Convert your turbojet program into a turbofan program by including calculations for a fan and a fan nozzle. The losses will

be given exactly as they were for the turbojet calculations. You should use values for gamma and c$\_($p$)$ calculated using the average

temperature for each of the components. Once your program is operating correctly, use the program to solve the problem

given below. Your solution to the problem should include values for gamma for each component and the thermodynamic

conditions downstream of each component.

Problem:

A turbofan operates at $38,000$ft at Mach $0\mathrm{.}8.$ The compressor pressure ratio is $22\mathrm{.}0$ and the fan pressure ratio is $1\mathrm{.}8.$ The

fuel has a heating value of $17,900$BTU$//$lbm and the exit total temperature from the burner is $2480^($@$)$R$.$ The core airflow is

$155$lbm$//$sec and the bypass ratio is $5\mathrm{.}50.$ The aerodynamic efficiencies of the compressor, fan, turbine, shaft, burner,

primary nozzle and fan nozzle are $88\%,87\%,91\%,99\mathrm{.}5\%,92\%,95\%$ and $95\%$ respectively. The total pressure ratio

$($recovery factor$)$ for the diffuser is $0\mathrm{.}93.$ Assume the total pressure ratio for the combustor is $0\mathrm{.}97.$ The nozzles are both

converging but might not be choked. Find the developed thrust and TSFC$.$ Rerun this problem with a specific heat ratio of

$1\mathrm{.}4$ and work the problem out by hand to compare with your code results.

$**$I just need a MATLAB code for this problem$**$