The mass flow through a rocket nozzle is therefore proportional to the throat area At and the chamber (stagnation) pressure px \ it is also inversely proportional to the square root of T/W. and a function of the gas properties. For a supersonic nozzle the ratio between the throat and any downstream area at which a pressure px prevails can be expressed as a function of the pressure ratio and the ratio of specific heats, by using Eqs. 3-4, 3-16, 3-21, and 3-23, as follows:
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