system due to budget limitations. If EM had the money of JB then I bet SpaceX would have dev. a larger Raptor for BFR with a smaller version for BFS. Dev. costs do not equate to op. costs.
Interesting Quote from Elon Musk:
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"
I was talking to Baldusi about that.
There are lots of factors to how heavy you make various engine parts, but the simplest rule to start with is the square/cube thing. If you make an engine twice as big in all dimensions, it weighs 8 times as much but only has 4 times the thrust.
Of course, it's way more complicated since the thickness of various components won't always exactly double, but it's a good basis for understanding why lots of smaller engines can be better than a few big ones.
The physics of rocket weight is primarily the physics of pressure vessels. Pressure vessel mass is proportional to volume and pressure: Mass = c*vol*press where c is a constant which depends on the shape and material properties. The mass of main combustion chamber, pre-burners, plumbing including turbo-pump bodies all are primarily stressed by pressure. Larger engines require less cooling due to the square-cube scaling difference. Larger engines also have proportionately smaller combustion chambers due to dwell time combustion physics. Smaller engines operate at lower Reynolds number which has negative effect on performance of small engines. All the physics points to less performance and less thrust to weight for smaller engines. Having said that, performance and thrust to weight are relatively flat from 100,000 - 2,000,000 lbf thrust range.
The performance and thrust to weight of extremely small rocket engines as shown in an earlier post will suffer significant Reynolds number induced losses. Turbo-pumps of that size will have very poor efficiency. The mass of flat slab sides compared to cylindrical pressure vessels of the same volume and pressure are much higher as is the wetted area that has to be cooled. Do we have a link to the paper? Their numbers are hard to fathom and would like to read the original paper.
John
Goes to show that SpaceX sized Raptor to minimize dev. costs of BFR/BFS system by having just one engine design throughout the system due to budget limitations. If EM had the money of JB then I bet SpaceX would have dev. a larger Raptor for BFR with a smaller version for BFS. Dev. costs do not equate to op. costs.
"Thrust to weight is optimizing for a surprisingly low thrust level, even when accounting for the added mass of plumbing and structure for many engines. Looks like a little over 230 metric tons (~500 klbf) of thrust per engine, but we will have a lot of them"
Could the progress be, that thrust to weight optimized further to an even lower thrust of 170 metric tons? 8)
OK, I skimmed the paper:AFAIK they never actually got this plan to work. The numbers looked impressive (100g thrust from a 1.2g structure that's 83.3:1. Not bad for V 0.1 tech) but as you noted they didn't include the pumps. AFAIK they never did. A thesis on why they failed indicate pumping losses were a big part of it and the structure could not take the pressure.
http://www.las.inpe.br/~jrsenna/AerospaceMEMS/Propulsao/S&Aav2997p1-7.pdf
They are weighing only the 1.2 gram main combustion chamber. No manifolds or plumbing. No pumps. No ignition. No nothing accept the chamber. The manifold it was mounted in for testing appears to be many times that weight. The above paper only tested it to 10% of its design pressure, so most of their conclusions are 10x extrapolations, though they may have done testing at higher pressures not reported in this paper. No attempt was made to predict turbo-pump efficiency, but I suspect they would be better of with a piston type pump for this size rocket because they are less effected by low Reynolds numbers. I see nothing in this paper relevant to large reusable rockets.
John
The fuel should be liquid though the pumps and regen cooling system, right up until it's partially combusted in the preburner. The only hot gas line is from the fuel preburner into the MCC injector, which is very short: only a few cm. See "Fuel to MCC" here: https://i.imgur.com/ld7z2Fn.jpg
At the expected pressures the fuel looks to me to become a super-critical fluid after it passes through the regen system. Once 'released' into the MCC, I'd expect it to flash into a vapour immediately.
My gut is telling me that with a pintle injector arrangement, having this SCF CH4 flow through the inner part should perhaps create better mixing when it impinges on an outer flow of still-liquid LOX.
Ross.
Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F. Also, it doesn't look like the MCC is using a pintle injector. Wrong shape. Looks more like a planar array of coaxial injectors, of course its hard to tell.
John
https://i.imgur.com/ld7z2Fn.jpg
.... Some things make me wonder..
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.
- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.
John
https://i.imgur.com/ld7z2Fn.jpg
.... Some things make me wonder..
* Why is there a pipe connecting "Lox to fuel preburner" and the lower fuel regen out pipe? Looks dangerous to connect the two.
- A small amount of Lox is needed by the fuel pre-burner to, well, burn. Opposite for the Lox pre-burner. I am pretty sure they are not connected. The Lox line is routed to the pre-burner's combustion zone were it is injected and mixed with some of fuel and burnt. The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
The labeling may not be perfectly correct, but is my best estimate of the layout. Also this is an early CAD rendering. We are not even sure the current layout is the same. It probably is, we just don't know.
John
It does look substantial but I don't think the thick 'pipe' connecting "Lox to fuel preburner" to the lower "Fuel Regen out" is part of the flow system.
It looks to be part of the sensing and control system used to control reagent supply to the three reaction chambers (preburners and main combustion chamber).
Seems the reagent supply valve controls take delta pressure(?) measurements using these (right to left):
> Heated methane pump output pressure before reaction chamber against cold oxygen pump output pressure before same reaction chamber.
> Cold oxygen pump output pressure [same pipe] against main combustion chamber hot oxygen* injection manifold pressure.
> Heated methane pump output pressure against either ??cold oxygen?? or ??cold methane?? pump inlet pressure
*that's actually oxygen rich gas mix between oxidizer preburner-turbine before main combustion chamber, not pure unreacted propellant like the other sensor inputs appear to be.
Ok, that makes sense. So the setting on this differential pressure measurement would control the LOX/LCH4 mixture ratio.
The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.
Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F.
The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F.
Please see the two quotes above. These are estimates of livingjw about the two pre-burner exit temperatures. Maybe before going on, I should ask.. how do you get to these? To an untrained person like me, these temperatures seem quite high. I have no experience with these and my thermodynamics is long ago (and not particularly great to begin with), so please be patient with me for the following.
The pre-burner burns LOX/LCH4 in just the right amount to run the turbine that drives the pumps and to provide the pressure AFTER the turbine that gets injected into the main combustion chamber. The pressure has to be higher than the pressure in the combustion chamber, so thats some considerable energy.
Here is the part that I am not comfortable with and could be completely wrong: My understanding is that the efficiency of the engine depends on the expansion ratio of the propellents, which depends on the temperature differential of the propellant going in the combustion chamber and the leaving temperature. So basically you want the incomming propellant as cold as possible and the outgoing propellant as hot as possible. Can you set this right please? I have the feeling there is some error in my understanding.
I one could compute the temperatures backwards. Starting with the chamber pressure and thrust of Raptor. One could compute the pressure after the turbines and the mass flow rate. This would inform about how strong the pumps would have to work and how much chemical energy has to be released inside the pre-burners, which would inform about the exit temperature and the mixture ratio inside the pre-burners. Maybe there is a better way of doing that?
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.
How common are single crystal alloy casting shops? Rene N5 and N6 are single crystal alloys. I presumed they were primarily only done in house at big OEMs like P&W and Rolls.
SpaceX is probably going to be forging or casting big portions of the turbine and nozzle with more conventional alloys like Inconel 718, Rene 41/Haynes R41, Haynes 188, 230, 282, ect.
Some parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.
N5 and N6 have significant high stress life up to 1799F and even higher in low stress locations like nozzle guide vanes.
How common are single crystal alloy casting shops? Rene N5 and N6 are single crystal alloys. I presumed they were primarily only done in house at big OEMs like P&W and Rolls.
SpaceX is probably going to be forging or casting big portions of the turbine and nozzle with more conventional alloys like Inconel 718, Rene 41/Haynes R41, Haynes 188, 230, 282, ect.
Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)QuoteSome parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.
Fuel rich operation temp of CH4 are limited by soot production, above certain temp CH4 decompose to C, yet below temp limit of 718 blades.The rest of the fuel is then mixed in to yield a uniform fuel rich gas mix at a temperature sufficiently low to be handled by un-cooled turbines and piping. Probably somewhere around 1400 F.Oxygen won't be liquid. It went through its own pre-burner and is most probably over 1000 F.
Please see the two quotes above. These are estimates of livingjw about the two pre-burner exit temperatures. Maybe before going on, I should ask.. how do you get to these? To an untrained person like me, these temperatures seem quite high. I have no experience with these and my thermodynamics is long ago (and not particularly great to begin with), so please be patient with me for the following.
The pre-burner burns LOX/LCH4 in just the right amount to run the turbine that drives the pumps and to provide the pressure AFTER the turbine that gets injected into the main combustion chamber. The pressure has to be higher than the pressure in the combustion chamber, so thats some considerable energy.
Here is the part that I am not comfortable with and could be completely wrong: My understanding is that the efficiency of the engine depends on the expansion ratio of the propellents, which depends on the temperature differential of the propellant going in the combustion chamber and the leaving temperature. So basically you want the incomming propellant as cold as possible and the outgoing propellant as hot as possible. Can you set this right please? I have the feeling there is some error in my understanding.
I one could compute the temperatures backwards. Starting with the chamber pressure and thrust of Raptor. One could compute the pressure after the turbines and the mass flow rate. This would inform about how strong the pumps would have to work and how much chemical energy has to be released inside the pre-burners, which would inform about the exit temperature and the mixture ratio inside the pre-burners. Maybe there is a better way of doing that?
There aren't many metals that can survive above 1400F. Personally I think the fuel rich outlet will be more around 1100F-1200F.
Fuel rich operation temp of CH4 are limited by soot production, above certain temp CH4 decompose to C, yet below temp limit of 718 blades.
Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)QuoteSome parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.
Yes, both CH4 and 718 works up to 1300F.Fuel rich operation temp of CH4 are limited by soot production, above certain temp CH4 decompose to C, yet below temp limit of 718 blades.
Isn't that temperature around 1300F?
Reddit AMA... (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/)QuoteSome parts of Raptor will be printed, but most of it will be machined forgings. We developed a new metal alloy for the oxygen pump that has both high strength at temperature and won't burn. Pretty much anything will burn in high pressure, hot, almost pure oxygen.
I was referring to the fuel turbomachinery when talking about the forging and cast materials.
Question, for those smarter then me on this... ???
How many grams of gaseous Methane and grams of gaseous O2 burned at a stoichiometric ratio in a "preburner/heater" would generate enough heat to then raise ONE whole Kilogram of either Liquid Methane OR Liquid Oxygen (likely are close but slightly different numbers) from a sub cooled at liquid nitrogen temperatures to a warm gas at say 204C (400F)...
My guess is... I don't know... ???
On edit...
My goal with this question is to back figure how many grams of prop is needed burned to then vaporize the entire full flow staged combustion prop flow to reach the declared thrust values they have claimed...
As a percentage... I am thinking that many may be surprised how small a % number it is... ;)
Point is... this is likely not like a typical rocket engine where a gas generator creates hot exhaust gases to drive the turbine (Merlin, F1, etc)
This is more like a power plant where heat is used to phase change something from a liquid to a gas and that then drives the turbine (much more mass at a much cooler temp)...
The nice part is the turbine exhausts can then be mixed and combusted to then make a real rocket out of it...
I may be crazy... but this is how I see Raptor... you got to forget SSME and every other rocket engine to date.
It's a different animal in many ways... ;)
Question, for those smarter then me on this... ???
How many grams of gaseous Methane and grams of gaseous O2 burned at a stoichiometric ratio in a "preburner/heater" would generate enough heat to then raise ONE whole Kilogram of either Liquid Methane OR Liquid Oxygen (likely are close but slightly different numbers) from a sub cooled at liquid nitrogen temperatures to a warm gas at say 204C (400F)...
My guess is... I don't know... ???
On edit...
My goal with this question is to back figure how many grams of prop is needed burned to then vaporize the entire full flow staged combustion prop flow to reach the declared thrust values they have claimed...
As a percentage... I am thinking that many may be surprised how small a % number it is... ;)
Point is... this is likely not like a typical rocket engine where a gas generator creates hot exhaust gases to drive the turbine (Merlin, F1, etc)
This is more like a power plant where heat is used to phase change something from a liquid to a gas and that then drives the turbine (much more mass at a much cooler temp)...
The nice part is the turbine exhausts can then be mixed and combusted to then make a real rocket out of it...
I may be crazy... but this is how I see Raptor... you got to forget SSME and every other rocket engine to date.
It's a different animal in many ways... ;)
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.
Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.
Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.
- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.
- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.
- Is there something else about the 5000+psi environment that you are concerned about?
John
Rocket turbines are usually single stage, uncooled and don't require such extended life.
It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.
Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.
- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.
- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.
- Is there something else about the 5000+psi environment that you are concerned about?
John
Rocket turbines are usually single stage, uncooled and don't require such extended life.
That used to be the case. In a point to point scenario as described by Gwynne Shotwell they will ramp up a number of uses equivalent to a years long distanse jet flights in a month. Even assuming jet engine equivalent service time they may need to swap out one engine for major overhaul every night. Or the whole set once a month.
At the pressure ratio and mass flow rate, the turbine will likely be pure impulse. On the "coking" issue, the main coking issue in with RP1 is the internal deposits that build up in the flow passages that have high heat flux. The coke can be removed on internal passages by caustic cleaning, a hassle that is best avoided. The methane will not have that issue and if the combustion is stratified in the vicinity of the walls Either fuel rich or O2 rich to remove the LSF caused distress, the liners should take it .
Edit: What about the combustion chamber and nozzle? What would be their life span with high pressure propellant pressed through their cooling channels?
It will take many years of flights and engineering advances to make the engines as long lived as needed.
1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.
John
I was including those items in my post , the highest heat flux will be in the throat of the MCC and any failure in the cooling system in the engine itself (not the external plumbing) would be nonfatal and be noticed on later inspection (as as happened on the SSME)Rocket turbines are usually single stage, uncooled and don't require such extended life.
That used to be the case. In a point to point scenario as described by Gwynne Shotwell they will ramp up a number of uses equivalent to a years long distanse jet flights in a month. Even assuming jet engine equivalent service time they may need to swap out one engine for major overhaul every night. Or the whole set once a month.
At the pressure ratio and mass flow rate, the turbine will likely be pure impulse. On the "coking" issue, the main coking issue in with RP1 is the internal deposits that build up in the flow passages that have high heat flux. The coke can be removed on internal passages by caustic cleaning, a hassle that is best avoided. The methane will not have that issue and if the combustion is stratified in the vicinity of the walls Either fuel rich or O2 rich to remove the LSF caused distress, the liners should take it .
Edit: What about the combustion chamber and nozzle? What would be their life span with high pressure propellant pressed through their cooling channels?
It will take many years of flights and engineering advances to make the engines as long lived as needed.
Yeah, high pressure oxidizing environment is extremely difficult to engineer. Pressure matters.It’s not just the temperature of the metal in steady state, it’s the fact it’s a complex mechanical system with seals and moving parts. Cycling it many times is likely the limiting factor. Also, it’s operating in a high pressure, oxidizing environment. This is different than mere jet engine engineering.
Also, rocket engines are actually a LOT cheaper than jet engines for the same power and even thrust.
- Mere jet engine engineering? Nothing mere about it. Rocket turbines are much simpler than turbines in a typical turbofan or turboshaft engine. Air breathing turbines are multi-staged, air cooled, withstand thousands of hours of life and thousands of cycles. They run at 2-3 times the rocket turbine entry temperature. Rocket turbines are usually single stage, uncooled and don't require such extended life.
- The LOX pre-burner turbine and ducting do have a more severe oxidizing environment, that is why alloys such as Mondaloy were developed.
- Is there something else about the 5000+psi environment that you are concerned about?
John
1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.
John
John, thanks for this. Reading above mentioned temperatures seems high, only for the normal life of most people. The uncooled temperature limit of usual alloys gives some much needed perspective. I didnt know at all that rocket engine turbines can withstand 1700F (925 C).
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.They're using a new, custom alloy that didn't exist before.
John
Obtainium?Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.They're using a new, custom alloy that didn't exist before.
John
Red matter.Obtainium?Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.They're using a new, custom alloy that didn't exist before.
John
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.They're using a new, custom alloy that didn't exist before.
John
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.
John
If the fuel side is limited to 1200F by soot forming (100F margin to 1300F), the oxygen side could go cooler than 1000F.1000F or even 1200F is hundreds of degrees away from the long life limit of nickel based high temperature alloys. Typical uncooled turbine temperature limits are around 1700F. Even regular steel alloys can handle 1000F. What metal do you think would be better and why? The cost of materials is almost always insignificant compared to the final cost of any complex system. Man hours and machine hours drives most of the cost. It would be pound foolish to use inferior materials. High temperature nickel based alloys are well know materials widely used through out industry.
John
John, thanks for this. Reading above mentioned temperatures seems high, only for the normal life of most people. The uncooled temperature limit of usual alloys gives some much needed perspective. I didnt know at all that rocket engine turbines can withstand 1700F (925 C).
That is true on the fuel rich side. I don't know about the Oxygen rich side. Probably lower. I don't have specs on Mondaloy type alloys. I am assuming lower. Trying to be conservative and starting with 1000F for the oxygen rich turbine.
John
Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.
John
Maybe what SpaceX has done is an adaptation of such an alloy and made one that is 3D printable.Are you concerned that Mondaloy type alloys in the LOX pre-burner will not be able to handle 1000F to 1400F temperatures over say 1000 start cycles? If not, how low do we have to go? I cannot find any specs on Mondaloy.
John
I found this:
"An alloy having the weight percent composition of
72.9 nickel, 16.6 cobalt, 8.1 chromium, 1.5 aluminum and
3.9 titanium was prepared. The alloy has been tested in high
pressure gaseous oxygen environments generally more
harsh than or Similar to a full-flow Staged combustion and
oxygen-rich Staged combustion rocket engine. The alloy
exhibited both high tensile and high burn resistance.
The entire patent is at :
https://patentimages.storage.googleapis.com/58/4f/d7/4d0e60f1762cd4/US20030053926A1.pdf
... What SpaceX needs is one that can be easily manufactured into complex parts/shapes probably through 3D printing.
I believe they have already settled on an specific one and are into producing parts. It is just we do not know what the alloy is and what that alloy's properties are other than that it meets the pressure, temperature, and long life that SpaceX want for its Raptor engine.... What SpaceX needs is one that can be easily manufactured into complex parts/shapes probably through 3D printing.
That sounds like they are still looking for it... are you sure they aren't already in production?
I am a Jet engine guy and I get this same argument. They find it hard to believe that the combustion Pressure is over 100 PSI less than the defused compressor discharge :o. and that the entire flow inside the engine is subsonic (except the Fan discharge). in this case it is likely a lot lower.
Is it known what the russian anti oxidation coatings are? The patent posted upthread was for an alloy that did not rely on a coating, which apparently is what Mondaloy also does. What then did the Russians use?
...Russians ORSC use standard cast nickel alloys that are more heat resistant than 718 but prone to oxygen fires, with fragile antioxidation coatings.
Is it known what the russian anti oxidation coatings are? The patent posted upthread was for an alloy that did not rely on a coating, which apparently is what Mondaloy also does. What then did the Russians use?
...Russians ORSC use standard cast nickel alloys that are more heat resistant than 718 but prone to oxygen fires, with fragile antioxidation coatings.
A. Some type of ceramic?
B. Some type of diffused aluminide with Pt/Rh?
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I have no idea what that really means, but it seems very relevant to this thread.
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I have no idea what that really means, but it seems very relevant to this thread.
Probably a hefty saftey margin. It's not like they can just whip up a better batch if they decide to increase the chamber pressure...https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I have no idea what that really means, but it seems very relevant to this thread.
800 ATM?! That's just a bit more than I might have expected...
Probably a hefty saftey margin. It's not like they can just whip up a better batch if they decide to increase the chamber pressure...https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I have no idea what that really means, but it seems very relevant to this thread.
800 ATM?! That's just a bit more than I might have expected...
In the creation of the RD-170 engine, the most powerful, many problems appeared. Pumps responsible for the supply of chambers in fuels were subjected to strong constraints, pump in one stage for the oxidizer and the pump in two stages for the fuel. The pump for the oxidizer work at 14 000 trs / min under a pressure of 600 atm, the pump of fuel of the first level was at 500 atm and the second stage one at 800 atm.
About this SX stuff: Single-crystal superalloys (https://en.m.wikipedia.org/wiki/Superalloy#Single-crystal_superalloys)https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I have no idea what that really means, but it seems very relevant to this thread.
https://en.m.wikipedia.org/wiki/Inconel
Single crystal (SX) superalloys have wide application in the high-pressure turbine section of aero and industrial gas turbine engines due to the unique combination of properties and performance.
About this SX stuff: Single-crystal superalloys (https://en.m.wikipedia.org/wiki/Superalloy#Single-crystal_superalloys)
run fuel at up to 36,000 psi (that 2,449 psi - to keep apples to apples).
12,000 psi isn’t totally insane when looking at fault tolerance of high pressure vessels... I’ve said it before - modern diesel common rail engines run fuel at up to 36,000 psi (that 2,449 psi - to keep apples to apples). No one seems to freak out with that...
About this SX stuff: Single-crystal superalloys (https://en.m.wikipedia.org/wiki/Superalloy#Single-crystal_superalloys)
SX300, SX500 will most certainly be for a proprietary SpaceX alloy, likely suitable for welding or laser sintering (like Inconel 718). Single crystal is only of use for getting utmost temperature capability (perhaps another 50-100K) in gas turbines inlet temperatures, it is finicky to make and crazy expensive to set up for it also needs cooled disk, firtree-root inserts for the single crystal blades and internal cooling passages. All of which are heavy and unlikely to feature in a super light weight Raptor turbine that operates at much lower temperatures. Creep life probably won't be a significant issue in Raptor turbines given only 10 hours of engine life.
The results reported here represent one of the few successes obtained in producing single-crystal epitaxial deposits through a powder bed fusion-based metal AM process and thus demonstrate the potential of SLE1 to repair and manufacture single-crystal hot section components of gas turbine systems from nickel-based superalloy powders.
https://twitter.com/elonmusk/status/1008385171744174080 (https://twitter.com/elonmusk/status/1008385171744174080)QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I have no idea what that really means, but it seems very relevant to this thread.
800 ATM?! That's just a bit more than I might have expected...
I believe the real kicker will be not the 300 Bar chamber pressure, but the much higher pre-turbine pressures that exist between the pump outlets and the turbine inlets on both sides... :o
Likely eye watering Bar numbers...
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:
There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.
So one scenario would be:
Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar
In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:
There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.
So one scenario would be:
Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar
In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:
There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.
So one scenario would be:
Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar
In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.
The test engine currently operates at 200 atmospheres, or 200 bar. The flight engine will be at 250 bar, and we believe that over time we can get that to a little over 300 bar.
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:
There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.
So one scenario would be:
Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar
In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.
- sub-scale demonstrator engine was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.
- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar
Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8 (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)
This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:
There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.
So one scenario would be:
Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar
In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.
- sub-scale demonstrator engine was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.
- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar
Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8 (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)
This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Best mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
I agree it won't be a box in a box and the tanks will obviously be an integral part of the structure, but if the tank structure they tested was not remotely like the final tank that will be used what was the point of that expensive test (with 1200 tons of liquid oxygen)? What evidence is there that the tank pressurization will be greater than 2 atmospheres?
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
In fact that test tank stood perfectly well without being pressurized.
I assumed ~3 bar because that is what tanks typically are designed for. It could be as low as ~2 but I think it will be closer to 3 for a couple of reasons. 1) makes designing the inducers and pumps easier, and 2) stabilizes the structure.
John
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
In fact that test tank stood perfectly well without being pressurized.
Umm...
I admit, this is after the rupture, but it still looks rather floppy.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
In fact that test tank stood perfectly well without being pressurized.
Umm...
I admit, this is after the rupture, but it still looks rather floppy.
I don't think the tank pressure can be 3 atm, in Elon Musk's 2017 presentation he showed the tank rupturing at just over 2.2 atm.
Remember the comment from the Reddit AMA (https://www.reddit.com/r/space/comments/76e79c/i_am_elon_musk_ask_me_anything_about_bfr/dodga3y):QuoteBest mass ratio is achieved by not building a box in a box. The propellant tanks need to be cylindrical to be remotely mass efficient and they have to carry ascent load, so lowest mass solution is just to mount the heat shield plates directly to the tank wall.
There wont be separate tanks. They will be an integral part of the body. To carry the load the tank walls will have to be much sturdier than the flimsy test tank that could not stand without being pressured.
In fact that test tank stood perfectly well without being pressurized.
Umm...
I admit, this is after the rupture, but it still looks rather floppy.
You can see that the dome is actually holding shape pretty well despite this clearly no longer being pressurized.
Remember that picture from the 2016 IAC where there was a bunch of people standing in front of the tank at the manufacturing site?
I know for a fact that the tank wasn't pressurized back then.
Autogenous pressurization is well understood technology. Shuttle used it.
John
I think Elon said they will be vented/at vacuum during transit to mars, so they will act as a dewar vessel aka thermos bottle for the landing propellant stored inside the header tanks. So unpressured tanks for transit and landing on Mars, i guess.Autogenous pressurization is well understood technology. Shuttle used it.
John
The discussion was about whether-or-not the empty tank would require constant pressurization to maintain its shape (a la the Atlas baloon tank).
I was responding to Wannamoonbase about autogenous pressurization.
- As far as tank design, normal practice is to make the tanks (without pressurization) stable enough to support their weight and anything that sits on top of it. Sometimes this includes the fuel, sometimes not. Also wind loads are considered.
- In all cases, it cannot fly without pressure stabilizing effect.
John
During a Mars flight I would expect the outer tanks to be vented to vacuum, for insulation purposes. But a small amount of the props from the inner tanks would be used to re-pressurise all tanks prior to any high-thrust activities, or the tanks would just not be strong enough for EDL or landing.
Elon Musk said they would subcool the propellant in the header tanks by venting to vacuum ahead of landing. I think they may vent to the main tanks, so the mass would not be wasted. Would the header tank fuel be cold enough before pressure inside the main tank becomes too high?<snip complex answer with maths>
Looking back on the subject of the cause of the green startup flame , I got to thinking that there is a good chance that on this "first go" on this design, it would make sense to fit both (TEA/TEB and "Electric") and due to the unknowns of the startup behavior the tried and true use of TEA/TEB would remove an unknown in the early testing.
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?
I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?
I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber. For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely. Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system. It's coming, no doubt.
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?
I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber. For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely. Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system. It's coming, no doubt.
Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.
And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
With the FFSC cycle of the raptor engine, i'm not sure wether a hard start would be an issue for the combustion chamber, because both fuel and oxidiser should enter the combustion chamber as a gas, so there's should not be the possibility of fuel pooling inside the engine and then ignite when hit by liquid oxygen. On the other side, Raptor needs 3 separate ignition sources, one for each preburner and one for the combustion chamber. So they might have started one or the other with TEA-TEB and used an augmented spark ignitor at the same time...My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?
