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Experimental studies on two dimensional shock boundary layer interactionsExperiments have been performed on the interaction of oblique shock waves with flat plate boundary layers in the 30.48 cm x 30.48 cm (1 ft. x 1 ft.) supersonic wind tunnel at NASA Lewis Research Center. High accuracy measurements of the plate surface static pressure and shear stress distributions as well as boundary layer velocity profiles were obtained through the interaction region. Documentation was also performed of the tunnel test section flow field and of the two-dimensionality of the interaction regions. The findings provide detailed description of two-dimensional interaction with initially laminar boundary layers over the Mach number range 2.0 to 4.0. Additional information with regard to interactions involving initially transitional boundary layers is presented over the Mach number range 2.0 to 3.0 and those for initially turbulent boundary layers at Mach 2.0. These experiments were directed toward providing well documented information of high accuracy useful as test cases for analytic and numerical calculations. Flow conditions encompassed a Reynolds number range of 4.72E6 to 2.95E7 per meter. The shock boundary layer interaction results were found to be generally in good agreement with the experimental work of previous authors both in terms of direct numerical comparison and in support of correlations establishing laminar separation characteristics.
Document ID
19840035094
Acquisition Source
Legacy CDMS
Document Type
Conference Paper
Authors
Skebe, S. A.
(Case Western Reserve Univ. Cleveland, OH, United States)
Greber, I.
(Case Western Reserve University Cleveland, OH, United States)
Hingst, W. R.
(NASA Lewis Research Center Cleveland, OH, United States)
Date Acquired
August 12, 2013
Publication Date
January 1, 1984
Subject Category
Aerodynamics
Report/Patent Number
AIAA PAPER 84-0099
Accession Number
84A17881
Funding Number(s)
CONTRACT_GRANT: NAG3-61
CONTRACT_GRANT: NAG3-102
Distribution Limits
Public
Copyright
Other

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