SEPARATION MEASUREMENTS OF SUPERSONIC TURBULENT BOUNDARY LAYERS OVER COMPRESSION CORNERS.
Abstract
To the designer of modern high-speed aerospace vehicles, separation of the boundary layer before macroscopic surface features is an important problem. A typical feature of interest is a compression corner, or ramp. To aid in the development of comprehensive theories concerning boundary layer separation before a compression corner, a series of experimental investigations were conducted for the adiabatic flow of a supersonic compressible gas over a compression corner. Tests were conducted at free-stream Mach Numbers of 2.00 to 4.00 in increments of 0.50, at flat plate Reynolds Numbers of 0.75 x 10 to the 7th and 1.5 x 10 to the 7th, and for compression angles of 10.3 deg, 20.1 deg, 30.5 deg, and 39.9 deg. Static pressure surveys of the flow ahead of and over the compression corner were made. These measurements were supplemented by high-speed schlieren photographs and shadowgraphs. The separation of the turbulent compressible boundary layer was found to have strong dependence on both the Mach Number and the Reynolds Number. For Mach Numbers less than 3.00, the separation distance ahead of the compression corner decreased with increasing Mach Number. For Mach Numbers of 3.50 and 4.00, the separation distance increased with Mach Number. At all Mach Numbers, an increase in Reynolds Number increased the separation distance. The Reynolds Number influence was greater at the higher Reynolds Numbers. Unsteadiness in the separation geometry occurred for separation distances greater than six or eight boundary layer thicknesses. The separation was found to result from a free interaction of the flow phenomena involved. (Author)
Document Details
- Document Type
- Technical Report
- Publication Date
- Jul 21, 1966
- Accession Number
- AD0644194
Entities
People
- Walter B. Gillette
Organizations
- University of Texas at Austin