Design of a Hypersonic Boundary Layer Transition Control Experiment Utilizing a Swept Fin Cone Geometry in Mach 6 Flow

Abstract

An experiment has been designed to study the boundary layer stability of a 7°half-angleright circular cone coupled with a 70°swept fin in hypersonic flow. The experiments are tobe conducted in the NASA Langley 20-Inch Mach 6 air tunnel. The hypersonic blowdowntunnel is capable of producing Mach 6 flow at unit Reynolds numbers up to 29 million/m (8.8million/ft) with run times on the order of minutes. The effectiveness of flow control techniqueson instabilities known to cause boundary layer transition will be tested. For this particularexperimental set up, there exist two major mechanisms of boundary layer transition. Thesemechanisms are the crossflow and second mode instabilities. Crossflow transition control willinvolve strategically placed surface roughness located on the fin portion of the model. Secondmode instability control will involve the use of varied nose tip bluntness. Infrared thermographywill provide a global view of the boundary layer development. Detailed spatial measurementswithin the boundary layer are made with a high-frequency off-wall total pressure probe. Resultswill be compared to a complimentary computational study performed by the ComputationalStability and Transition (CST) Lab at Texas A&M University. Additional experiments will beconducted at the University of Notre Dame AFOSR-ND Large Mach 6 Quiet Tunnel. Theseexperiments will involve the current geometry as well as a considerably larger model and willallow for the investigation of tunnel conditions and model size on boundary layer transitionand control.