The Effect of Splitter Vane Circumferential Location on the Aerodynamic Performance of a Supersonic Compressor Cascade.
Final technical rept. 1 Mar 76-1 Feb 77,
GENERAL MOTORS CORP INDIANAPOLIS IND DETROIT DIESEL ALLISON DIV
Pagination or Media Count:
This report describes the experimental investigation of a linear stationary supersonic compressor cascade incorporating splitter vanes, blades of constant spanwise geometry, and contoured sidewalls. Previous experimental studies of this cascade showed that the cascade performance concurred with the modeled rotor data and that the splitter vane location andor shape was not optimal. Hence, the overall objective of this program was to experimentally determine if a preferred circumferential position for the splitter vane existed. This was accomplished by modifying the original cascade hardware to permit the splitter vanes to be moved in the equivalent circumferential direction with respect to the principal blades. The aerodynamic characteristics of the cascade were then experimentally determined at 41 test conditions. These covered a range of static pressure ratios between 1.6 and the spill point at the design inlet Mach number, for each of eight splitter vane locations, one of which was the original 50 percent spacing location. Author
- Fluid Mechanics
- Jet and Gas Turbine Engines