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Experimental investigation of the performance of a supersonic compressor cascadeSupersonic cascade wind tunnel results are presented for a linear, supersonic compressor cascade derived from the near-tip section of a high-throughflow axial flow compressor rotor over the inlet Mach number range of 1.30-1.71. Laser anemometry was used to obtain flow-velocity measurements showing the wave pattern in the entrance region. Attention is given to the unique-incidence relationship for this cascade, which relates the supersonic inlet Mach number to the inlet flow direction. An empirical correlation is obtained for the influence of the independent parameters of back pressure, axial velocity density ratio, and blade element performance.
Document ID
19880067148
Acquisition Source
Legacy CDMS
Document Type
Conference Paper
Authors
Tweedt, T. L.
(NASA Lewis Research Center Cleveland, OH, United States)
Schreiber, H. A.
(NASA Lewis Research Center Cleveland, OH, United States)
Starken, H.
(DFVLR, Institut fuer Antriebstechnik, Cologne Federal Republic of Germany, United States)
Date Acquired
August 13, 2013
Publication Date
June 1, 1988
Subject Category
Aerodynamics
Report/Patent Number
ASME PAPER 88-GT-306
Accession Number
88A54375
Distribution Limits
Public
Copyright
Other

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