SCALING EFFECTS ON SHOCK-BOUNDARY LAYER INTERACTION IN TRANSONIC FLOW.
Abstract
Experiments were conducted to determine the effects of unit Reynolds number and boundary layer condition on shock-boundary layer interaction and the subsequent flow development for simulated airfoil contours. Tests were made with three circular-arc, half-thickness models mounted on the tunnel floor. The models had thickness ratios of 6, 8, and 10 percent. Boundary layer bleed was provided upstream of the models to eliminate the thick test-section wall boundary layer. Sandpaper wedges at three chordwise locations and the natural transition case were investigated for each model. The freestream Mach number range was varied from 0.70 to 1.2; the Reynolds number, based on model chord, ranged from 2,000,000 to 17,000,000. The pressure distribution and boundary layer profiles at six chordwise stations were measured throughout the above-mentioned test range on each model. Schlieren photographs were obtained to supplement the pressure measurements. Prior to the shock-boundary layer interaction measurements, tests were conducted on a flat plate with five different boundary layer trips and natural transition. Here, boundary layer profile measurements permit a comparison of the way in which the mode of transition affects various characteristics of the boundary layer. (Author)
Document Details
- Document Type
- Technical Report
- Publication Date
- Mar 01, 1968
- Accession Number
- AD0830030
Entities
People
- E. Stanewsky
- J. G. Hicks
Organizations
- Lockheed Martin