3. A ramjet engine is designed using 1-D parametric cycle analysis principles. The engine is designed to fly at a flight Mach number of Ma-3.2 and at an altitude of 18 km, where the temperature and pressure are -56 C and 7.43 kPa, respectively. The maximum stagnation temperature in the combustion chamber due to material considerations is Tos 3000 K. The heating value of the fuel (hydrogen) is QR 100 MJ/kg. The burner and nozzle stagnation pressure ratios are 0.90 and 0.85, respectively. The exhaust is under expanded, and as a result, the nozzle exit pressure is twice the ambient pressure (P.-2P.). According to US Military Specs/Standards/Handbooks MIL-E-5008B, the shock recovery of supersonic inlets can be estimated as: Poz/Poa 1 -0.075(M-1)1.35 Compute, 1 < M, < 5 a. The stagnation pressure ratio across the inlet b. The fuel-to-air ratio c. The flight and exhaust speeds MIL-E-5008B d. The size (frontal area) of the engine if the thrust requirement per engine is 100 kN and if the engine exit area is 0.6 m c. The thrust specific fuel consumption (TSFC) [unit should be Ns f. The overall efficiency. [40 Marks]