A rocket motor is designed to give 10, 000 N thrust at 10, 000 m altitude. The combustion chamber pressure and temperature are 2 × 106 P a and 2800°K, respectively. The gases exit the combustion chamber through a Laval nozzle. Find the exit Mach number, and the cross-sectional areas of the exit and the throat of the nozzle. Assume the nozzle flow is isentropic and one-dimensional, and that the ratio of specific heats γ for the combustion gases is 1.32.