Aerospace Propulsion Assignment 4 — Turbofan 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
Consider a turbofan engine flying at 800 km/hr and a total pressure of 48.622 kPa (with the total pressure being the sum of the dynamic and the static pressure). It is known that the turbine generates a power of 0.6 MJ per kg of flow in the turbine. We know that the fan nozzle is a converging duct and that the pressure at the exit of the fan nozzle is of 60 kPa. The fan efficiency is known to be of $\eta_{\rm fan}=0.9$ and the compressor efficiency is of $\eta_{\rm comp}=0.8$. We also know that the stagnation temperature at the entrance of the combustor is 400 K higher than in the freestream and that the fuel air ratio within the combustor is of 0.04.
​ Do the following:
(a)  Draw a schematic of the turbofan and indicate clearly the station numbers at all important locations.
(b)  Find the freestream stagnation pressure and temperature.
(c)  Find the fan stagnation pressure ratio.
(d)  Find the stagnation temperature of the air at the entrance of the fan nozzle.
(e)  Find the bypass ratio.
(f)  Find the stagnation pressure at the entrance of the combustor.
Use a gas constant $R=287$ J/kgK, a specific heat at constant pressure $c_p=1000$ J/kgK, and a specific heat ratio $\gamma=1.4$ throughout.
Answers
1.  261 K, 50.45 kPa, 2.25, 337 K, 2.95, 830 kPa.
09.06.23
Due on Tuesday October 8th at 23:59.
09.30.24
PDF 1✕1 2✕1 2✕2
$\pi$