A non-afterburning turbojet is being designed for operation at an altitude of 15 km and a Mach number of 1.8. The maximum stagnation temperature at the inlet of the turbine is 1500 K. The fuel is a jet fuel having a LHV of 43124 kJ/kg and fst is 0.06. The following efficiencies apply at this Mach number: ηd= 0.09
ηc=0.9
ηb= 0.98
rb= 0.97
ηt= 0.92
ηn= 0.98

Use a gamma value of 1.4 up to and including the burner inlet, and a value of 1.3 for the rest of the engine. Assume R = 0.287 kJ/kgK throughout the engine (note that specific heats may be calculated from knowing gamma and R). Plot the specific thrust, TSFC, thermal, propulsion and overall efficiencies as a function of the total pressure ratio across the compressor, r_c. Consider arrange of r_c from 2 to 60. Also plot the nozzle area ratio (exit to throat) as a function of r_c.