Consider an ideal turbojet with an afterburner operating at flight Mach number of 2.0 at an altitude where the ambient static temperature: Ta = −70 ◦C and static pressure: pa = 15 kPa. The compressor pressure ratio is 10, the maximum temperature at the entrance to the nozzle is 2250 ◦K while the turbine outlet temperature is 700 ◦K. We assume that the exit Mach number Megxit = Mf light with the flow expanding in the nozzle to the ambient conditions at the exit. If the total fuel-to-air ratio for = 0.0519 for a fuel heating value of 45,000 kJ/kg, then determine: