An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with a low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of Detailed flow visualizations on the surfaces and five-hole probe measurements inside the blading and in the wake region showed clearly a three-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3-D shock system at blade passage entrance. The experimental data have been used to validate and improve the 3-D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distributions on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results. An analysis of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow near the walls is seen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the sidewall are in qualitatively good agreement to numerical surface streaklines. A considerable improvement of the numerical results could be achieved by a gradual grid refinement, especially in the corner region and by successive code development.
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e-mail: anton.weber@dlr.de
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July 2002
Technical Papers
3-D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall
Anton Weber,
e-mail: anton.weber@dlr.de
Anton Weber
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
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Heinz-Adolf Schreiber,
e-mail: heinz-a.schreiber@dlr.de
Heinz-Adolf Schreiber
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
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Reinhold Fuchs,
Reinhold Fuchs
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
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Wolfgang Steinert
Wolfgang Steinert
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
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Anton Weber
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
e-mail: anton.weber@dlr.de
Heinz-Adolf Schreiber
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
e-mail: heinz-a.schreiber@dlr.de
Reinhold Fuchs
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
Wolfgang Steinert
German Aerospace Center (DLR), Institute of Propulsion Technology, 51170 Ko¨ln, Germany
Contributed by the International Gas Turbine Institute and presented at the International Gas Turbine and Aeroengine Congress and Exhibition, New Orleans, Louisiana, June 4–7, 2001. Manuscript received by the IGTI, October 2, 2000. Paper No. 2001-GT-345. Review Chair: R. A. Natole.
J. Turbomach. Jul 2002, 124(3): 358-366 (9 pages)
Published Online: July 10, 2002
Article history
Received:
October 2, 2000
Online:
July 10, 2002
Citation
Weber, A., Schreiber, H., Fuchs , R., and Steinert, W. (July 10, 2002). "3-D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall ." ASME. J. Turbomach. July 2002; 124(3): 358–366. https://doi.org/10.1115/1.1460913
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