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Control plate for shock-boundary layer interactionPaper describes tests and computations for a relatively unique technique to greatly reduce/eliminate the separation region for shock-boundary layer interactions. A number of studies have shown that the usual effects of such interactions include increased local heating and wall pressures, thickening of the boundary layer and a decrease in the momentum of the flow and, for stronger waves, flow separation. This flow situation is particularly prevalent in supersonic and hypersonic inlets where severe performance degradation can occur due to flow separation. High performance engine design generally requires a uniform entering flow field with little stagnation pressure loss. Previous approaches to the problem involved primarily active devices (e.g., suction or blowing); the present paper considers a passive device. The boundary layer separation control technique considered herein involves the placement of an embedded plate in the outer portion of the boundary layer and parallel to the wall. This control plate is situated such that the incident shock impinges upon and reflects from its surface, thus greatly lessening the pressure gradient in the low momentum near wall region.
Document ID
19850045726
Acquisition Source
Legacy CDMS
Document Type
Conference Paper
Authors
Goodman, W. L.
(NASA Langley Research Center Hampton, VA, United States)
Morrisette, E. L.
(NASA Langley Research Center Hampton, VA, United States)
Hussaini, M. Y.
(NASA Langley Research Center Hampton, VA, United States)
Bushnell, D. M.
(NASA Langley Research Center Hampton, VA, United States)
Date Acquired
August 12, 2013
Publication Date
March 1, 1985
Subject Category
Aerodynamics
Report/Patent Number
AIAA PAPER 85-0523
Accession Number
85A27877
Distribution Limits
Public
Copyright
Other

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