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(20 points) Consider the flow through a rocket engine nozzle. Assume that the gas flow through the nozzle is an isentropic expansion of a calorically perfect gas. In the combustion chamber, the gas which results from the combustion of the rocket fuel and oxidizer is at a total pressure and temperature of 15 atm and 2500 K, respectively. The gas (rocket fuel and oxidizer mixture) expands to supersonic speed through the nozzle, with a temperature of 1350 K at the nozzle exit. The ratio of the specific heats for this gas is y = 1.2. Calculate: a) static pressure at the exit b) nozzle exit Mach number.