Problem 1 A rocket exhaust nozzle has a ratio of exit to throat area of 4.0. The gases in a combustion chamber are generated with stagnation pressure equal to 3MPa (3000 kPa) and stagnation temperature equal to 1500 K (see Figure). Assume the exhaust gas has properties of air 500 K 3 MIPa Determine the exhaust Mach number, the rocket exhaust temperature and the rocket exhaust velocity for isentropi nozzle flow. Determine the exhaust Mach number, rocket exhaust temperature and the rocket exhaust velocity for the case where a normal shock is located just inside the nozzle exit plane. a. b.