Test of a Supersonic Axial Compressor Stage Incorporating Splitter Vanes in the Rotor
Interim rept. Dec 1972-31 Dec 1974
AEROSPACE RESEARCH LABS WRIGHT-PATTERSON AFB OH
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Complete experimental results are presented from tests of an axial- compressor stage designed for a tip speed of 1600 ftsec, a stage total pressure ratio of 3.06, and an inlet hubtip radius ratio of 0.75. The rotor had been redesigned to incorporate a splitter vane between each pair of principal airfoils. At design speed, the compressor passed 88 percent of design flow, achieved a stage total pressure ratio of 2.77, and achieved isentropic efficiencies of 0.846 for the rotor and 0.674 for the stage. This represented a major improvement over the preceding configuration tested without rotor-splitter vanes. Future tests are to include various types of boundary-layer control.
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