Call this a thought experiment.

Spacecraft to Earth orbit and beyond are traditionally launched by means of one or more rocket engines. The fundamental principle is that a rocket moves by expelling the vapour from a high-energy combustion. This energy is adequate to move the rocket as well as anything (within reason!) in physical proximity….

The obvious limitation of this approach is that the spacecraft a rocket may carry is a miniscule fraction of the combined mass of the rocket and its fuel.

Say that a spacecraft launched is capable of human-rated re-entry. What is the typical proportion of its re-entry propellant and heat-shield as a fraction of

  • total payload, and
  • launch vehicle propellant mass?

(The figures will probably vary from one launch engine to another; an approximate range will suffice in the answer. E.g. 3% to 5% of payload, and 0.9% to 1.2% of propellant mass.)

  • 3
    $\begingroup$ The NASA page The Tyranny of the Rocket Equation discusses this, and includes fuel fractions for some common propellant types for an SSTO mission. $\endgroup$
    – ToxicFrog
    Commented Dec 24, 2013 at 1:16

2 Answers 2


You are looking for the Propellant mass fraction.

Of course it depends on the components and will vary from one engine to another like you suggest, but for a quick answer, on average:

Propellant mass fractions are typically around 0.8 to 0.9.

  • $\begingroup$ Note that that's the propellant fraction for the takeoff propellant. The question was about the reentry. At reentry, only a small amount of delta-V is required so the propellant fraction is much lower. $\endgroup$
    – Hobbes
    Commented Aug 9, 2015 at 11:00

It depends on a few choices made by the designer of the spacecraft. If it follows a lifting skipping reentry trajectory, the spacecraft can potentially get away without a heat shield (since the airframe serves as the heat sink) and only have a small amount of propellant to deorbit. (around 10% of the spacecraft mass is the reentry propellant in this case)

If a ballistic reentry trajectory is used, then the heat fluxes rise enormously because all of the spacecraft kinetic energy is being dumped into the air via shock wave heating. Because of this, with existing materials, ablative heat shields are the only designs that work for this reentry trajectory, meaning the heat shield gets eaten away by reentry. (here it could be almost 25% of the spacecraft mass could be the heat shield/deorbit propellant. It depends on the specific heat and heat of vaporization of the heat shield material. Currently the state of the art here is the PICA-X material SpaceX uses on its Dragon capsules)

If a powered reentry trajectory is used, where the spacecraft uses reverse thrust to slow down throughout the whole trajectory, the heat loads experienced during reentry can be greatly reduced since if done right, most of the spacecraft kinetic energy can be dumped at altitudes above 200,000 ft, turning a hypersonic reentry into a slightly hypersonic reentry. However, the propellant cost of such a trajectory is high. Current rocket engines, with a maximum Isp of 465, would drive the spacecraft to be nearly 90% propellant if this reentry mode is picked.


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