The Table Below Details The Engine Operating Parameters Of An Aircraft Powered By A Single Turbojet Which Has Just Reach 1 (76.44 KiB) Viewed 44 times
The table below details the engine operating parameters of an aircraft powered by a single turbojet which has just reached the take-off speed. Use the data below to calculate the specific thrust (J/m) at take-off. Freestream static temperature and pressure (T, and p.) Take-off speed (u) Compressor inlet area (A) Compressor pressure ratio (Poz/Poz) Turbine inlet temperature (TO) Afterburner ON (nozzle inlet temperature) (1964) Exhaust nozzle choked at exit 288 K and 101325 Pa 68 m/s 0.785 m2 20 1500 K 1700 K Take the following values to be constant through the engine; y = 1.36 R=287J/kg.K Cp=1084J /kg.K Qx = 45x100 J/kg The following assumptions may be made; • All isentropic efficiencies are equal to one (na = n =n=n. = 1.0); • No stagnation pressure loss in the burner ( Po4 = Poz); • Combustion efficiencies in the burner and the afterburner are equal to one (n = na = 1.0); and No stagnation pressure loss in the afterburner (Poo = Pos).
A multi-stage axial compressor is to be designed such that it operates with an inlet stagnation temperature (Tol) of 293K and an outlet stagnation temperature (Ton) which must not exceed 663K The design is to consist of an unknown number of 0.5 reaction stages operating with the following conditions; • Rotor-blade inlet angle (B2) = -60° • Rotor-blade outlet angle (B2) = -40° • Constant axial velocity (CZ) = 125 m/s; y=1.3, c, = 1300 J/kg.K a) Assuming the total shaft work input is shared equally by each stage, calculate the maximum number of stages required to satisfy the design temperature criteria stated above. b) If the multi-stage polytropic efficiency is 0.8, calculate the compression ratio (Pon/Poi) which can be achieved with the number of stages calculated in part-a. Accompany your calculations with a labelled sketch of the rotor-blade velocity diagram.
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