Aerospace Propulsion Assignment 3 — Afterburner Turbojet Cycle Analysis  
Instructions
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.
05.22.14
Problem #1
A turbojet is flying at an altitude of 11 km and a Mach number of 0.9. 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 air mass flow rate in the engine is of 29.17 kg/s. After the turbine, we measure a stagnation temperature of 1280 K and a stagnation pressure of 248.9 kPa. The heating value of the fuel is of 44 MJ/kg. An afterburner is installed to increase the thrust of the engine. The afterburner is essentially a second combustor located in-between the turbine and the nozzle. Both in the afterburner and in the combustor, the ratio between the injected fuel mass flow rate and the air mass flow rate is of 0.03. Knowing that the burning efficiency in the afterburner is of 0.8 and that the thrust of the engine (including external drag) is of 34 kN, do the following:
(a)  Find the stagnation temperature after the afterburner, $T_{\circ 6}$.
(b)  Find the percent decrease in stagnation pressure within the afterburner, $(P_{\circ 5}-P_{\circ 6})/P_{\circ 5}$.
(c)  Find the percent increase in thrust due to turning on the afterburner.
(d)  Find the percent decrease in efficiency due to turning on the afterburner.
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.
Answers
1.  2240 K, 20.5%, 44.5%, 27.8%.
09.06.23
Due on October 1st at 11:59 pm.
09.17.24
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