Experiments have been performed in a Supersonic cascade facility to elucidate the fluid dynamic phenomena and loss mechanism of a strong shock-wave turbulent boundary layer interaction in a compressor cascade. The cascade geometry is typical for a transonic fan tip section that operates with a relative inlet Mach number of 1.5, a flow turning of about 3 degrees, and a static pressure ratio of 2.15. The strong oblique and partly normal blade passage shock-wave with a pre-shock Mach number level of 1.42 to 1.52 induces a turbulent boundary layer separation on the blade suction surface. Freestream Reynolds number based on chord length was about 2.7×106. Cascade overall performance, blade surface pressure distributions, Schlieren photographs, and surface visualisations are presented. Detailed Mach number and flow direction profiles of the interaction region (lambda shock) and the corresponding boundary layer have been determined using a Laser-2-Focus anemometer. The obtained results indicated that the axial blade passage stream sheet contraction (axial velocity density ratio) has a significant influence on the mechanism of strong interaction and the resulting total pressure losses.
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ASME 1991 International Gas Turbine and Aeroengine Congress and Exposition
June 3–6, 1991
Orlando, Florida, USA
Conference Sponsors:
- International Gas Turbine Institute
ISBN:
978-0-7918-7898-9
PROCEEDINGS PAPER
An Investigation of a Strong Shock-Wave Turbulent Boundary Layer Interaction in a Supersonic Compressor Cascade
H. A. Schreiber,
H. A. Schreiber
Institut für Antriebstechnik, Linder Höhe, Köln, Germany
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H. Starken
H. Starken
Institut für Antriebstechnik, Linder Höhe, Köln, Germany
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H. A. Schreiber
Institut für Antriebstechnik, Linder Höhe, Köln, Germany
H. Starken
Institut für Antriebstechnik, Linder Höhe, Köln, Germany
Paper No:
91-GT-092, V001T01A040; 11 pages
Published Online:
March 10, 2015
Citation
Schreiber, HA, & Starken, H. "An Investigation of a Strong Shock-Wave Turbulent Boundary Layer Interaction in a Supersonic Compressor Cascade." Proceedings of the ASME 1991 International Gas Turbine and Aeroengine Congress and Exposition. Volume 1: Turbomachinery. Orlando, Florida, USA. June 3–6, 1991. V001T01A040. ASME. https://doi.org/10.1115/91-GT-092
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