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This topic is offered for NewMars members who would like to collaborate in design of an SSTO using LH2 fuel and LOX
We assume at the outset that the process of developing a design for a specific mission will be an iterative process.
To try to make the process as free of distractions as possible, I'm offering the opening scenario as:
1) Payload in addition to the vehicle itself, delivered to LEO (300 km) is 100,000 kg (50 metric tons)
2) Fuel is LH2
3) Oxidizer is LOX
4) Location of launch is the equator to maximize the Earth's momentum contribution.
5) Other specifications can be added as they are requested
The business case is the LEO RV market. The system will be designed to convert into a habitat upon arrival in LEO.
It should be noted that the habitat volume will be much greater with this design.
Historical note: 100,000 kg is a payload that GW Johnson used for his calculations showing that an LH2 SSTO is feasible.
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Last edited by tahanson43206 (Yesterday 06:19:04)
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Post #3 kbd512
https://newmars.com/forums/viewtopic.ph … 39#p232839
Analysis showing LH2 possible at 100,000 kg payload
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Last edited by tahanson43206 (Yesterday 06:23:22)
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We finally arrive at an energy economics result favorable to LH2:
Delta IV Heavy Energy Economics Improve Upon Delta IV Medium
447,150,942N-s * 3 (Common Booster Cores) + 124,353,610N-s (DCSS Upper Stage) = 1,465,806,436N-s Total Impulse (both stages)
1,465,806,436N-s / 28,790kg of payload = 50,914N-s/kg of payload to LEO
Falcon 9 Block V Heavy Energy Economics
1,178,870,316N-s * 3 (Booster Cores) + 363,937,025N-s (Upper Stage) = 3,900,547,973N-s Total Impulse (both stages)
3,900,547,973N-s / 63,800kg = 61,137N-s/kg of payload to LEO
50,914N-s/kg / 61,137N-s/kg = 0.83
1. I take this to mean that the vehicle is very sensitive to thrust performance. Tahanson43206 set the payload target to orbit at 100,000kg. Therefore, we want to know the energy economics after we scale-up the LH2 powered SSTO to lift 100t. There is at least some potential here which indicates favorable energy economics associated with using Hydrogen fuel. At the 25t payload level or less, RP1 fueled solutions or LH2 paired with solid rocket motors appear to result in much smaller / lighter / equally performant vehicles in the realm of energy economics. This partially explains why vehicles like Ariane V, Atlas V, and Delta IV Heavy have all failed to compete with Falcon 9 Block V. For small launchers, no economy of scale is possible. You either have highly favorable energy economics or the cost of your engines and fuels override whatever minor performance advantages exist.
2. This is an actual desirable result showing a clear advantage in energy consumption required to attain orbit. Whereas the Delta IV Medium is little better than a Falcon 9 Block V, Delta IV Heavy consumes measurably LESS energy than Falcon 9 Block V Heavy. Therefore, it must be the case that using LH2 as the fuel of choice doesn't produce the desirable energy economics until the launch vehicle size is scaled-up quite substantially.
3. To make LH2 more competitive, we need better performing engines, such as RDEs, made from somewhat exotic but now much more affordable materials like RCC, in order to minimize engine mass or improve engine TWR, however you prefer to think about it. RDE alone would result in 150:1 TWR, and using RCC instead of stainless or Nickel-Copper alloys could increase that to 600:1, making LH2 engines suitable for SSTOs.
4. Any reusable SSTO must use CFRP propellant tanks. No other material is suitable because TPS mass will be added for reentry protection. However, I was considering the propellant mass fraction of the stainless steel balloon tanks used by the Centaur Upper Stage. Recent advances in Mangalloy welding make me think that Mangalloy, which has double the yield strength of 304L stainless, could be a lower cost and stronger alternative to fabricate much larger SSTO propellant tanks. 5.1m diameter 304L tanks are clearly doable, but ideally we want 10m diameter tanks. The Japanese and Koreans have done extensive testing with this material for storage of LNG, LN2, and LH2. Since they have test data to back up their assertions, I'm going to take them at their word. Some kind of coating, perhaps Silicon-based, would be required to protect the alloy from LOX, perhaps LH2 as well, but the end result could be relatively inexpensive, strong enough, and light enough. I was thinking about using steel cable with weldments inside the tank that keep the structure in tension and pull double duty as slosh baffles. I've never seen this done for this application, but we use it in bridges and other applications.
5. Someone needs to run an analysis to determine what flight trajectories are optimal for LH2 SSTOs, given their lower TWR. Is there a payload performance benefit to adding some forward / perpendicular velocity during part of the flight within the lower atmosphere, so long as the aerodynamic heating and drag loss is not too severe? If we're going to use steel balloon tanks, this might be more tolerable. All LH2 tanks have some external thermal insulation applied so that heat transfer rates remain tolerable. What kinds of lightweight thermal insulation do we have today? Perhaps we have something based upon spray-on aerogels that minimize weight and do a better job than polyurethane foam.
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