Why were different fuels used for different stages of Saturn V? I read that the first stage used kerosene (RP-1) and LOX combination, but liquid hydrogen and LOX for the second and third stages.

With my limited knowledge, I would expect that the fuel which gives the maximum specific thrust would be used in all the stages.

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    $\begingroup$ Ugh, while they’re closely related I don’t think this is a duplicate. A bunch of QAs touch on it but don’t answer it explicitly. $\endgroup$ – Russell Borogove Jan 10 '19 at 13:32
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    $\begingroup$ The third stage is also using LH2/LOX just as the second stage. The same rocket engine J-2 is used for both the second and third stage. The service module may be seen as a fourth stage using another hypergolic and storable fuel oxidator combination. LH2/LOX is neither storable for the whole mission nor hypergolic (self igniting on contact). $\endgroup$ – Uwe Jan 10 '19 at 16:37
  • $\begingroup$ @Russell Borogove: Thanks Russell, you have pointed out my concern correctly. If it is the administrator who is "Closing" the question, my suggestion is that there is no point in "closing" some question just because some "CUE" words are matching with a different question (which is asked earlier). One needs to evaluate if both the questions are the same, or is the answer to the previous question really clearing the doubt in the new question. $\endgroup$ – Niranjan Jan 11 '19 at 2:50
  • $\begingroup$ @ Uwe: Hi there, the first stage used RP1 and not LH2. But I could make out that different fuel is needed since the firing is out of atmosphere, and reliability and repeatability of firing is of paramount importance. For example, the ascent stage of Lunar module could not have depended on any system which would have "ignited" the fuel externally. This introduces an element of uncertainty. use of an hypergolic fuel combo. reduced the risk of malfunction. $\endgroup$ – Niranjan Jan 11 '19 at 2:55

RP-1/LOX (also known as kerolox) is run at a mixture ratio of about ~2.5-2.6, while LH2/LOX (also known as hydrolox), has optimal mixture ratios ranging from 4.13 at sea level to 4.83 in vacuum. The STS ran with a ratio of about 6.

At 4.12, the bulk density of hydrolox is 0.29g/cm³, at 4.83 its 0.32g/cm³. Bulk density for kerolox is about 0.81–1.02g/cm³, depending on mixture ratio.

This means a first stage using hydrolox would have to be about 2.5 times to 3.5 times as big as a stage using kerolox.

Furthermore, while hydrolox has a higher Isp then kerolox, building high-thrust kerolox engines is easier then building high-thrust hydrolox engines. The the RS-68A, which is the most powerful hydrolox engine ever constructed (this was not until the 90's), only produces about 3.5 million newtons of thrust, compared to about 7.7 million for the F-1 and almost 9 million for the F-1A.

For the first stage, the higher energy density of kerolox means the stage can be kept at a reasonable size and producing engines with high enough trust is feasible. A bigger first stage would not have been feasible.

For the same reasons, hydrolox was selected for the upper stages. The increased volume was not problematic, but the increased performance was sorely needed.

For the SPS (the SM's main engine), there were other concerns. Cryogenic propellants such as LH2 need to be kept cool, otherwise the fuel would just boil off. The added weight for a cryoplant and energy requirements were prohibitive. Simply ignoring boiloff over such long periods of time was also infeasible, as too much of the fuel would evaporate.

The SPS engine thus needed a fuel that is more stable, and is easy to store and handle, and is reliable. The SPS engine needs to fire multiple times, up to ten or more, with absolute reliability. Lighting a rocket engine is surprisingly difficult to do. Using a hypergolic propellant for the SPS engine means that the propellant ignites when fuel and oxidizer mix, making it easy to handle. Furthermore, the propellant needed to be storable in zero-gravity. The SPS was pressure-fed, so using a propellant that does not mix with the pressure gas was necessary. More information about the SPS subsystem can be found in the APOLLO EXPERIENCE REPORT - SERVICE PROPULSION SUBSYSTEM by Cecil R. Gibson and James A. Wood.

So, in summary:

  • The S-IC used kerolox because of the higher energy density and thrust then hydrolox,
  • the S-II (and S-IVB) used hydrolox for the better Isp and thus delta-v,
  • the SPS used Aerozine50/N2O4 for long-term storability and ease of ignition and
  • the LM used the same Aerozine 50 / N2O4 combination for both RCS (attitude control) as well as for the APS and DPS engine for the same reasons

Each stage used the fuel that was most suitable for the task at hand, resulting in different fuels being "optimal" for each stage. Raw performance, e.g. thrust or Isp, is not the sole deciding factor. Energy density and storability as well as ease of use for the given task also play major roles in selecting the propellant. Another great overview over the different fuels and their pros and cons can be found on Rocket propellants by Robert A. Braeunig.