I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber. For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely. Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system. It's coming, no doubt.
Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.
And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
With the FFSC cycle of the raptor engine, i'm not sure wether a hard start would be an issue for the combustion chamber, because both fuel and oxidiser should enter the combustion chamber as a gas, so there's should not be the possibility of fuel pooling inside the engine and then ignite when hit by liquid oxygen. On the other side, Raptor needs 3 separate ignition sources, one for each preburner and one for the combustion chamber. So they might have started one or the other with TEA-TEB and used an augmented spark ignitor at the same time...My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?
I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber. For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely. Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system. It's coming, no doubt.
Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.
And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts.
My thinking is "RUD prevention". In the early going they could use the electric torch for a few milliseconds then revert to TEA/TEB to prevent a hard start.
How does TEA/TEB ignition prevent hard starts compared to an augmented spark igniter?
I'd think augmented spark igniter would be what the flight engine will have. There's no point in going to TEA/TEB and back to spark. You have access to easily vaporized propellants (methane and oxygen), which makes the spark ignition easy compared to RP1.
Spark igniters can be tricky; you have to be sure that it's putting enough heat energy into the chamber. For a suitable bolus of TEA/B, that's not a problem – it will light the engine safely. Also, they may just have been in a rush to get off first firings and didn't want to take the time to develop a spark ignition system. It's coming, no doubt.
Are we sure we see a TEA/TEB ignition and not a camera white balance issue? The evidence I have seen so far is inconclusive. Also, how is a TEA/TEB ignition simpler when a spark ignited blow-torch is the design choice? You would have to feed the two liquids into the combustion chamber which would require some significant plumbing, not sure this would be faster than blow torches. I agree its possible, but I dont think its simpler.
And regards to spark ignition. I hope I understand that correctly, if not please correct me:
If the combustion chamber starts to fill with fuel and oxidizer, you need to ignite that very fast, otherwise there is a potential that after a few milliseconds, too much of both is inside the chamber before ignition and you get a hard start, meaning the engine explodes. This is particular a problem of spark ignition because it is very localized source and if not the right mixture of CH4 and O2 are at the spark source, a delayed ignition can lead to a hard start. But I dont think this can happen in Raptor. If I understand correctly, they use a blowtorch with small amount of CH4 and O2 which gets ignited ahead of the chamber. Like a has stove. Once confirmed burning, the main fuel flow starts. The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
There is a relevant comment by warp99 on Reddit about this Tweet. I cannot speak for its accuracy, but I am sure that people here can:
There will be a significant pressure drop across the injectors usually at least 50% of the combustion chamber pressure to ensure combustion stability.
So one scenario would be:
Tank pressure 2 bar
Pump output 800 bar
Preburner output 790 bar
Turbine output 400 bar
Injector input 390 bar
Combustion chamber 250 bar
In this case the mass flow through the turbine and pump section is the same (full flow) and the pressure drop across the turbine section and the pump section is the same at around 400 bar. If they are confident of meeting the pressure requirements this simplifies the mechanical design of the turbopump with no gearing requirement and similar turbine and pump chamber volumes.
300bar was the target chamber pressure of the flight engine... 250bar on sub-scale version.
- sub-scale demonstrator engine was said to be ~200 bar MCC.
- first iteration of full scale engine was said to be ~250 bar MCC.
- later iterations to hit 300 bar MCC.
- Tank pressure probably closer to 3 atmospheres.
- My calculations for maximum pump output pressure to be ~440 bar (for MCC = 250 bar)
- Pre-burner output (turbine inlet) pressure ~370 bar (fuel side, includes 41 coolant & 28 injector bar drop)
- Turbine output pressure ~278 bar
Current assumptions:
- all injector drops ~28 bar
- coolant drop ~41 bar
- pump efficiencies ~.8 (lox pump might be higher, fuel pump lower)
- turbine efficiencies ~.8 (probably higher)
This is a work in progress. I am currently designing the pumps and turbines. 2 stage fuel pump, 1 stage lox. I might have to add a boost pump. Both turbines appear to be single stage.
I'm a jet engine guy and I can tell you that in that application the fuel from the start nozzle must spray directly on the plug to get ignition. In this Raptor application the spark igniter will likely ignite a "small rocket engine" which will exhaust into either the pre-burners or MCC. The the ignition must propagate to all the the injectors (assuming coaxial). The velocity of the mixed fuel exiting the injectors needs to be lower than the flame propagation velocity (in the pre-burners) to allow each injector to ignite, "nothing could go wrong"
I'm a jet engine guy and I can tell you that in that application the fuel from the start nozzle must spray directly on the plug to get ignition. In this Raptor application the spark igniter will likely ignite a "small rocket engine" which will exhaust into either the pre-burners or MCC. The the ignition must propagate to all the the injectors (assuming coaxial). The velocity of the mixed fuel exiting the injectors needs to be lower than the flame propagation velocity (in the pre-burners) to allow each injector to ignite, "nothing could go wrong"
The blowtorch is too small to cause a hard start and the existing flame makes a hard start of the remaining propellants virtually impossible.
Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.
That strikes me as the sort of Elon statement that is liable to be over parsed (are ignition fluids counted as "propellants", etc). Or the type of engineering issue that is liable to end up changing when the final design actually makes it into production, even if it was one of the earlier goals.When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.
Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.
No, Raptor cannot use TEA/TEB, helium, or nitrogen. They add major operational headaches both on Earth and especially on Mars.Hasn't Elon insisted that the rocket would only use 2 propellants (methane and oxygen)? Meaning not using inert gasses for pressurization and for cold gas thrusters. That also precludes TEA and TEB from being used on Raptor.
That strikes me as the sort of Elon statement that is liable to be over parsed (are ignition fluids counted as "propellants", etc). Or the type of engineering issue that is liable to end up changing when the final design actually makes it into production, even if it was one of the earlier goals.
When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.Humans do not mind some nitrogen in the air. Well, it is not absolutely necessary. Helium does the job also, but... Pure oxigen is not ideal, and not only because of the fire. I am not sure if we have discussed this..
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.Humans do not mind some nitrogen in the air. Well, it is not absolutely necessary. Helium does the job also, but... Pure oxigen is not ideal, and not only because of the fire. I am not sure if we have discussed this..
When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.
When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.
If nitrogen and argon were not available on Mars, what would be the alternative for thrusters? Could you use cold methane or oxygen as thruster gas? At least then you would be unlikely to run out.
When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.
If nitrogen and argon were not available on Mars, what would be the alternative for thrusters? Could you use cold methane or oxygen as thruster gas? At least then you would be unlikely to run out.
I'm reading that there's some Nitrogen and Argon in the atmosphere, so once the CO2 freezes out, you have them for free.When he says this, isn't Elon thinking of what can be made on Mars. If you can't make TEA and helium or nitrogen on Mars, he's does not want to depend on those items. Every pound of thruster nitrogen you carry to Mars is a pound less of payload. I would think SpaceX will try very hard to avoid that.Humans do not mind some nitrogen in the air. Well, it is not absolutely necessary. Helium does the job also, but... Pure oxigen is not ideal, and not only because of the fire. I am not sure if we have discussed this..
-----
ABCD: Always Be Counting Down
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.
John
The next question is the starting method , Bootstrap or high pressure turbine impingement spinup or, something else.I'm betting on them using some of the RCS high pressure methane and lox for spinning up the respective turbines.
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.
John
Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.
John
Ok, so not quite 600 C. What do you think the temp drop across the turbine and injector might be?
I am currently calculating a turbine total pressure ratio of about .7-.75 and a total temperature out of about 500 C. Total temperature will not change through the injector, but static temperature drops with the pressure drop across the injectors according to adiabatic flow.
John
I am currently using 1000 F (540 C) for my turbine inlet temperatures for the 250 bar MCC. This gives me temperature margin for the 300 bar MCC growth version. All this is subject to change, since I am not done with my trades, but it is in the ball park.
John
Created Raptor pre-burner combustor designs to achieve deep throttling as well as a spark-ignited torch igniter that was capable of starting the engine from on-board autogenous blow-down propellants.
Some posts back there was speculation about the ignition method for Raptor. Found this on LinkedIn:
John Bucknell (https://www.linkedin.com/in/john-bucknell-pe-7111a47/):QuoteCreated Raptor pre-burner combustor designs to achieve deep throttling as well as a spark-ignited torch igniter that was capable of starting the engine from on-board autogenous blow-down propellants.
He left SpaceX in 2012, so this might have changed...
"Designed and built the subscale Raptor rocket for proof of concept testing able to test eighty-one configurations of main injector."
Oh wow, that's a lot of prototypes.
I’ve been working on Mars for the last four years, so I’m not going to take any credit for the Block 5 engine and all the upgrades that have happened
>
So, it seems the real work on Raptor started in 2014...
Slight change in subject.
Looking at the proximity of McGregor (4 miles) 45 deg. to the sub scale test stand and Oglesby (2.72 miles) Which is aligned with what looks like the production test stand (at least for full scale development and certification testing) I would think that the Raptor will cause some noise issues. Any talk of some suppressed test stands for the Raptor?
The only saving grace is the characteristics of a supersonic jet noise propagation
From "Ingeniare. Revista chilena de ingeniería, vol. 14 Nº 3, 2006"
"The maximum radiation of a jet exhaust, which is highly directional and has maximum
intensity at angles of between 30 Deg. and 45 Deg."
Which sort leaves, at least the center of town, in the cone of silence :)
The article was for ground test of rocket engines.Says jets in the text you gave, is this a translation issue?
I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.
Says jets in the text you gave, is this a translation issue?
I got it now. In my defense, the stuff at the left doesn't look very "engine-ey" in that the (apparent) bell seems smaller than the (apparent) combustion chamber.I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.
Not quite sure to whom that is addressed, but my post was in response to:Says jets in the text you gave, is this a translation issue?
I got it now. In my defense, the stuff at the left doesn't look very "engine-ey" in that the (apparent) bell seems smaller than the (apparent) combustion chamber.I would have thought that the illustration would have made it clear I was referring to rocket engines. But i like your clarification.
Not quite sure to whom that is addressed, but my post was in response to:Says jets in the text you gave, is this a translation issue?
This issue is known as denotation vs. connotation. The denotation is the original literal meaning of a word. The connotation is the way people use it in common vernacular. The present connotation of jet obviously is an airplane powered by turbojet or turbofan engines.
I would think the latter...
I would think the latter...
I am confused as the Wiki shows 2 different Vac Isp's. Is 356s the Isp reached, or is it due to a redesign from the previous announcement?
Does today's test video show a full scale Raptor running?
Does today's test video show a full scale Raptor running?
That’s my assumption. It looked like a different test stand (?), so I assume it was the full size engine. I hope we get some clarification.
I have a feeling that they ran into a big issue with designing the vacuum regen cooled nozzle.Could be. There were two alternate reasons given though:
It needed to be very large while still being light enough to not chew up the ISP gains, and still able to handle the thermal flux of the exhaust, which previously Elon has described as 'nuts'.
Does today's test video show a full scale Raptor running?
Does today's test video show a full scale Raptor running?
In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.
Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.
Sounds like we have a 300 bar.
Does today's test video show a full scale Raptor running?
In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.
Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.
Sounds like we have a 300 bar.
To be slightly more specific, EM called the Raptor a "200-ton class" engine. He also stated that the chamber pressure is "300 bar, approximately".
I have a feeling that they ran into a big issue with designing the vacuum regen cooled nozzle.Could be. There were two alternate reasons given though:
It needed to be very large while still being light enough to not chew up the ISP gains, and still able to handle the thermal flux of the exhaust, which previously Elon has described as 'nuts'.
(1) Get started with just one engine design instead of two (with the second only used for 4 out of 38 engines)
(2) More survivable options in the case of engine out.. can survive 4 engines failing and still land.. could be particularly useful early on while working out the bugs!
I updated my scale draw with 20180917 announced size.
;)
Titus
Does today's test video show a full scale Raptor running?
That’s my assumption. It looked like a different test stand (?), so I assume it was the full size engine. I hope we get some clarification.
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?
45m28s
So this is the Raptor engine that will power BFR, both the Ship and the Booster, it's the same engine. And this is approximately 200 tonne thrust engine. That's aiming for a roughly 300 bar or three hundred atmosphere chamber pressure.
And - depending upon ... if you have it at a high expansion ratio - has the potential to be, to have a specific impulse above 380.
And it's a staged combustion, full flow, gas-gas. [Praise for the propulsion and other teams.]
(https://www.youtube.com/watch?v=zu7WJD8vpAQ?t=45m28s)
1h34m50s (https://youtu.be/zu7WJD8vpAQ?t=5690)
Hey Elon, Tim Dodd the Everyday Astronaut here. I see that you changed the engine configuration for the BFS. Can you talk a little about, you know, if there's still engine out capability? Is it vacuum optimized, but still landable on sealevel? Can they function as an abort system? Can you just kind of tell us about your new decision making on that?
Oh yeah, actually you noticed that. That's a good thing to notice, good eye.
So in order to minimize the development risk and cost we decided to commonize the engine between the Booster and the Ship.
So a future upgrade path for BFS would be to have a vacuum optimized nozzle. [pointing at diagram on wall] (https://i.imgur.com/jduFRIg.jpg) So these nozzles are kind of a sea level sized nozzle so they are able to operate well at sea level, they are essentially the booster sized nozzle.
Where you see that cargo around the perimiter, you can actually switch out those cargo sections for a vacuum nozzle version of Raptor. And the vacuum nozzle can go all the way to the perimeter, basically the skin of the vehicle. So you can have something which has maybe 3 or 4 times the exit diameter of the raptors that you see there as engines in the perimeter, and the exchange would be you that you'd loose basically two of those cargo racks in exchange for every vacuum engine, but then your total payload performance to Mars would increase significantly.
But we can do the 100 tonnes to surface of Mars with those engines, but I think version 2 would have the vacuum engines most likely in place of those cargo racks.
Having those engines in that configuration with 7 engines means it's definitely capable of engine out at any time including two engine out in almost all circumstances. So you could loose two engines and still be totally safe. In fact some cases you can loose up to 4 engines and still be totally fine. It only needs 3 engines for landing, 3 out of 7.
Does today's test video show a full scale Raptor running?
In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.
Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.
Sounds like we have a 300 bar.
To be slightly more specific, EM called the Raptor a "200-ton class" engine. He also stated that the chamber pressure is "300 bar, approximately".
This would also mean a (very roughly) 1.1-1.2 TWR for SSTO, meaning they will have much more commonality between the early hopper and final craft. Earlier on they wanted to use a different engine configuration depending on the tests they're going to run, I could see them moving straight to final production processes and tooling now.
Does today's test video show a full scale Raptor running?
In the stream Musk said it was 200 tons, which is 1993 kN or 448,000 lbf.
Wiki has 1,700 kN or 380,000 lbf for the 250 bar engine.
Sounds like we have a 300 bar.
To be slightly more specific, EM called the Raptor a "200-ton class" engine. He also stated that the chamber pressure is "300 bar, approximately".
This would also mean a (very roughly) 1.1-1.2 TWR for SSTO, meaning they will have much more commonality between the early hopper and final craft. Earlier on they wanted to use a different engine configuration depending on the tests they're going to run, I could see them moving straight to final production processes and tooling now.
Not sure. Bet the longer more voluminous BFS plus 3 big fins/wings/legs masses well above the old 85 tonnes
Would a vacuum nozzle extension constructed in such a way make any sense? I've seen mention of Horizontal Flow engine bells (http://www.aerospaceweb.org/design/aerospike/outflow.shtml)
So you can have something which has maybe 3 or 4 times the exit diameter of the raptors that you see there as engines in the perimeter, and the exchange would be you that you'd loose basically two of those cargo racks in exchange for every vacuum engine,
Is there something odd with methlox and vacuum nozzles overall? First Blue Origin shelved development of a vacuum-optimized BE-4, and now SpaceX pushed VacRaptor, in the future...
Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.Looks still slightly overexpanded in image. Look at flame dia. at mach disk, it looks smaller than nozzle dia. So looks like flight spec. nozzle on Raptor being tested.
This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).
(added image)
Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.Looks still slightly overexpanded in image. Look at flame dia. at mach disk, it looks smaller than nozzle dia. So looks like flight spec. nozzle on Raptor being tested.
This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).
(added image)
At startup I didn't see the green flash.
Looks like they now have spark ignition working. ;)
Please watch the full video of the Raptor test. You see it throttle down a few seconds after start and the mach disk dia. decreases with decreasing chamber pressure. Mach disk to nozzle dia. ratio is dependent on the nozzle exit pressure if Pe does not equal Pa. Optimal expansion will yield no shock diamonds. During throttle down shock diamonds are always present indicating that the nozzle is still overexpanded at full thrust. If nozzle is underexpanded at full thrust then it will reach optimum expansion at a specified thrust setting then go overexpanded as throttle is further lowered. Going from underexpanded to optimal to overexpanded during throttle down you would very briefly see the shock diamonds disappear before returning again.Measuring off the video and assuming the handrails are a standard 42" height, the Raptor shown is 1.1 meters diameter and slightly underexpanded.Looks still slightly overexpanded in image. Look at flame dia. at mach disk, it looks smaller than nozzle dia. So looks like flight spec. nozzle on Raptor being tested.
This fits with the full-size Raptor being 1.3 meters and optimally expanded or slightly overexpanded at SL, and with this being a full pressure or near full pressure test (since at lower pressure it would look overexpanded).
(added image)
This is of great general education interest. The simple public illustrations of over versus under expansion by visual inspection overlook this detail of complex plume features.
How does mach disc diameter relate to nozzle exit diameter and pressure?
Maybe this question and answer belong in the questions and answers forum for better discoverability?
At startup I didn't see the green flash.Laser ignition was AFAIK still on the table.
Looks like they now have spark ignition working. ;)
At startup I didn't see the green flash.
Looks like they now have spark ignition working. ;)
Here is the Raptor startup in GIF form...
At startup I didn't see the green flash.
Looks like they now have spark ignition working. ;)
Here is the Raptor startup in GIF form...
Is that a pilot methalox torch before the vapor from the full propellant flow? That would indicate spark ignition
After Mr Musk showed the Raptor engine test and praised the SpaceX propulsion team, he said this:My bet: staging timing choice have direct impact on difficulty of recovering.
"I don't think most people, even in the aerospace industry, like, know what question to ask, like it took us a long time to even frame the question correctly. Look, once we could frame the question correctly, the answer was, I wouldn't say easy, but, the answer flowed once the question could be framed with precision. Framing that question with precision was very difficult."
Anyone here have any idea of what he was talking about? Is this about fundamentals; propellant selection, cycle choice, on what scale the engine will be built?
My guess would be it took a long time come up with a way to have the project coupled with economics size the engine correctly. Not a confident guess.
Matthew
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
BFS and Raptors Optimized for Point to Point Travel, and not Mars.
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?
Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?
Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.
The base of the BFB flares out. The diameter is close to 10m around the bottom.
With the new chamber pressure and thrust details, does any one have any good ideas on what the new ISP might be if it changes at all?I too wonder about the Isp for these "sealevel like" engines at SL and in an a vacuum. Musk said that the thrust is now 200 tonnes (up from 170) and that the chamber pressure has increased to 300 bar from 250.
We also know from Musk that they would be 380 (or more?) with a vacuum optimized nozzle, but what is it with the current nozzle? Assuming the nozzle is still the exact same size as the SL version from a year ago, can any of your rocket engine gurus estimate the Isp increase?
Also Musk's wording about the nozzle diameter was a bit peculiar. Could it indicate a nozzle that this slightly bigger than a normal SL optimized nozzle?
Just saw on twitter Elon said BFR booster was starting at be 31 Raptors, but could fit 11 more. 42 in 9 metres means max Raptor nozzle OD >=1.2m.
The base of the BFB flares out. The diameter is close to 10m around the bottom.
Missed that. 42 @1.3m fit in 10m
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018
https://twitter.com/elonmusk/status/1042525258899550209?s=19
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018
https://twitter.com/elonmusk/status/1042525258899550209?s=19
I wonder why they “kind of have to” add more engines later.
I wonder why they “kind of have to” add more engines later.Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk <span class="emoji-outer emoji-sizer"><span class="emoji-inner" style="background: url(chrome-extension://immhpnclomdloikkpcefncmfgjbkojmh/emoji-data/sheet_apple_32.png);background-position:97.94359576968273% 95.94594594594594%;background-size:5418.75% 5418.75%" data-codepoints="2714-fe0f"></span></span> @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018
https://twitter.com/elonmusk/status/1042525258899550209?s=19
I wonder why they “kind of have to” add more engines later.
Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.
I wonder why they “kind of have to” add more engines later.Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.
As one specific example, I wondered how this drop from 150 to "100+" would affect the target of being able to perform all F9 and FH missions. You need huge margins to be able to do what an expendable FH does, while remaining reusable, without refueling.
Actually, it just occurred to me, maybe this drop in performance could actually be a preparation to exceed the 150 ton figure. They did increase the volume after all. What if they had decided to oversize the upper stage, even though this would drop performance in the short term, so that more iterative performance could be delivered later while not making any future radical changes to the very difficult upper stage.
Missed it originally.... Brilliant!I wonder why they “kind of have to” add more engines later.Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.
As one specific example, I wondered how this drop from 150 to "100+" would affect the target of being able to perform all F9 and FH missions. You need huge margins to be able to do what an expendable FH does, while remaining reusable, without refueling.
Actually, it just occurred to me, maybe this drop in performance could actually be a preparation to exceed the 150 ton figure. They did increase the volume after all. What if they had decided to oversize the upper stage, even though this would drop performance in the short term, so that more iterative performance could be delivered later while not making any future radical changes to the very difficult upper stage.
Isn't Elon just referencing Hitchhiker's Guide to the Galaxy here? 31+11 = 42 engines. If he's going to name his Mars ship Heart of Gold, that's why he has to add more engines to make it 42. It's not a technical reason, it's an Elon reason.
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk ✔ @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018
https://twitter.com/elonmusk/status/1042525258899550209?s=19
BFS and Raptors Optimized for Point to Point Travel, and not Mars.
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.
A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.
Note 100+ includes 150.I wonder why they “kind of have to” add more engines later.Laymans guess: could be tied to why they are now saying "100+ tons" instead of "150 tons".. maybe they are not hitting the performance they were designing for in this first iteration.
I am using prior generations figures, as we do not have a useful dry mass for the present version. If you calculate it strictly according to the things said in the announcement, as outlined above, you get around 60, which seems very, very optimistic.BFS and Raptors Optimized for Point to Point Travel, and not Mars.
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.
A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.
No, it cannot.
If the normal BFS LEO capacity is 100 tonnes, a tanker which is exactly 25 tonnes lighter but has exactly 25 tonnes bigger tanks lifts exactly 125 tonnes of fuel to orbit. And if the tanks are of same size, then it's ls than 125 tonnes.
Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk <span class="emoji-outer emoji-sizer"><span class="emoji-inner" style="background: url(chrome-extension://immhpnclomdloikkpcefncmfgjbkojmh/emoji-data/sheet_apple_32.png);background-position:97.94359576968273% 95.94594594594594%;background-size:5418.75% 5418.75%" data-codepoints="2714-fe0f"></span></span> @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018
https://twitter.com/elonmusk/status/1042525258899550209?s=19
I wonder why they “kind of have to” add more engines later.
Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.
x33’s TPS was pretty lightweight.Christian Daniels @CJDaniels77
Replying to @elonmusk
Did the number of engines on the booster change at all?
|
Elon Musk <span class="emoji-outer emoji-sizer"><span class="emoji-inner" style="background: url(chrome-extension://immhpnclomdloikkpcefncmfgjbkojmh/emoji-data/sheet_apple_32.png);background-position:97.94359576968273% 95.94594594594594%;background-size:5418.75% 5418.75%" data-codepoints="2714-fe0f"></span></span> @elonmusk
31 engines, but with room to add 11 more down the road. Kinda have to.
5:26 PM - Sep 19, 2018
https://twitter.com/elonmusk/status/1042525258899550209?s=19
I wonder why they “kind of have to” add more engines later.
Future iterations of BFB/BFS will likely need more lifting power, I'm guessing.
One guess is the heat shield material weighing more than planned.
If they go for non ablative shields they are all heavier than ablative.(correct me if I am wrong).
Examples:
1. carbon-carbon
2. ceramic tiles
3. metallic x-33 style
BFS and Raptors Optimized for Point to Point Travel, and not Mars.
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.
A 1185 ton initial mass tanker can deliver one half its fuel mass (550 tons) to a tanker in an orbit with 1800m/s delta-v, and 350 ISP and then burn back for home at a little more than 1800m/s.
No, it cannot.
If the normal BFS LEO capacity is 100 tonnes, a tanker which is exactly 25 tonnes lighter but has exactly 25 tonnes bigger tanks lifts exactly 125 tonnes of fuel to orbit. And if the tanks are of same size, then it's ls than 125 tonnes.
Has anyone analyzed the new raptor firing video? Does it look similar to the one from last year or can you tell any differences?It appears to be the same test stand with a different engine
I have zero expertise on rocket engines, hence the question.
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.Isp is not unimportant if you are interested in whether there could be a version of the BFS that could do SSTO with a meaningful payload. Given that the market for large GEO sats is in decline and the market for small LEO sats is increasing, this could be relevant to SpaceX in the future. The 7 sea level engines would be able to get the fully loaded thing off the ground at over 1.2g. But now the vacuum Isp may end up being the limiting factor. This is why it is relevant.
Most recently, photos captured earlier this summer showed that a new prototype was installed on SpaceX’s horizontal Raptor test stand in McGregor, Texas, looking nearly identical to the deep black Raptor nozzle shown in Monday’s presentation. Previous Raptor prototypes seen during testing or at the test stand appeared to have a nozzle closer to SpaceX’s silver Merlin 1Ds, whereas this newest iteration’s nozzle doesn’t seem to reflect the powerful spotlights surrounding it.
Is the new Raptor engine design similar to the SSME that can go from sea level to vacuum? I know it looses ISP at altitude, but is that the trade off for cost and simplicity.Yes, and basically every sea level engine can do that...
ISP is unimportant if fuel transfer works, and BFS is reliable. Major changes in architecture cost can utterly swamp it.Isp is not unimportant if you are interested in whether there could be a version of the BFS that could do SSTO with a meaningful payload. Given that the market for large GEO sats is in decline and the market for small LEO sats is increasing, this could be relevant to SpaceX in the future. The 7 sea level engines would be able to get the fully loaded thing off the ground at over 1.2g. But now the vacuum Isp may end up being the limiting factor. This is why it is relevant.
