Experimental Investigation of a Supersonic Compressor Cascade
Technical Report,01 Jul 1971,30 Apr 1973
GENERAL MOTORS CORP INDIANAPOLIS IN INDIANAPOLIS United States
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This report describes in detail the experimental investigation of a stationary, linear, supersonic compressor cascade with blades of constant spanwise geometry and constant thickness linear sidewalls. The selected blade element was representative of streamline 19 of an advanced compressor configuration resulting from the Aerospace Research Laboratories axial compressor research program. The investigation covered the range of inlet relative Mach numbers of 1.535 - 1.683 and a range of static pressure ratios of approximately 1.1 - 2.3 and included laser velocimeter measurements of the flow within and around the cascade at the design Mach number.
- Fluid Mechanics
- Jet and Gas Turbine Engines