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RESEARCH PAPERS

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

[+] Author and Article Information
C. Hah

NASA Lewis Research Center, Cleveland, OH 44135

D. C. Rabe

Wright-Patterson AFB, Dayton, OH 45433

T. J. Sullivan, A. R. Wadia

GE Aircraft Engines, Cincinnati, OH 45215

J. Turbomach 120(2), 233-246 (Apr 01, 1998) (14 pages) doi:10.1115/1.2841398 History: Received February 01, 1996; Online January 29, 2008

Abstract

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of eight periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier–Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20 percent of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Copyright © 1998 by The American Society of Mechanical Engineers
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