Agreed. All rocket engines increase Isp with altitude. I have calculated that the exhaust pressure of the SL raptor is likely about 0.6 atm. This is equivalent to the engine being optimized for an altitude of ~4000 m or ~13,000 ft.Is the new Raptor engine design similar to the SSME that can go from sea level to vacuum? I know it looses ISP at altitude, but is that the trade off for cost and simplicity.Yes, and basically every sea level engine can do that...
Property | 2017 | 2018 |
Chamber Pressure | 250 bar | 300 bar |
Thrust at sea level | 1,700 kN | 2,095 kN |
Thrust in vacuum | 1,834 kN | 2,229 kN |
Specific Impusle at sea level | 330 sec | 335 sec |
Specific Impusle in vacuum | 356 sec | 356 sec |
I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...
I know that the vac Isp largely depends on expansion ratio, but not solely. Chamber pressure should have at least some effect.I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...
Vac isp depends almost completely on expansion ratio. This is why throttling an engine in vacuum does not really hurt it's specific impulse.
There's more metrics to a rocket engines performance when in an actual rocket than ISP alone. Going for 300 bar should optimise another metric, which is the thrust to weight ratio.True and that is visible in the increased thrust. On the other hand, vehicle dry mass seems to have increased by at least 2 tonnes due to the increased length and increased fin size plus extra fin.
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers. The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon. The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design. It is assumed that all else has remained the same. It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.
Predicted Raptor Sea Level Engine Performance
Property 2017 2018 Chamber Pressure 250 bar 300 bar Thrust at sea level 1,700 kN 2,095 kN Thrust in vacuum 1,834 kN 2,229 kN Specific Impusle at sea level 330 sec 335 sec Specific Impusle in vacuum 356 sec 356 sec
I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing. He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS. He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged. For these reasons, predictions on the vacuum Raptor would be too speculative and therefore to worth publishing. If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).
envy887 is right. Vacuum Isp is a product of the vacuum thrust coefficient and the characteristic velocity. The vacuum thrust coefficient reflects the exhaust expansion properties and the design of the nozzle; I assumed neither of these changed. The characteristic velocity is basically a function of the propellants, which also did not change. I double checked andI know that the vac Isp largely depends on expansion ratio, but not solely. Chamber pressure should have at least some effect.I am surprised that the vacuum ISP has not improved at all compared to the 2017 version, despite the 20% higher chamber pressure...
Vac isp depends almost completely on expansion ratio. This is why throttling an engine in vacuum does not really hurt it's specific impulse.
With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers. The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon. The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design. It is assumed that all else has remained the same. It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.
Predicted Raptor Sea Level Engine Performance
Property 2017 2018 Chamber Pressure 250 bar 300 bar Thrust at sea level 1,700 kN 2,095 kN Thrust in vacuum 1,834 kN 2,229 kN Specific Impusle at sea level 330 sec 335 sec Specific Impusle in vacuum 356 sec 356 sec
I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing. He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS. He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged. For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing. If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).
John,With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers. The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon. The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design. It is assumed that all else has remained the same. It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.
Predicted Raptor Sea Level Engine Performance
Property 2017 2018 Chamber Pressure 250 bar 300 bar Thrust at sea level 1,700 kN 2,095 kN Thrust in vacuum 1,834 kN 2,229 kN Specific Impusle at sea level 330 sec 335 sec Specific Impusle in vacuum 356 sec 356 sec
I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing. He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS. He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged. For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing. If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).
Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.
John
John,With the available data on the Raptor sea level engine, I have reverse engineered some other likely performance numbers. The performance numbers are based on the past presentations by SpaceX on the ITS and BFR systems as well as the September 17th, 2018 announcement of the first private crewed flight around the Moon. The only new information on the Raptor from the Moon announcement was the chamber pressure is now 300 bar instead of 250 bar from 2017 and no vacuum engine is in the design. It is assumed that all else has remained the same. It should be reasonable to assume the engine size has not changed since the diameter of the booster (9 m) and the number of engines (31) is the same.
Predicted Raptor Sea Level Engine Performance
Property 2017 2018 Chamber Pressure 250 bar 300 bar Thrust at sea level 1,700 kN 2,095 kN Thrust in vacuum 1,834 kN 2,229 kN Specific Impusle at sea level 330 sec 335 sec Specific Impusle in vacuum 356 sec 356 sec
I did not include the vacuum engine because (1) it has been removed from the design and (2) the comments Elon Musk made at the Moon announcement about the potential vacuum engine were very confusing. He said the vacuum engine would have "3 or 4 times the exit diameter of the Raptors," however this would be too large to fit on the BFS. He also said the vacuum engine could have an Isp of 380 sec; this may fit on a BFS however the other engines may have to be rearranged. For these reasons, predictions on the vacuum Raptor would be too speculative and therefore not worth publishing. If people are interested I could make estimates on the thrust chamber dimensions like I did for the Blue Origin BE-4 (https://forum.nasaspaceflight.com/index.php?topic=45518.0).
Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.
John
Great work! I calculated an ER of 32.0-33.6. My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL. There is definitely some uncertainty in my calculation. Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be? My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.
Thanks, Mike
Wow! Thank you, this is amazing! It appears you provided the output for the 2016 raptor. I honestly do not understand why the output Isp does not match the Isp SpaceX advertised, but I am excited to learn more about NASA's CEA. Thank you.John,
Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.
John
Great work! I calculated an ER of 32.0-33.6. My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL. There is definitely some uncertainty in my calculation. Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be? My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.
Thanks, Mike
I use NASA's CEA. It is free and on line. I have attached a typical output summary for the main chamber and pre-burners.
John
Wow! Thank you, this is amazing! It appears you provided the output for the 2016 raptor. I honestly do not understand why the output Isp does not match the Isp SpaceX advertised, but I am excited to learn more about NASA's CEA. Thank you.John,
Mike,
I did the engine geometry last year and its the same this year. BTW, my estimates match yours within a second or two. The numbers shown are last years, so just substitute this years numbers. Threw in my numbers just for comparison.
John
Great work! I calculated an ER of 32.0-33.6. My lower ER would explain the my higher calculated Isp since the engine is over-expanded at SL. There is definitely some uncertainty in my calculation. Do you have a good reference for what the heat capacity ratio and characteristic velocity of the Raptor should be? My estimate was 1.211-1.314 for heat capacity ratio and 1879-1975 m/s for characteristic velocity.
Thanks, Mike
I use NASA's CEA. It is free and on line. I have attached a typical output summary for the main chamber and pre-burners.
John
@livingjw
Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?
@livingjw
Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?
I believe that within a certain small % range of pressure increase, thrust and Isp all go up quite linearly with pressure, that does not whole true for example for a 100% increase.
@livingjw
Just out of curiosity. If the Raptor have higher chamber pressure (like for example 315 bar or 330 bar) in the future and the rest of the engine is mostly unchanged. What would be the changes to the thrust and the ISP of the engine?
I believe that within a certain small % range of pressure increase, thrust and Isp all go up quite linearly with pressure, that does not whole true for example for a 100% increase.
Linearly OK, but at what rate of change, isp/delta-bar? I hope it is 42!
Chamber Pressure | 250 bar | 300 bar | 315bar | 330bar |
Thrust at sea level | 1,700 kN | 2,095 kN | 2,206 kN | 2,318 kN |
Thrust in vacuum | 1,834 kN | 2,229 kN | 2,340 kN | 2,452 kN |
Specific Impusle at sea level | 330.0 sec | 334.6 sec | 335.6 sec | 336.5 sec |
Specific Impusle in vacuum | 356 sec | 356 sec | 356 sec | 356 sec |
. Vacuum Isp is not affected by chamber pressure.
Chamber Pressure 250 bar 300 bar 315bar 330bar Thrust at sea level 1,700 kN 2,095 kN 2,206 kN 2,318 kN Thrust in vacuum 1,834 kN 2,229 kN 2,340 kN 2,452 kN Specific Impusle at sea level 330.0 sec 334.6 sec 335.6 sec 336.5 sec Specific Impusle in vacuum 356 sec 356 sec 356 sec 356 sec
Jon, how did you estimate the fuel:oxy mass ratios for the preburners? Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....
. Vacuum Isp is not affected by chamber pressure.
Chamber Pressure 250 bar 300 bar 315bar 330bar Thrust at sea level 1,700 kN 2,095 kN 2,206 kN 2,318 kN Thrust in vacuum 1,834 kN 2,229 kN 2,340 kN 2,452 kN Specific Impusle at sea level 330.0 sec 334.6 sec 335.6 sec 336.5 sec Specific Impusle in vacuum 356 sec 356 sec 356 sec 356 sec
What about improvements in combustion efficiency when running the engine further still from stoichiometric? Does higher chamber pressure enable running even more methane-rich?
Other question: does increased chamber pressure mean decreased throat area? For a theoretical const-mdot engine?
Jon, how did you estimate the fuel:oxy mass ratios for the preburners? Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....
I chose 1000 F to minimize the pressure drop across the turbines needed to drive the pumps. The higher the turbine inlet temperature, the lower the pressure drop, the lower the overall pressure rise required of the pumps. Turbine materials are able to handle 100's of degrees higher temperatures, uncooled.
John
Here are some rough estimates. I expect these are accurate to within 2%. Basically the vacuum thrust scales linearly with chamber pressure, something like ~7.4 kN/bar. The sea level thrust is going to be 134 kN less than the vacuum thrust; this is based on the 1.3 m nozzle exit diameter. Vacuum Isp is not affected by chamber pressure.What would 300, 315 and 330 bar look like with 2.4 m and 1.7 m exit diameter bells?
Chamber Pressure 250 bar 300 bar 315bar 330bar Thrust at sea level 1,700 kN 2,095 kN 2,206 kN 2,318 kN Thrust in vacuum 1,834 kN 2,229 kN 2,340 kN 2,452 kN Specific Impusle at sea level 330.0 sec 334.6 sec 335.6 sec 336.5 sec Specific Impusle in vacuum 356 sec 356 sec 356 sec 356 sec
Isn't 1.7 m too large?I thought they would have nozzles with up to 2.4 meters diameter for the vac version?
Isn't 1.7 m too large?
There is obviously a lot of scope for further optimisation of Raptor nozzles for both booster and spaceship. I think the Ø9m booster suffers badly from being too small a diameter.The booster tail seems to not be 9m.
I think the Ø9m booster suffers badly from being too small a diameter. If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.
I think the Ø9m booster suffers badly from being too small a diameter. If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.
I agree on the problems with the smaller diameter. There appears to be advantages to increasing the diameter beyond 9 meters.
Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.
Maybe make the 1.0 model of the BFS at 9 meters and get flying. But why not order another tool.
That's my 2 cents from my armchair.
I think the Ø9m booster suffers badly from being too small a diameter. If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.
I agree on the problems with the smaller diameter. There appears to be advantages to increasing the diameter beyond 9 meters.
Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.
Maybe make the 1.0 model of the BFS at 9 meters and get flying. But why not order another tool.
That's my 2 cents from my armchair.
This is SpaceX. The minute the 9 m flies, they'll be working on the 12, or the 15.
Why would it be "decades" if even the first design didn't take even a single decade?
Is the planned expansion ratio still the same as in 2017 version? Or did they make the nozzle of the "sea level optimized" raptor slightly bigger in the 2018 version? (optimizing it for slightly higher than sea level to be better comphromize for vacuum, or optimizing it for higher chamber pressure even on sea level)
Increasing the nozzle size might explain the width increase in the base of the rocket?
Because their Mars plans are insanely large and they only went to 9 since the leap was too large, required extra infrastructure, etc.I think the Ø9m booster suffers badly from being too small a diameter. If not constrained by the tooling they would probably prefer to go for a much larger diameter ~6x the optimised nozzle diameter + spacing pitch (assuming outer ring directly aligned with tank walls) to allow 37 engines, thicker tank walls, a shorter stack, reduced booster surface area, and create more drag on re-entry to lower terminal velocity and reduce landing burn fuel requirements.
I agree on the problems with the smaller diameter. There appears to be advantages to increasing the diameter beyond 9 meters.
Yes tooling is expensive, but they are just starting and are making foundational decisions they may live with for decades.
Maybe make the 1.0 model of the BFS at 9 meters and get flying. But why not order another tool.
That's my 2 cents from my armchair.
This is SpaceX. The minute the 9 m flies, they'll be working on the 12, or the 15.
Why would it be "decades" if even the first design didn't take even a single decade?
Did SpaceX immediately start working on 4 and 5m diameter Falcon rockets after F9 flew in 2010?
Why would they when it is far easier to stretch the 9m? There is no real benefit for just going to 12m from 9m.
There are ground infrastructure issues that makes going larger than 9-10m difficult. If they plan on using existing launch sites, like 39A. And perhaps ANY land launch site in the continental US. I don't think they will be able to go bigger without a large floating launch infrastructure.
The 9m is not what they wanted to build, it's what they could practically build. Some parallels to F1.If you are looking at a rapid build up, launching more and more per synod, you cannot assume steady state economics - amortising vehicles over 10 synods is a nonsense if you're doubling every synod.
I guess we'll see... Chat again about this 4 years after first BSR Mars-bound flight?The 9m is not what they wanted to build, it's what they could practically build. Some parallels to F1.If you are looking at a rapid build up, launching more and more per synod, you cannot assume steady state economics - amortising vehicles over 10 synods is a nonsense if you're doubling every synod.
This means that capital cost is a major limit - and while in principle a 14m (say) rocket might have half the capital cost per ton, this does not make it meaningful in the context of launches costing sub $10/kg to LEO.
(P2P cost of 9m).
The vast majority of supplies to Mars do not (after the first several launches) need to go in 14m rockets, at half the capital cost per ton.
They are just fine in a 8.5m Aluminium tank, at perhaps 5PSI, with passive thermal control and no ECLSS, and just enough attitude control to remain nice and stable for a few hours once near Mars. (launched from, perhaps even herded by, and caught by a 'barn door' BFS.)
You may want large vessels for bulk passenger transport at some time in the future, but these are going to remain a lot more expensive than bulk cargo on-off delivery tanks.
(This is all post 2026, and first ISRU crew return on the nominal plan)
3) What will the first Raptor improvements be? Vacuum version perhaps? Increased reliability/reduced strain from reuse?I think one of the first long poles in the tent would be adjustments for rapid production.
... It will not work with methane.Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?... It will not work with methane.Says who?
Can you keep methane cold enough in the hydraulics to remain liquid?... It will not work with methane.Says who?
The methane in the tanks will be liquid. A hydraulic pump can bring the the methane up to a pressure that should prevent boiling. It is not ideal, but I think it is workable.Can you keep methane cold enough in the hydraulics to remain liquid?... It will not work with methane.Says who?
Yeah, but can you guarantee that all hydraulic fluid will circulate fast enough? One stagnant region near an edge of a cylinder, and you have gas pocket.The methane in the tanks will be liquid. A hydraulic pump can bring the the methane up to a pressure that should prevent boiling. It is not ideal, but I think it is workable.Can you keep methane cold enough in the hydraulics to remain liquid?... It will not work with methane.Says who?
There are a pile of interrelated problems in this area that smell like they might be amenable to a really clever combined solution.The methane in the tanks will be liquid. A hydraulic pump can bring the the methane up to a pressure that should prevent boiling. It is not ideal, but I think it is workable.Can you keep methane cold enough in the hydraulics to remain liquid?... It will not work with methane.Says who?
Instead of hydraulics, could they use a pneumatic system and use the methane in gaseous form?Not easily, because you want the TVC system to be stiff. Hydraulic fluid is incompressible, so when you put a load on the end of an actuator the pressure will increase, but the volume (and therefore position) stays the same. In a pneumatic system the actuator will move until the actuator pressure balances the applied load. Theoretically, if you had an extremely fast active system balancing the pressure on each side of the actuator piston using valves that have all the flow area and instant response time, it might be possible. It would also be the embodiment of "all you have to do is just...".
Yup. But as Geza says, electro-mechanical has potential. Also electro-mechanical with local closed hydraulic amplification.Instead of hydraulics, could they use a pneumatic system and use the methane in gaseous form?Not easily, because you want the TVC system to be stiff. Hydraulic fluid is incompressible, so when you put a load on the end of an actuator the pressure will increase, but the volume (and therefore position) stays the same. In a pneumatic system the actuator will move until the actuator pressure balances the applied load. Theoretically, if you had an extremely fast active system balancing the pressure on each side of the actuator piston using valves that have all the flow area and instant response time, it might be possible. It would also be the embodiment of "all you have to do is just...".
This is also why using liquid methane won't work. Even with magic insulation to keep things from boiling, the cylinders are dead-headed volumes where the fluid will not be circulated out constantly. The methane will go supercritical if it gets too hot and then the density becomes very nonlinear with respect to pressure and the system loses stiffness.
Jon, how did you estimate the fuel:oxy mass ratios for the preburners? Your preburner outputs look quite hot, with a full flow design that's an absurd amount of thermal power going into the turbines....
I chose 1000 F to minimize the pressure drop across the turbines needed to drive the pumps. The higher the turbine inlet temperature, the lower the pressure drop, the lower the overall pressure rise required of the pumps. Turbine materials are able to handle 100's of degrees higher temperatures, uncooled.
John
Evolution of Rolls-Royce air-cooled turbine blades and feature analysis (https://ac.els-cdn.com/S1877705814038065/1-s2.0-S1877705814038065-main.pdf?_tid=2a885133-3a4d-4cc5-b8ea-0f5c8b4d6a3b&acdnat=1538489242_c8deeea0cade4657c4a01d8d68b7ac9e)
Even the materials in the very earliest turbojets could handle 1000K (1340F) turbine inlet temperature. To me this feels like one of the big advantages of full-flow -- the temperatures are low compared to gas turbines.
I wonder how shops/vendors are coping with the increased demand from these megalithic engines. I'm talking about forging, and casting shops that already have typically long lead times. Has SpaceX bought up forges like they have done for machine shops? SpaceX is talking about putting a dozen engines per BFR, and I presume they will still be getting business making the Merlins.
I see a lot of the turbomachinery being castings, and forgings, especially with their proprietary alloys I presume only select few forges will be producing.
Blue's also producing their rocket engines, Aerojet Rocketdyne is as well (well, not so much with the announcement, but SLS RS-25 is still going on)
Forges don't pop up overnight either, and they have demand from other industries as well.
I'm fairly certain that SpaceX builds the Raptor engine inhouse and a lot of parts are 3D printedThe nozzle is only the large part, in addition, and it's probably not going to be cast.
I'm fairly certain that SpaceX builds the Raptor engine inhouse and a lot of parts are 3D printedThe nozzle is only the large part, in addition, and it's probably not going to be cast.
https://twitter.com/nasaspaceflight/status/1076615264043757569QuoteWhile we have you, Elon.... How well is Raptor performing during test stand firings at McGregor? On track to support your Super Heavy/Starship schedule?
https://twitter.com/elonmusk/status/1076616737020231681QuoteYes. Radically redesigned Raptor ready to fire next month.
Yes, full flow, gas-gas, staged combustion. Will take us time to work up to 300 bar. That is a mad level.
You def don’t want electric pumps on a rocket engine! Raptor turbopumps alone need 100,000 horsepower per engine. That’s not a typo.
Well what can I say - I'm getting the redesign blues here. :(
Possiblities:
- relocation of the lox pump
- addition of boost pumps
- use of lox for some of the cooling. This would remove the need for a separate lox autogenous heat exchanger
Any other ideas?
John
Well what can I say - I'm getting the redesign blues here. :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later? That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it. No room for complacency.
100,000 HP Whoa 😮I don't think that's the highest ever.. I was pretty sure the RS-25 was higher, but ... "The fuel pump alone delivers as much as 71,000 horsepower, the oxygen pump delivers about 23,000"[1]
https://twitter.com/elonmusk/status/1076684059827302400QuoteSpaceX metallurgy team developed SX500 superalloy for 12000 psi, hot oxygen-rich gas. It was hard. Almost any metal turns into a flare in those conditions.
https://twitter.com/elonmusk/status/1076686201061404672QuoteOur superalloy foundry is now almost fully operational. This allows rapid iteration on Raptor.
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most, but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.Well what can I say - I'm getting the redesign blues here. :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later? That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it. No room for complacency.
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most, but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.Well what can I say - I'm getting the redesign blues here. :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later? That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it. No room for complacency.
On the other hand if the breakthrough radically changes the internal layout and geometry at the heart of the engine as well as introducing new materials then they will need to start testing again to gain experience. The more radical the redesign the greater the loss of experience with that design. Still here's hoping, I guess it’s just the SpaceX way.
... Still here's hoping, I guess it’s just the SpaceX way.
True and somewhat reassuring, I hope your right.Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most, but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.Well what can I say - I'm getting the redesign blues here. :( Isuppose it doesn't matter as long as it does what it says on the tin in the end, but every major redesign announcement gives me that uncomfortable feeling that they are not fully in control of the situation. I hope I'm wrong and worrying needlessly but... :-\Is it really not plausible that an unforeseen breakthrough happened, that was worth implementing now rather than later? That there's no setback. Like Robotbeat said, we're going to Mars and basically anyone who qualifies to work at SpaceX and is a believer will be there pouring everything they have into it. No room for complacency.
On the other hand if the breakthrough radically changes the internal layout and geometry at the heart of the engine as well as introducing new materials then they will need to start testing again to gain experience. The more radical the redesign the greater the loss of experience with that design. Still here's hoping, I guess it’s just the SpaceX way.
They are well aware of any such risks.
I'm not sure if you were around back when SpaceX was upgrading the F9v1.0 to the v1.1. New engines, basically a brand new rocket. "They are changing too much!", was the constant complaint on this forum. Over and over. Some posters insisted that they needed to fly out the CRS contract with the v1.0 - anything else was just breach on contract and the sky was falling too.
But time has proved SpaceX right. It was the new design that allowed them to actually reach the performance and flight rate they originally hoped for, to reach the point now where they have caught up with their order backlog.
So I wouldn't worry. Assume they know what they are doing. And surely it is better to make drastic(?) changes before production has already started?
Sure it's plausible that they have made a breakthrough and I hope they have, I'm a bigger SpaceX amazing people than most...
...but they seem to be having so many breakthroughs, redesigns and changes that I'm a little nervous. With luck the breakthrough will turn out to be "just" a new super alloy to build the Raptor from and most of the experience gained in 1200 seconds of hot fire testing will still be somewhat relevant.
Well what can I say - I'm getting the redesign blues here. :(
I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.
You may well be right in which case excellent news! It’s just that scaling a third scale development engine to a full sized engine doesn't sound like a radical redesign, especially when IIRC it was stated that this was always the plan and would not be "that difficult".Well what can I say - I'm getting the redesign blues here. :(
I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.
This.
From what I hear from sources: The previous version of Raptor, which was extensively tested over the past two years was a DEVELOPMENT engine.
During the extensive testing campaign Tom's team learned many, many new things. They also found dozens upon dozens of items to improve on the next, closer-to-flight-design.
That next, closer-to-flight-design is what will be on the test-stand soon.
But the basic principle behind the engine is still the same: full flow, gas-gas methane-lox rocket engine. And yes: the improvements - as confirmed by Elon - are also designed to reach that magical 300 bar number.
This is what Agile rocket engine development is all about: build a minimal viable product (the initial Raptor we have seen) and test the hell out of it. Learn all you can and build-in all the improvements. Test the hell out of it. Learn some more and build an even better one. Than test the hell out of that as well.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because it takes the most amount of time to get to something that is ready to do the job.
Rocket Engines Are Hard (TM).
Don't expect the current version of Jeff's BE-4 to be the one that will be on the first New Glenn and Vulcan. It will be a thoroughly improved and enhanced BE-4 as well.
You may well be right in which case excellent news! It’s just that scaling a third scale development engine to a full sized engine doesn't sound like a radical redesign, especially when IIRC it was stated that this was always the plan and would not be "that difficult".
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.
John
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.
John
If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.
John
If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.
Are you talking about boost pumps?
Radical redesign, in my mind, implies some sort of configuration change. Transients might dictated changes to the pre-burners. Vibrations and combustion instability might have caused them to relocate the lox pump, etc... Just throwing things out there.
John
If they've moved up to multiple turbopumps for each fuel, I'm going to have to quit guessing what SpaceX is doing altogether.
Are you talking about boost pumps?
More than two, parallel turbopumps.
You may well be right in which case excellent news! It’s just that scaling a third scale development engine to a full sized engine doesn't sound like a radical redesign, especially when IIRC it was stated that this was always the plan and would not be "that difficult".
It's not just scaling. The development engine(s) we've seen had a very much not flight-like thrust to weight ratio and had a lot of optimization in that department yet to be done. There was some comment about that by Elon within the last year or so, but I'm too lazy to go dig it up.
>
I never, ever expected a flying water tower boilerplate as the engine proof of concept vehicle.
100,000 HP Whoa 😮I don't think that's the highest ever.. I was pretty sure the RS-25 was higher, but ... "The fuel pump alone delivers as much as 71,000 horsepower, the oxygen pump delivers about 23,000"[1]
1 - https://www.nasa.gov/missions/highlights/webcasts/shuttle/sts111/ssme-qa.html
100,000 HP Whoa 😮= 74.6MW. From this looks like Raptor thrust is back towards 2016 ITS design levels. Would allow Super Heavy engine no. to be reduced to 19.
You are right, 170MW but a bit OT here.100,000 HP Whoa 😮I don't think that's the highest ever.. I was pretty sure the RS-25 was higher, but ... "The fuel pump alone delivers as much as 71,000 horsepower, the oxygen pump delivers about 23,000"[1]
1 - https://www.nasa.gov/missions/highlights/webcasts/shuttle/sts111/ssme-qa.html
I would bet the RD-170 is the king in terms of Turbopump power since it has both the highest thrust and chamber pressure.
Are the nozzles of the Raptors built taking into account the need to fly engines forward when entering the atmosphere at the first or second space velocity? That is, will the Raptors be a kind of heat shield?
No, Raptors must be shielded during atmospheric entry. Although, maybe not ...
That's the plan- "Raptors must be shielded during reentry."
"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."
Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
That's the plan- "Raptors must be shielded during reentry."
"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."
Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.
That's the plan- "Raptors must be shielded during reentry."
"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."
Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.
But can they take the dynamic pressure?
That's one of MANY questions for the engineers, when he isnt busy with twitter.That's the plan- "Raptors must be shielded during reentry."
"No, Raptors must be shielded during atmospheric entry. Although, maybe not ..."
Um. I thought reentry was meant to be "body first, skydiver style", maybe that's what he's talking about. Some renderings show the engine bells not extending beyond a shroud at the bottom. Possible room for cargo there as well. The engines and bells designed for surface level operation, suboptimal for vacuum. I wonder if in the end they will tweak the nozzles a bit to improve vac ISP.
However, as he typed it, Elon remembered that the engine bells already have cooling channals. "Although mayby not..." gonna need to run some numbers with the engineers.
But can they take the dynamic pressure?
That's one of MANY questions for the engineers, when he isnt busy with twitter.