  • $\begingroup$ SM did not use only Aerozine50/N2O4, this combination was used by the main engine. But the attitude control thusters used monomethylhydrazine/N2O4. See this question. $\endgroup$ – Uwe Jan 11 '19 at 22:41
  • $\begingroup$ @Uwe Good catch, I updated the summary to reflect better that I was talking about the SPS ;) $\endgroup$ – Polygnome Jan 11 '19 at 22:48
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    $\begingroup$ Yes, the fuel cost is quite low, but if RP-1 is hlaf the price of LH2, then that would shave 2% of the launch cost. NASA has a large budget, but not an unlimited one. Are you sure about the $93 price, it seems high. $\endgroup$ – user20636 Jan 12 '19 at 21:23
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    $\begingroup$ you will note, @sean, I said "whereas the entire propellant load for a F9 is supposed to be around 200k" While the entire fuel load is RP-1, its propellant load is RP-1 and LOX $\endgroup$ – user20636 May 20 '19 at 13:15
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    $\begingroup$ @JCRM: Oh, whoops! :-S $\endgroup$ – Sean May 21 '19 at 1:53

RP-1/LOX was selected for the S-IC first stage for the simple reason of size: LH2 is half as dense as RP-1, and the resulting first stage would be aerodynamically and structurally infeasible.

LH2/LOX was selected over RP-1/LOX for the S-II second stage and S-IVB third stage for the reason you'd expect: the 25% or so increase in efficiency. The smaller size of these stages means the reduced fuel density isn't as much of a problem as it would have been for the first stage.

The SPS engine has two requirements that greatly constrain the fuel choice: it needs to ignite possibly ten times or more (up to three mid-course corrections on the way to the Moon, lunar orbit insertion, descending to the LM release orbit, docking with the LM after ascent, trans-Earth injection, and up to three mid-course corrections on the way back). Further, it needs absolute reliability: if any of the other engines fail to ignite, you can abort, but most of the time, firing the SPS engine is the abort mode.

These requirements mean the only practical choice is a hypergolic propellant: since the fuel will spontaneously ignite on contact with the oxidizer, firing the engine is simply a matter of opening two valves. The Aerozine 50/dinitrogen tetroxide combination was selected because it was well-understood, having been used in a number of rockets and missiles.

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    $\begingroup$ RP-1/LOX was used for the first stage because the F-1 engine was the only one available engine delivering the necessary thrust. About 40 J-2 LH2/LOX engines as used for second and third stage would have been neccessary to deliver the same thrust as 5 F-1. Development of the F-1 was started in 1955 before NASA was founded. It took a long time to solve the problem of combustion instabilities. To test the second stage of Saturn V in flight, a thoroughly tested first stage was needed. Not using the F-1 engine would have delayed the Apollo mission substantially. $\endgroup$ – Uwe Jan 12 '19 at 10:29
  • $\begingroup$ @Mark. Thanks Mark but with this logic, selecting RP-1/LOX combination even for the second and third stage would have reduced the size of the entire stage 2 &3 and Saturn V on the whole, as well. Unless, the larger size was acceptable because it gave some other advantages like being hypergolic etc... Further the term "efficient" appears to have been used only in terms of higher "specific impulse / thrust". Since if the size would have been taken in account, the importance of being "higher efficient" would have diminished somewhat. (the correct word would have been - suitable) $\endgroup$ – Niranjan Jan 14 '19 at 4:15
  • $\begingroup$ @Niranjan, RP-1/LOX for the third stage would indeed have reduced the size of the third stage, but it would have increased the mass of that stage, and thus required a larger first and second stage to lift it. Similar reasoning applies to the second stage. It's only the first stage where the mass can be increased without penalty. $\endgroup$ – Mark Jan 14 '19 at 4:20

RP-1/LOX is a lot easier to handle than LH2/LOX--note that the Falcon 9 doesn't use LH2 in any stage to make life simpler. The efficiency of your fuel isn't as important in the first stage, they traded ease of handling for less efficiency. There's also the advantage that RP-1 is a lot more dense than LH2, the rocket doesn't need to be as big. Another advantage to the truly massive first stage of the Saturn V.

  • $\begingroup$ given that two of the six stages were using hydrogen there would have been little additional complexity. While true, the second sentence isn't a reason for any choice. $\endgroup$ – user20636 Jan 11 '19 at 22:59
  • $\begingroup$ If it were just the hassle of handling the LH2 on the ground everybody would use it. It's the hassle of handling the LH2 in the stage itself. $\endgroup$ – Loren Pechtel Jan 12 '19 at 2:30
  • $\begingroup$ @ Loren Pechtel. Hi Loren, perhaps your opinion about trading w.r.t ease of handling V/S efficiency for the first stage seems more likely. Thanks. It seems H2L+LOX was used in 2nd and 3rd stages for being Hypergolic. $\endgroup$ – Niranjan Jan 14 '19 at 4:21
  • $\begingroup$ @Niranjan LH2 + LOX isn't hypergolic. Watch a video of the engine start on the space shuttle--a big shower of sparks to ensure there's no buildup of unburnt fuel during ignition. There would be no need of this if it were hypergolic. It's the stuff used in deep space that was hypergolic. $\endgroup$ – Loren Pechtel Jan 14 '19 at 5:38
  • $\begingroup$ @LorenPechtel: Hypergolic, no. Very easily ignited, yes. Which provides almost as much advantage in the simplicity-of-ignition department. $\endgroup$ – Sean May 18 '19 at 22:57

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