The exhaust gases from a rocket engine can be assumed to behave as a perfect gas with a specific heat ratio of 1.3 and a molecular weight of 32. The gas is expanded from the combustion chamber through the nozzle. At a point in the nozzle where the cross-sectional area is 0.2 m2 , the pressure, temperature, and Mach number are 1500 kPa, 800°C, and 0.2, respectively. At some other point in the nozzle, the pressure is found to be 80 kPa. Find the Mach number, temperature, and crosssectional area at this point. Assume a one-dimensional, isentropic flow.