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.Quoting in full instead of linking, because twitter embeds can confuse which post is being linked by scrolling your browser view.
Well what can I say - I'm getting the redesign blues here. :(
I read "radically redesigned" merely to mean that they are going to fire the flight ("light and tight") version of the Raptor.
This.
From what I hear from sources: The previous version of Raptor, which was extensively tested over the past two years was a DEVELOPMENT engine.
During the extensive testing campaign Tom's team learned many, many new things. They also found dozens upon dozens of items to improve on the next, closer-to-flight-design.
That next, closer-to-flight-design is what will be on the test-stand soon.
But the basic principle behind the engine is still the same: full flow, gas-gas methane-lox rocket engine. And yes: the improvements - as confirmed by Elon - are also designed to reach that magical 300 bar number.
This is what Agile rocket engine development is all about: build a minimal viable product (the initial Raptor we have seen) and test the hell out of it. Learn all you can and build-in all the improvements. Test the hell out of it. Learn some more and build an even better one. Than test the hell out of that as well.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because it takes the most amount of time to get to something that is ready to do the job.
Rocket Engines Are Hard (TM).
Don't expect the current version of Jeff's BE-4 to be the one that will be on the first New Glenn and Vulcan. It will be a thoroughly improved and enhanced BE-4 as well.
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.Star Hopper will most likely be using the dev. version of Raptor not the new version that SpaceX plans to fire next month. Not possible to go from 1st test fire to flight qualification within 3 months.
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.
Is anyone else concerned like me that there still has not been a full scale Raptor fired and Elon is saying that BFR is going to fly in 2019? The way things are I don't see how they can qualify a full-scale raptor in 2019 and place it on a rocket and fly that rocket even if it is sub-orbital. I don't see how it could be done.Star Hopper will most likely be using the dev. version of Raptor not the new version that SpaceX plans to fire next month. Not possible to go from 1st test fire to flight qualification within 3 months.
Woa, there is a huge difference between firing a 2/3 scale engine and a full scale. Then there is a huge jump from firing an engine, learning something, then building a modified engine and testing that and so forth. Difference between a Merlin 1C and 1D was minor. Has anyone seen/heard a full scale Raptor firing?Full scale Raptor not fired yet. EM tweeted new Raptor which I assume to be full scale to be fired next month.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
This other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
Also when awarded, the contract was for an upperstage for Falcon 9 and Falcon Heavy. Later, SpaceX decided to use it on BFR/Spaceship only:
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>
That story link was 2016. SpaceX was working on the powerhead at Stemmis in 2014, so decisions were made well before then.
April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)
Which quotes an announcement of intent 6 months earlier, so Q4 2013.
Long before Air Force funding.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>
The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.
April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)
Which quotes an announcement of intent 6 months earlier, so Q4 2013.
Long before Air Force funding.
Edit: typo, clarity
Mars colonization vehicle/BFR was ALWAYS the goal. Raptor was simply the first part of that vehicle to be developed. It was always intended for the Mars vehicle.
But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. (snip)
Mars colonization vehicle/BFR was ALWAYS the goal. Raptor was simply the first part of that vehicle to be developed. It was always intended for the Mars vehicle.
But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. (snip)
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>
The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.
April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)
Which quotes an announcement of intent 6 months earlier, so Q4 2013.
Long before Air Force funding.
Edit: typo, clarity
But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. Raptor enabled BFR development, but BFR wasnt the cause of Raptor development. Air force funding and SpaceX matching investment turned the engine from a TRL 3 design into a TRL 7 functioning prototype, but the BFR system only crystallized as Raptor engine development accelerated, not starting at the same point.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
>
The first Air Force money was January 2016. SpaceX was working on the powerhead at Stennis in 2014, so decisions were made well before then.
April 2014... (http://spaceref.biz/company/spacex-set-to-test-raptor-engine-components-at-nasa-stennis.html)
Which quotes an announcement of intent 6 months earlier, so Q4 2013.
Long before Air Force funding.
Edit: typo, clarity
But that is exactly my point, the Raptor engine was around and well defined (and funded) long before it was the BFR engine. Raptor enabled BFR development, but BFR wasnt the cause of Raptor development.
>
The current Raptor concept “is a highly reusable methane staged-combustion engine that will power the next generation of SpaceX launch vehicles designed for the exploration and colonization of Mars,”
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?was merely 'cause
Because the USAF gave them $61.4 million to develop it.
Why the h*ll do you characters think that SpaceX began with development of the BFR/BFS engine first?
Because the USAF gave them $61.4 million to develop it:
https://spacenews.com/air-force-adds-more-than-40-million-to-spacex-engine-contract/
Also when awarded, the contract was for an upperstage for Falcon 9 and Falcon Heavy. Later, SpaceX decided to use it on BFR/Spaceship only:QuoteThis other transaction agreement requires shared cost investment with SpaceX for the development of a prototype of the Raptor engine for the upper stage of the Falcon 9 and Falcon Heavy launch vehicles.
https://dod.defense.gov/News/Contracts/Contract-View/Article/642983/
And a 6 year development cycle for such an engine is really fast!
New superalloy foundry could allow the next version of Raptor to be dev. in as little as one year. EM said it would allow for rapid iteration of Raptor.And a 6 year development cycle for such an engine is really fast!
In 1988 I had the great privilege of attending an "Introduction to Rocket Engines" lecture by a retired rocket engineer in Huntsville. He was German and worked on the V2 before being brought to the United States after the war to work on engines for the US. Really fascinating as he walked us around to several different engines in the museum and could talk of the issues and fixes they worked through.
I can still hear him saying in his German accent: "It takes 10 years to develop a new engine. People always want it faster but it takes 10 years."
You clearly have never heard of the Falcon X, Falcon X Heavy and Falcon XX concepts from 2010. Those preliminary design concepts were for SHLVs capable of putting humans on Mars.
And you have clearly not read my posts where I talked about Falcon X. I have been on this forum for over a decade, and heard the first rumblings about a SpaceX SHLV since 2010.
I consider the Falcon X to be almost a completely different design to the BFR architecture unveiled in 2016, and that was developed on the basis of Raptor.
I could keep on, but I think that this has gone far enough. I think there is a fundamental disagreement here on the difference between concepts and actual system development, and will leave it at that.
“I’ve been working on Mars for the last four years, so I’m not going to take any credit for the Block 5 engine and all the upgrades that have happened,” he said. “I’ll take credit for developing the team that developed the Merlin 1D engine.”
Mueller told GeekWire that he’s been mulling over the Raptor for about a decade. The engine doesn’t make use of the Merlin design, but goes instead with a full-flow, staged-combustion system that requires a clean-sheet design.
I can still hear him saying in his German accent: "It takes 10 years to develop a new engine. People always want it faster but it takes 10 years."
Merlin developement has been quite fast. We could assume Raptor will take about the same.
It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.But probably wouldn't be good enough for a Mars architecture -- not at a size that SpaceX can afford to build.
Do we have a solid, official explanation of why SpaceX is going for both LOX/methane and full-flow staged combustion without easier intermediate steps, such as gas-generator LOX-methane and/or fuel-rich staged combustion?
There have been rumours and a lot of speculation that LOX/methane is needed for reuse without a lot of maintenance, since kerolox suffers from coking problems. Full-flow staged combustion is of course intended to give maximum Isp and T/W.
But how necessary is LOX/methane for reusability and how important are staged combustion and full-flow staged combustion in particular for performance? Hasn't Musk said that current Merlins can be used up to ten times without a lot of maintenance? I'm getting the feeling that people are skeptical about this, but do we have any hard facts?
It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.
As for the full flow staged combustion cycle, I believe this is the most efficient cycle so an obvious choice for a billionaire who has hired some of the best smarts in the world and wants the highest performance engine for his Mars spaceship.
he answer might be that coking isn't as urgent a problem as people are suggesting, because the current Merlins can indeed already handle ~10 launches.
As for the full flow staged combustion cycle, I believe this is the most efficient cycle so an obvious choice for a billionaire who has hired some of the best smarts in the world and wants the highest performance engine for his Mars spaceship.
Sure, that must be the reason to want it eventually. But given that it's such a long journey to get there, why not do something simpler first? The answer might be that coking isn't as urgent a problem as people are suggesting, because the current Merlins can indeed already handle ~10 launches. On the other hand, there seems to be an urgent desire to get SH and BFS operational as soon as possible.
Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.
Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.
You're not answering my question. I didn't say full-flow staged combustion was impossible, just that it was hard and (for whatever reason) they appear to be in a hurry. We already know why LOX/methane was chosen and why full-flow staged combustion is awesome. The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.
Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious what that reason is.
The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry?
Methane was preferred because it doesn't coke and can be manufactured fairly easily on Mars unlike RP1. Good luck trying to Persuade Mr Musk to do something simpler first, he would want you to explain the physics of why it was not possible.
You're not answering my question. I didn't say full-flow staged combustion was impossible, just that it was hard and (for whatever reason) they appear to be in a hurry. We already know why LOX/methane was chosen and why full-flow staged combustion is awesome. The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.
Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious what that reason is.
The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry? Fine if you don't know or don't care, but simply repeating what we already know doesn't seem useful.It is unlikely any CH4 engine would ever be made “fuel-rich” as it would still be susceptible to some coking in the pre-burner. Blue Origin’s BE-4, for example, is oxidizer rich for this reason. The time to develop a ORSC engine is neither shorter nor less expensive. The FFSC engine offers slower and cooler running turbo pumps, ideal for reliability and reuse. In addition, they do not suffer from seal failure and introduction of hot, oxygen rich gas being injected into the fuel plumbing via the shared turbine axis.
Musk is of course under no obligation to work incrementally, but it does appear to be his MO in other areas and there's probably a rational reason for the path he has chosen. I'm curious to know what that reason is.
This idea that Musk only (or even mostly) only works incrementally has little or no basis in fact. Do you think that the BFR architecture is just an incremental step?
The question is why not start with gas-generator or fuel-rich staged combustion LOX/methane instead if you're in a hurry?
'Cause he hired, and then built over time, a super top-notch propulsion engineering team, and they thought they could build a methalox FFSC engine in about 7 years if adequately resourced. They obviously had that argument put forward by October 2012, when Musk announced the next-gen engine would not use Merlin's fuel, and confirmed it as methane in Nov 2012. Based on the testing of the Raptor dev engine(s) since first fire in Sept 2016, and it's now Dec 2018, I'd say they are right on target. Improvements to the dev Raptor test engine, ostensibly, the initial production engine, hits the test stand in early 2019. First flight using three (3) Raptors by 2Q2019 on the Starship BFH1.
This is a competent propulsion team!
Why methane? Because Mars! No oil on Mars to make RP-1 for the return flight. Musk explained this in detail in his IAC talk in Sept 2017.
Why FFSC? Because Efficiency! Higher ISP, which is needed in the long run for the sorts of objectives SpaceX has for themselves.
Why these choices, and not the choices you want? (or anybody else on NSF wants, or the US government might "want" if they were to have been asked?) Because private capital, and choices that can be made by actors in a world of economic incentives rather than that political incentives that drove all nation state/government space efforts for the past six+ decades.
This is not your father's, nor your government's, space technology development effort.
So why do we keep comparing what a private company like SpaceX can do to these legacy ways of thinking, and then being surprised by the results? ::)
Edit: fixed a typo
Why these choices, and not the choices you want? (or anybody else on NSF wants, or the US government might "want" if they were to have been asked?)
This is not your father's, nor your government's, space technology development effort. So why do we keep comparing what a private company like SpaceX can do to these legacy ways of thinking, and then being surprised by the results? ::)
No, my question was a different one, why ...
Do we have a solid, official explanation of why SpaceX is going for both LOX/methane and full-flow staged combustion without easier intermediate steps, such as gas-generator LOX-methane and/or fuel-rich staged combustion?
It would seem that a gas generator LOX/methane engine wouldn't have coking problems, could have excellent T/W, and Isp that is substantially above that of Merlin.
[...]
As such, if your goal is a FFSC engine, then building anything else is just a waste of time and money.
Methane and kerosene have very close theoretical max ISPs, and while methane's is indeed a couple percent higher,
Thanks, I didn't know that, I thought the Isp difference was much bigger.
Is it possible to get full flow without having both an oxygen-rich preburner for the oxygen turbopump and a fuel-rich one for the fuel? Or will Raptor only have oxygen-rich preburners as I believe RD-180 does?This very thread has the answers. The first few pages go into great detail. In short summary, using pure CH4 in the fuel rich pre-burner of a FFSC engine produces little coking because the pressures and temperatures are kept low enough. If one were to run methane or in Blue Origin’s case, LNG, through a FRSC cycle, it would run high and hot enough to produce coking. The greater mass of O2 also makes ORSC more desirable than FRSC. The SSME is FRSC because H2 does not coke and NASA didn’t believe it was possible to reuse an engine that had been exposed to high pressure, hot O2.
For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.Absolutely. The only advantages of CH4 out of Earth’s dense atmosphere and gravity well is its potential to be carbon neutral and its effect on reusability, including clean burning, spark ignition and autogenous tank pressurization.
Methalox needs staged combustion to beat kerolox on performance.
You can also use the super efficient engine cycles that hydrogen can use: FFSC and expander.For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.Absolutely. The only advantages of CH4 out of Earth’s dense atmosphere and gravity well is its potential to be carbon neutral and its effect on reusability, including clean burning, spark ignition and autogenous tank pressurization.
Methalox needs staged combustion to beat kerolox on performance.
Another advantage, at least for Boca Chica, is the ability to tap directly into nearby CNG feed line and refine their own methane. That's got to save a few bucks.For a booster engine, where density and thrust are paramount, kerosene actually has considerably better performance. Falcon 9 is already volume-limited (diameter by road transport and length by bending fineness) and would lose quite a bit of payload, on the order of 10%, if they tried to switch it from M1D to a GG methalox engine.Absolutely. The only advantages of CH4 out of Earth’s dense atmosphere and gravity well is its potential to be carbon neutral and its effect on reusability, including clean burning, spark ignition and autogenous tank pressurization.
Methalox needs staged combustion to beat kerolox on performance.
Upthread it said raptor has the same shaft for methane preburner and oxygen reburner. I think they have different shafts?
Also coking in the methane preburner:
What temperature does coking happen at?
I thought I remember reading that the methane preburner was relatively cool and could almost be just be an expander cycle. Which I assume to mean is that just vaporizing the methane is enough to provide the horsepower.
Do I have that right?
EDIT: Sounds like point 2 is already answered.
why not do something simpler first?
perhaps it would not in fact have gained them much timeAbsolutely certainly it would have lost them time. It's not a good question to ask.
This idea that Musk only (or even mostly) only works incrementally has little or no basis in fact. Do you think that the BFR architecture is just an incremental step?
No, you're right, the whole BFR project, not just Raptor, appears to be following a different philosophy. Tesla Model 3 too, come to think of it. And, as you say, we don't know the full story. I'm sure that if and when we ever hear it, it will make sense. It's just fun to speculate what the reasons might be.
I know it's fun to speculate here on NSF but you're overthinking this. The reason Spacex is following this route is because they can! Just that simple.
Upthread it said raptor has the same shaft for methane preburner and oxygen reburner. I think they have different shafts?
Also coking in the methane preburner:
What temperature does coking happen at?
I thought I remember reading that the methane preburner was relatively cool and could almost be just be an expander cycle. Which I assume to mean is that just vaporizing the methane is enough to provide the horsepower.
Do I have that right?
EDIT: Sounds like point 2 is already answered.
Pixel measuring the latest pictures of the testbed shows that the Raptor bell diameter is 1/6.89 the diameter of the test article as a whole... That works out to 1.3 meters if it is 9m in diameter.
Pixel measuring the latest pictures of the testbed shows that the Raptor bell diameter is 1/6.89 the diameter of the test article as a whole... That works out to 1.3 meters if it is 9m in diameter.
and remind me of what the different size estimates of the bell have been over the years?
Could be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.
John
Which helps underscore the point that SpaceX and propulsion head Tom Mueller have assembled a really top-notch team of rocket engine designers, technicians, testers, and the ops team that keeps it all running. They had very good reasons to go the way they chose to plan to go, and by all accounts, are achieving on their development milestones.They have suffered personnel losses to several other engine development teams (mostly NewSpace) ... significant ones. Enough to make those other teams quite good (that's a good thing, IMHO). And yet I suspect they are still the best.
Probably the best and most experienced rocket engine development team today globally, given their achievements and recent iterative and incremental real-world experience in many rocket engines with many propellant mixes.
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.No, the new radically redesigned Raptor will have it's 1st test fire next month. I think the Raptors on Star Hopper may be the 2017 spec. ones with 1.3m dia. nozzles which are likely the 1st revision from the original dev. Raptor. The new radically redesigned Raptor will most likely be revision 2 and likely have somewhat greater thrust than the 2017 spec. judging from EM's tweet about it's pump power.
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.No, the new radically redesigned Raptor will have it's 1st test fire next month.
>
Stennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.
Think those 3 Raptors are final scale, radically redesigned, and especially flight ready versions.No, the new radically redesigned Raptor will have it's 1st test fire next month.
>
Not trying to be too picky, but Musk didn't say 1st/first test fire,
Elon Musk ✔ @elonmusk
Yes. Radically redesigned Raptor ready to fire next month.
6:13 PM - Dec 22, 2018
which could mean the hopper with radically redesigned Raptors doing it's first hop. RR Raptor's tests could have been at Stennis.
https://www.nasa.gov/centers/stennis/about/stennis/index.htmlQuoteStennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.
--- snip ---
which could mean the hopper with radically redesigned Raptors doing it's first hop. RR Raptor's tests could have been at Stennis.
https://www.nasa.gov/centers/stennis/about/stennis/index.htmlQuoteStennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.
--- snip ---
which could mean the hopper with radically redesigned Raptors doing it's first hop. RR Raptor's tests could have been at Stennis.
https://www.nasa.gov/centers/stennis/about/stennis/index.htmlQuoteStennis also is partnered with companies such as Space Exploration Technologies Corp. (SpaceX) to test engines and engine components for their private space efforts.
Do you have a source for any testing by SpaceX at Stennis after 2015?
>
I know for a fact that Stennis would like to have SpaceX back; but I've never seen a source since very early 2016 that even discusses SpaceX and the US government rocket engine test facilities at Stennis.
Space Exploration Technologies Corp., Hawthorne, California, has been awarded a $40,766,512 modification (P00007) for the development of the Raptor rocket propulsion system prototype for the Evolved Expendable Launch Vehicle program. Work will be performed at NASA Stennis Space Center, Mississippi; Hawthorne, California; McGregor, Texas; and Los Angeles Air Force Base, California; and is expected to be complete by April 30, 2018.
>
Could be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.
John
It's not the wrong size if the purpose is deep throttle at SL, instead of vacuum optimization for which the dual bell was traditionally proposed.
@ CT SpaceGuyCould be a dual bell but if it is 1.3 m diameter, it's the wrong size. I suspect it is a quick mockup. I still think this build is a static prototype. Lots of reasons to build a static prototype. I would like to be wrong.
John
It's not the wrong size if the purpose is deep throttle at SL, instead of vacuum optimization for which the dual bell was traditionally proposed.
I get the ideal of a duel bell. Any notion of how effective this is.The space shuttle used a de Laval nozzle which helped with flow separation at sea level.
Has there ever been an operational duel bell used before?
Ken
I get the ideal of a duel bell. Any notion of how effective this is.The space shuttle used a de Laval nozzle which helped with flow separation at sea level.
Has there ever been an operational duel bell used before?
Ken
https://en.wikipedia.org/wiki/Space_Shuttle_main_engine (https://en.wikipedia.org/wiki/Space_Shuttle_main_engine)
Cross posting this:
What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.
John
Cross posting this:
What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.
John
Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.
Cross posting this:
What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.
John
Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.
Cross posting this:
What they might have done is to add a radiation cooled extension to the existing prototype engine bell. That could be done without messing with the cooling passages. They probably have at least three of those. This would allow them to better characterize the altitude at which the exhaust attaches to the extension.
John
Those nozzles have terrible sight lines for radiative cooling, especially the middle one, and even more so if they put that skirt back on the vehicle. And Raptor had always been planned, even in the highly expanded vacuum version, to have full regen nozzles. Musk confirmed this several times.
The production version will surely have regenerative cooled nozzle, but this may not be the production version or anything close to that. I had the same thought as livingjw when I first saw the nozzle images, it is as if they welded a MVac nozzle extension to the nozzle of the dev Raptor. This could be just a temporary test setup, radiation cooling should work with just 3 of them, similar to Delta-IV Heavy.
Except the claim is it is the outside bell (the extension) that is SL optimized.
So the inner bell would would be attached at above Earth SL pressure.
Which makes no sense, unless people are mis estimating the sizes of both and/or the operating pressure of the engine.
Except the claim is it is the outside bell (the extension) that is SL optimized.
So the inner bell would would be attached at above Earth SL pressure.
Which makes no sense, unless people are mis estimating the sizes of both and/or the operating pressure of the engine.
It makes a lot of sense, when you consider that the engine needs to be able to deep throttle, and expansion ratio is one of the hard limits on deep throttling. Assuming the numbers put forwards by others in this thread, it provides 1 to 50 expansion ratio when going full throttle, and yet allows throttling down to 15% or even less. That's something that a standard bell really couldn't do.
It also sort of explains how they intend to deal with some of the hard problems of steeped nozzles: The nozzle is never intended to spend any real time in the dangerous zone where it's overflowing the inner nozzle but not really fully filling up the outer. On startup at full power, the full nozzle is always in use. When returning, the 3 engines are spun up and fired at a throttle level that just barely fills out the inner one.
The sizes of the bells and the raptor's known characteristics dont line up with your claim of a vac-optimized outer bell. They DO line up with Tuna's claim of a SL optimised outer bell, deep throttle optimized inner bell.Except the claim is it is the outside bell (the extension) that is SL optimized.
So the inner bell would would be attached at above Earth SL pressure.
Which makes no sense, unless people are mis estimating the sizes of both and/or the operating pressure of the engine.
It makes a lot of sense, when you consider that the engine needs to be able to deep throttle, and expansion ratio is one of the hard limits on deep throttling. Assuming the numbers put forwards by others in this thread, it provides 1 to 50 expansion ratio when going full throttle, and yet allows throttling down to 15% or even less. That's something that a standard bell really couldn't do.
It also sort of explains how they intend to deal with some of the hard problems of steeped nozzles: The nozzle is never intended to spend any real time in the dangerous zone where it's overflowing the inner nozzle but not really fully filling up the outer. On startup at full power, the full nozzle is always in use. When returning, the 3 engines are spun up and fired at a throttle level that just barely fills out the inner one.
- The claim is that the inside bell is SL optimized, not the outer.
- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.
- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.
- In the case of the hopper, the lower bell can be radiation cooled, even with the some radiation blockage, because it would not be exposed to attached flow at low hopper altitudes.
John
- The claim is that the inside bell is SL optimized, not the outer.
- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.
- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.
- The claim is that the inside bell is SL optimized, not the outer.
I think that claim is wrong, and does not line up with published performance numbers.Quote- To our best ability to determine, the demo engine(s) had expansions around 25 and pressures around 2000 psi and had a diameter a little under a meter. Someone recently estimated the diameter of the bells on the hopper were about 1.3 m and looked like a dual bell. I just speculated that they might have added an uncooled dual bell like nozzle extension. The size would be about right.
For the latest performance numbers, 1.3m would be rather small for a vacuum nozzle, and since the engines are on the rocket apparently to perform fit tests, it would not make sense for the final articles to have larger ones.Quote- At sea level at full throttle the lower/outer bell flow is designed to be separated. That is the whole point of a dual bell nozzle. You get more thrust out of your engine at low altitudes when this flow is separated.
That is one of the things you can get by designing dual nozzles. However, it is not the only possible design goal.
Raptor engines differ from most other rocket engines by having an additional requirement: The ability to deep throttle down when right at sea level altitude. This raises from the need to support multi-engine-out vertical landing. (Engine startup is much slower than raising throttle on an already running engine, so it's necessary to be able to land on at least two less engines than are typically lit, and preferably still at much less than 100% throttle so you have reserve for the "oh crap" situation where you lose two engines and only realize it when their thrust falls to 0 and you need to very rapidly compensate.)
Traditionally designed engine nozzles cannot support this, unless they are very low expansion ratio and therefore very low performance. Using a dual nozzle where the outer nozzle is sea level optimized for full throttle and at maximum possible expansion ratio, and the inner nozzle is sea level optimized but for a very low power output is an elegant solution to their problem.
And if feel that the inner nozzle is sea level optimized and the outer one is vacuum optimized, how do you believe Raptor will achieve <20% throttle with the inner nozzle?
People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?Mish mash doesn't mean they concatenated two nozzles they had lying around the scrapyard.
I don’t expect such a nozzle on Raptor. But I could be wrong.
People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?Mish mash doesn't mean they concatenated two nozzles they had lying around the scrapyard.
I don’t expect such a nozzle on Raptor. But I could be wrong.
These nozzles came.from somewhere. They look dual-belled, and they look hi fidelity. Why would mock-up bells be so complex? Mock-up bells would be conical...
We can only see a tiny bit of it - can't tell if conical or not.People need to dial down their expectations about a double/triple/hybrid/whatever nozzle. There is no evidence for it. (Other than the mishmash of Raptor parts hanging from the BFH, which has other plausible explanations) Wishful thinking is not evidence - remember the mega vacuum deployable nozzle?Mish mash doesn't mean they concatenated two nozzles they had lying around the scrapyard.
I don’t expect such a nozzle on Raptor. But I could be wrong.
These nozzles came.from somewhere. They look dual-belled, and they look hi fidelity. Why would mock-up bells be so complex? Mock-up bells would be conical...
Indeed they would be - the upper part on these nozzles is exactly that - conical. The bottom part appears to be a real nozzle (or shaped like a real nozzle), but the upper part is the “fake bit” IMO. This part is what people see and think dual/hybrid. But it is strictly conical, and thus IMO just a conical cover over where real engine detail/plumbing will be.
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.
But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
What are the disadvantages of a dual-bell nozzle?
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.
But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.
The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?
And from the looks of it, that separation ring is naturally reinforcedSince they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?
The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.
The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.And from the looks of it, that separation ring is naturally reinforced
Because, irrespective of your angst over these specific nozzles, any location where the curvature of the generatrix of a bell is discontinuous will be naturally reinforced, as you'd like it to be if it is where you want the flow to separate.The problem with flow separation in a continuous nozzle is the separation point can jump around, which causes high stress in the nozzle. The "kink" at the transition between the small and large nozzles gives the flow a defined, constant place to separate. This results in much lower stress on the nozzle.And from the looks of it, that separation ring is naturally reinforced
Huh? Wha... where do you get that from?
