Accession Number:

ADA014732

Title:

Test of a Supersonic Axial Compressor Stage Incorporating Splitter Vanes in the Rotor

Descriptive Note:

Interim rept. Dec 1972-31 Dec 1974

Corporate Author:

AEROSPACE RESEARCH LABS WRIGHT-PATTERSON AFB OH

Report Date:

1975-06-01

Pagination or Media Count:

432.0

Abstract:

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.

Subject Categories:

  • Jet and Gas Turbine Engines

Distribution Statement:

APPROVED FOR PUBLIC RELEASE