1) Simple turbojet operating with compressor pressure ratio: 8.0, turbine inlet temperature: 1200 K and mass flow: 15kg/s, when the aircraft is flying at: 260m/s at the altitude: 7000m. Suppose the component efficiencies, and I.S.A conditions given, compute propelling nozzle area, net thrust developed and s.f.c Polytropic efficiencies of compressor and turbine 0.87 Isentropic efficiency of intake 0.95 Isentropic efficiency of propelling nozzle 0.95 Mechanical efficiency 0.99 combustion chamber pressure loss is 6 percent compressor delivery pressure. Combustion efficiency 0.97 Atmospheric conditions: 7000m are 0.41 bar and 242.5 K. Suppose calorific value of fuel 42000 kJ/KgK.

2) Deduce thrust equation for turbojet engine.

3) What do you mean by propulsive and nozzle efficiencies?

4) Describe starting problem related with supersonic inlet. prepare down the solution for this problem.

5) Deduce the relation between area ratio Amas/Ai and external deceleration ratio ui/ua. Subscript I indicates inlet and a indicates free stream.

Use the values given below to solve the numerical:

For air Cp = 1.005 KJ/KgK, γ = 1.4
For combustion gas : Cp = 1.147 KJ/KgK,  γ = 1.333
For air and combustion gas: R = 0.287 KJ/KgK