1. 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 fat is 0.06. The following efficiencies apply at this Mach number: na = 0.9 nc = 0.9 nb = 0.98 Tb = 0.97 nt = 0.92 nn = 0.98 Use a gamma value of 1.4 up to the burner, and a value of 1.3 for the rest of the engine. Assume R is 0.287 kJ/kgK throughout the engine. Plot the specific thrust, TSFC, nth, np and no as a function of rc, the total pressure ratio across the compressor. Is there an optimum rc that minimizes TSFC? Is there an optimum re that maximizes specific thrust? Consider a range of rc from 2 to 60. Assume the exhaust is ideally expanded. Also plot the nozzle area ratio as a function of rc. 2. Add an afterburner to the above situation. The maximum stagnation temperature downstream of the afterburner is 2000 K. The afterburner combustion efficiency nab is 0.95 and the total pressure ratio is rab is 0.97. All other efficiencies remain the same. Use a gamma value of 1.4 up to the primary burner, and a value of 1.3 for the rest of the engine. Assume R is 0.287 kJ/kgK throughout the engine. Plot the specific thrust, TSFC, nth, np and no as a function of r, the total pressure ratio across the compressor. Consider a range of re from 2 to 60. Assume the exhaust is ideally expanded. Also plot the nozzle area ratio as a function of rc. Comment on the changes that occur due to the addition of the afterburner (i.e. compare the plots to that of problem 1). Note that the overall fuel-to-air ratio can not exceed the stoichiometric value (fb+fab<fst).

Fig: 1