These “engine” pictures really are the ultimate Rorschach test. People see whatever the heck they want. Double/triple/quintuple nozzles, natural reinforced points, you name it. Pretty soon someone will argue that this is proof of Raptor being an aerospike engine.
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.
But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.
The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
I wonder if this will give them another advantage:
On launch from an unprepared surface (Mars) you could potentially ignite all the engines at a low throttle setting and allow them a short time to stabilize before moving to full thrust.
What are the disadvantages of a dual-bell nozzle?
They are:
(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;
(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).
What are the disadvantages of a dual-bell nozzle?
They are:
(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;
(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).
I believe there’s one more:
(3) there’s a transient instability spike at a certain atmospheric pressure where it transitions from bell to the other. This could damage or destroy the bell. However this may not be an issue because during ascent the booster will transition this atmospheric pressure and it won’t (I assume) have these bells. The Starship will ignite in vacuum so avoids the issue. On descent there may be concern for the Starship, but careful engine shutdown and reignition selection could avoid this issue, as could throttling.
What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
And they only need 3 engines that work at sea level for landing. The rest can (and ultimately probably will) be optimised entirely for vacuum.
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
And they only need 3 engines that work at sea level for landing. The rest can (and ultimately probably will) be optimised entirely for vacuum.
No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.
The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.
Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?
I think these bells are a lot larger tho...Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?
Pretty sure 31 is readily doable based on past discussion. https://forum.nasaspaceflight.com/index.php?topic=41363.msg1729558#msg1729558
It is quite possibe to fit 31 1.3m wide engines on the bottom of a 9m wide booster. Outer ring of 12, middle ring of 12 and inner ring of 6 with 1 in the middle or example.Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
And they only need 3 engines that work at sea level for landing. The rest can (and ultimately probably will) be optimised entirely for vacuum.
No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.
The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.
Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
Agreed - however are we positive the three engines we’re seeing on the BFH are 1.3 meters in diameter? With the advent of a dual bell design are we sure the diameter hasn’t increased? Someone cut out one of the workers in an image with the engines showing and turn him sideways...It is quite possibe to fit 31 1.3m wide engines on the bottom of a 9m wide booster. Outer ring of 12, middle ring of 12 and inner ring of 6 with 1 in the middle or example.Side note - your “30-40 engines in the booster” statement made me try to visualize the reality of the. The BFH may not be full height, but it’s full diameter to the actual BFS/Starship, and therefore the BFB/SH. Personally, based on the three engines/mock-ups we see installed on the BFH, I can’t imagine fitting more than a dozen of those things in a 9 meter diameter circle. Certainly not 30-40, or am I off the mark?Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
And they only need 3 engines that work at sea level for landing. The rest can (and ultimately probably will) be optimised entirely for vacuum.
No. According to the “dear moon” presentation, all 7 engines are the same on Starship. And likely the same as the 30-40 engines in the booster. This is what really allows them to accelerate (or maintain) the aggressive schedule. Propulsion is almost always the long pole for any launch vehicle.
The first “real” Starship that has begun early fabrication in San Pedro will have 7 (plain) Raptors, to allow SSTO or near SSTO testing before the booster is ready. If the real prototype has enough performance for getting into orbit (with zero payload), they will probably try it.
Any vacuum Raptor modifications are possible down the line, but it is not part of the initial scope.
9m booster cross sectional area = 64 sqm 31 raptos at 1.3m cross sectional area 41 sqm. Of course whether this is a good idea or not is another matter. I assume it is as they tend to be quite a smart bunch.
Agreed - however are we positive the three engines we’re seeing on the BFH are 1.3 meters in diameter?
Thanks - looks pretty spot on.Agreed - however are we positive the three engines we’re seeing on the BFH are 1.3 meters in diameter?
Including some spare pixel "padding", I've got 6.8 engines within 9m for 1.32m
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Ok, seems I am the only one who doesnt know how that works on a dual nozzle. Suppose the engine fires at sea level with low thrust, so that the inner nozzle is optimal but the outer nozzle is too large. Then, there would still be flow separation on the outer nozzle. How comes that the outer nozzle doesnt shake itself appart in this case?
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.
But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.
The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
Also, renders of the booster show that it widens to ~10 meters near the base.
Given how many things have changed, I wouldn't rule the vacuum raptor out just yet.
It was a temporary compromise even as presented... So with the switch to metal and change in time-table - who knows.
>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)
What are the disadvantages of a dual-bell nozzle?
They are:
(1) slight loss of performance since the nozzle contour isn't optimized for a single flow condition;
(2) risk of flow tripping irregularities between a multiple engine configuration (such as BFR).
I believe there’s one more:
(3) there’s a transient instability spike at a certain atmospheric pressure where it transitions from bell to the other. This could damage or destroy the bell. However this may not be an issue because during ascent the booster will transition this atmospheric pressure and it won’t (I assume) have these bells. The Starship will ignite in vacuum so avoids the issue. On descent there may be concern for the Starship, but careful engine shutdown and reignition selection could avoid this issue, as could throttling.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)
A few months ago Musk tweeted something about 42 engines as a BFR upgrade.
Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.
Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for Starship.
A smaller nozzle generates more thrust at SL because it can produce a higher exhaust velocity than an overexpanded nozzle.
But this doesn't matter for Raptor since the flow will not detach from either a 50:1 or a 19:1 nozzle at SL and full throttle. Throttling to the point where the flow detaches from the larger nozzle will cost more thrust than is gained from the higher exhaust velocity of the small nozzle.
The reason they want a small inner nozzle is not higher thrust (or efficiency) at deep throttle when at sea level, it's to make the engine survive. The reason you can't use vacuum nozzles at sea level is not that they are inefficient, it's that when the exhaust jet separates from the nozzle before the end, the whole system will typically catastrophically shake, to the point where it will tear itself apart.
Since they want to deep throttle at sea level, and have an efficient nozzle, they have two very conflicting design constraints. A dual bell nozzle solves this for them, giving them both the nozzle that their engine can use at deep throttle, and the second, outer one that gives them a high expansion ratio.
I wonder if this will give them another advantage:
On launch from an unprepared surface (Mars) you could potentially ignite all the engines at a low throttle setting and allow them a short time to stabilize before moving to full thrust.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)
A few months ago Musk tweeted something about 42 engines as a BFR upgrade.
Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.
Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for Starship.
It is an SSFO.What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
That's reasonable.I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)
A few months ago Musk tweeted something about 42 engines as a BFR upgrade.
Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.
Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for Starship.
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.Higher chamber pressure equals lower thrust?
If you hold pump power constant.I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.Higher chamber pressure equals lower thrust?
If you hold pump power constant.I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.Higher chamber pressure equals lower thrust?
EM is looking at 25MPa Pc for Raptor at IOC then gradually raising it to at least 30MPa later. If the new Raptor produces 3.25MN SL thrust at 25MPa Pc, it will produce around 3.9MN SL thrust at 30MPa Pc assuming everything else is constant.If you hold pump power constant.I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.Higher chamber pressure equals lower thrust?
But to vary chamber pressure during development, assuming you're not modifying the chamber geometry - wouldn't you be changing pressure by varying pump power?
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa. Dual bell design could allow throttling below 20% allowing engine count on SS to be reduced to 5 while still giving engine out capability on landing. EM also said that 30MPa will take some time so I am betting on the TP power he tweeted being for 25MPa. Also 19 engines gives the perfect fit under the base which 31 does not and can do away with the tapered bottom skirt. I am also betting that future larger dia. SH/SS systems will keep the same engine count as the 9m dia. one with the superalloy foundry allowing Raptor to be scaled up to larger sizes through multiple iterations. Hopefully EM will clarify everything including a major Raptor update when he does his full technical overview of the system after Star Hopper flies.>
We might see a shorter larger diameter higher engine count superheavy announced during the next year (and yes I know that Elon tweeted that outline shape would stay about the same with material change)
A few months ago Musk tweeted something about 42 engines as a BFR upgrade.
Then SpaceX images showed the 9m core having a ~10m engine skirt, which ISTM makes moving to 10m much easier.
Which in turn begs if SH will stage higher/faster, allowing smaller tanks/lower mass for Starship.
This ^^^
Long felt that 31 engines even with the add on mass of the 10m skirt was a design patch not an enhancement. Whether it's 19, 21 or whatever, I also look for raptor thrust upscaling.
What's interesting about this is let's say it's 19 engines. Even if initial Super Heavy's have interim not quite full thrust raptors, the full stack could easily loft ~60+ tonnes to LEO. Launches of Starlink could commence with gradual payload upratings, just like with the Merlin family. Earlier useful flight starts and experience gained while learning and enhancing.
Schedule acceleration.
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
Why would the booster have dual bell Raptors?that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile? This transition zone with associated flow separation would have to be avoided, correct? It would seem that even if it was quick, the rapid transitions of all of the many engines would cause serious vibration. So what is the solution? The only one I could think if is throttling down significantly, which I understand is already done late in the booster burn on an F9 to avoid over-acceleration with an almost empty rocket.
I was assuming (as others have?) that there's now a single raptor design for both the booster and the starship. I can see how the dual bell would work for the Starship but it would make more sense to me if there was a sea level bell on the booster version ...I think perhaps there is a common engine design but the bells are different - kind of like the Merlin 1D SL and Vac (although I think this is an oversimplification as I believe the M1D Vac has other optimizations / modifications from the sea level version).
I think perhaps there is a common engine design but the bells are different - kind of like the Merlin 1D SL and Vac (although I think this is an oversimplification as I believe the M1D Vac has other optimizations / modifications from the sea level version).
that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile?
Why would the booster have dual bell Raptors?
So you're saying the booster, which doesn't need a vac optimized engine (as apposed to the Starship), needs to have a dual bell nozzle, rather than just the smaller diameter one? I'm not questioning your statement, rather trying to understand for myself. It would seem that unless absolutely required that the inherent risks of a dual bell over a single diameter design would dictate that the booster has a sea level optimized bell (like the M1D SL) and the Starship would have the dual bell (analogous to the F9 Upper Stage using the M1D VAC).that means the problems are bypassed for this application, but the question was what the problems we're in a more general sense.What are the disadvantages of a dual-bell nozzle?From some of the explanations I've seen (not all consistent) there's a transition at altitude when the lower part of the flow expands to the walls of bottom part of the bell. It seems like there could be some uncertainties there.
Just a note about requirements.
They have to work at maximum efficiency at vacuum AND they have to ignite, deep throttle and work at decent efficiency at sea level for landing burn. No other altitudes and no transitions, this is not a SSTO.
So, if I'm not mistaken, wouldn't the super heavy booster with the dual bell raptors start at sea level on the inner nozzle but, just after max Q, it would transition to operating in near vacuum conditions which would cause the flow to try to expand to the larger or second nozzle profile?
This depends on the exact chamber pressure and expansion ratios, but in all likelihood, no. Any Raptor at full chamber pressure is going to have attached flow all the way out any expansion ratio up to 75:1 or thereabouts.
The smaller inner bell, (if that's what it is), is only ever relevant at SL AND at deep throttle. It has to be at BOTH conditions, or flow won't detach from the larger nozzle. The small bell (if that's what it is) plays no part at all in ascent, which is all at either wide-open throttle or reduced pressure.Why would the booster have dual bell Raptors?
Because it has to deep throttle to land at SL on Earth. That's the only flight regime where a 15 or 20 ER nozzle on Raptor plays any role at all. In all other flight regimes the flow will be fully attached out to an ER of 40 or greater.
So you're saying the booster, which doesn't need a vac optimized engine (as apposed to the Starship), needs to have a dual bell nozzle, rather than just the smaller diameter one? I'm not questioning your statement, rather trying to understand for myself. It would seem that unless absolutely required that the inherent risks of a dual bell over a single diameter design would dictate that the booster has a sea level optimized bell (like the M1D SL) and the Starship would have the dual bell (analogous to the F9 Upper Stage using the M1D VAC).
I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.I don't think so. 2000kN thrust at 330s Isp gives 617kg/s propelleant flow, with 930kg/m³ densified 3.8:1 methalox mix that is 0.66m³/s, and with 80MPa outlet pressure get 53MW ideal pumping power. But as I recall rocket turbopumps sacrifice a lot of efficiency for weight savings, and are generally only in 70-80% range. With 70% pump efficiency would have about 74MW shaft power. That is close enough to quoted figure that I don't think we are looking at a massive change in thrust.
That's likely what SpaceX is doing. We, on the other hand, are estimating Raptor thrust based on fixed numbers (like turbopump power) tweeted by Elon that are probably still being tweaked...But to vary chamber pressure during development, assuming you're not modifying the chamber geometry - wouldn't you be changing pressure by varying pump power?If you hold pump power constant.I am betting on a lower engine count 9m dia. SH with 19 engines following EM's tweet of 74.6MW(100,000HP) TP power for the new Raptor. TP power gives about 2.75MN SL thrust for Pc of 30MPa or about 3.25MN for Pc of 25MPa.Higher chamber pressure equals lower thrust?
So you're saying the booster, which doesn't need a vac optimized engine (as apposed to the Starship), needs to have a dual bell nozzle, rather than just the smaller diameter one? I'm not questioning your statement, rather trying to understand for myself. It would seem that unless absolutely required that the inherent risks of a dual bell over a single diameter design would dictate that the booster has a sea level optimized bell (like the M1D SL) and the Starship would have the dual bell (analogous to the F9 Upper Stage using the M1D VAC).
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.
SpaceX don't yet have a vacuum test stand do they? Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.
SpaceX don't yet have a vacuum test stand do they? Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.
You are assuming that SpaceX hasn't already sufficiently tested dual bell engines and retired the risks to the point where they consider them completely safe. Also, assuming that the outer bell has a 50:1 ratio, the exhaust would be fully attached at sea level and at full throttle, and would remain attached for the entire ascending trajectory of the booster, effectively rendering the bell into just a fancy single nozzle for that part of flight, with no additional risks. The only difference is that unlike an efficient sea level nozzle, during the landing burn they can run at a low enough throttle that they can survive multiple engine failures.
SpaceX don't yet have a vacuum test stand do they? Without that it is pretty hard to do a full assessment of a dual bell or over-expanded nozzle that may experience destructive transient nozzle flow re-attachment side forces during ascent.
Thus a major rationale for The hopper to fly ASAP so they can get on with 1st generation engine pre-production
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.At that point, it's less "deep throttling" and more "prelighting" backup engines in case there is an engine failure during descent. By having a couple spare engines idling at >5%, they can be rapidly throttled up if needed, more easilly than doing a fresh turbopump ignition>Main chamber ignition>full power landing.
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.
You could fire 5 engines at max thrust, then cut out 4 of them. But you better have a really good handle on the shutdown transients, and have a high confidence that the single engine won't die on you at the worst possible moment.
1) This is NOT a vacuum nozzle, and doen't need a vac stand for testing.
2) There is no detachment nor re-attachment during ascent, for a 1.3 meter nozzle on Raptor at full throttle. This nozzle will have fully attached flow for all flight regimes except terminal landing. This is assured by the amount of mass Raptor has to flow.
The only time the nozzle flow will be detached from the wall is during startup (see SSME ringing) and terminal landing. Both can be tested at sea level.
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.
You still need throttling. Because while turning off engines can be reasonably predictable, turning them on is not. And so if you ever hope to increase thrust during a landing, all the necessary engines need to be already running.
With 42 engines or even 31, how much throttling do you need to land on one engine (or even 3)? I am thinking none. Dual bell is cool but I don't think it's needful for SuperHeavy.
You don't need deep throttling, but you do need to rapidly go from ~6 times the vehicles weight in thrust down to 1.5 times or less. Anything else leads to longer landing burns and wastes a lot of heavy fuel that you just accelerated to Mach 8 and back.
You could fire 5 engines at max thrust, then cut out 4 of them. But you better have a really good handle on the shutdown transients, and have a high confidence that the single engine won't die on you at the worst possible moment.
How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?
I suspected that too but then I have more questions:How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?
I think they are hydrodynamic bearings which use the cryo propellants.
John
Pics of the bells and the bottom of a fin. Taken on 1-8-19 during the fit test excitement.
Using BocaChicaGal's photo, focusing on left 2 Raptors: Enhanced to show additional Raptor components through gap in the leg.
Note** that the left bell and expander section are mismatched in O.D. ... whereas center is much smoother transition.
Peeking around leg, we can see part of the thrust chamber and lack of plumbing. This confirms EM's statement that thee were not flight ready units but development test articles being used as boiler plate fit check and alignment units...
Thanks for the great work BCG/Nomadd/Austin... and many others..
I suspected that too but then I have more questions:How are Raptors turbopumps bearings lubricated?
What's the cold start process regarding that lubrication?
I think they are hydrodynamic bearings which use the cryo propellants.
John
1. How do they start? Hydrodynamic bearing has to have some static pressure untill it it transits to hydrodinamic mode.
2. Cryo proppelants are fluid, but also low pressure. How's the sealing to between the bearing and hot, high pressure side of pump? How do cryo fluids in bearings don't start to boil?
Does anyone have any rocket turbopump cross section? I dissassembled large turbines with additional pumps for hydrostatic or hydrodynamic bearings but how does this work without additional pump?
Latest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:Hydrodynamic bearings have rubbing until the speed is high enough, thus they have a lot of wear with start/stop cycles. Using a hydrostatic bearing pump means you can eliminate the rubbing before start-up, eliminating that source of wear and tear from start/stop cycles, increasing the cycle life.
https://twitter.com/blueorigin/status/950365085091811330QuoteLatest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:Hydrostatic and hydrodynamic bearings are different.
https://twitter.com/blueorigin/status/950365085091811330QuoteLatest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
At that point, it's less "deep throttling" and more "prelighting" backup engines in case there is an engine failure during descent. By having a couple spare engines idling at >5%, they can be rapidly throttled up if needed, more easilly than doing a fresh turbopump ignition>Main chamber ignition>full power landing.My recollection (which admittedly could be wrong) was that Elon's justification for landing on 3 engines at low throttle was to be able to land on 2 at higher throttle, if there was an engine issue. I therefore, wouldn't expect any 'idling' engines.
Yes, plenty of work has been done in that area. The best primer is probably the paper that kicked off the whole hydrostatic bearing discussion (this is what you mean by "some high pressure gas"), "Reddecliff and Vohr, 1969". A lot of work has been done since then. The French had a hydrostatic bearing test stand for the never-completed European staged combustion engine (TPX/TPTech). Pratt & Whitney wanted to replace the pump side bearing in one of the Space Shuttle turbopumps with a hydrostatic one, but nothing came of this either.
IHI, the Japanese rocket turbopump manufacturer, has released information about a turbopump that they want to sell overseas that uses hydrostatic bearings, so they are probably the furthest along. Unless of course SpaceX uses hydrostatic, but for all we know their bearings run on fairy dust.
Superconducting magnetic bearings have been studied as well.
Here are the three largest problems for these bearings:
The stiffness is tiny compared to rolling element bearings. Hydrostatic bearings have at most a tenth the stiffness, magnetic bearings maybe only a hundredth or a thousandth. This means that rotordynamic instability is a huge development risk, since it can't be well predicted.
They require an additional supply of propellant, which lowers the efficiency of the turbopump and increases the complexity.
Rolling element bearings have made huge strides, so expendable and even some reusable rockets have really no need. Silicon nitride bearing balls together with induction-harded Cronidur steel races have made some other components in the Space Shuttle look really bad in terms of life time.
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:
https://twitter.com/blueorigin/status/950365085091811330QuoteLatest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
Snip...I suspect it probably does. Although this is probably moderated by the fact that Blue have not reached orbit yet.
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
SpaceX set out to build a cheap reliable Mars transportation system. The size of the engines was determined by the engineering trades made in the (apparently ongoing) design process. If they were dumb enough to get into a pissing match over engine size, they would have the biggest engine powering a more expensive or less capable or incapable spacecraft.
Matthew
Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor? Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?
Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor? Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?My recollection is that modeling showed that using more smaller engines counterintuitively resulted in a better T/W even including all of the plumbing.
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:Hydrostatic and hydrodynamic bearings are different.
https://twitter.com/blueorigin/status/950365085091811330QuoteLatest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
https://en.m.wikipedia.org/wiki/Fluid_bearing
I thought that Raptor uses hydrodynamic ones because I couldn't see the piping for hydrostatic bearings in that Raptor render from IAC.
A really high stakes way of testing a nozzle, with more limited instrumentation potential. In particular long term vac nozzle reliability will be hard to prove out without an altitude test facility.Are there any facilities other than Arnold Engineering Development Complex that can altitude test a vacuum expansion nozzle on a Raptor? Was their ~2.4MN thrust limit a deciding factor in sizing the Raptor?Yes. The other facility will economical soon, the entirety of outer space.
EM tweeted that Raptor's TP power will be 74.6MW (100,000HP) so the new radically redesigned Raptor may have more thrust than BE-4. We will not know until EM announces the specs. for the new version of Raptor. I doubt EM bothers about what size engines BO develop.Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
BE-4's thrust has nothing to do with Raptor. If SpaceX cared about having a bigger engine they'd build one.
Is there any informed opinion that Raptor has hydrodynamic bearings, or is this entirely speculation? We know BE-4 has them, though Blue called them hydrostatic bearings:Hydrostatic and hydrodynamic bearings are different.QuoteLatest BE-4 engine test footage where we exceeded our Isp targets. We continue to exercise the deep throttling of our full scale 550,000 lbf BE-4, the reusability of our hydrostatic pump bearings and our stable start/stop cycles. More to follow from ongoing tests. #BE4 #NewGlenn
https://en.m.wikipedia.org/wiki/Fluid_bearing
I thought that Raptor uses hydrodynamic ones because I couldn't see the piping for hydrostatic bearings in that Raptor render from IAC.
Hydrostatic have less wear so better for multiple starts the downside is lower performance. For RLV the lower performance is worthwhile trade for longer engine live.
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!
EM tweeted that Raptor's TP power will be 74.6MW (100,000HP) so the new radically redesigned Raptor may have more thrust than BE-4. We will not know until EM announces the specs. for the new version of Raptor. I doubt EM bothers about what size engines BO develop.Wonder if it irks Elon that Bezos's methalox BE-4 has more thrust than Raptor?
BE-4's thrust has nothing to do with Raptor. If SpaceX cared about having a bigger engine they'd build one.
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!
Well to be fair, a dual nozzle is still a possibility. (and I say that as someone who very much doubts it) It's just that the supporting evidence for it has withered away.
Don't worry, speculation will now move from OMG dual bell! to OMG retractable nozzle!Calm down, the real flight Raptors may well just have conventional nozzles after all. The loosely fitted nozzle extensions on the placeholders gave the visual appearance of dual bells. We should find out within the next couple of months when the flight Raptors get installed. The nozzle extensions on the placeholders may have only been held by friction.
*ghasp* and all the pixel counting and numbers crunching w.r.t. low thrust capable dual nozzles and its implication was a misinterpretation of ambiguous things we see in images? No way!
Well to be fair, a dual nozzle is still a possibility. (and I say that as someone who very much doubts it) It's just that the supporting evidence for it has withered away.
Still a possibility, my question being "Why are these placeholder engines equipped with a nozzle skirt that isn't even attached?"
Photoshoot optics of the Hopper for how it will *approximately look with the upcoming production engines?
Protection of the placeholder engines from the elements? (Why protect an engine that most likely will not be flown on this test article? And how would they be more protective than the usual blue coverings used on Merlins?)
Some type of test of a bell extension for gritty Boca Chica conditions?
UFO DefenceTM?
Lacking a vacuum testing facility could a Raptor vacuum expansion nozzle be adequately tested for mechanical durability and cooling adequacy using a partial conical base filling plug held inside the nozzle? The (ablatively covered or transpiration cooled?) cone would deflect flow outwards to hug the nozzle walls without separation that would otherwise occur due to over-expansion.
Is fuel rich the norm in rocket engines?
Is fuel rich the norm in rocket engines?
It is.
Is fuel rich the norm in rocket engines?
It is.
At least in the west. Russia invested heavily in oxygen-resistant alloys so they have some fairly advanced engines running oxygen rich.
Correct me if I am wrong but I think the reason for "all" rocket engines being fuel rich is that they run cooler that way and minimal impact to the isp.
Correct me if I am wrong but I think the reason for "all" rocket engines being fuel rich is that they run cooler that way and minimal impact to the isp.
Yeah, and in fact fuel-rich typically has slightly better Isp than stoichiometric, in addition to being cooler. Also, fuel-rich is more compatible with ordinary alloys. At least, that's my understanding.
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.
Someone may correct me, but I understand it as this:
Water has a worse molecular mass (18) than methane (16). So I would assume that water does not increase ISP as much as methane. Also, you never get a stochiometric combustion anyway because the combustion process is not perfectly uniform within the engine. As a result, you would get oxygen rich local environments in the exhaust. I can imagine that this would potentially damage the engine bell. By burning fuel rich, the non-uniformity of combustion will always (or say to much much higher likelihood) stay fuel rich, which is non-corrosive.
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.
Someone may correct me, but I understand it as this:
Water has a worse molecular mass (18) than methane (16). So I would assume that water does not increase ISP as much as methane. Also, you never get a stochiometric combustion anyway because the combustion process is not perfectly uniform within the engine. As a result, you would get oxygen rich local environments in the exhaust. I can imagine that this would potentially damage the engine bell. By burning fuel rich, the non-uniformity of combustion will always (or say to much much higher likelihood) stay fuel rich, which is non-corrosive.
Yep, anything containing oxygen will likely be your heaviest combustion product, that's why you're running fuel rich in the first place. You can do that up to the limit where going richer would mean exhausting too much solid C which reduces efficiency while not increasing isp. High efficiency tripropellant concepts typically use hydrogen as the working fluid for that reason.
I know it is OT here but it fits into the discussion. Would it be possible to do a tri-propellant engine, burning LOX and methane at stochiometric ratio for maximum chemical energy and add water as a third medium to limit temperatures and increase ISP? Especially on Mars produced propellant it might maximise the use of energy put into production.If instead of water you inject more fuel and oxidiser, you get the Thrust Augmented Nozzle (https://selenianboondocks.com/2007/11/thrust-augmented-nozzles/). As far as I am aware, no examples have been flown.
Teslarati: SpaceX sends “radically redesigned” Starship engine to Texas for hot-fire tests (https://www.teslarati.com/spacex-radically-redesigned-starship-engine-shipped-texas-hot-fire-testing/)
Most notable was an obvious secondary preburner/turbopump stack and the lack of any exhaust port, whereas M1D relies on a single turbopump and exhausts the gases used to power it.So does Raptor no longer have the oxygen tubopump integrated above the MCC?
Teslarati: SpaceX sends “radically redesigned” Starship engine to Texas for hot-fire tests (https://www.teslarati.com/spacex-radically-redesigned-starship-engine-shipped-texas-hot-fire-testing/)QuoteMost notable was an obvious secondary preburner/turbopump stack and the lack of any exhaust port, whereas M1D relies on a single turbopump and exhausts the gases used to power it.So does Raptor no longer have the oxygen tubopump integrated above the MCC?
