How to determine the amount of combustion energy to provide the thrust in turbojet and turbofan engine? and in the following example How much of combustion Energy is Provided for the thrust? A jet aircraft moves with a velocity of 200 m/s where the air temperature is 20°C and the pressure is 101 kPa. The inlet and exit areas of the turbojet engine of the aircraft are 1 m2 and 0.6 m2, respectively. It is known that the exit jet nozzle velocity is 1522 m/s (from lab calculation) if the exhaust gases expand to 101 kPa at a temperature of 1,000°C. The mass flow rates of the inlet and exhaust flow are 240 kg/s and 252 kg/s, respectively. As a thermal engineer, your task is to (a) determine if the temperature of the exhaust gases is too high for the turbine blades as they exit from the combustion chamber. (b) Determine the amount of combustion energy necessary to provide the thrust. The maximum tolerable temperature of the blades is 3,000 K. It is known that the pressure ratio of the multi-stage compressor is 8 to 1. Assumptions and simplifications: neglect all losses and irreversibilities. All processes are isentropic. Neglect all kinetic energy components except at the inlet and the nozzle. Air and fuel mixture behaves as an ideal gas and has the same thermal properties as the air. All shaft works produced by the turbine are used to drive the compressor. Air (& mixture) has a constant CP=1 kJ/kg.K, and k=1.4