As I stated in the answer to another question, LH2 suffers some serious drawbacks compared to other fuels.

Off the top of my head, they are

  • Extremely low density, resulting in:

    • Lower mass-fraction because of high tank mass
    • High aerodynamic drag due to tank volume
    • Larger and heavier vehicle structure due to tank volume
  • Boiloff and tank seepage, resulting in:

    • Even higher tank mass because of the need for insulation
    • More complicated tank arrangements because of the inability to butt LH2 tanks against relatively warm cryopropellants like LOX.
    • Long-term storage issues from propellant boiloff

The only possible advantages I can think of are

  • Marginally higher specific impulse
  • Producing only water as a byproduct, keeping environmentalists happy.

In light of all this, what rationale is there for a LH2/LOX Rocket?

  • $\begingroup$ LH2 is extremely cold, a lot of materials become very brittle at those temperatures. Construction of valves and hoses are difficult. LH2 can't build deposits in the cooling channels of the rocket engine, it can't polymerize or even carbonize like kerosene. $\endgroup$ – Uwe Oct 21 '16 at 11:35
  • $\begingroup$ Con: GH2 can embrittle metals causing them to lose strength over time. Pro: H2/O2 combustion is a very quick and simple chemical process (very much unlike hydrocarbon combustion). It is usually very stable. Pro: The specific impulse is not marginally higher, it's a lot higher, and in the end that's what counts. $\endgroup$ – Rikki-Tikki-Tavi Jan 20 '17 at 12:49

It's not a "marginally higher" specific impulse. High performance hydrogen engines typically have a vacuum Isp of around 420-450 seconds, compared to 310-350 for hypergolics or kerosene. That's about 30% delta v advantage, ton for ton, which more than offsets the structural volume penalty.

The drag penalty is mostly irrelevant for upper stages as long as their diameter is no larger than lower stages.

Another drawback to hydrogen is that low propellant density yields lower thrust from a similar chamber size; again not a big problem for upper stages.

The clean exhaust is nice, but not a huge advantage over kerosene. It is much more attractive than toxic hypergolics, of course.

Consider two 22-ton upper stages each pushing a 5 ton payload.

  • Stage H is 20 tons hydrogen-LOX, 2 tons dry mass, 450s Isp.
  • Stage K is 21 tons kerosene-LOX, 1 ton dry mass, 350s Isp.

H delivers 5956 m/s of delta-v versus K's 5162 m/s -- a 15% improvement despite hauling twice the dry structural mass.

  • $\begingroup$ It's worth pointing out that one of the drivers in fuel selection for optimal Isp is to have the exhaust gases be as light as possible, as the maximum exhaust velocity is inversely proportional to the square root of molecular weight. A hydrogen-based engine will have almost exclusively water as its exhaust gas. Long chain hydrocarbons like kerosene will have (in addition to the water) carbon dioxide, which is much heavier, as well as other constituents due to incomplete combustion, like carbon monoxide (which is actually better than CO2) and residual unburned hydrocarbons (worse). $\endgroup$ – Tristan Jul 29 '16 at 15:24
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    $\begingroup$ Saturn/Apollo is a terrific example of the tradeoffs in fuel selection. Kerosene first stage to meet the colossal liftoff thrust requirement and keep the stage diameter somewhat sane; hydrogen second and third stages to make the delta-v needed for orbital and translunar injection; hypergolics on the CSM and LM to store fuel for a two-week mission and start reliably, repeatedly, over many short burns. $\endgroup$ – Russell Borogove Jul 29 '16 at 16:01
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    $\begingroup$ Does hydrolox = LOX/LH2? $\endgroup$ – uhoh Aug 3 '16 at 12:02
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    $\begingroup$ Yes. Kerolox = kerosene/LOX, methalox = methane/LOX. In casual rocket-engine-discussion use, these are synonymous with the name of the fuel by itself (because other oxidizers are rarely used with those fuels). $\endgroup$ – Russell Borogove Aug 3 '16 at 14:41
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    $\begingroup$ @UIDAlexD: Thermal nuclear rockets may work on pretty much damn everything as long as its boiling temperature is less than the chamber melting temperature. They are way more efficient on hydrogen than anything else, but they won't suffer too badly if other, not too heavy elements are introduced into the mix. $\endgroup$ – SF. Jan 19 '17 at 12:54

Another advantage (relating to efficiency) is that it's easier to build a staged combustion rocket engine using LOX/LH2. Specifically, you can use LH2 for a fuel-rich staged combustion engine, like the Space Shuttle main engines (and the Energia main engines). Most other common rocket fuels don't work for fuel-rich combustion, I believe due to the risk of coking. Oxidizer-rich cycles don't have that problem, but require very advanced metallurgy to safely handle the (extremely corrosive) hot oxidizer. Russia / the USSR figured out how to do that in the 60s, but the USA never built a working oxygen-rich (or full-flow) staged combustion engine until the 2000s.

In the end, efficiency is efficiency; some new propellants that could get 600s ISP with a good TWR but could only be used in a gas generator cycle (or something similarly wasteful) would still be more efficient than the best staged combustion engines available today. Nonetheless, the fact that LH2 makes it easier to build higher-efficiency rockets (not just because of the inherent suitability of the propellants for high ISP) is - or at least was - significant.

  • $\begingroup$ fuel-rich combustion with methane seems like it has some advantages over the alternatives. $\endgroup$ – Philip Ngai Jun 17 '18 at 18:13

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