Accession Number : AD0731830

Title :   Boundary Layers on Airfoils in Transonic Flow and the Control of Shock-Induced Separation.

Descriptive Note : Final rept. 1 Jun 67-30 Dec 70,

Corporate Author : OHIO STATE UNIV RESEARCH FOUNDATION COLUMBUS

Personal Author(s) : Lee,John D.

Report Date : AUG 1971

Pagination or Media Count : 72

Abstract : The transonic terminal shock wave is always oblique (rather than normal) at the surface and the associated deflection of the boundary layer corresponds to the maximum deflection permitted by the Mach number upstream of the shock. When the Mach number is near 1 the deflection is small, and the separated boundary layer will usually reattach leaving a bubble separation. At higher Mach numbers the separation becomes more severe and reattachment may be affected only by extremes in boundary layer controls, e.g., streamwise blowing and vortex generators were found to be useful. In many cases, a controlled attachment simply results in a postponement of the separation to a higher Mach number with an increased deflection and higher drag. The study was performed with the OSU 12-inch transonic wind tunnel using 6-inch chord airfoil models in the Mach number range from 0.4 to 0.9. Data taken were in the form of surface static pressures, wake pitot pressures, force balance outputs and schlieren photographs.

Descriptors :   (*AIRFOILS, TRANSONIC CHARACTERISTICS), (*FLOW SEPARATION, CONTROL), BOUNDARY LAYER CONTROL, SHOCK WAVES, DEFLECTION, WIND TUNNEL MODELS, MODEL TESTS, FLOW VISUALIZATION, DRAG, AREA SUCTION, BLOWERS, GEOMETRIC FORMS, SPOILERS, VORTEX GENERATORS

Subject Categories : Aerodynamics
      Fluid Mechanics

Distribution Statement : APPROVED FOR PUBLIC RELEASE