UNIFIED ENGINEERING Fall 2004

Ian A. Waitz

Problem T10 (Unified Thermodynamics)

The compressor in a turbojet engine has a pressure ratio (ratio of exit stagnation pressure to inlet stagnation pressure) of 12. The ratio of combustor exit stagnation temperature to compressor inlet stagnation temperature is 6. The atmospheric temperature is 220 K and the pressure is 50kPa

and the jet is flying at a speed of 250m/s. Assume that the flow through the inlet, the compressor and the turbine can all be assumed adiabatic and quasi-static, with no friction or other losses, and the air can be treated as a perfect gas with constant specific heats (R=287 J/kg-K, c_p=1003.5 J/kg-K, c_v=716.5 J/kg-K).

a) What is the stagnation temperature at the compressor inlet?

b) What is the ratio of the stagnation pressure at the inlet to the atmospheric pressure?

c) What is the stagnation temperature ratio across the turbine?

d) What is the stagnation pressure ratio across the turbine?