Random Thoughts: Which is a Better ISRU Propellant on Venus/Mars–LOX/LH2 or LOX/CH4?

I’m not sure if someone has already run the analysis, but I’m kind of curious about which ISRU-derivable propellant combination is better for locations like Venus or Mars where there is plenty of CO2 available in the atmosphere, but limited water.

Assume for a second that water isn’t available in useful quantities. I’m not sure yet if the concentration in the Venusian atmosphere is high enough to be useful, and it’s not yet clear that there’s easily accessible water at most places on Mars–there might be, but it’s far from clear1. Assume for now that you have to bring your hydrogen with you, from Earth.

A few quick observations:

  1. One immediately obvious point is that if you have to BYOH2 you’ll probably want to bring the hydrogen as LH2, not water. While everyone complains about how hard it is to store LH2 for long durations in space, water is still ~89% Oxygen, which is a material you can get almost anywhere, especially if you have CO2 available, so for every kg of hydrogen you bring tied up in water, you’re lugging around an unnecessary ~9kg of oxygen. You can definitely make a dewar with active cryogenic cooling that masses far less than 9x the mass of LH2 you want to bring with you–it may be “harder” to do it that way, but is far, far more efficient.
  2. A typical O/F ratio for LOX/CH4 is probably around 2.8-2.9:13, which means that about ~6.5% of your propellant mass is hydrogen for LOX/CH4. For LOX/LH2, you’re probably looking at an O/F ratio of around 5-6 typically, which would yield ~14-17% hydrogen. So for every kg of hydrogen you bring along, you could get4 ~15.4kg of LOX/CH4, or you could get 6-7kg of LOX/LH2.
  3. If you’re limited by hydrogen you can bring, rather than dry mass, or volume, or other things, it’s not yet clear which of those will result in more payload in orbit, since the two have significantly different bulk densities and Isp values. That’s the analysis that would be fun to run. My guess is a lot will depend on the required delta-V5, whether you’re looking at 1, 2, or 3 stages, if you assume on-orbit refueling before the earth-return, etc.
  4. One way to cheat a little with LOX/LH2 would be to use a LOX-rich Thrust Augmented Nozzle (TAN). Basically, you have a core running at the more traditional 5.5-6:1 O/F ratio, while initially running the afterburning portion of the engine at a much higher mixture-ratio, possibly even higher than stoichiometric! As the rocket accelerates, you could throttle down this element and then shut it off. This is probably more useful for Venus ascent than Mars, but would allow you to get not only a much higher engine T/W ratio than you could realistically get normally with LOX/LH2 engines, but also give you more propellant per kg of brought hydrogen, because you’re shifting your mission-averaged mixture ratio to a very, very lean range.

I honestly don’t know the answer, and don’t have time yet to run the numbers, but I’m genuinely curious. If you have a fixed supply of hydrogen, which ISRU propellant method (using a Sabatier reactor to convert H2 and CO2 to LOX/CH4, or using a solid electrolysis cell to crack O2 out of the CO2 to make LOX/LH2) actually yields the most mass delivered to orbit or to an Earth return trajectory from Mars or Venus? Has anyone already done this analysis? If not, I may try to find some time at some point to run the numbers.

[Update 9:58pm on 9/5/2016: in case you’re curious what brought this on, I was thinking about Venusian Rocket Floaties again, and was wondering whether a Venusian launcher first stage would want to be LOX/CH4 or LOX/LH2.]

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Jonathan Goff

Jonathan Goff

President/CEO at Altius Space Machines
Jonathan Goff is a space technologist, inventor, and serial space entrepreneur who created the Selenian Boondocks blog. Jon was a co-founder of Masten Space Systems, and is the founder and CEO of Altius Space Machines, a space robotics startup in Broomfield, CO. His family includes his wife, Tiffany, and five boys: Jarom (deceased), Jonathan, James, Peter, and Andrew. Jon has a BS in Manufacturing Engineering (1999) and an MS in Mechanical Engineering (2007) from Brigham Young University, and served an LDS proselytizing mission in Olongapo, Philippines from 2000-2002.
  1. Except to Mars optimists who always assume the most favorable definition of maybe
  2. Bring Your Own Hydrogen
  3. This is based on a ton of assumptions about chamber pressure, expansion ratio, etc. I used this chart, and eyeballed 1500psi with a decent expansion ratio.
  4. Assuming perfect chemical conversion efficiency and no losses along the way
  5. More delta-V like a Venus return will probably favor hydrogen more relative to methane, less delta-V like just getting to Mars orbit would likely favor methane more relative to hydrogen.
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14 Responses to Random Thoughts: Which is a Better ISRU Propellant on Venus/Mars–LOX/LH2 or LOX/CH4?

  1. Chris Stelter says:

    Neither. CO/O2 for first stage. H2/O2 for second stage, unless storability is too annoying, then CH4/O2 for second stage.

    Also, you could use raw liquified CO2 for thrust augmentation.

  2. Chris Stelter says:

    If you’re oxygen-rich, then past the certain point where the extra oxygen can help improve combustion completeness, you might as well use something like CO2 or nitrogen instead of ISRU oxygen, as the oxygen is just serving as reaction mass. (The argument against this is the complexity of having 3 tanks instead of 2, but I think that long-term that may be small compared to the energy advantage of not splitting off that oxygen.)

