Arocket engine uses a h2-o2 mixture as a propellant. the mean molar mass of the propellant is 11.58 kg/kmol with ratio of specific heats g = 1.22. the stagnation pressure at the throat of the nozzle is 8.26 mpa and the static temperature at the throat is 3305 k. the throat area at = 750.4 cm2 . the ratio of the exit plane area to the throat area is 39.8 and the static pressure at the exit plane of the nozzle is 18.1 kpa. (a) calculate the mass flow rate of propellant throug