Aerospace Propulsion Assignment 2 — Turbojet Cycle Analysis  
Write your solutions in single column format, with one statement following another vertically. Write your solutions neatly so that they are easy to read and verify. Don't write one line with two equal signs. Highlight your answers using a box. Solve the problems starting from the Aerospace Propulsion Tables. Do not use any other document to solve the problems.
Problem #1
Consider a turbojet engine flying at Mach 0.9 at an altitude of 11 km. The mass flow rate of air coming in the inlet is of 29.17 kg/s. The nozzle is such that it is made of a converging section only which is designed such that it leads to the largest thrust. We know that the compressor efficiency is of 85%, that the turbine efficiency is of 90%, and that the stagnation pressure drops by 2% in the combustor. We also know that the compressor stagnation temperature ratio is of 2.512, that the stagnation temperature at the combustor exit is of 1650 K, that the heating value of the fuel is of 44 MJ/kg, and that the fuel-air ratio is of 0.03. Do the following:
(a)  Find the freestream temperature, and the stagnation pressure and temperature at the inlet entrance.
(b)  Find the power input to the flow by the compressor.
(c)  Find the burning efficiency in the combustor.
(d)  Find the stagnation pressure at the nozzle exit.
(e)  Find the stagnation temperature at the nozzle exit.
Assume that the gas constant $R=287~$J/kgK, the specific heat ratio $\gamma=1.4$, and the specific heat at constant pressure $c_p=1000$ J/kgK throughout.
1.  217 K, 252 K, 38.39 kPa, 11.11 MW, 0.808, 248.9 kPa, 1280 K.
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