Transition of Hypersonic Boundary Layer

Abstract

An investigation was conducted in a hypersonic shock tunnel to study the boundary layer on a 10° cone of 4 ft length having a wall to stagnation temperature ratio of 0.214. An impact pressure survey of the various types of boundary layer flows was made at 34.5 in. from the cone tip. The free-stream conditions at this location for Mach number and stagnation temperature were approximately 10° and 1400°K. The tunnel operating conditions and cone tip configurations were combined such that laminar, transition, and turbulent flows were obtained at this location. The impact pressure data was reduced to velocity, temperature, and density profiles for comparison. Surface heat transfer results were also obtained at this location along with schlieren photographs of the boundary layers. All three types of information showed that the various boundary layer flows actually existed. The heat gauge showed the passage of turbulent bursts in the transition flow, as observed on a flat plate at subsonic velocities. From the velocity profiles, the local skin friction coefficient was determined for the laminar and turbulent flows, and it agreed reasonably well with the corresponding theories. For the smooth cone, it was not possible to develop fully turbulent flow at the 34.5-in. location for any of the present stagnation conditions but with tip roughness, the transition Reynolds number was sufficiently reduced to force this location into completely turbulent flow.