An analysis was made to determine the turbine cooling-air require¬ments of a particular turbojet-engine design believed to be representative of future low-specific-weight engines designed for supersonic flight at high altitudes. The engine, which was operated over a wide range of flight Mach numbers and altitudes, had an eight-stage transonic compres¬sor and a high-stress two-stage turbine having a turbine-inlet temperature of 2500° R. The mode of compressor operation considered was a combination of constant equivalent and mechanical speeds.
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