Flow Separation in Shock Wave Boundary Layer Interactions at Hypersonic Speeds

INTRODUCTIONShock boundary layer interactions are usually the cause of flowseparation at hypersonic and supersonic speeds. In addition to increasingthe drag and the aerodynamic heating on the aircrafts and missiles, theyadversely affect the inlet and control surface performance. The conditionsleading to flow separation and the subsequent changes in the flow have beenthe subject of extensive research efforts over the past 30 years. Theresults of these investigations have been compiled and discussed in severalexcellent reviews [i-4]. However, in spite of these efforts, the combinedinfluence of viscous-inviscid interactions, turbulence, and compressibilityon the resulting complex flow fields is not fully understood at high speeds.Owen [5] discussed the problems associated with the experimentalmeasurements of flows at hypersonic speeds, and the inadequacy of existingexperimental techniques in shock boundary interactions involving separation,time dependent flow reversal, and high levels of turbulence.The focus of the present work is the assessment of the experimentaldata on separated flow in shock wave turbulent boundary layer interactionsat hypersonic speeds. The data base will consist mainly of two dimensionaland axisymmetric interactions of shock wave and turbulent boundary layer.Only two configurations will be considered, namely compression corner orcylinder-flare, and externally generated oblique shock interactions withboundary layers over flat plates or cylindrical surfaces. Interactions inforward or backward steps will not be included. This choice is based on theabsence of any characteristic length, with the exception of the incomingboundary layer thickness, in the two configurations shown in Fig. i.This work was supported under NASA Contract NASI-18458, Task 19, by theTheoretical Flow Physics Branch, Fluid Mechanics Division, with The GeorgeWashington University.