EDIT: So I misunderstood the question. I believe the design I saw had oxy in line and integrated with the combustion chamber and the fuel turbopump off to the side. I remember someplace a picture of the engine and annotations labelling everything. I think livingjw did it.That picture of Raptor is in the article.
I believe this (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1700053#msg1700053) is the post with the labelled version mentioned above.
Also this flow diagram (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1826715#msg1826715) might be helpful in visualizing how it works.
I believe this (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1700053#msg1700053) is the post with the labelled version mentioned above.
Also this flow diagram (https://forum.nasaspaceflight.com/index.php?topic=41363.msg1826715#msg1826715) might be helpful in visualizing how it works.
The one difference from that flow diagram is that I expect is for the fuel line for regenerative cooling to start from after the first pump stage and return to the second stage. Given their insane chamber pressures, pushing the coolant through the entire cooling circuit at full pressure would require making the structure of the rocket a lot heavier than it needs to be.
That rocket engine is a piece of art. 8)
Finally we can take the crown from the Russians!
Am I the only one who can't tell where that engine is in relationship to the people in the picture?Nope.
Am I the only one who can't tell where that engine is in relationship to the people in the picture?
Ok, here is an eyeball comparison assuming the height of the people is about the same. I do not guarantee the accuracy.
So what type of nozzle is that? An inverse duel bell? It looks reversed compared to a duel bell nozzle.
Full flow staged combustion cycle still?
Yes
Wow, cool! 8)
One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.
Wow, cool! 8)
One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.
This is the first production engine, according to Musk's December tweet. I wonder how much scope they have for simplification, since all the lessons from Merlin's will be incorporated?
Wow, that engine looks chunky compared to Merlin. Since thrust is 200t, maybe 250t later I doubt TWR will increase much over Merlin (~100t), but it might be equal.
“The Merlin holds the thrust-to-weight record for now,” he said. “But the Raptor’s coming.”
Wow, that engine looks chunky compared to Merlin. Since thrust is 200t, maybe 250t later I doubt TWR will increase much over Merlin (~100t), but it might be equal.
If the nozzle is 1.3m, here is a rough size comparison using pixel measuring between Raptor, Merlin, and a 1.78m man.Nice drawing! Thanks for the comparison.
(https://i.imgur.com/osNnhWQ.jpg)
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing! (edit not so big after all)
...
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
I'm pretty sure BE-4 has a 1.8m nozzle
That rocket engine is a piece of art. 8)
Finally we can take the crown from the Russians!
The turbo machinery is arranged so it all fits above the footprint of the bell, probably so they can cram 31 of these things under the "Super Heavy" booster.
I'm curious what the large metal piece with the slots in the side at the back is though.
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
I'm pretty sure BE-4 has a 1.8m nozzle
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
I'm pretty sure BE-4 has a 1.8m nozzle
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
I'm pretty sure BE-4 has a 1.8m nozzle
Using ImageJ (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.
So the initial common 200 tons Raptor is going to be less powerful than BE-4, while the sea-level optimized ~250 tons Raptor (so about 2,452 kN) should be a tiny more powerful than BE-4 (2,400 kN).Yes that is bit insane... Look at the size of that thing!
...
Could you add SSME?
Using ImageJ (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.
That's a manlet.
Using ImageJ (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.
That's a manlet.
Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).
Using ImageJ (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.
That's a manlet.
Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).
No but a 2.4m nozzle does not match up with the BE-4 reverse engineering thread that I linked above. Those calculations get you something in the neighborhood of 1.8m. I think there is a bit less uncertainty in that than in the man shaped outline somebody stuck on that picture.
Using ImageJ (https://imagej.nih.gov/ij/) with the image from Blue's website and assuming 180cm for the human outline in the image I get: 247 cm.
That's a manlet.
Wikipedia (https://en.wikipedia.org/wiki/List_of_average_human_height_worldwide) lists the average height of a US male as 175.7 cm. 180cm is slightly above average, so not a manlet (https://www.urbandictionary.com/define.php?term=Manlet).
No but a 2.4m nozzle does not match up with the BE-4 reverse engineering thread that I linked above. Those calculations get you something in the neighborhood of 1.8m. I think there is a bit less uncertainty in that than in the man shaped outline somebody stuck on that picture.
Man, SpaceX played a mean game of Tetris with that plumbing. It is much more compact then I was expecting considering the combustion cycle type.
The thrust per cubic meter is for fun.
it is the volume calculated from the diameter of the nozzle by the height.
of note I calculate the nozzle of the BE4 closer to 1.9m than 1.8m based on Blue's 6' scale.
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)
Fuel line leading from the fuel turbopump to the regen cooling system.
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)
Fuel line leading from the fuel turbopump to the regen cooling system.
- I think that might be coming from a fuel valve downstream of the pump. Hard to tell, but I believe the turbo-pump and pre-burner is on the backside and not shown on either photo.
- black coating might a little insulation, or not.
John
What is the black coating on the piping coming out on the right? (Bottom of the hanging Raptor)
Fuel line leading from the fuel turbopump to the regen cooling system.
- I think that might be coming from a fuel valve downstream of the pump. Hard to tell, but I believe the turbo-pump and pre-burner is on the backside and not shown on either photo.
- black coating might a little insulation, or not.
John
Could be. Looks to me like the fuel pump is mounted transversely now, and that line is coming done off the end of it. Most of the pump and all of the preburner would still be hidden behind the engine.
is it actually both turbo pumps stacked on each other on top? or inside each other?
look at the ropes. they hold engine close to the center. if there was huge massive machinery it will be below engine that photo
It hits my eye as if I'm looking at a V12 engine sat vertically on a display stand.
Man, SpaceX played a mean game of Tetris with that plumbing. It is much more compact then I was expecting considering the combustion cycle type.
If they are going to stick so many of them close to each other on the booster, the compact volume is a necessity.
Wow, cool! 8)
One note about all the piping around it. This is still a development engine, the final version will presumably be simplified further.
This is the first production engine, according to Musk's December tweet. I wonder how much scope they have for simplification, since all the lessons from Merlin's will be incorporated?
While true, the first engines on a test stand would always have extra instrumentation and margins. Compare some of the first M1Ds on the test stand (image 1) to a recent production version (image 2). I think they already have plans to simplify it. But this first version may be good enough to mount on the Hopper.
There is precedent for this - Grasshopper also used a pre-production M1D (similar to picture 1), before they optimized it for the engine clustering in F9v1.1.
I think the nozzle outline does appear a little odd, though it may be nothing. The shape of the nozzle might be due to trying to limit separation at part throttle at sea level similar to this discussion taken from "Nozzle Design"Here's the link for that Rao (https://en.wikipedia.org/wiki/G._V._R._Rao) paper from 1958 "Exhaust Nozzle Contour for Optimum Thrust":
by R.A. O'Leary and J. E. Beck, Spring 1992
....
The first choice for an SSME nozzle contour would obviously be one that maximized
nozzle thrust; that is, a Rao optimum contour. ...
I think the nozzle outline does appear a little odd, though it may be nothing. The shape of the nozzle might be due to trying to limit separation at part throttle at sea level similar to this discussion taken from "Nozzle Design"Here's the link for that Rao (https://en.wikipedia.org/wiki/G._V._R._Rao) paper from 1958 "Exhaust Nozzle Contour for Optimum Thrust":
by R.A. O'Leary and J. E. Beck, Spring 1992
....
The first choice for an SSME nozzle contour would obviously be one that maximized
nozzle thrust; that is, a Rao optimum contour. ...
https://doi.org/10.2514/8.7324
(I don't know if it's available anywhere for free.)
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?
Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?
I predict even though Raptor is more compact and higher thrust, it will do better at this, because this is not their first reusable engine. Yes, Blue has some experience from New Shepard, to be sure. But not quite the same, with dozens of reused Merlin engines flown now.Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?
That’s one of the things that will be fascinating to find out. Having said that, if you look at projected flight rates it definitely appears that SpaceX is looking to reuse Raptors both more times and more frequently.
I predict even though Raptor is more compact and higher thrust, it will do better at this, because this is not their first reusable engine. Yes, Blue has some experience from New Shepard, to be sure. But not quite the same, with dozens of reused Merlin engines flown now.Overall its amazingly compact for producing close to the same thrust as BE4. I wonder how they will compare on reusability?
That’s one of the things that will be fascinating to find out. Having said that, if you look at projected flight rates it definitely appears that SpaceX is looking to reuse Raptors both more times and more frequently.
Lovely diagram. Where's the heat exchanger to gasify the methane for tank pressurant? Or is it going to gasify on its own when introduced since it is hot (from the engine main combustion and nozzle heat exchange) and high pressure liquid which when released to a lower pressure environment, gasifies?
At @SpaceX Texas with engineering team getting ready to fire new Raptor rocket engine
Is that a pen in his hand? Are the team all signing the nozzle perhaps? I am s ure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.
Is that a pen in his hand? Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.
Is that a pen in his hand? Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.
Aren't there strict sound limits at McGregor this time of night? I wouldn't expect this to light up until daylight.
Aren't there strict sound limits at McGregor this time of night? I wouldn't expect this to light up until daylight.
I don't think they're supposed to test after 2200 Local. Betting they'll try for Sunday so it'll make the Monday news cycle.
Aren't there strict sound limits at McGregor this time of night? I wouldn't expect this to light up until daylight.
I don't think they're supposed to test after 2200 Local. Betting they'll try for Sunday so it'll make the Monday news cycle.
SpaceX most certainly does not schedule test fires based on hitting news cycles. Musk flew in tonight, so they're almost certainly aiming for a static fire tonight (nominally).
Is that a pen in his hand? Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.
...Yes, it is part of the test stand, engine (with sensitive parts) is front of it:
Note the hexagonal plate behind the engine, that protects the sensitive parts of the engine. Could this be something added just for this engine test cell.... or, is this the load bearing plate for the engine hexagonal bay? The equivalent of the top plate in this M1D image:
Cross posting this Raptor schematic from the autogenous pressurization thread. Better place for it. Started with the SSME schematic and changed it to reflect Raptor as I currently understand it.Shouldn't the O2 for CH4 precombustion be taken after the GO2 heat exchanger and not after the LO2 pump?
John
Lovely diagram. Where's the heat exchanger to gasify the methane for tank pressurant? Or is it going to gasify on its own when introduced since it is hot (from the engine main combustion and nozzle heat exchange) and high pressure liquid which when released to a lower pressure environment, gasifies?
Yes, methane is hot after cooling MCC. Will gasify during pressure drop.
John
IIUC, only the CH4 has a heat exchanger (the nozzle jacket)Cross posting this Raptor schematic from the autogenous pressurization thread. Better place for it. Started with the SSME schematic and changed it to reflect Raptor as I currently understand it.Shouldn't the O2 for CH4 precombustion be taken after the GO2 heat exchanger and not after the LO2 pump?
John
O2 would be gasseous when injected to CH4 precombustion chamber which is better for combustion.
That's flashlight for you Yanks, not a pen.Is that a pen in his hand? Are the team all signing the nozzle perhaps? I am sure it's exciting as hell, but I do hope they are not doing critical fitting work with a very expensive engine using a crew that has been at it for perhaps 16 hours or more at 2:15 in the morning.
Elon is holding a torch light. You can see the light shining in the bell.
...IIUC, only the CH4 has a heat exchanger (the nozzle jacket)How they then pressurize GO2 tank, electric heater? With Webasto?
...
https://twitter.com/elonmusk/status/1091958352513425408QuoteAt @SpaceX Texas with engineering team getting ready to fire new Raptor rocket engine
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.
But you can try to estimate the size by this palm print here:
; P
The pictures above show the nozzle to be smooth on the inside and smooth on the outside. How happens cooling channels within that? Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough). Does that imply that its 3D printed? Not liquid cooled? Other?
The pictures above show the nozzle to be smooth on the inside and smooth on the outside. How happens cooling channels within that? Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough). Does that imply that its 3D printed? Not liquid cooled? Other?
The pictures above show the nozzle to be smooth on the inside and smooth on the outside. How happens cooling channels within that? Obviously not braised tubes but I don't know what constructions are left (not because its an earth shattering mystery, only because I personally don't know enough). Does that imply that its 3D printed? Not liquid cooled? Other?
The same way it is done for Merlin 1D (see image)... with internal channels. There is basically three layers... the inside, middle, and outside. The middle copper(?) or other alloy layer is milled to have lots of small channels.
We have a (brief) test fire!!! Only caught the audio @Erdayastronaut @johnkrausphotos
First firing of Starship Raptor flight engine! So proud of great work by @SpaceX team!!
https://twitter.com/elonmusk/status/1092268892339273730A) You quick!QuoteFirst firing of Starship Raptor flight engine! So proud of great work by @SpaceX team!!
What can we glean from the exhaust plume in relation to the nozzle?
And is the green from the copper in the cooling channels? Was there some burn-through to the inner layers?
Either way it was way cool!
https://twitter.com/elonmusk/status/1092272377889738753Judging by the response of the observers, I'd say yes.
Was that supposed to happen...
Wow - that was some bright ignition! What do you suppose that vapor cloud was behind the bell just before ignition. If you step thought the video I thought it was Going to be a RUD...
Anyone else wondering about the greenish color in the flame at the end of the burn? Curious how well their advanced materials are holding up.Engine-rich combustion? Copper?
Saw the same thing with the development engine and IIRC the consensus was it's an oversaturated CCD in the camera.People keep abusing the word "consensus."
https://twitter.com/elonmusk/status/1092280273599979520?s=09Dan Guisinger
Gaseous CH4/O2 & heavy duty spark plugs. Basically, a 💨 of insane power 😀
Saw the same thing with the development engine and IIRC the consensus was it's an oversaturated CCD in the camera.People keep abusing the word "consensus."
Consensus means that basically everyone agreed.
There DEFINITELY wasn't consensus there.
It's green.
BTW, I find it odd that Elon doesn't seem to know whether or not it was engine-rich combustion or saturation (and this doesn't look like saturation, FWIW).
Either way, maybe we should worry about it if engines prove short-lived because of it?
Would Elon post it so quickly if the result wasn’t what they expected and hoped for? C’mon.Sure, because he's Elon Musk and he can do that.
With regards to the green:I agree. Folk's should just refer to Musk's tweet. And if it was actually chamber erosion that's not hard to find or to fix really, adjusting the burn/mix ratio can fix things like that.
Over-analysis is running wild here (again).
BTW, I find it odd that Elon doesn't seem to know whether or not it was engine-rich combustion or saturation (and this doesn't look like saturation, FWIW).
There is no defined graph or table or exact timeline on how long it takes to develop a new rocket engine from scratch. Aside from nuclear reactor design, nuclear submarine design, supersonic and hyper-sonic aircraft design, and sub-sea infrastructure design, this is basically the hardest thing to create and build from scratch in the human world right now. It takes as long as it takes depending on a wide variety of factors including but not limited to LV requirements first, cash available second, materials third, and breakage in testing fourth. With that said certain kinds of breakage are worse than others, burning up your combustion chamber liner during startup is definitely not good, but it's a test article. Also, it is often *semi easy* to fix problems like this in early test engines by changing things like mixture ratio or making other small tweaks. Not at all a show stopper.If I can’t concern troll myself, can someone else do it please? Is that what you are asking for?
No. Please re-read the actual words I wrote. I think I was very explicit about what I was not interested in. I wasn't terribly helpful about what I was interested in, though, so let me correct that.
I am interested in the process of bringing a rocket engine up. What things typically go wrong, what things typically go right, how long does it take to run through a test program, what things are you looking at and looking for, etc. Creating a new rocket engine isn't the sort of thing that's done every day, so it's interesting. I'm interested in precisely the sort of thing one might expect someone that reads "the evolution of the spacex raptor engine" to be interested in. So yes, I'm interested in "green", but I'm not interested in "oh gods, everything is coming apart at the seams."
I hope that's really clear now.
If not, well, don't worry, this is the last I'll write about it.
\Engine probably still cooling down+people need to sleep. They can come back in the morning and stick a camera in there to find out.I was actually surprised they ran this at all on Superbowl Sunday (in TX?!)
You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's.
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.
But you can try to estimate the size by this palm print here:
; P
Perfect! Is the hand what you used to set your scale in your Merlin/BE4 image?
So during the burn video there are 2 green flame episodes, one near the beginning, one at shutdown. Since these are not actually at startup, and supposedly Raptor uses spark ignition anyway, the only explanation I can think of is a bit of copper chamber or bell vaporizing.
The first frame of the early incident at 5:44 you just see a little streak by the bell, the next frame it's partway down the jet, then it's at the end of the jet. For some reason this frame is doubled. Then the next frame, no more green. However the vapor patterns along the ground don't suggest that any video is missing, they change from frame to frame in a consistent way.
Also, is it possible to tell the exhaust velocity from the spacing of the mach diamonds? They are extremely consistent right until shutdown starts around 6:20. You can see from one frame to the next that they are sliding to the right at that point.
Under certain mix conditions methane in combustion can produce a greenish tint or a green flash. Depends on temperature pressure % O2 in the burn and also other atmospheric gases interfering. The wide field video with sound shows much less green than the close in shot. Still good possibility exists it could also be due to internal erosion of some kind in the chamber. Really easy way to find out stick a camera or scope up there after it's cooled down and see if there is damage.I am sure that the throat diameter is measured to a um after each firing, plus microscopy shots to examine the microstructure, plus any number of techniques to ascertain surface composition... It's so easy these days, and those are such obvious questions...
If no visible damage continue test campaign see if engine RUD's itself after multiple firings. Reminder the last SSME test saw a large hydrogen flame shooting out of part of the power-head or test stand and that was not really a massive show stopper, without the government shutdown test campaign would probably still be proceeding. Also engine did not undergo RUD despite failure. Chamber erosion is very different however, but again, if no visible damage is found with camera/scope then just keep firing and see what happens.
Easy way to find the flaws in a piece of machinery is push the test article until it breaks and then see where it broke and why. Like with the windblown starhopper fairing these things are test articles and are meant to be broken. Do not be surprised if they explode break fall over or otherwise fail.
What you don't want is to be on a flight to the moon and discover a flaw the hard way then, as Apollo 13 did. Or worse during launch or re-entry.
What made you decide to not use TEA-TEB for ignition? Is the idea that during engie chill, CH4/O2 are released to ignite for liftoff? I don't think I've ever heard of spark plugs being used for rockets.
Spark plugs ignite dual blow torches that ignite preburners & main chamber
With that said certain kinds of breakage are worse than others, burning up your combustion chamber liner during startup is definitely not good, but it's a test article. Also, it is often *semi easy* to fix problems like this in early test engines by changing things like mixture ratio or making other small tweaks. Not at all a show stopper.
You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's. This motor fired and did not explode or destroy itself on the first run, and its being developed at rapid pace, with many rapid changes, and probably not much cash. Doing pretty good so far, as far as anyone knows the sub-scale raptor never had an RUD during last year's campaign and this variant didn't on the first test either.
With that said certain kinds of breakage are worse than others, burning up your combustion chamber liner during startup is definitely not good, but it's a test article. Also, it is often *semi easy* to fix problems like this in early test engines by changing things like mixture ratio or making other small tweaks. Not at all a show stopper.
You should really take a look at how hard development of the Saturn V F1 and SSME engines was. Many many RUD's. This motor fired and did not explode or destroy itself on the first run, and its being developed at rapid pace, with many rapid changes, and probably not much cash. Doing pretty good so far, as far as anyone knows the sub-scale raptor never had an RUD during last year's campaign and this variant didn't on the first test either.
That's why I don't think this is a massive issue. This is the first firing of a new engine, not a final design they have built a hundred of already. Making slight changes to the mixture or bell isn't the worst thing that could have happened and the hopper might be fine to run as is or with an added ablative liner.
Quote from: @psalman03What made you decide to not use TEA-TEB for ignition? Is the idea that during engie chill, CH4/O2 are released to ignite for liftoff? I don't think I've ever heard of spark plugs being used for rockets.
https://twitter.com/elonmusk/status/1092320321229643776QuoteSpark plugs ignite dual blow torches that ignite preburners & main chamber
The green is a reflection of the envy that other rocket engines felt right that very moment.In my HIGHLY amateur opinion, I though the green exhaust looked cool. 8)
Under certain mix conditions methane in combustion can produce a greenish tint or a green flash. Depends on temperature pressure % O2 in the burn and also other atmospheric gases interfering. The wide field video with sound shows much less green than the close in shot. Still good possibility exists it could also be due to internal erosion of some kind in the chamber. Really easy way to find out stick a camera or scope up there after it's cooled down and see if there is damage.
If no visible damage continue test campaign see if engine RUD's itself after multiple firings. Reminder the last SSME test saw a large hydrogen flame shooting out of part of the power-head or test stand and that was not really a massive show stopper, without the government shutdown test campaign would probably still be proceeding. Also engine did not undergo RUD despite failure. Chamber erosion is very different however, but again, if no visible damage is found with camera/scope then just keep firing and see what happens.
Easy way to find the flaws in a piece of machinery is push the test article until it breaks and then see where it broke and why. Like with the windblown starhopper fairing these things are test articles and are meant to be broken. Do not be surprised if they explode break fall over or otherwise fail.
What you don't want is to be on a flight to the moon and discover a flaw the hard way then, as Apollo 13 did. Or worse during launch or re-entry.
Due unknown distance between Mr. Musk and the nozzle, pixel counting is meaningless. Also according to the metadata; picture is taken with iPhone, e.g. with wide-angle lens which making the perspective error even worst.
But you can try to estimate the size by this palm print here:
; P
Perfect! Is the hand what you used to set your scale in your Merlin/BE4 image?
Isn't that bit rich, coming from a *rebellious colony which still uses dead king's feet for measurements. Anyway here more of The True Size of Things (https://forum.nasaspaceflight.com/index.php?action=dlattach;topic=42585.0;attach=1542450;image):
* : P
This is just an example of this exact discussion starting up after previous video releases earlier in the Raptor development process. People probably should relax.And, that engine ran for over a thousand seconds of burn time during the test campaign, if I recall correctly. So A) we've seen this before and B) the engine ran for a really long time. And it is literally the first firing of the first production engine. What we saw was a great success. It doesn't mean it's ready to fly tomorrow.
hit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.
John
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.To assist folks finding John's schematic, here's a link to the post:
John
...Here's the scale references:
No sarcasm from me, I was thrilled at the ingenuity.
I'm happy to see our scale references so far holding up.
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.
John
John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?
Quotehit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power
What does "power" mean in this context?
I'll go out on a bit of a hopeful limb here until someone that knows corrects me back to reality. If we're at 60% of max thrust and 60% of Pch then wow, it would be;
Pressure @ 100% = 283 bar, 4104 psi
Thrust @ 100% = 193.2 mt, 426,000 lbf
Don't run with these numbers, they are more of a question than a statement and are likely to be stricken from the record when someone with real knowledge sets me straight.
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.
John
John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?
? The Raptor has never been shown with boost pumps.
John
An engine that can be transported in a cargo van that produces 7GW, enough power to run Switzerland, with all that power passing through a hole the size of a toilet seat. Rocket engineering is best engineering. :)
I changed source of methane to pump outlet for higher pressure to Lox pre-burner. Probably could use either source, but pressure after cooling MCC is only 10 -15 bar higher than lox pump pressure and and I would like to have more like more like 30 ish. Schematic is on page 79 of this thread.
John
John what do you think is driving the boost pumps? The outlet of the main pumps? Separate preburners? One on the shaft of the main one running off of outlets of the main?
? The Raptor has never been shown with boost pumps.
John
So you are saying it's a two stage pump. The first stage would be akin to a "boost pump" in my books.
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.
Also... 350/280 (bar) = 250/200 (tonnes force)
;)
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.
Also... 350/280 (bar) = 250/200 (tonnes force)
;)
I am more inclined to think the current rating of 200-ish tonnes is 100% rated at 250 bar. That would make the 250-ish tonne version 100% rated around 310 bar. Obviously a small increase in throat size would be needed which would reduce the expansion ratio a little. I am getting ready to remodel the Raptor with the latest data.
John
Does anyone know what very loud bark is at the end of the run?
Not that it's relevant but it sounds like the barking dog chemistry experiment.Does anyone know what very loud bark is at the end of the run?Not unusual. It may happen with a rapid decrease in pressure be it the engine or feed lines.
Quotehit 170 bar and ~116 metric tons of force – the highest thrust ever from a SpaceX engine and Raptor was at ~60% power
What does "power" mean in this context?
I'll go out on a bit of a hopeful limb here until someone that knows corrects me back to reality. If we're at 60% of max thrust and 60% of Pch then wow, it would be;
Pressure @ 100% = 283 bar, 4104 psi
Thrust @ 100% = 193.2 mt, 426,000 lbf
Don't run with these numbers, they are more of a question than a statement and are likely to be stricken from the record when someone with real knowledge sets me straight.
Looking at the numbers more and more, an increase in chamber pressure from 250 bar to 280 bar for the initial Raptor engine seems to make sense. It explains pretty much all the increase in thrust from the 1.7MN version while keeping the engine bell the same size and expansion ratio.
Also... 350/280 (bar) = 250/200 (tonnes force)
;)
I am more inclined to think the current rating of 200-ish tonnes is 100% rated at 250 bar. That would make the 250-ish tonne version 100% rated around 310 bar. Obviously a small increase in throat size would be needed which would reduce the expansion ratio a little. I am getting ready to remodel the Raptor with the latest data.
John
Fuel rich exhaust is yellow, sufficient or oxygen rich exhaust is blue. Mix that amount of them and you get a green flash at the switch from one condition to the other. In my humble opinion !
And you just can't mix them; colors depend on the chemical mix of the flame. The electrons "jump" from their ground state to a higher energy level in high temperature. As they return to their ground state they emit visible light. There needs to something that emits green...
Because we are talking about light emission, yellow plus blue makes white. In the pigment model, yellow plus blue makes green. Since green is a primary light color, it is correct that to get green something must be emitting in the green band. http://learn.leighcotnoir.com/artspeak/elements-color/primary-colors/And you just can't mix them; colors depend on the chemical mix of the flame. The electrons "jump" from their ground state to a higher energy level in high temperature. As they return to their ground state they emit visible light. There needs to something that emits green...
What about different portions of the plume having different chemical mixes? is it reasonable to have a "skin" of yellow fuel rich combustion surrounding a core of blue combustion, at least during startup and shut down?
Because we are talking about light emission, yellow plus blue makes white. In the pigment model, yellow plus blue makes green. Since green is a primary light color, it is correct that to get green something must be emitting in the green band. http://learn.leighcotnoir.com/artspeak/elements-color/primary-colors/
Because we are talking about light emission, yellow plus blue makes white. In the pigment model, yellow plus blue makes green. Since green is a primary light color, it is correct that to get green something must be emitting in the green band. http://learn.leighcotnoir.com/artspeak/elements-color/primary-colors/
Don't tell the RGB screen you are looking at that emitted colors don't mix. You will lose all those wonderful yellow, oranges, purples and whites. /snark
Poll: When will full-scale hot-fire testing of Raptor begin?