  3. Paul D. says:

    It depends on the launch architecture. If (as I posited on twitter) it has a pressure-fed first stage with steel tanks, the stage could be recovered by simply allowing it fall to an altitude where it would float (not something that’s as easy to do on Earth, where there’s a density discontinuity between atmosphere and ocean).

    Ideally you’d want to launch from a position where the launch platform is in winds moving east faster than the recovery depth of the first stage, so one could just wait for the wind to carry you to the stage to winch it up again. Not sure if the winds are suitable for this on Venus.

    A pressure fed stage prefers dense, non-cryogenic propellants. So, some sort of dense hydrocarbon (with a C:H ratio around 1) and N2O4 for the oxidizer. Pressurize the tanks with helium (which is recovered and reused).

  4. Bob Steinke says:

    @Chris Stelter There is a tiny advantage to using oxygen in excess beyond stoichiometric rather than inert mass like CO2 or N2. An excess of one reactive species will drive more complete combustion because the other species has lots and lots of potential partners.

  5. Paul D. says:

    You might also consider chlorine – based propellants on Mars where the regolith is maybe 1% perchlorate.

  6. Chris Stelter says:

    @Bob: Indeed, that’s why I prefaced my post with “past the certain point where the extra oxygen can help improve combustion completeness.” 🙂

  7. John hare says:

    With stochiometric about 5.3, would there be some density impulse advantage to a mixture leaner than 2.9?

  8. David S says:

    With a little simplification, RP-1 is C1H2. RP-1 has a density of 0.81, but if you chill it you get down to 1 g/ml. Carbon has a mass of 12, hydrogen has a mass of 1. So 1 ml of chilled RP-1 contains 0.14 grams of hydrogen.

    Liquid hydrogen has a density of 70.8 kg/m3. So 1 ml of liquid hydrogen contains 0.071 grams of hydrogen.

    So if you want to send hydrogen to mars, you send it as RP-1. Converting it to something else when you get to Mars seems unnecessary…

  9. David S says:

    Your tankage may vary… ;-}

  10. John hare says:

    I ran a note on 6,000 m/s from the surface of Mars. I came up with 0.514 tons of LH2 per ton of payload. With CH4, 0.28 tons of hydrogen

  11. John hare says:

    BOTE not note. Autocorrect the

  12. Jonathan Goff Jonathan Goff says:

    The problem is that bringing the hydrogen in the RP-1 means that for those 0.14g/mL of hydrogen you bring, you’re hauling 0.67g/mL of carbon from earth that you could’ve easily gotten from the Mars atmosphere. That is far, far heavier than a LH2 tank, insulation, and if needed cryocooler. Yes, LH2 is bulky, but for the rocket equation you usually care more about the mass you’re having to move around, not the volume.


  13. gbaikie says:

    It seems to me that H2 is most valuable rocket fuel per kg – on Mars, Moon or Venus
    surface. I assume on lunar surface if LOX is worth 1, then H2 worth 4 [or more].
    With LEO it’s 1 LOX and H2 is worth 2 [or more]. And with Venus or Mars surface, 1 to 5 [or more].
    “Which is a Better ISRU Propellant on Venus/Mars–LOX/LH2 or LOX/CH4?”

    It seems to me that with Venus, one going to use 2 [or three] stage rocket, and LOX/LH2 on upper stage might be better. With Mars it seems one is probably going to use a one stage rocket. One stage rocket to low Mars orbit, refuel it, and land it on Mars surface to be reused. From Low Mars orbit back to Earth, I think LOX/LH2 could probably be better.
    With Venus one may or may not try to reuse the the first stage- but recovering the first stage in general should be easier on Venus as compared to Earth- of course the needed added infrastructure may be too expensive and a distraction to exploration effort.
    I assume ISRU is meant to apply to NASA type exploration, as compared to mining which is more applicable to commercial/settlements type activity. Or ISRU purpose is solely related to lowering the cost of governmental exploration.
    Also for same reasons reusing the rocket to leave Mars, may or may not done in earlier part of Mars exploration.

    It seems to me that with NASA exploration of Venus or Mars, one going to need a fair
    amount of fuel in orbit. And a good orbit for either could be L-2.
    L-2 can be shaded, and/or L-2 could have 24 hour sunlight- depending where or what kind of orbit in L-2.
    A general idea of using High orbits is one use ion engines and/or aerobraking to get to low orbits. With Venus you going to have twice as sunlight as Earth and one doesn’t have the Van Allen belts, so using ion engines would better as compared to using ion engine with Earth’s LEO to high orbit. And with Mars one simply has less delta-v needed [though less sunlight as compared to Earth].
    Also one may using Ion rocket to move cargo to Mars [or Venus] and once ion engines with it’s solar panels are at Mars- the solar panels maybe more valuable left in Mars orbit as compared to returning them to Earth orbit. In such situation is might be better to ship water to Mars [or Venus] high orbit, where they use a surplus of solar panels to make rocket fuel.
    One might also consider having high pressure and cryogenic Hydrogen stored at Mars [or Venus] L-2. Stores as gas, and puts into spacecraft as liquid. And added mass of tankage if not going to the surface, might worth the added mass.

  14. gbaikie says:

    Re: “And with Mars one simply has less delta-v needed [though less sunlight as compared to Earth].”
    Also I forgot to mention, that one has longer orbital period- this makes easier to pulse hohman transfer type trajectory, rather spiraling out and wasting delta-v.

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