In a given rocket engine, a mass flow of propellants equal to 87.6 lbm /s is pumped into the combustion chamber, where the temperature after combustion is 6000°R. The combustion products have mixture values of R = 2400 ft • lb/(slug)(°R) and γ = 1.21. If the throat area is 0.5 ft2, calculate the pressure in the combustion chamber.
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