Test of a Supersonic Axial Compressor Stage Incorporating Splitter Vanes in the Rotor
Abstract
Complete experimental results are presented from tests of an axial- compressor stage designed for a tip speed of 1600 ft/sec, a stage total pressure ratio of 3.06, and an inlet hub/tip 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.
Document Details
- Document Type
- Technical Report
- Publication Date
- Jun 01, 1975
- Accession Number
- ADA014732
Entities
People
- A. J. Wennerstrom
- R. D. Derose
- W. A. Buzzell
Organizations
- Air Force Research Laboratory