The Apollo Lunar Module LM used the same fuel/oxidizer combination (Aerozine 50 fuel / nitrogen tetroxide (N2O4) oxidizer) for both descent and ascent stage engines and reaction control system. This combination was also used for the engine of the Service Module SM.

The reaction control system of the SM used the same oxidizer but the different fuel monomethylhydrazine MMH. The Command Module CM also used MMH.

Both fuel/oxidizer combinations are hypergolic and non cryogenic storable liquids. What was the reason to choose two different fuels for the SM?

  • 4
    $\begingroup$ Sounds like yet another case of giving everyone a slice of the pie. Different contractors, different designs, the only reason being keeping all the contractors in business so that their experts wouldn't be left unemployed and seeking employment in other countries. $\endgroup$
    – SF.
    May 27 '18 at 13:12

This is especially interesting considering that the Service Module and LM RCS used the same thruster hardware (Marquardt R-4D). The R-4D was originally designed for MMH and first flew on Lunar Orbiter 1:

Marquardt experimented with a variety of liquid storable propellants. They selected NTO and MMH for their thrusters. However, government requirements led Marquardt to also use Aerozine 50... In the last couple of decades, Marquardt learned how to use the same thruster with several different propellants, namely, hydrazine, UDMH, MMH, or a mixture of any of these.

Both fuels deliver nearly identical specific impulse; Aerozine-50 is a couple of percent denser than MMH, so would produce very slightly more thrust for the same volume flow rate, but that kind of marginal performance difference wouldn't have been a consideration for the RCS thrusters.

MMH has a much lower freezing point (-52ºC) than Aerozine-50 (-7ºC). (Aerozine was developed for the Titan II, an ICBM usually sited in heated silos, thus its density impulse advantage was more important than its thermal range.)

The LM could pump fuel back and forth between the ascent engine and RCS tanks, which offered some contingency options, but there would be very few situations where it would be necessary or useful.

If I had to guess, I'd say the freezing point made MMH generally preferable, but that the LM designers preferred a single type of fuel -- remember that the CSM and LM were developed by different contractors (North American and Grumman respectively) with different engineering priorities. I didn't find any insight into the fuel selection in Kelly's book on the LM.

  • 1
    $\begingroup$ The R-4D thruster seemed to be used only for SM and LM but not for CM. The Wikipedia article lists only the propellant NTO/MMH for this thruster. But obviously two different fuels were used with this thruster for the Apollo mission. The SM should use the AJ10 engine and this engine was developed for Aerozine 50. The engine and the fuel were both developed by Aerojet. $\endgroup$
    – Uwe
    May 27 '18 at 15:57
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    $\begingroup$ May be the designers of the SM prefered to use the main engine and the RCS thrusters with the fuel they were designed for. Later the designers of the LM prefered to use the same single fuel due to the extreme weight limits. Both decisiones were sucessful as we know now. $\endgroup$
    – Uwe
    May 27 '18 at 16:35
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    $\begingroup$ The LM was designed much later than the CM and SM. There was enough time to test the RCS thrusters thorughly with another fuel. $\endgroup$
    – Uwe
    Aug 2 '18 at 9:11
  • $\begingroup$ The "marginal performance difference" between A50 and MMH could well be critical if the CSM's RCS thrusters had to be used to deorbit the spacecraft (the standard backup deorbit method for Earth-orbit missions in the event of an SPS failure) or for the final orbital insertion (a Mode V abort, which was available for the ASTP in the event of an S-IVB failure during the last 1.5 seconds of the insertion burn). Which would seem to make it better to use A50 for the CSM RCS thrusters and MMH on the LM, rather than the other way around, but I didn't design the thing... $\endgroup$
    – Vikki
    Nov 22 '18 at 5:13
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    $\begingroup$ @Sean It wouldn't be critical for deorbit; that wouldn't be time-critical. The window in which Az50 would get you a safe Mode V abort but MMH wouldn't would be about 30 milliseconds long: the kind of marginal performance difference that wouldn't be a consideration. $\endgroup$ Nov 22 '18 at 5:35

The decision to use different fuels for the main engine and the RCS was made before contractors were assigned, back when various NASA centers were conducting feasibility studies. At the time, the mission mode was expected to be a direct descent to the moon in one spacecraft. The main propulsion was expected to be solid rocket motors! Service propulsion system. Early requirements for the service module included vernier and main propulsion systems for a direct lunar landing profile. The main propulsion system was to consist of several identical solid-propellant motors which would provide thrust for translunar abort and lunar ascent. A separate module was to be designed that would provide for terminal descent. These requirements were changed early in 1962 to specify a single service module engine. Earth-storable liquid hypergolic propellants were to be used by the new system, which could include single or multiple thrust chambers. The service propulsion system was to be capable of providing for abort after jettison of the launch escape system, for launch from the lunar surface, and for midcourse corrections during earth return.

Apollo Program Summary Report

This is also confirmed in The Apollo Spacecraft: A Chronology:

The Group for Onboard Propulsion reviewed the three contractors' work on the Apollo feasibility studies. Among studies being undertaken by the NASA Centers and reported on at this meeting were: an STG consideration of an all-solid fuel propulsion system for a circumlunar flight, determination of midcourse and abort propulsion system requirements based on Saturn trajectories (MSFC), experimental evaluation at zero gravity of expulsion bag techniques for cryogenic propellants (Lewis), analysis and experiments on solid propellant rocket motors of very high mass fraction (Langley), methods of achieving thrust vector control by secondary injection of gases and the design of a highly reliable and versatile bipropellant spacecraft propellant system using hydrogen tetroxide and hydrazine or hydrazine derivatives (JPL), and a contract to examine hardware requirements for space missions and lunar landings (NASA Headquarters).

1961 January 6

By 1961 November 27, main propulsion was changed to a not-yet-specified hypergolic propellant. However, RCS development was already progressing along a separate path:

A single-engine service module propulsion system would replace the earlier vernier and mission propulsion systems. [...] Earth-storable, hypergolic propellants would be used by the new system, which would include single- or multiple-thrust chambers with a thrust-to-weight ratio of at least 0.4 for all chambers operating (based on the lunar launch configuration) and would have a pressurized propellant feed system.

The reaction control systems for the command and service modules would now each consist of two independent systems, both capable of meeting the total torque and propellant requirements. The fuel would be monomethylhydrazine and the oxidizer would be a mixture of nitrogen tetroxide and nitrous oxide.

The change to a lunar orbital rendezvous didn't occur until 1962 July 11. By then, too much work had been put into the two engine systems to justify consolidating their fuel.


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