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RESEARCH PAPERS

Experimental and Analytical Investigation of the Effects of Reynolds Number and Blade Surface Roughness on Multistage Axial Flow Compressors

[+] Author and Article Information
A. Schäffler

Motoren- und Turbinen-Union, München GmbH, Dachauer Str. 665, 8000 München 50, Germany

J. Eng. Power 102(1), 5-12 (Jan 01, 1980) (8 pages) doi:10.1115/1.3230232 History: Received November 06, 1978; Online September 28, 2009

Abstract

The general effect of Reynolds Number on axial flow compressors operating over a sufficiently wide range is described and illustrated by experimental data for four multistage axial compressors. The wide operating range of military aircraft engines leads in the back stages of high pressure ratio compression systems to three distinctly different regimes of operation, characterized by the boundary layer conditions of the cascade flow: • laminar separation, • turbulent attached flow with hydraulically smooth blade surface, • turbulent attached flow with hydraulically rough blade surface. Two “critical” Reynolds Numbers are defined, the “lower critical Reynolds Number” below which laminar separation occurs with a definite steepening of the efficiency/Reynolds Number relation and an “upper critical Reynolds Number” above which the blade surface behaves hydraulically rough, resulting in an efficiency independant of Reynolds Number. The permissible blade surface roughness for hydraulically smooth boundary layer conditions in modern high pressure ratio compression systems is derived from experimental data achieved with blades produced by grinding, electrochemical machining and forging. A correlation between the effect of technical roughness and sand type roughness is given. The potential loss of efficiency in the back end of compression systems due to excessive blade roughness is derived from experimental results. The repeatedly experienced different sensitivity of front and back stages towards laminar separation in the low Reynolds Number regime is explained by boundary layer calculations as a Mach Number effect on blade pressure distribution, i.e. transonic versus subsonic flow.

Copyright © 1980 by ASME
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