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I'm obviously not a rocket scientist, so this might rank among the stoopidest questions around here. I wonder if it is feasible to design a rocket engine which could be reused in space by refilling it with different kinds of rocket fuels.

In some imaginary future of a developed space economy, could the same rocket engine sometimes use LOX+H2 extracted from an asteroid, and then sometimes replace that fuel with LOX+CH4 produced in Mars' atmosphere? Or any other combination of more than one relatively easily available fuels.

EDIT: I don't refer only to the "engine", but to the rocket as a whole. This includes the tanks for fuel or other propellant and the plumbing which goes with it. I'm not even sure about the difference between an engine and the rest of a rocket.

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  • $\begingroup$ I was sure I remembered that the RL-10 had been modified to run on Methane, as well as Hydrogen, but I cannot find the reference now. Maybe it was another engine? $\endgroup$ – geoffc Jun 10 '14 at 17:55
  • $\begingroup$ I recall a Russian engine working on LOX, kerosene and LH2. astronautix.com/engines/rd0120td.htm $\endgroup$ – Deer Hunter Jun 10 '14 at 19:45
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    $\begingroup$ One issue you'd run into is tank sizes: H2 takes up much more volume than CH4, so if you went from H2 to CH4 you'd end up with lots of unused space. $\endgroup$ – Hobbes Jun 11 '14 at 8:48
  • $\begingroup$ I'm curious about a different case: Optional oxidizer. An upper stage that would reach orbit on UDMH/N2O4 and then maneuver around (change orbits) on UDMH as monoprop. $\endgroup$ – SF. May 30 '18 at 9:32
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Typically the difference between an engine of a rocket and a rocket is the engine only encompasses the injectors and control of those ejectors. Arguably, any control over the nozzle direction would also come under "engine" but it's a fuzzy line.

A rough breakdown of a rocket propulsion system is: fuel storage, piping, injectors, combustion chamber, nozzle. So lets go through them 1 by 1.

Fuel storage: this is your tanks. Now different fuels have different densities, and for a given pressure rated tank you will be able to hold a different amount of liquid from one fuel to the next. Also different fuels can react with the tanks inner walls in different ways. Further to this, helium (I know this isn't hydrogen but it's for example purposes) can actually leak through aluminium. So you can see there are a lot of issues with this, but none of them rule out using different fuels.

Piping: very similar to your fuel tank, we'll bundle valves into this section as well. Aside from the solid/fluid interactions that we discussed with the tank there's a couple of other issues with the piping. It needs to be big enough to let the right amount of fuel through, which means making it big enough for the fuel that requires the largest diameter piping, but then you are potentially over flowing the other fuels. Still, using the engine at some reduced efficiency would still qualify as using it!

Injectors: as I understand it the injectors are to mix and ignite the fuel. I would imagine (could be wrong here) that these things are designed bespoke for the fuel. The fluid mechanics of it are likely to be a nice complicated task, but again if we don't mix well but still ignite then we'll just have a reduced efficiency!

Combustion chamber: this is very similar to the issues related to the injectors, just to a lesser extent I would imagine.

Nozzle: this is a tricky one. The purpose of the nozzle is to turn the heat into speed. Now an ideal nozzle converges up to the point where the the exhaust reaches mach 1, and then diverges (subsonic fluid flow increases in speed through a convergent nozzle, supersonic is the opposite). I see no reason why you couldn't make the convergent part of your nozzle long enough for the exhaust that requires the maximum length, since you would just get a reduced efficiency from the other exhausts!

So in summary, I see no reason why it wouldn't work. You might just get a large reduction in efficiency, meaning more propellant being required for a specific delta V.

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    $\begingroup$ I read your answer and come to the exact opposite conclusion. For all the reasons you state rockets tend to be designed around the propellants, which is one of the most important early decisions to make in designing a rocket and engine. Your entire answer is a good summary of the reasons it probably wouldn't work very well. $\endgroup$ – Adam Wuerl Jun 19 '14 at 6:02
  • $\begingroup$ I'd go as far to say it almost definitely wouldn't work very well. However it's still possible. The dry mass of the system would be higher than desirable for sure. $\endgroup$ – ThePlanMan Jun 19 '14 at 8:11
  • $\begingroup$ "if we don't mix well but still ignite then we'll just have a reduced efficiency" Wouldn't that impure mixing lead to (potentially very serious) pogo oscillations? $\endgroup$ – a CVn Jan 8 '16 at 14:26
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    $\begingroup$ For all practical purposes, this is not feasible, and you left out pumps. $\endgroup$ – kert Nov 1 '16 at 17:22
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    $\begingroup$ More missing parts : igniters $\endgroup$ – kert Nov 1 '16 at 20:36
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A rocket engine type could use different propellants but it needs modifications. I can't say if astronauts could do such modifications in space for your scenario and if they can how easy it is, but maybe theoretically it could be done. Using different propellants in a rocket engine type has been proved by Aerojet LR-87 rocket engine. It was the only rocket engine in the world that with modifications has tested and operated using all three major propellant combinations: LOX/RP-1, LOX/LH2, N2O4 /Aerozine-50. It tested even an N2O4/Alumizine version. For the LOX/CH4 which was not popular at that time and not studied or invested for LOX/CH4 rocket engine projects, I don't know if they could, maybe yes.

