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Transonic Compressor Rotor Cascade With Boundary-Layer Separation: Experimental and Theoretical Results PUBLIC ACCESS

[+] Author Affiliations
R. Fuchs, W. Steinert, H. Starken

Deutsche Forschungsanstalt für Luft- und Raumfahrt e.V., Köln, Germany

Paper No. 93-GT-405, pp. V03CT17A066; 9 pages
doi:10.1115/93-GT-405
From:
  • ASME 1993 International Gas Turbine and Aeroengine Congress and Exposition
  • Volume 3C: General
  • Cincinnati, Ohio, USA, May 24–27, 1993
  • Conference Sponsors: International Gas Turbine Institute
  • ISBN: 978-0-7918-7892-7
  • Copyright © 1993 by ASME

abstract

A transonic compressor rotor cascade designed for an inlet Mach number of 1.09 and 14 degrees of flow turning has been redesigned for higher loading by an increased pitch-to-chord ratio. Test results, showing the influence of inlet Mach number and flow angle on cascade performance are presented and compared to data of the basic design. Loss-levels of both, the original and the redesigned higher loaded blade were identical at design condition, but the new design achieved even lower losses at lower inlet Mach numbers.

The computational design and analysis has been performed by a fast inviscid time-dependent code coupled to a viscous direct/inverse integral boundary-layer code. Good agreement was achieved between measured and predicted surface Mach number distributions as well as exit-flow angles.

A boundary-layer visualization method has been used to detect laminar separation bubbles and turbulent separation regions. Quantitative results of measured bubble positions are presented and compared to calculated results.

Copyright © 1993 by ASME
This article is only available in the PDF format.

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