Boundary-layer transition on sharp cones at hypersonic Mach numbers.

TLDR: In this paper, an experimental investigation to locate the beginning of transition from laminar to turbulent boundary-layer flow has been conducted at zero angle of attack on sharp, smooth cones having semiapex angles of 2.87°, 5°, and 10° in the contoured nozzle of the Langley 22-in. helium tunnel at a freestream Mach number of about 20.8.

Abstract: An experimental investigation to locate the beginning of transition from laminar to turbulent boundary-layer flow has been conducted at zero angle of attack on sharp, smooth cones having semiapex angles of 2.87°, 5°, and 10° in the contoured nozzle of the Langley 22-in. helium tunnel at a freestream Mach number of about 20. Local Mach number at the boundarylayer edge was thus varied from 7.4 to 16.6. The data indicate that local transition Reynolds number increases rapidly with local Mach number. Techniques used to detect onset of transition included surface pilot tube, drag force, boundary-layer pitot-pressure surveys, schlieren photographs, and hot-film anemometer measurements. Skin-friction coefficient, displacement thickness, momentum thickness, and velocity ratio profiles were determined for laminar, transitional, and turbulent hypersonic boundary layers. A hot-film anemometer survey of the model boundary layers showed disturbances originating within the boundary layer at much lower Reynolds numbers than the Reynolds number for which transition is felt at the model surface. The maximum disturbance level occurred at a location corresponding to about 0.845 (boundary-layer thickness) with the disturbance speed being subsonic relative to the local edge velocity. In addition, source flow effects on transition Reynolds number were examined at a local Mach number of 15.8.