Inside a rocket, tanks are put one above the other. This make sense as it may be a good compromise between tanks shape and the aerodynamicity of the whole rocket. In some rocket stage, the oxygen tank is above the fuel tank, in other it is the opposite. Here are some illustrative examples: 

What are the criteria taken into account to decide which tank is put above the other when designing a rocket stage?

  • $\begingroup$ this is probably petty of me, but I'm curious why the accepted answer was changed from mine to Marbles' (admittedly also very good answer) after four years? $\endgroup$
    – Erin Anne
    Commented Oct 19, 2019 at 0:31
  • $\begingroup$ @ErinAnne It was a misclick. redone. That is true that both answer are good. $\endgroup$
    – Manu H
    Commented Oct 21, 2019 at 11:29

6 Answers 6


"Stages to Saturn, A Technological History of the Apollo/Saturn Launch Vehicles" by Roger E Bilstein (also NASA SP-4206, available http://history.nasa.gov/SP-4206/contents.htm and elsewhere) is a great reference on design decisions in the Saturn V rockets. One of the things it notes multiple times is that the cryogenic propellants cause freezing problems with the other propellants.

Stage I of the Saturn V is LOX and RP1. From Chapter 7 (pg 191 in the book):

The special problem of the LOX tank involved the feed lines leading to the thirsty engines about 15 meters below the fuel tanks. To do the job, the S-IC used five LOX suction lines, which carried oxidizer to the engines at 7300 liters (2000 gallons) per second. To achieve such high rates of flow, the lines could not be bent around the outside of the fuel tank; therefore, designers ran them right through the heart of the fuel tank. This in turn caused considerable fabrication problems, because it meant five extra holes in both the top and bottom of the fuel tank and presented the difficulty of avoiding frozen fuel around the super-cold LOX lines. The engineering fix on this included a system of tunnels, each one enclosing a LOX line, especially designed to carry an effective blanket of insulating air. Even so, the warmer fuel surrounding lines created some thermal difficulties in keeping the LOX lines properly cool. So the S-IC used some of its ground-supplied helium to bubble up through the LOX lines, and kept the liquid mixed at a sufficiently low temperature to avoid destructive boiling and geysering, or the creation of equally destructive cavities in the LOX pumps.

So on the S-I (Saturn V Stage 1), you want to avoid turning the RP-1 into a slushie while still delivering the LOX to the engines as liquid. The S-II (Saturn V Stage 2) actually has similar thermal design in its way, but the colder element (now LH2) in the stage is again stacked atop and flowed past the warmer element (now the LOX).

I could swear there's phrasing to this effect that I'm just not finding in this read, but the basic idea is that if you freeze the stuff in the lines, the rocket is done for, but if you're flowing cold lines past a warmer tank the balance of the heat transfer is likely to keep your cold lines slightly warm, and your warm tank slightly cold, and everything keeps on working.


From Sutton, "Rocket Propulsion Elements", 1976 Edition:

Tanks can be arranged in a variety of ways and the tank design can be used to exercise some control over the change in the location of the center of gravity.

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As in all aerospace development, the booster tank arrangement is a design tradeoff. Minimization of structural weight and overall system cost would be traded off. I seriously doubt that thermal considerations play a major role in the arrangement of the tanks; the booster is only fueled for a short time and the flight time is even shorter. More important considerations are what happens to the c.g. during propellant drain and the effects on the overall system stability. The overall size of the stage and its ability to be used with existing or planned facilities is also a factor.

In short, there is no simple answer, and the different configurations found in existing boosters show the results of these tradeoffs.


One factor is the relative density of the propellants; putting the denser one higher in the stack gives a higher center of gravity, which is advantageous for aerodynamic stability (think of an arrow or throwing dart with its heavy head). LOX is denser than kerosene or liquid hydrogen, so you want it on top in general.

In the case of the Saturn V S-II and S-IVB stages, unlike your other examples, there's a single vessel with a common bulkhead separating the LOX and LH tankage. There might be a structural load reason for putting the lighter hydrogen propellant on top, or a temperature management reason for putting the colder one on top, further from the rocket exhaust.

  • 1
    $\begingroup$ "putting the denser one higher in the stack gives a higher center of gravity, which is advantageous for aerodynamic stability" - but it also provides a vector of spectacular failure. So, is it still a win? $\endgroup$ Commented Aug 14, 2015 at 21:00
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    $\begingroup$ Putting the denser one higher in the stack also reduces the necessary ullage pressure in that tank to avoid cavitation at that tank's turbopump inlet, as you can use the hydrostatic pressure head associated with the height of the tank to your advantage. $\endgroup$
    – Tristan
    Commented Aug 14, 2015 at 21:54
  • $\begingroup$ What vector of spectacular failure are you talking about? Rockets don't generally just fall over for no reason. $\endgroup$ Commented Aug 14, 2015 at 22:08
  • 3
    $\begingroup$ @RussellBorogove Yes. They typically fall over for a variety of small and difficult-to-test-or-anticipate reasons. $\endgroup$
    – geometrian
    Commented Aug 15, 2015 at 5:16

In order to locate the center of gravity toward the top, in the case of cryo genic propellants it is theoretically desirable to position the hydrogen tank closer to the engine in order to exploit the taller tank height of the hydrogen and the greater weight of the oxygen for effective thrust vector control. Nevertheless, for cryogenic upper stages the opposite is usually done (e.g., for the Centaur and ESC-A upper stages), since a heavy oxygen tank mounted below the hydrogen tank yields a smaller dimensioning loadcase and thus a lower mass for this stage. (Handbook of space technology, Wilfried,Ley,Klaus,Wittman,Willi


CG, Center of Gravity is usually the reason to put LOx ahead of fuel but in the Saturn with the F-1 engine thrust structure you wouldn't want to get cold next to the LOx if the LOx was aft. With the 2nd/3rd stages as aforementioned the LH2 was ahead of the LOX but LH2 being so much lower in density and temperature than the LOx would be the same reason, keeping the thrust structure from the cold, in this case 421 below zero of LH2 vs only 300 below of LOx.


Within the atmosphere large forces are possible due to wind shear so a lot of thrust vector control is need and having the CG forward reduces the needed gimbaling angle of the engine to maintain control of the vehicle by the forward CG reducing the aerodynamic instability of the vehicle. Once outside of the atmosphere (i. e., the upper stages) there are no aerodynamic forces and a smaller gimbal angle can control the stage so the lighter weight of the structure produced by putting the stage CG closer to the load of the engines is now possible. This I learned from an engineer who was asked to do a 6D simulate of a vehicle which used just the first stage of a multistage solid missile, but found that the TVC (gimbal angle) of the first stage was insufficient to make first stage flyable because the CG for the first stage alone was insufficient forward without the upper stages to be controllable in the atmosphere. Note the Saturn 1B where the lox was not place forward of the fuel used movable fins to augment the maneuverability due to gimbaling of the engines. Also only the outer 4 engines were gimbaled limiting control force.


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