LR-87 rocket engine versions:

LR-87-3, LOX/RP-1 was used in the Titan I missile.

LR-87-5, N2O4 /Aerozine-50 used in Titan II. Was lighter and simpler than its predecessor and did not need independent ignition system.

According astronautix:

LR-87 Alumizine, N2O4/ Alumizine tested in laboratories using a metallized fuel (for greater impulse density) and gelled propellants (to facilitate in-space starts after a period of coasting). The Aerozine 50 was slurried with aluminum powder (using Carbopol 904 gelling agent), and the engine was operated without any modifications, but could not attain stable long duration operation in that configuration. This was the first time a liquid rocket booster engine had ever been run on aluminized propellant. Many years later Aerojet operated small thrusters on metallized storable propellants and achieved satisfactory results.

LR-87 LH2, LOX/LH2 was in competition with J-2 for Saturn V. It was the first large Lox/LH2 engine in the world an LR-87 regeneratively cooled thrust chamber and nozzle, with a modified injector, and redesigned fuel pump. In comparison with LR87 Lox/RP-1, major changes were made on the injectors, and the RP-1 fuel pump was replaced by a single stage hydrogen pump specially designed for the purpose. The oxygen pump and its gearbox were the same.

Also are the tripropellant rocket engines that probably would be more practical since you don't have to do major modifications by changing parts. Considered two engines in one, with a common engine core with the engine bell, combustion chamber and oxidizer pump, but two fuel pumps and feed lines. In the 1960s, Rocketdyne fired an engine which mixed three separate streams of propellants, using a mixture of liquid lithium, gaseous hydrogen, and liquid fluorine to achieve highest specific impulse for a chemical rocket motor by 542 seconds.

RD-701 lox/lh2/kerosene. Uses one oxidizer and two fuels, switching them in mid-flight. Proposed for reusable MAKS space plane project. Built but cancelled later due to lack of funding. The mixing block of the main combustion chamber has three groups of injectors, one for each of three components.

Other versions RD-704, RD-0120TD, RD-0750.

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Specifically for the oxidizer/fuel combinations you specified, this appears to be unlikely for the simple reason of reaction kinetics.

We have

O2 + 2 H2 → 2 H2O

and

2 O2 + CH4 → 2 H2O + CO2

Note how in the first reaction we burn 1 part oxidizer with 2 parts fuel; in the second it 2 parts oxidizer with 1 part fuel (i.e. four times the oxidizer in relation).

Even taking different densities into account, this would require the turbo pumps to deal with such different mixture rates. I'm not a rocket scientist but this appears to be an engineering challenge. Consider that often the pumps for both fuel and oxidizer are driven by a single shaft, i.e. you can't simply open a valve to get some more of that oxidizer in, and less of that fuel.

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I think the where of the question is pretty important. Earth has a fairly strong gravity well and a poor efficiency rocket may not even be able to launch itself to orbit let alone an actual payload. Someplace like Mars with it's lower gravity or already in space is a different story however. For example for SpaceX's mars plan the launch from earth is a ship on top of a rocket which then gets refueled in orbit to get to mars but the return trip is just the ship launching straight from mars and not refueling at all.

I am not a rocket scientist either but from what I've ready you basically design your rocket around your engine and choice of fuel. Every little aspect ends up being impacted by those choices and each domino impacts other pieces. As was mentioned by multiple other posts the choice of fuel affects the ratios of the fuels and therefore the sizes of the tanks. Even something as simple as SpaceX's change to use sub-cooled fuels resulted in a change in the sizes of the two tanks. They were able to cool and densify the LOx more than the JP1 so they had to increase the relative size of the JP1 tank compared to the LOx tank to compensate.

I assume similar adjustments had to be made to the pumps because you ultimately want the same ratio of JP1 to LOx but need a different flow rate to accomplish that.

I'm pretty sure you could build a rocket that could run on different fuels but it would undoubtedly be more complex and/or less efficient than a single fuel rocket and for launching from earth that might be a complete no go. Just look at how difficult SSTO (single stage to orbit) designs are when trying to get something with usable capacity.

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