Compressible Flow Project: Turbojet Design

I would like the tasks on the attached PDF to be completed in a clear professional manner that includes all methods, equations, and MATLAB/ Excel code (where applicable) used to arrive at final solution. Thank you. A propulsion system uses a turbojet engine. The combustion gas drives the turbine and is then accelerated to high speeds through a converging-diverging nozzle. The temperature at the exit of the combustion chamber (inlet of the turbine) is 1313 K. It is assumed that the combustion process takes place at very low Mach numbers (< 0.2), and therefore is a constant pressure process (no pressure loss). The thermodynamic processes in the compressor and in the turbine are considered isentropic (do not confuse this isentropic process with that in variable area flows) and the Mach numbers at their inlets and outlets are low. Consequently, in the calculations the thermodynamic properties in the combustion chamber, compressor, and the turbine can be considered as stagnation properties (as we described in the class). The flow in the nozzle is also assumed to be isentropic and the flow at the nozzle exit is perfectly expanded. Assume the mass flow rate of the fuel is negligible. Tasks to be completed: 1 Determine the optimal compressor pressure ratio to achieve the maximum specific thrust on the ground (temperature=293K, pressure=100kPa, and negligible flight Mach number, e.g., during take off). Determine the air mass flow rate needed to achieve a thrust of 55kN. If an afterburner is used, determine the afterburner temperature in order to achieve a thrust of 70kN. 2 If the turbine inlet temperature in Part 1 is increased to 1600 K, determine the optimal compression ratio and the specific thrust. 3 Determine the optimal compressor pressure ratio (for a turbine inlet temperature of 1313 K) to achieve the maximum specific thrust for a flight Mach number of 2.1 at 15000 m altitude, where the atmospheric temperature and pressure are 216K and 12.11kPa, respectively. You need to do calculations for several compression ratios to find the optimal value. Assume that the inlet is isentropic. 4 Design a two-dimensional inlet with three oblique shocks and one normal shock for operation at M=2.1. Minimize the stagnation pressure loss by varying the deflection angles. A good starting point is to use three identical deflection angles and find the best value. The angles can then be varied around this value to achieve the best values. For simplicity, use the (constant) properties of air in all calculations. The specific thrust for a perfectly expanded nozzle can be calculated as R/m=ue-ui. Important: Obtain ue using the pressure and temperature just upstream of the nozzle. Perform step-by-step calculations of the temperature and pressure (from the inlet to the compressor, combustion chamber, turbine, and nozzle). Do NOT use the efficiency formula. keywords: aerospace engineering, mechanical engineering, aerodynamics, Matlab