A supersonic compressor stage has been designed for a high pressure ratio at a tip relative inlet Mach number of 2. The stage was operated in the original configuration, but serious inlet stall occurred at part-speed operation. An inlet blockage ring, a bleed system and a variable geometry inlet guide vane have been analyzed and applied to this configuration. The results obtained with the bleed system in the complete stage are presented. The rotor performance is discussed and compared with the stage performance.
The Mach 2 Axial Flow Compressor Stage
Breugelmans, FAE. "The Mach 2 Axial Flow Compressor Stage." Proceedings of the ASME 1975 International Gas Turbine Conference and Products Show. Volume 1A: General. Houston, Texas, USA. March 2–6, 1975. V01AT01A021. ASME. https://doi.org/10.1115/75-GT-22
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