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Quote from: hkultala on 10/04/2016 11:41 amQuote from: MATTBLAK on 10/03/2016 09:27 pmHas a version of the Merlin ever seriously been considered that runs on LOX/CH4? Even without all the full flow, staged combustion features of the Raptor; with subcooled propellants, what kind of performance could be squeezed out of them?Something like 15-20 second(<10%) increase in isp over Merlin, but T/W would be worse due methane needing bigger pipes and bigger pumps.Would require redesigning too many parts of the engine, that not worth doing.Well, the Airforce is paying for 1/3rd of the development cost, so they apparently hope it will be used to carry their payloads.That 10% boost in Isp (348->382 sec) on the second stage will give a 23% boost in LEO payload, and a 64% boost for GTO payloads (assuming the wet&dry weights are the same, according to my calculations). This would let the F9 match the Atlas 551, even with booster RTLS. For the first stage though, if the tank volume stays the same, the lower fuel density (therefore lower gross weight) will offset some of the Isp advantage.
Quote from: MATTBLAK on 10/03/2016 09:27 pmHas 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.
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?
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?
Quote from: Dante80 on 10/04/2016 11:52 amHere 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. 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 mVacuum Engine: - Expansion Ratio = 200 - Isp vacuum = 382 - Thrust Vac = 3.5 MN - Diameter Exit = 3.79 mBooster 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.cfgEngine size is defined by required thrust in vacuum.Results:http://lpre.de/upload/Raptor_performance.txthttp://lpre.de/upload/Raptor_nozzle.txthttp://lpre.de/upload/Raptor_cycle.txtO/F = 3.8Ae/At = 200Isp vac = 383 sThrust vac = 3.50 MNDe = 3.8 mModel for engine with sea-level nozzle:https://github.com/lpre/RPA-Examples/blob/master/Configs/Cycle%20Analysis/Raptor_SL.cfgEngine size is defined by throat diameter obtained from analysis of engine with vacuum nozzle.Results:http://lpre.de/upload/Raptor_SL_performance.txthttp://lpre.de/upload/Raptor_SL_nozzle.txthttp://lpre.de/upload/Raptor_SL_cycle.txtO/F = 3.8Ae/At = 40Isp vac = 356 sThrust vac = 3.26 MNIsp SL = 330 sThrust SL = 3.02 MNDe = 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
Quote from: livingjw on 10/08/2016 02:46 amWe 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. JohnQuestion, 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.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.
Quote from: RedLineTrain on 10/03/2016 06:08 pmIt 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 MINDMOLESTERHi 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.Biggest question right now is sealing the carbon fiber tanks against cryo propellant with hot autogenous pressurization. The oxygen tank also has an oxidation risk problem as it is pressurized with pure, hot oxygen. Will almost certainly need to apply an inert layer of some kind. Hopefully, something that can be sprayed. If need be, will use thin sheets of invar welded together on the inside.https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94tbej/?context=3&st=iuo8s2ur&sh=8d4dc7b8
They could use 200 - 300 deg F nitrogen instead.
Quote from: livingjw on 10/24/2016 04:29 pmThey 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.
Quote from: rsdavis9 on 11/03/2016 06:42 pmI'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.