You can (almost) stop the poll now.
Does anyone know what very loud bark is at the end of the run?
Does anyone know what very loud bark is at the end of the run?
If you are referring to the video in Musk's tweet, the audio and video are not synced because it was filmed from some distance away. The loud noise just before shutdown is actually the ignition.
This tweet chain made me look like :o
https://twitter.com/bluemoondance74/status/1092615576210550786
But there is no way that was a Raptor burn already, or is it? Must be a Merlin test?
Does anyone know what very loud bark is at the end of the run?Anyone that has worked with pulse jets especially large ones will recognize that noise.
This tweet chain made me look like :o
https://twitter.com/bluemoondance74/status/1092615576210550786
But there is no way that was a Raptor burn already, or is it? Must be a Merlin test?
Does anyone know what very loud bark is at the end of the run?Anyone that has worked with pulse jets especially large ones will recognize that noise.
The result of "ambient pressure" ignition of residual fuel in a "pipe". On this engine, as it was shutting down "rich" at some point the "fresh" air enters the nozzle and combustion chamber till the mixture leans to the point where rapid combustion occurs and induces an resonant oscillation in the "pipe" (not unlike the "bark" of the hydrogen in the test tube)
livingjw's schematic as vector graphic:This is really great, and it's reminded me of a question I've had about the oxygen preburner since the first reveal of the Raptor: what goes on in a torus-shaped combustion chamber? What's the injector have to be like?
In regards to the green emissions seen in raptor test vids, is it possible that the first burn of the rocket does indeed "burn off" some initial residue, or a first layer of copper, or something to that effect, as part of a completely normal process? If it's a residue left from manufacture, it'd be burnt off and not present later. If it's a "first layer" of copper, this could be bad, but could it also be normal - as in, after heating the alloy in this specific way, it changes the structure or something somewhat, making it more resistant to this kind of torture in the future? It seemed a pretty bright green and highly localized to be merely one of the emission lines of methalox combustion, but I could obviously be wrong about that.
I know nothing about metallurgy, so please be kind to my dumb speculations ;)
In any case I'm not trying to amplify any sense or worry or anxiety, as I'm fairly confident these guys know what they're doing. I'm just curious! I want to know more about how rockets and stuff work!
The design of the oxygen turbopump on the Raptor actually reminds me of the early centrifugal compressor type jet engines, with the compressor, combustion chamber, and turbine all in a compact line like that. Those had individual compressor cans, but some modern jet engines have toroidal combustion chambers instead, and I suspect that those have some design similarity to the Raptor's oxygen preburner.livingjw's schematic as vector graphic:This is really great, and it's reminded me of a question I've had about the oxygen preburner since the first reveal of the Raptor: what goes on in a torus-shaped combustion chamber? What's the injector have to be like?
The design of the oxygen turbopump on the Raptor actually reminds me of the early centrifugal compressor type jet engines, with the compressor, combustion chamber, and turbine all in a compact line like that. Those had individual compressor cans, but some modern jet engines have toroidal combustion chambers instead, and I suspect that those have some design similarity to the Raptor's oxygen preburner.livingjw's schematic as vector graphic:This is really great, and it's reminded me of a question I've had about the oxygen preburner since the first reveal of the Raptor: what goes on in a torus-shaped combustion chamber? What's the injector have to be like?
In regards to the green emissions seen in raptor test vids,I'm not understanding how copper is exposed unless there is a significant erosion of another material. Isn't it true that the copper is clad on both sides with another material? Perhaps nickel and cobalt if I understand right? Or maybe I got that material wrong but looking into the nozzle its certainly not copper colored inside and [?]that material extends through the throat and into the combustion chamber[/?]
In any case I'm not trying to amplify any sense or worry or anxiety, as I'm fairly confident these guys know what they're doing. I'm just curious! I want to know more about how rockets and stuff work!
Raptor just achieved power level needed for Starship & Super Heavy
https://twitter.com/elonmusk/status/1093423297130156033QuoteRaptor just achieved power level needed for Starship & Super Heavy
Design requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
Did you folks track down the source of the green hue?
Vaporized some copper
250 bar pressure already? I vaguely remember that 300 was the target for future iterations of the engine and I got the impression that even reaching 250 will be a challenge. Things are moving pretty quickly.Your member name says it all...
https://twitter.com/elonmusk/status/1093424663269523456QuoteDesign requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
..
(curious as to the duration of this test)
Isn’t therefore the energy density higher? Which in turn leads to higher chamber pressure?https://twitter.com/elonmusk/status/1093424663269523456QuoteDesign requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?
Also, if chamber pressure grows linearly with that 10-20% thrust increase, this particular Raptor would experience chamber pressures of at least ~280 bar (and perhaps more than 300 bar) at full thrust. In mid-December, he suggested that it would "take [SpaceX] time to work up to 300 bar...that is a mad level".
https://twitter.com/elonmusk/status/1076618077301665793
https://twitter.com/elonmusk/status/1093424663269523456QuoteDesign requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?
Also, if chamber pressure grows linearly with that 10-20% thrust increase, this particular Raptor would experience chamber pressures of at least ~280 bar (and perhaps more than 300 bar) at full thrust. In mid-December, he suggested that it would "take [SpaceX] time to work up to 300 bar...that is a mad level".
https://twitter.com/elonmusk/status/1076618077301665793
https://twitter.com/elonmusk/status/1093424663269523456QuoteDesign requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.https://twitter.com/elonmusk/status/1093423297130156033QuoteRaptor just achieved power level needed for Starship & Super Heavy
And not a hint of green on it :) wow.. that was soon! Look at the flame surrounding the central column of exhaust with mach diamonds. What does this, film cooling of the nozzle with methane?
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.
Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.
Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
Technically speaking, the whole Starship will be film cooledLooks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
https://twitter.com/elonmusk/status/1093424663269523456QuoteDesign requires at least 170 metric tons of force. Engine reached 172 mT & 257 bar chamber pressure with warm propellant, which means 10% to 20% more with deep cryo.
Perhaps I'm just out of the loop but is anyone able to fully explain why colder propellant would translate into 10-20% greater thrust? I'm familiar with subcooling for density and improved mass ratios but not with cooler prop = more thrust. Perhaps cooling-related?
I assumed it's related to mass flow. The same machinery can transport 10-20% more propellant if it's denser. Your point about chamber pressure might work the opposite way - they would theoretically reach 280 bar from increased pump capability alone but might have to dial it back because the chamber can't take it.
I heard the test from 30 miles away!
Maybe a little longer. My guess is under 10 seconds.
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
A little googling shows that film cooling is ubiquitous for high chamber pressure engines on top of the regenerative cooling. Doesn't cost that much isp when using a low molecular weight high specific heat fuels like (in particular) hydrogen or methane as the film coolants.
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!
The gist, assuming thrust is roughly proportional to chamber pressure:
~60% power = 116t thrust, 170 bar
~90% power = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!
The gist, assuming thrust is roughly proportional to chamber pressure:
~60% power = 116t thrust, 170 bar
~90% power = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar
I suggested earlier ~280bar for initial Raptor and ~350bar for full thrust version.
Honestly, given everything we know so far the math works out the best with those numbers, and they so far seem to fit better and better as we learn more... But they seem crazy. :o
.@SpaceX employees appear to be working over-time tonight @ the McGregor, TX facility, as the burst and brief, low rumble of an engine test was heard @ 10:51 pm CST. Note: There are Noise Ordinances in neighboring towns. I checked, and 11pm is the cut-off. Just in-time! 🔥👍
(Amendment: The most recent Ordinance I found for McGregor was a few yrs old: 9 pm limit, except for prior-approved cases, w/ 115-decibel max. Surrounding towns have later cut-off times, much lower decibel limits. *McGregor is the only town which can enforce an Ord. on SpaceX.)🧐
Only a few seconds
Musk has also conveniently created an invaluable second point of data for Raptor performance combined with SpaceX's Instagram update!
The gist, assuming thrust is roughly proportional to chamber pressure:
~60% power = 116t thrust, 170 bar
~90% power = 172t thrust, 257 bar
~100% power = 193t thrust, 285 bar
Curious what the informed opinion is on next test steps. They have apparently reached their required thrust targets. Do they continue increasing thrust to see what they can get to or now begin increasing test duration at target thrust levels to simulate full flight profile?
My model to date (work in progress):
- Keeping the 2017 Raptor (er=35, Aexit=1.3m) dimensions and increased the chamber pressure from 250 to 270 bar. This takes the engine from 1.7 sl MN / 1.83 vac MN to 1.83 / 1.96 MN (200 tonnes vacuum).
pressure, thrust
----------------------
- 170 bar, 113 mt SL
- 257 bar, 177 mt SL
- 270 bar, 183 mt SL 200 mt vacuum
John
>
Does anyone think they will achieve 300+ bars?
>
Does anyone think they will achieve 300+ bars?
I've lost count of the "SpaceX will/could never..." proclamations which ended up as roadkill.
So, what's up with the weird-looking exhaust plume? It's like there's a detached sheath of methane flowing out of the nozzle, igniting and creating a secondary, much slower plume. I wonder whether this image was taken during steady operation or ignition/shutdown transient.Your question about Raptor's exhaust plume was answered in previous posts. They increased the film cooling with methane to avoid overheating the thrust chamber allowing them to test at full thrust with current thrust chamber design. They will learn from this and make the next iteration of Raptor with redesigned cooling channels to reduce the amount of film cooling required.
I can't notice a similar detached plume developing at any point in any the previously released video.
Crazy thought: Is there any reason whatsoever you'd ever want to drill some small holes at the nozzle bottom on a test engine and release some of the regen cooling methane out into the void? I can't think of one, but this almost looks like that to me. ???
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.
I found https://www.sciencedirect.com/science/article/pii/S2212540X1830004X which talks about ISP reduction, but without figures for methane. It also discusses coking problems using RP-1 as the coolant, which is applicable to Merlin. That article links to https://arc.aiaa.org/doi/pdf/10.2514/6.2010-6721, but I sadly don't have institutional access. If someone does and indicates it's worth it, I'll be happy to create an account (:sigh:) and purchase it. I can't even find the price without an account apparently though.
[the really interesting pointer appears to be to https://iris.unicampania.it/handle/11591/172095#.XFxBW89KiGR, but it wasn't immediately clear to me how to even get access to that one.]
For an outfit that has professed a desire for economy and not performance, SpaceX is attempting a no compromise engine, near the top of all performance categories, reusable and cheap.
I found https://www.sciencedirect.com/science/article/pii/S2212540X1830004X which talks about ISP reduction, but without figures for methane. It also discusses coking problems using RP-1 as the coolant, which is applicable to Merlin. That article links to https://arc.aiaa.org/doi/pdf/10.2514/6.2010-6721, but I sadly don't have institutional access. If someone does and indicates it's worth it, I'll be happy to create an account (:sigh:) and purchase it. I can't even find the price without an account apparently though.
[the really interesting pointer appears to be to https://iris.unicampania.it/handle/11591/172095#.XFxBW89KiGR, but it wasn't immediately clear to me how to even get access to that one.]
In the spirit of democratizing research, here are your mentioned papers.
Does anyone think they will achieve 300+ bars?
What I find interesting is the speed with which they test. It's been a couple of days since the last test. Just a couple. They've sorted out the green issue, presumably, and are ramping up the power.
Yet when you see reports from other engine makers, its weeks between each test, or at least that is the impression one gets. What on earth takes the time between tests? It it simply bureacracy? Or rigid timescales? Seems to me the iteration time really makes SpaceX stand out.
SpaceX may test Raptor briefly at 30MPa soon although it will likely be a few years before Raptor is routinely running at that during normal operational missions.At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
I'm reminded of Intel's tick-tock model (https://en.wikipedia.org/wiki/Tick%E2%80%93tock_model) from the mid 2000's. SpaceX is so far forward of the rest of the launchers (ULA/BO/ESA...) in their research and development it's not even funnyhmm, tick-tock is not directly applicable here - it's holding one thing constant while changing another (in intel's case, semiconductor manufacturing process vs chip design). In chip design -- where you can have relatively clean interfaces that let you change one thing without changing the other -- you can pull it off, and it lets you debug your unproven manufacturing process using a chip design that's known to work, and your unproven chip designs with a manufacturing process known to work.
I'm reminded of Intel's tick-tock model (https://en.wikipedia.org/wiki/Tick%E2%80%93tock_model) from the mid 2000's. SpaceX is so far forward of the rest of the launchers (ULA/BO/ESA...) in their research and development it's not even funnyhmm, tick-tock is not directly applicable here - it's holding one thing constant while changing another (in intel's case, semiconductor manufacturing process vs chip design). In chip design -- where you can have relatively clean interfaces that let you change one thing without changing the other -- you can pull it off, and it lets you debug your unproven manufacturing process using a chip design that's known to work, and your unproven chip designs with a manufacturing process known to work.
But you don't have that in rockets - you can't trivially replace the Merlin engines on a Falcon 9 with Raptors without changing a bunch of other things as well..
100,000 Hp just to run the fuel pumps. For one individual engine. 100,000 Hp to power the fuel pumps. There might be 31 of them going off at once.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.
BE-4 is fairly far along.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.
BE-4 is fairly far along.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.
BE-4 is fairly far along.
Respectfully disagree on the state of BE-4.
Blue and X use different terms to describe their development engines.
BE-4 has long firing of "full scale engine". This enging was full size but incapable of more than 70% of full thrust.
X has long firing of "sub-scale raptor". This engine was full size but only able to reach 250 of the 300 bar pressure goal.
These two development engines appear the same full size not full thrust.
BE-4 is designing a new engine that will be able to fire at full thrust.
X just unveiled a "flight engine" that has briefly firied at lowest acceptable thrust level for system to work.
Agree regarding BE-4s lower thrust testing, but AFAIK the Raptor tests that happened before this week were done with an engine that was physically smaller, not just lower thrust. I think this is the first time that a full-size Raptor has been fired and it hasn't done any long duration burns yet.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.
BE-4 is fairly far along.
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
Yeah, I agree, they may do so soon, but they'll need to do a firing at full power for a long duration before they pass Blue.
BE-4 is fairly far along.
BE-4 is pretty far along, but Blue has yet to fire the full-thrust version of the engine. SpaceX has fired their flight thrust version of Raptor.
Unless Blue is planning a suborbital test vehicle, that will be no contest.BE-4 is pretty far along, but Blue has yet to fire the full-thrust version of the engine. SpaceX has fired their flight thrust version of Raptor.Let's see which one flies first!
Okay, but a month before that, they were still seemingly using carbon fiber for Starship, and they called it BFR.At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
Okay, but a month before that, they were still seemingly using carbon fiber for Starship, and they called it BFR.At the rate SpaceX are going with Raptor, please don't be surprised if Raptor is tested at 30MPa Pc within the next few weeks. Probably won't be running at that operationally for some time though.Musk's month-old tweet (https://twitter.com/elonmusk/status/1076618077301665793) suggests it will take some time to reach 300 bar. Doesn't sound like something achievable in a few weeks.
Now, they've switched to stainless, built a mockup/fairing, had a photo-op, had it blow over, and are now fitting out the mockup to do flights.
A heck of a lot changes in a months' time nowadays.
How about we not turn this into yet another BO vs SoaceX thread?(mod) Yes, let us not do that, thanks
So at this point it definitely appears that Raptor has passed BE-4 in terms of its development progress.
Raptor has fired for a few seconds.
]100,000 Hp just to run the fuel pumps. For one individual engine. 100,000 Hp to power the fuel pumps. There might be 31 of them going off at once.
Back to my bogglement with the power of the Raptor fuel pumps (propellant pumps actually, both sides)...
100,000 hp x 31 engines = 3,100,000 hp total for a SH.
Now some stats for Hoover Dam:
Flow through turbines - 906 m^3/s
Head - 158.5 m
Max power - 3,000,000 hp or 2,000,000 kW.
So about the same.
Keep in mind I'm not talking about the power of the Raptor engines themselves, only the liquid pumps that keep them fed.
Agree regarding BE-4s lower thrust testing, but AFAIK the Raptor tests that happened before this week were done with an engine that was physically smaller, not just lower thrust. I think this is the first time that a full-size Raptor has been fired and it hasn't done any long duration burns yet.
First time it has been publicly announced. But do we really know there were no full-sized test engines in the last year and a half? It's hard to imagine that SpaceX has just now produced the first full-size "radically-redesigned" test article and fired it at near full power a couple of days later, or that Musk would so confidently predict the hopper would be flying on these in the next two months, unless this were a smaller revision of an earlier full-size development engine.
]100,000 Hp just to run the fuel pumps. For one individual engine. 100,000 Hp to power the fuel pumps. There might be 31 of them going off at once.
Back to my bogglement with the power of the Raptor fuel pumps (propellant pumps actually, both sides)...
100,000 hp x 31 engines = 3,100,000 hp total for a SH.
Now some stats for Hoover Dam:
Flow through turbines - 906 m^3/s
Head - 158.5 m
Max power - 3,000,000 hp or 2,000,000 kW.
So about the same.
Keep in mind I'm not talking about the power of the Raptor engines themselves, only the liquid pumps that keep them fed.
So what you are saying is, that the turbopumps on the Super heavy could push the Colorado river back up into lake meade.
Edit: somehow quoted the wrong post.
Where did the term gas-gas come from? The Wikipedia article on FFSC references an NSF article, and Elon tweeted that the Raptor is gas-gas, but I can't find other references. I have a very poor understanding of chemistry, but aren't both the fluids coming out of the preburners supercritical? The pressure should be well over 3000psi.
On a related not, what would be the upper and lower bounds for O/F ratios on the preburners? I'm curious how the individual preburners' environments differ.
Where did the term gas-gas come from?
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD).
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.
John
SX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.
John
Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".
https://twitter.com/elonmusk/status/1008385171744174080QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD).
Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.
Jeff was stationed at Edwards AFB, CA where he joined the liquid rocket engine branch at the Air Force Research Laboratory and worked several component and engine technology programs that included his leadership of the joint Air Force-NASA Integrated Powerhead Demonstration engine which was the world’s first hydrogen full-flow staged combustion cycle rocket engine.
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.
John
>
Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.
....Raptor turbopumps alone need 100,000 horsepower per engine. That’s not a typo.
100,000 horsepower is about 75 MW. And he says per engine, so could that be 37.5 MW per pump? All this unit willy-nilliness hurts my head.>
Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.
Musk said in a Dec 22, 2018 tweetstorm,Quote....Raptor turbopumps alone need 100,000 horsepower per engine. That’s not a typo.
100k HP = 74.57 MW
https://twitter.com/elonmusk/status/1076618886932353024
I would imagine that oxygen pump is more powerful.Can you elaborate? Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.
I tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD).
Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.
Well, they hired Jeff Thornburg:QuoteJeff was stationed at Edwards AFB, CA where he joined the liquid rocket engine branch at the Air Force Research Laboratory and worked several component and engine technology programs that included his leadership of the joint Air Force-NASA Integrated Powerhead Demonstration engine which was the world’s first hydrogen full-flow staged combustion cycle rocket engine.
NopeI tried to find the source of the rumor that SpaceX accuired both data and hardware of the Integrated Powerhead Demonstration (IPD).
Were you able to find any evidence for this rumor? I think it's only mentioned by someone in this thread.
...Do you know what happens people who post traced art online?
Either way, HVM and livingjw, I'll be happy to help get a properly licensed and up-to-date version of your graphic up on Wikipedia....
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.
John
That's great. Do you have updated Isp figures to go with the new thrust and pressure values?
I would imagine that oxygen pump is more powerful.Can you elaborate? Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.
On the shuttle, the fuel pumps were about 3x as powerful. Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.
John
Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".
https://twitter.com/elonmusk/status/1008385171744174080QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.
I would imagine that oxygen pump is more powerful.Can you elaborate? Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.
On the shuttle, the fuel pumps were about 3x as powerful. Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.
The volumetric flow of methane is a little lower, but the pressure is higher since it has to flow through the regen system while the oxygen goes straight to the MCC. I would expect the fuel pump to need slightly more power, but they should be roughly similar.
I would imagine that oxygen pump is more powerful.Can you elaborate? Intuitively it would seem like pumping the oxygen would take more work because it's heavier and there's more of it, but (in my enthusiast's understanding) pumps do pressure-volume work, and the fuel has more volume.
On the shuttle, the fuel pumps were about 3x as powerful. Methane is more dense than hydrogen, and I've heard that the LOX and CH4 tanks will have similar volume, so I'd imagine that the two pumps would need similar power.
Added estimated pressures based on 1000 F turbine inlet temperature and 300 bar MCC. These are estimates assuming typical injector and coolant pressure drops and pump efficiencies. Total pump power is around 50 MW.
John
Perhaps I'm missing something but Musk stated that Raptor's "hot, oxygen-rich turbopump" needed new custom alloys to survive "~800 atm".
https://twitter.com/elonmusk/status/1008385171744174080QuoteSX 300 & soon SX 500. Kind of a modern version of Inconel superalloys. High strength at temperature, extreme oxidation resistance. Needed for ~800 atmosphere, hot, oxygen-rich turbopump on Raptor rocket engine.
Beats me. He also said pumps needed 100 MW. I'm calculating 50ish MW. I'm currently cross checking my model with the RD180. Stay tuned.
Here are a few possible guesses/suggestions that may help reconcile the numbers:
-The peak pressure may be much higher on the fuel side because it will be sized for a pressure drop to feed a much larger vacuum nozzle.
-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it.
-The Oxidizer line into the fuel rich preburner may need a boost pump. I think the RD-180 boosts fuel pressure to 700+bar in order to feed into the preburners. SSME has to boost ox pressure to 472 bar to feed into the preburners. Fuel line may not need one if my first suggestion is true.
(https://pbs.twimg.com/media/CL2pTAAUYAAuQGf.jpg)
Boost pumps in general might be the culprit for the apparent increase in plumbing in the new Raptor model.
-Liquid-gas preburners seem to need much larger pressure drops on the liquid side to feed well.
Just some guesses.
Here are a few possible guesses/suggestions that may help reconcile the numbers:
...
-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it....
Here are a few possible guesses/suggestions that may help reconcile the numbers:
...
-The 800atm figure number and 100,000hp (75kw) may refer to eventual numbers for the high thrust (330+ bar?) variant. The 800atm figure may also have a safety factor in it....
Another guess: 800 bar is the very local dynamic pressure due to pressure waves generated by the turbine blades, which is chemically relevant in local areas of the turbine, but is not the mean pressure at a larger scale.
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
If the amount of film cooling is so easily tunable that it could be altered for a nearly immediate subsequent test, and if they had this problem at the start of testing of the previous Raptor, then I wonder why they didn't start with more film cooling and then tune it back to a point where a hint of copper started to show in the spectrum?
Maybe the film cooling theory is wrong.
The copper may have just been traces from production and burned off in the first firing. Elon seemed not to worry about it at all.
Possible Raptor test?There are regular rocket engine tests in McGregor so the test heard may not necessarily be a Raptor test. Perhaps someone can find out if the noise was a Raptor test or not. A regular M1D test can sometimes sound a lot louder than normal due to a temp. inversion.
https://twitter.com/DJSnM/status/1094701660390121479
Looks like film cooling with methane to me. They must have increased film cooling over the previous test to protect the thrust chamber. Excess film cooling will reduce Raptor's Isp so they will no doubt find out the optimum balance of film cooling over time.
You think that thin halo around the core exhaust stream is near-pure methane? From distant memory don't almost all modern rocket engines do that to some extent to keep the walls cooler?
That's called film cooling, where propellant is used to cool the interior of the combustion chamber and/or nozzle by literally creating a film with boiling propellant that partially insulates them from the combustion gases. Nominally, Raptor (being a full-flow staged-combustion engine optimized in part for efficiency) would not use film cooling, as it's much less efficient than regenerative cooling.
If the amount of film cooling is so easily tunable that it could be altered for a nearly immediate subsequent test, and if they had this problem at the start of testing of the previous Raptor, then I wonder why they didn't start with more film cooling and then tune it back to a point where a hint of copper started to show in the spectrum?
Maybe the film cooling theory is wrong.
So what's the benefit or the lox hardware being directly atop The chamber?
And why is it but done before?
Is the idea directly from the integrated power head demonstrator?
So what's the benefit or the lox hardware being directly atop The chamber?
And why is it but done before?
So what's the benefit or the lox hardware being directly atop The chamber?
One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.QuoteAnd why is it but done before?
That's a good question.
This engine schematic just keeps getting better. I see you are still in the 'no boost pumps' camp - I do not know enough high pressure fluid dynamics to have an informed opinion.
>
So what's the benefit or the lox hardware being directly atop The chamber?
One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.QuoteAnd why is it but done before?
That's a good question.
https://twitter.com/elonmusk/status/1094782854007910400
Yes, aiming for 380 sec Isp with vac nozzle. Maybe 382 if we get lucky.
But not an extendable nozzle though, as that just saves length. Nozzle diameter is limited by body diameter.
Propellant was not deep cryo. CH4 & O2 were just barely below liquid temp at 1 bar. In theory, Raptor should do ~300 bar at deep cryo, provided everything holds together, which is far from certain. However, only 250 bar is needed for nominal operation of Starship/Super Heavy.
Much above 300 bar main chamber pressure means extreme oxygen preburner pressure of 700 to 800+ bar. Definitely pushing the limit of known physics.
Is the SL engine's expected vac ISP relatively unchanged from IAC 2017?
Close
This will sound implausible, but I think there’s a path to build Starship / Super Heavy for less than Falcon 9
Wow! I assume the switch to Stainless Steel is a big factor in this?
Yes
Stainless steel is CHEAP! Welding it is easy!
Super Heavy & Starship will eventually use gaseous methane & oxygen to pressurize the tanks so no helium is needed. They'll use the same gasses for the maneuvering thrusters as well, so no nitrogen or hydrazine will be needed either.
True
Could this be something that was made (more) possible by 3D metal printing?So what's the benefit or the lox hardware being directly atop The chamber?
One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.QuoteAnd why is it but done before?
That's a good question.
268.9 Bar equates to 3900.0 psi. I think someone at McGregor is working in English units and Elon had to translate before tweeting. I don't think those are Elon's favorite units.
-----------------Could this be something that was made (more) possible by 3D metal printing?So what's the benefit or the lox hardware being directly atop The chamber?
One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.QuoteAnd why is it but done before?
That's a good question.
268.9 Bar equates to 3900.0 psi. I think someone at McGregor is working in English units and Elon had to translate before tweeting. I don't think those are Elon's favorite units.
