Today it might seem to be an hypothetical question. I read that Surveyor mission confirmed "re-igniting" a LOX+LH2 engine in space is successful. What if the use of H2 was not successful ? - specially re-igniting in space - !! Further, use of hypergolic propellants in LEM ascent module, was selected to ensure lift off from moon surface, which was verified during Apollo 10. Had that been unsuccessful, was NASA prepared to launch all Apollos with LOX+RP1 (or any other, PROVEN propellant configuration) configuration on all stages?

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    $\begingroup$ Are you asking about being able to switch from (for example) LOX+LH2 to LOX+RP1 on the fly, or are you asking about design decisions that are made years before the vehicle actually flies? $\endgroup$ Jun 1, 2021 at 16:37
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    $\begingroup$ LOX/LH2 and hypergolics were hardly "unproven" at the time. $\endgroup$ Jun 1, 2021 at 16:53
  • $\begingroup$ @David & Christopher, all the engines (- mainly LOX + RP1/LH2), were found found to ignite & perform satisfactorily while on earth i.e. at launch. I believe the performance of LOX with RP1 was proven in multi stage rockets in space also. I was not aware if LH2 was also proven or not. Just in case it had not been, would RP1 have been tried out on LEMs (probably the only risk was about its ignition)? $\endgroup$
    – Niranjan
    Jun 3, 2021 at 13:25

4 Answers 4


Using RP-1 and LOX for the Service Module and the LM was impossible.

Replacing the fuel by a less energetic combination would require larger tanks and different engines designed for another propellant and more thrust for the heavier SM and LM.

A heavier LM and SM would require another much bigger Saturn V, with larger tanks and more engines. Eight or even ten engines instead of five for the first and second stage.

So only the CM could be used, but another and totally different version of the SM, the LM and the Saturn V would be needed.

But LOX could not be stored in the SM and CM for more than a week between launch and splashdown.


If the Surveyor flights had failed to re-ignite the Centaur's RL10, the engine would have been redesigned until it could re-ignite.

While it was theoretically possible to build a moon rocket with kerosene instead of hydrogen upper stages (like the USSR's N1 program), the US had decided early on that hydrogen was more practical for upper stages leaving LEO. In fact, Von Braun favored kerosene for the Saturn family, but his colleague Krafft Ehricke, working at Convair/General Dynamics in the mid-50s, led the push for a hydrogen engine for the Centaur upper stage.

  • $\begingroup$ Perfect. Thanks a lot. Many times you catch the "essence" of my question, and provide a pinpoint clarification. $\endgroup$
    – Niranjan
    Jun 3, 2021 at 13:30

There are lots of things on spacecraft that are built redundantly. Some are even built to have dissimilar redundancy such as redundant sensors that measure the same thing but are built by different manufacturers. Another example is that some spacecraft have a backup flight software system written by a group of developers who are firewalled from the developers who write the primary flight software system.

There are however some things in spacecraft that have zero redundancy, and others that do not use dissimilar redundancy. The Apollo return capsule itself had zero redundancy. A backup return capsule was not launched in case the first capsule failed.

Propulsion systems oftentimes employ some form of redundancy. Some vehicles have multiple cross-fed propellant tanks, but all hold the same fuel or oxidizer. Some vehicles have redundant thrusters, but all thrusters are designed to use the same propellant. I have yet to see a vehicle that uses dissimilar redundancy in its propulsion system. The mass penalty for dissimilar redundancy in a propulsion system would be huge.

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    $\begingroup$ While at least some part of NASA appear to really like dissimilar redundancy, there are strong arguments against it. One is that the devices / software won't be all that dissimilar. The builders of the dissimilar systems went to the same colleges, read the same books. The dissimilar systems most likely will have similar behaviors and similar failure points. Another is that dissimilar redundancy oftentimes is ridiculously expensive. Some argue that it would be far better spend a fraction of the money needed to build a dissimilar system to make the primary (and only) system that much better. $\endgroup$ Jun 1, 2021 at 16:32
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    $\begingroup$ I haven't finished my coffee yet, but I think OP is asking about switching the design from LH/LOX to kerosene/LOX entirely rather than about a backup system. $\endgroup$ Jun 1, 2021 at 16:34
  • $\begingroup$ @RussellBorogove I'll ask the OP to clarify the question. $\endgroup$ Jun 1, 2021 at 16:34
  • $\begingroup$ @ David, as I have indicated in my comment (on your comment) above, I am not looking for switchover to a different fuel. I was, in a way looking into the possibility of using the much proven RP1-LOX combination for upper stages and LEM. But your answer to my question had been useful to me in knowing many more things I dint know. Thanks. Russel has given a short & specific answer to my question, where as Dr. Sheldon and Uwe, have gone into details of "WHY RP1 WAS NOT BE USED" on upper stages & CM, LEM. That is also quite interesting. Thanks to all. $\endgroup$
    – Niranjan
    Jun 3, 2021 at 13:44
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    $\begingroup$ OP is an abbreviation for “Original Poster”, that is, “the person asking the question”. $\endgroup$ Jun 3, 2021 at 13:52

There are many problems with your suggestion.

  1. The exact hypergolic combination (Aerozine 50 and nitrogen tetroxide) used by Apollo was already used and proven during Gemini. There was no concern that it would fail. (That was a major purpose of the Gemini program: to prove the technologies needed for Apollo.)

  2. The specific impulse of the RP-1 S-IC first stage was 263 seconds; of the hydrolox S-IVB third stage was 421 seconds; and of the hypergolic CM and LM were about 310 seconds. You are replacing each propellant with an inferior one.

  3. Your question seems to suggest using RP-1 for the Apollo third stage, CSM, or LM. By the time these parts of the spacecraft are used, the first stage -- which is the only stage that used RP-1 -- has already been discarded. So you would either need to keep the dead weight of the first stage with you, or add extra tankage to these later spacecraft components.

  4. You can't just put RP-1 into a hydrolox or hypergolic engine and expect it to work; they are very different designs. So you would need to add additional engines to the third stage, CSM, or LM.

  5. RP-1 is one of the dirtiest-burning rocket fuels, and can leave deposits in the engine. This makes it extremely unreliable to restart an engine. In contrast, hydrolox and hypergolic engines burn cleanly, and in the case of Apollo were used for re-ignitable engines.

Items #2 through #4 each imply adding extra weight and extra cost, for something that is less reliable than what was already used.

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    $\begingroup$ SpaceX seems to have figured out how to restart RP1 engines. $\endgroup$
    – RonJohn
    Jun 2, 2021 at 6:05
  • $\begingroup$ @Dr. Sheldon: Point no. 1 explains it all...... Thanks. Without any intention to contradict your points 2, 3, 4 & 5, I was merely trying to find out (as if I am the designer/project manager) - If not LH2, what else can be used? We have to go to moon within this decade, regardless of efficiency. Also, RonJohn.. Your input was a new information. Great. Thanks. $\endgroup$
    – Niranjan
    Jun 3, 2021 at 13:55
  • $\begingroup$ 1. Most of the Saturn/Apollo design decisions were made pre-Gemini, and Gemini mainly helped with training and operations development experience fonApollo. The AJ10 used on the SM flew in 1958 albeit with a different hypergolic fuel mix. $\endgroup$ Jun 3, 2021 at 14:00
  • $\begingroup$ 2. You’re comparing the kerosene sea level specific impulse through a modest nozzle expansion to one of the highest expansion ratio hypergolic rocket engines ever built. Kerosene and hypergolics typically have very similar specific impulse figures. $\endgroup$ Jun 3, 2021 at 14:02

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