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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 (2025-07-15 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 (2025-07-15 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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https://en.wikipedia.org/wiki/Single-stage-to-orbit
https://selenianboondocks.com/2008/01/o … ched-ssto/
https://repository.gatech.edu/server/ap … b6/content
https://ntrs.nasa.gov/api/citations/199 … 024861.pdf
https://forum.nasaspaceflight.com/index … ic=36003.0
https://apps.dtic.mil/sti/tr/pdf/ADA321348.pdf
https://en.wikipedia.org/wiki/Single-stage-to-orbit
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What I showed was an all-expendable SSTO could be built, with the best-performing version using LOX-LH2, and plain-vanilla metal tankage of about 5% stage inert mass. I also showed that the dominant effect is raw Isp, not density impulse (although that does act in the right direction).
The biggest problem is getting engines of a chamber pressure high enough to reduce their dimension enough so that enough of them would actually fit behind the stage at both an acceptable stage L/D ratio (something in the 6 to 10 range or else you need to double the 5% drag loss you assumed), and an acceptable takeoff thrust/vehicle weight ratio (something above 1.5, or you need to double the 5% gravity loss you assumed). I could not achieve believable engines that would fit, with methane.
This thing had roughly the same payload fraction (about 6-7%) as an all-expendable TSO using LOX-LH2 in the upper stage, and pretty much any combination you desire in the lower stage. The real problem is the all-expendable design. You lose your engines every time you fly, and these will have to push engine state-of-the-art VERY hard, so they will be very expensive things to lose. The economics would be horrible, when by payload reduction to about half, you can recover the booster of a TSTO, leaving the rest of the design expendable. That is a fairly easy way to get far better economics, as we have all seen, here in recent years.
There is no "lower thrust/weight" ascent trajectory for vertical launch. You either dawdle at lower net effective acceleration and use most of your propellant in the first 30,000 feet, or you accelerate much harder and use your propellant more efficiently. You need at least 0.5 net gee upward off the pad, for a thrust/weight ratio of at least 1.5. We've already seen it many times since the 1950's. That effect is just no longer debatable.
Making an SSTO into a reusable craft is just NOT going to happen at a stage inert fraction in the 5-10% range! Neither will be making a TSTO upper stage reusable (something SpaceX is attempting with its Starship upper stage, and they are not yet successful). These things, to be reusable, must also be fully-qualified reentry vehicles, and must also be fitted out to land, either vertically or horizontally. That costs added mass, period! It is unlikely in the extreme to even happen nearer 15%. Most aircraft have inert mass near about 40% of max takeoff mass. Using composites reduces that below that notional 40%, but you cannot use them everywhere. Nowhere that gets really hot, for sure!
You run the rocket equation for yourself, at any Isp you think models a propellant combination you like, but with a ~25% inert mass fraction. If you get a positive payload fraction without using some sort of gas core nuclear, I'd like to hear about it. I typically get around 1300 sec Isp min. Higher if you want significant payload fraction.
Airbreathers reaching a practical staging speed or transition-to-rocket speed in the 1-2 km/s range, are simply not a practical way to approach this. The air at the requisite staging or transition altitude is just too thin, you are 40+ km up! Maybe 60+ km up! ALL airbreathers (no matter WHAT they are!!!) have a thrust level that is more-or-less proportional to the atmospheric pressure in which they are flying. Up there, 5-to-15 times nothing is still nothing! No significant thrust to push considerable mass! You will neither accelerate nor climb! It is called the "service ceiling" effect. Only rockets are immune to it. Because they do not breathe air.
GW
Last edited by GW Johnson (Yesterday 09:10:13)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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GW, A while back here, we looked into the possibility of building some kind of vertical launch assist. This would accelerate a rocket to ~Mach 1, vertically, over a vertical distance of a few km. The idea was to eliminate the propellant mass needed to reach Mach 1 from the launch pad, when propulsive efficiency is low. The concept appeared to show promiss. The most practical option appeared to be a steam cannon, which would accelerate a rocket sitting atop a sabot. The cannon would be about 3km long and would be located within a deep pit, with the rocket exiting the barrel a few hundred metres above ground level.
This concept was aimed specifically at reducing the required mass ratio for an SSTO to reach orbit. Do you have any methodology for determining how effective it would be?
Last edited by Calliban (Yesterday 09:53:41)
"Plan and prepare for every possibility, and you will never act. It is nobler to have courage as we stumble into half the things we fear than to analyse every possible obstacle and begin nothing. Great things are achieved by embracing great dangers."
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