-----------------Could this be something that was made (more) possible by 3D metal printing?So what's the benefit or the lox hardware being directly atop The chamber?
One of the heaviest things in the engine is the high-pressure piping that leads from the turbopumps to the chamber. The weight of a pipe is linearly dependent on both flow rate and peak pressure, and these pipes need to survive both >500bar pressures and massive flow rates. Correspondingly, the pipe walls are really thick. By making the oxidizer turbopump exhaust directly into the chamber, they can minimize the length of the heaviest pipe in the engine, thus greatly reducing weight.QuoteAnd why is it but done before?
That's a good question.
268.9 Bar equates to 3900.0 psi. I think someone at McGregor is working in English units and Elon had to translate before tweeting. I don't think those are Elon's favorite units.Just look at the graph he showed - it is just the value reached, not a translation.
This engine schematic just keeps getting better.... One slight suggestion:How's about showing a pintle injector in the main chamber? I'm not certain but I think we know that to be the case? and if it is pintle was that what you assumed in your pressure drop calculations? How's about the injector type in the preburners, are they pintle a well or is that not applicable in that application? I know what you (livingjw) said about having local areas of stoichiometric combustion in a sea of oxygen or a sea of methane but does that mean it has to be multiple areas or would they go for just one if the preburners are pintle-ly injected?
So what's the benefit or the lox hardware being directly atop The chamber?
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
- Added pressures, temperatures and pump powers to engine schematic.
- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.
- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power
- At 320 bar I get 2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.
Added some more flow detail.
John
I'm not sure where this "knowledge" comes from, but my understanding is that pintle-type injectors are excellent at mixing liquid propellants - something that would be required in the preburners, but not the main combustion chamber. If the propellants are in the gas phase, they should mix well enough not to need a complicated injector. I have no idea if this distinction still applies with supercritical fluids, as I know next to nothing about them and their characteristics relative to liquids or gasses, so on that I must defer to someone with more expertise.This engine schematic just keeps getting better.... One slight suggestion:How's about showing a pintle injector in the main chamber? I'm not certain but I think we know that to be the case? and if it is pintle was that what you assumed in your pressure drop calculations? How's about the injector type in the preburners, are they pintle a well or is that not applicable in that application? I know what you (livingjw) said about having local areas of stoichiometric combustion in a sea of oxygen or a sea of methane but does that mean it has to be multiple areas or would they go for just one if the preburners are pintle-ly injected?
How's about showing a pintle injector in the main chamber? I'm not certain but I think we know that to be the case? ...I'm not sure where this "knowledge" comes from, but my understanding is that pintle-type injectors are excellent at mixing liquid propellants - something that would be required in the preburners, but not the main combustion chamber. If the propellants are in the gas phase, they should mix well enough not to need a complicated injector.
Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you 😁How's about showing a pintle injector in the main chamber? I'm not certain but I think we know that to be the case? ...I'm not sure where this "knowledge" comes from, but my understanding is that pintle-type injectors are excellent at mixing liquid propellants - something that would be required in the preburners, but not the main combustion chamber. If the propellants are in the gas phase, they should mix well enough not to need a complicated injector.
That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.
That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you
https://twitter.com/elonmusk/status/1094782854007910400I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).
That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you
Yes but that does not have to mean it uses a pintle injector. There are many kinds.
https://twitter.com/elonmusk/status/1094782854007910400I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).
At 320 bar chamber pressure, model predicts a Lox pump pressure of 683 bar. CH4 pump 698 bar. 320 chamber pressure vacuum engine, ER=119, matched stated 250 tonne max value.
Seems to me the white horizontal line is RD180 operating pressure and the wiggly line is data from the Raptor test fire, showing a ramp up over timehttps://twitter.com/elonmusk/status/1094782854007910400I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).
I think based on the graph’s legend, Musk is showing the RD180 pressures. So not space cows harmed in the making of that graph.
Seems to me the white horizontal line is RD180 operating pressure and the wiggly line is data from the Raptor test fire, showing a ramp up over timehttps://twitter.com/elonmusk/status/1094782854007910400I hope that the graph stopped there by selected window, and not by "loss of signal", (if it did I hope the pressure sensor or other parts didn't hit any cows).
I think based on the graph’s legend, Musk is showing the RD180 pressures. So not space cows harmed in the making of that graph.
This engine schematic just keeps getting better.... One slight suggestion:How's about showing a pintle injector in the main chamber? I'm not certain but I think we know that to be the case? and if it is pintle was that what you assumed in your pressure drop calculations? How's about the injector type in the preburners, are they pintle a well or is that not applicable in that application? I know what you (livingjw) said about having local areas of stoichiometric combustion in a sea of oxygen or a sea of methane but does that mean it has to be multiple areas or would they go for just one if the preburners are pintle-ly injected?
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
- Added pressures, temperatures and pump powers to engine schematic.
- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.
- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power
- At 320 bar I get 2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.
Added some more flow detail.
John
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?
At 320 bar chamber pressure, model predicts a Lox pump pressure of 683 bar. CH4 pump 698 bar. 320 chamber pressure vacuum engine, ER=119, matched stated 250 tonne max value.
250 tonne thrust was stated for a sea-level: https://twitter.com/elonmusk/status/1091156245132673024
I agree and I tried to say the same but it wasn't obvious from my comment.That is my impression as well - A pintle injector does not make sense for a gas injector. Neither SpaceX, Elon, nor Muller have never described Raptor as having an pintle injector.Travel time of CH4 and O2 through combustion chamber is very short and the speed of combustion is finite and relatively slow so to guaranteee that most of combustion happens in combustion chamber you have to help flows mix and I believe livinjw's idea of lots of parallel injectors for CH4 and O2 is the best because they shorten the distance that gasses need to travel so that they can mix. You can't count on turbulent flow to do everything for you
Yes but that does not have to mean it uses a pintle injector. There are many kinds.
This may be a newbie question, but when the engine is throttled, is it controlled simply by changing flow via valves from the tank inlets?
A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
Can't be exactly similar since they are different cycles, but one could throttle the gas generator maybe?A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
Can't be exactly similar since they are different cycles, but one could throttle the gas generator maybe?A safer way is to control the flow of the second fluid into the preburners. (That is, in the oxygen-rich side controlling the flow of fuel, and in the fuel-rich side controlling the flow of oxygen.) Reducing that flow reduces power in the turbine, reducing the force at which fuel is pumped through and therefore total power.Do we know if that is how the Raptor does it? Is Merlin throttled in a similar fashion?
lack of any exhaust port
LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?
This engine schematic just keeps getting better. I see you are still in the 'no boost pumps' camp - I do not know enough high pressure fluid dynamics to have an informed opinion.
>
Not sure if this applies; boost pumps = multi-stage? but from 2016...
David Ki Sun Yoon v@DavidKYoon
@elonmusk Sweet Jesus, that means you are pumping to 45-50 MPa... Surely this will be using multiple stage pumps?
|
Elon Musk ✓ @elonmusk
@DavidKYoon yes
11:41 PM - 25 Sep 2016
https://twitter.com/elonmusk/status/780296159315173376
LO2 pump one stage pumping ratio would be 1:210 (3bar:630bar).
Wouldn't it be more realistic tlfor LO2 pump to be two stage with ratios ~1:15 eg. first stage 3:50bar and second stage 40:700bar?
That's what inducers (the helical screw shaped part of the impeller at the front) are for - to gently accelerate the inlet fluid (which is already flowing at ~15m/s) into the main part of the pump impeller.
https://hackaday.com/2019/02/13/the-impossible-tech-behind-spacexs-new-engine/That's a very clear and thorough description of the point of staged-combustion engines, and why full-flow is so beneficial.
https://hackaday.com/2019/02/13/the-impossible-tech-behind-spacexs-new-engine/That's a very clear and thorough description of the point of staged-combustion engines, and why full-flow is so beneficial.
...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Tweaked CH4 pump to clear turbine exhaust path. Changed thrust to 2.05 MN. Changed Title to 12 Feb 2019.
Tweaked CH4 pump to clear turbine exhaust path. Changed thrust to 2.05 MN. Changed Title to 12 Feb 2019.
Quoting from the schematic thread. Where do you think they would fit the main LOX valve? upstream from the pump? Squished between pump and gas generator?
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.
I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!
What are the approximate dimensions of the Raptor engine manifold?
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.
What are the approximate dimensions of the Raptor engine manifold?
Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?
Not only that, you can also directly print negative molds:Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.
Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.
...
What are the approximate dimensions of the Raptor engine manifold?
Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?
Thank you, John. I was mainly interested in understanding the size of the item that was being casted in Musk's tweet.
What are the approximate dimensions of the Raptor engine manifold?
Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?
A very modern take on the very ancient "lost-wax" process.Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.
Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.
Demo video (https://www.youtube.com/watch?v=UJvjIB0rAUs)
It's a shame more people don't take the time to learn about manufacturing processes. It isn't all blacksmithing anymore! I guess that is a big positive about SpaceX...they are making manufacturing cool...how many on this site have learned about casting, welding, spin forming, and CNC machining? BTW a lot job openings in the trades to do this type of work.A very modern take on the very ancient "lost-wax" process.Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.
Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.
Demo video (https://www.youtube.com/watch?v=UJvjIB0rAUs)
What are the approximate dimensions of the Raptor engine manifold?
Throat is around .22 meter diameter (~9 inches). Assuming 4 to 1 contraction yields 2 to 1 diameter, so main chamber will be about .44 meters (~18 inches). Raptor engine main injector diameter will be about the same, or did you want another dimension?
I think main injector to nozzle throat contraction is generally far less than 2:1 diameter ratio.
SSME contracts far less than that, looks to be more like about 50% (~2:1 area ratios)
https://upload.wikimedia.org/wikipedia/commons/8/88/SSME_powerhead.jpg
F1 had main injector thurst chamber face diameter of 39inches, and a throat diameter of something like 36inches.
http://heroicrelics.org/info/f-1/f-1-thrust-chamber/f-1-cut-away-thrust-chamber-sm.jpg
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.
I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!
That's the explanation I was looking for. Thanks CorvusCorax.So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.
I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!
Close, but no cigar. A "full flow" staged combustion engine implies, ALL of the propellant moves through pumps, preburners and turbines, before entering the main combustion chamber.
As you wrote, this turns all the propellant gasseous, but thats not the main thing.
Because all of the propellant moves through the turbines, you essentially get a "low pressure, high flowrate" turbine, as opposed to a " high pressure, low throughput" turbine with the same turbine power - which would be used in a gas generator engine to minimize the amount of propellants wasted.
Even a staged combustion engine like RD180 or SSME - which is not full flow - needs to pump all the propellant with just a fraction of it passing through preburner and turbine. And that means for the same pump horsepower, they need a larger pressure differential in the turbine.
Physics and metallurgy define the maximum pressure in your preburners. Thats the limiting factor. Less pressure drop in the turbines mean more pressure in your main combustion chamber. Which usually gives you a more efficient engine, both in regard to (sea level) ISP and thrust 2 weight metrics.
So why would Raptor's CH4 turbopump run at lower pressure than either of SSMEs? Is maybe Raptor's higher Isp and Pc due to gas/gas mixing rather than liquid/gas in the MCC?...Maybe full-flow gets better ISP at sea-level due to higher achievable MCC pressures simply due to lower stress on the turbo pumps?
Yes.
My understanding is that because the SSME is a fuel-rich staged combustion engine, a single turbine must have the power to run both pumps. In this case, it has to handle all the power for both. In a FFSC engine, there are two turbines, one for each propellant. Thus, each turbine requires power sufficient only for its own propellant, and thus can run at less extreme pressures and temperatures. However, combined, they provide more power than a similarly-sized single ox- or fuel-rich turbine could provide, given material/engineering limitations. This results in the ability to have higher chamber pressure and Isp.
I'm not a engineer, though, so I could be wrong about this. Please correct me if that's the case!
Close, but no cigar. A "full flow" staged combustion engine implies, ALL of the propellant moves through pumps, preburners and turbines, before entering the main combustion chamber.
As you wrote, this turns all the propellant gasseous, but thats not the main thing.
Because all of the propellant moves through the turbines, you essentially get a "low pressure, high flowrate" turbine, as opposed to a " high pressure, low throughput" turbine with the same turbine power - which would be used in a gas generator engine to minimize the amount of propellants wasted.
Even a staged combustion engine like RD180 or SSME - which is not full flow - needs to pump all the propellant with just a fraction of it passing through preburner and turbine. And that means for the same pump horsepower, they need a larger pressure differential in the turbine.
Physics and metallurgy define the maximum pressure in your preburners. Thats the limiting factor. Less pressure drop in the turbines mean more pressure in your main combustion chamber. Which usually gives you a more efficient engine, both in regard to (sea level) ISP and thrust 2 weight metrics.
I actually started (https://en.wikipedia.org/w/index.php?title=Raptor_(rocket_engine_family)&dir=prev&action=history) the Raptor page on Wikipedia (https://en.wikipedia.org/wiki/Raptor_(rocket_engine_family)) in 2009. Back then it was thought to be a Hydrogen powered upper stage rather than an engine. Amazing that after 10 years the name has stayed but the meaning behind it is almost 180 degrees different.
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?
The methane pump has multiple stages. Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?
The methane pump has multiple stages. Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?
This confuses me a bit because I thought pumps increase pressure as well as move mass. So I would expect pump inlet pressures to be lower than the output.
The methane pump has multiple stages. Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?
Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure, and thus the cooling loop's inflow is highest of all. There couldn't be an even higher-pressure path going directly from the last pump stage to the preburner, because then you'd have backpressure through the preburner and into the cooling loop's outflow.
(If I'm wrong, I'd really appreciate a gentle education on why.)
This confuses me a bit because I thought pumps increase pressure as well as move mass. So I would expect pump inlet pressures to be lower than the output.
The methane pump has multiple stages. Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?
Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure, and thus the cooling loop's inflow is highest of all. There couldn't be an even higher-pressure path going directly from the last pump stage to the preburner, because then you'd have backpressure through the preburner and into the cooling loop's outflow.
(If I'm wrong, I'd really appreciate a gentle education on why.)
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?
The methane pump has multiple stages. Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?
Does the cooling loop have to be roughly at that pressure to stay liquid while cooling the nozzle? Would the heat otherwise vaporize the methane?Methane should turn supercritical in the cooling loop, at the high pressures they have there's no phase change between liquid and supercritical fluid IMO. So no explicit boiling, the fluid just gets hotter and more compressible.
The methane pump has multiple stages. Could it be possible (and would it make any sense) to run the cooling loop between these stages, at lower pressure?
Absolutely not my area of expertise, but my expectation would be that to avoid backflow every stage would be at a higher pressure than the one after it. So if the MCC hits 300 bar, then the preburner turbine's outflow must be greater than 300 bar, and so the turbine's inflow must be even higher pressure than that, and the cooling loop's outflow must be still higher pressure
Looks positively medieval. Half expected to see a burly fellow in a leather apron start hammering away at it.
Investment casting, or as I like to call it "old school 3D printing". Been around for centuries and still used for intricate casting including air cooled turbine blades.
Interestingly, investment casting works quite well with 3D printed models. I would expect in this case, the model has been 3D printed with some kind of plastics.
Doesn't the regenerative cooling loop have to significantly exceed main combustion channel pressure for purposes of film-cooling to begin with?
Add on top of it lower pump temperatures, is it really that bad an idea to have the coolant loop after all main pump stages?
Thought exercise... What are the ramifications of using regenerative cooling + film cooling?Basically the same as the heat shield- the film is too thin to stop radiative heating, but it can minimize convective heating, while the regenerative cooling handles the radiative heating from the plasma fireball it's containing.
*Not necessarily advocating that film cooling is actually being used.
The point RosKosmos is trying to make is that in your tweet you compared a liquid/gas engine to a gas/gas engine, and that the operating parameters of those two are not comparable. (Just translating the article, I definitely don't know enough to make a judgement.)
Not true. Limiting factor in any staged combustion rocket engine, liquid/gas or gas/gas, is pressure & temperature in oxygen preburner
Like was said before in this thread, all the modern staged combustion engines use both regenerative and film cooling, but they are separated systems. Coolant (fuel) for the film cooling is tapped off before the regenerative cooling loop in RD-180.
[1] Space Transportation System, Training Data, Space Shuttle Main Engine Orientation
[2] Atlas V Launch Services User’s Guide
SpaceX Merlin architecture is simpler than staged combustion (eg SSME or RD), but it has world record for thrust/weight & thrust/cost engine. Raptor has better Isp, but I’m worried it may fall short on those two critical metrics.
"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
Sorry if this was already talked about somewhere else:
643 bars in the manifold towards the nozzle? Isn't that a bit much pressure for the cooling jacket?
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.
- Added pressures, temperatures and pump powers to engine schematic.
- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.
- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power
- At 320 bar I get 2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.
Added some more flow detail.
John
Are there component differences between SN1 and SN2 that would prevent damage?
SN2 has changes that should help
https://twitter.com/elonmusk/status/1098653939141009408
Godspeed, Raptor SN1...
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.
- Added pressures, temperatures and pump powers to engine schematic.
- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.
- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power
- At 320 bar I get 2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.
Added some more flow detail.
John
Something I've been thinking about; perhaps the "full thrust" raptor drops the ER slightly while increasing chamber pressure?
how much would ISP change going from 270-280 bar with and ER of 35 (~0.22m throat) to ~320 bar with an ER of 32 (~0.23m throat)? Those numbers should get you 250 tonnes of thrust too.
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.
Big engines? He hasn't drawn the F1 yet. ;^)
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
Insane :D In case you're bored, how about adding an F1-engine?
Been working on my cycle model and have results consistent with data. I have to make assumptions about pressure drops across the coolant channels and combustion injectors. I used RD180 and SSME to get in the ball park.
- Added pressures, temperatures and pump powers to engine schematic.
- Still using dimensions and performance from 2017 but analyzing cycle at 300 bar.
- At 300 bar I get 2.05 MN at SL, and have a total of 62 MW total turbine power
- At 320 bar I get 2.2 MN at SL, and have a total of 72 MW or 96.6 k hp.
Added some more flow detail.
John
Something I've been thinking about; perhaps the "full thrust" raptor drops the ER slightly while increasing chamber pressure?
how much would ISP change going from 270-280 bar with and ER of 35 (~0.22m throat) to ~320 bar with an ER of 32 (~0.23m throat)? Those numbers should get you 250 tonnes of thrust too.
Note, ER limited by 1.3 m exit diameter constraint.
pc (bar) Dthroat ER IspSL (sec) IspVac (sec) Thrust SL (MN) Thrust Vac (MN)
270 .22 35 332 356 1.84 1.97
270 .22 120 ---- 375 ---- 2.08
320 .23 32 335 354 2.41 2.54
320 .23 120 ---- 375 ---- 2.70
Would have to reduce ER to 32.
New Drawing comparison:Vector graphics:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
New Drawing comparison:Vector graphics:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
Interesting Merlin vs Raptor comparison by Elon:
https://twitter.com/elonmusk/status/1098613993176850432QuoteSpaceX Merlin architecture is simpler than staged combustion (eg SSME or RD), but it has world record for thrust/weight & thrust/cost engine. Raptor has better Isp, but I’m worried it may fall short on those two critical metrics.
In May last year Tom Mueller seemed to be expecting a different outcome:
https://twitter.com/spacecom/status/999691403172036608Quote"Merlin holds the thrust to weight record now, but Raptor's coming." Mueller #ISDC2018
I removed copyrighted parts and added NASA Public Domain F-1...New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.
Big engines? He hasn't drawn the F1 yet. ;^)New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
Insane :D In case you're bored, how about adding an F1-engine?
Not enough room left to fit it on the paper... ;D
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”
It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.
Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.
It's also a sustainer stage, hence the big nozzle. The thrust chamber itself is only about a meter in diameter (don't quote me on that, just eye balling the hardware)
New Drawing comparison:
BE-4, Raptor, Merlin, SSME, NK-33, RD-180, Prometheus, and yes even a Rutherford for good measure 8)
That Rutherford looks cute compared to all the big engines. Also interesting to see how big the RS-25 was, considering its thrust & chamber pressure.
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”From a previous tweet:
It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.
Propellant was not deep cryo. CH4 & O2 were just barely below liquid temp at 1 bar. In theory, Raptor should do ~300 bar at deep cryo, provided everything holds together, which is far from certain. However, only 250 bar is needed for nominal operation of Starship/Super Heavy.
You do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”
It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.
Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.
I removed copyrighted parts and added NASA Public Domain F-1...
I'm sorry to say that. You know a lot about rocket engines guys but it's really hard to follow a thread about the Raptor and the news about it when it's full of diagrams which a lot of cases aren't related to the Raptor.
I'm sorry to say that. You know a lot about rocket engines guys but it's really hard to follow a thread about the Raptor and the news about it when it's full of diagrams which a lot of cases aren't related to the Raptor.
Tom Müller and other Elven-smiths are forging Twi~Essɛn right now, but it takes some time before that bird of prey expels its fire from its gullet and out of its mouth... So we need to entertain ourselves some how.
My guess would be they learned something important on the first runs that immediately got incorparated into sn2 and would make further testing of that part of the design less useful so they decided to stress some other parts hard since sn2 is coming soonYou do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”
It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.
Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.
Maybe. But there is a reason to find out what it can't do early on, rather than taking a lot of time characterizing what it can do: if it ultimately falls short, you've wasted that time.
With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.
There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.
With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.
There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.
With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.
There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.
There is zero chance that SpaceX was expecting an ISP of 380 for this engine & expansion ratio. The target would be more 340 ish. Chamber pressure is more key in getting the T/W they want, they are already pretty far up the ISP vs. chamber pressure curve above 200 bar so that ISP is more incremental. Others correct me if I am wrong please, but the ISP is not linear to chamber pressure, whereas thrust is more proportional to chamber pressure in that is is the first order determining factor in the mass flow rate.
Way too much concern for such limited information.
With regard to the end of the SN1 engine and the testing campaign, this is somewhat speculative but bear with it.
There were several issues, at least that is the perception, with the SN1 engine. Some were expected, some weren't. Right now the big one, at least from what we can gather externally and what many on this forum observed, is ISP. That is, the fuel efficiency. SpaceX was looking to reach 380 or even higher if possible, it now appears what they actually got was quite a bit lower but by how much we don't know. This is due to the film cooling problem. On the first test they did not have adequate film cooling and vaporized a small amount of the copper chamber liner. This would have become a large amount in a sustained MDC length test, but it was not terribly significant since these tests were very short.
There is zero chance that SpaceX was expecting an ISP of 380 for this engine & expansion ratio. The target would be more 340 ish. Chamber pressure is more key in getting the T/W they want, they are already pretty far up the ISP vs. chamber pressure curve above 200 bar so that ISP is more incremental. Others correct me if I am wrong please, but the ISP is not linear to chamber pressure, whereas thrust is more proportional to chamber pressure in that is is the first order determining factor in the mass flow rate.
Way too much concern for such limited information.
I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.No unburned fuel is dumped.
I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.No unburned fuel is dumped.
I wonder how fast can SpaceX iterate a design? Could they have tested version 1.0 then incorporate design changes into the next one and be testing version 2.0 the next week or two?Super allow foundry allows for rapid iteration of Raptor. So it is entirely possible that SpaceX have made changes to Raptor SN2 based on data from SN1 firings. Modern CAD/CAM + 3D printing of mold patterns for cast parts can allow for very rapid iteration of designs.
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.
They can't do that unless they also make the upper stage a lot heavier. The upper limit on Raptor size is set by the requirements to be able to hover and to be able to safely land even if they lose 2 engines during landing. This means that the thrust of the raptor at min safe throttle at sea level cannot be more than 1/3rd of the mass of the upper stage when near-empty.
They might relax those requirements a bit for the initial, unmanned versions, but I really don't see them making raptors bigger than the current design without also going to a larger diameter craft. Besides, Merlins are the champions of thrust/cost already, even with Raptors being a little worse than that, so long as they are reusable enough times, they should be fine.
Future iterations of Raptor could have even deeper throttle capability which could allow the no. of Raptors on SS to be reduced from 7 to 5 which would allow Raptor to be sized for 19 on SH. Also 19 is one of the nos. that gives max. packing density on SH while 31 is not. OTOH, SpaceX could make Raptor smaller for 9 on SS and 37 on SH which would also give max. packing density on SH. Still think that smaller Raptor may cost more than larger one for entire SS/SH propulsion system.EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.
They can't do that unless they also make the upper stage a lot heavier. The upper limit on Raptor size is set by the requirements to be able to hover and to be able to safely land even if they lose 2 engines during landing. This means that the thrust of the raptor at min safe throttle at sea level cannot be more than 1/3rd of the mass of the upper stage when near-empty.
They might relax those requirements a bit for the initial, unmanned versions, but I really don't see them making raptors bigger than the current design without also going to a larger diameter craft. Besides, Merlins are the champions of thrust/cost already, even with Raptors being a little worse than that, so long as they are reusable enough times, they should be fine.
EM said he was worried about Raptor meeting it's thrust/cost goal. Perhaps SpaceX could scale up Raptor to increase it's thrust/cost ratio. Larger engines should be cheaper per unit of thrust than smaller ones with everything else being constant. 19 larger Raptors for SH should be cheaper than 31 smaller Raptors for same installed thrust.he is saying here that Raptor might not be a record-breaker, not that it won't meet its design goals.
I think the concern about having to dial up film cooling is that it means you're dumping more unburned fuel out of your nozzle and this is diametral to maximizing ISP.No unburned fuel is dumped.
Superheavy would work fine with Merlin...
...or a modified Merlin burning Methalox if that turned out to be more cost effective...
My guess would be they learned something important on the first runs that immediately got incorparated into sn2 and would make further testing of that part of the design less useful so they decided to stress some other parts hard since sn2 is coming soonYou do have to wonder if the unstated part of “(as expected)” is “but not as hoped.”
It’s doubtful they actually already intended to test to destruction after so few runs, especially when they need three tested to operational readiness for BFH in just a couple months.
Elon does appear to be worried that performance for Raptor may not improve at fast enough rate to overcome any weight budget challenges Raptor or Starship has getting built.
Maybe. But there is a reason to find out what it can't do early on, rather than taking a lot of time characterizing what it can do: if it ultimately falls short, you've wasted that time.
Superheavy would work fine with Merlin, or a modified Merlin burning Methalox if that turned out to be more cost effective (and at this stage who knows, it might, as per-flight engine depreciation is quite possibly the dominant cost for the whole SS/SH system).
Regarding film cooling: methane has low molecular mass and high specific heat. At point it enters the chamber it is already at 700-800K and even it it only stays at that temperature without combusting with LOX will expand and be ejected with a velocity of nearly 2500m/s (Isp 250s). In reality it will be heated further by radiation, so that no-oxygen thin film region next to the wall costs very little in reduced average Isp.
Hydrogen film cooling is even better- it can potentially have higher Isp than the main combusted flow.