Hypervelocity dissociating flow over a spherically blunted cone

Recently several hypersonic vehicles are being developed in several countries. For the design of these vehicles, understanding the flow physics is necessary. Recently, the free piston driver for large shock tunnels became practical and it enables us to simulate the hypervelocity flow in the ground based facilities. Also the computing resources have grown dramatically and it enables us to compute the hypervelocity flow which is chemically and thermally nonequilibrium in a reasonable computation time. In this thesis the combined approach of experiment and computation has been applied to the hypervelocity flow on a spherically blunted cone. The experiments are conducted in the newly developed free piston shock tunnel called T5 at the Graduate Aeronautical Laboratories, California Institute of Technology. Three kinds of the gases, nitrogen, air and carbon dioxide are used. The flow fields are computed by a CFD code using the two temperature model by Park. Since the flow field in the experiments is visualized with the differential interferogram, the computed density field is used to generate the differential interferogram. It can be concluded that the two temperature model CFD code can reproduce the basic flow feature such as the inflection point in the shock wave. Heat transfer at the stagnation point is then examined. It correlates well with the equation by Fay and Riddell. Subsequently, the after body heat flux can be predicted by Lees' theory very well. The heat flux on after body is well correlated with Stanton number and local Reynolds number for each gas but the difference between the gases are significant. This is partly because the recombination plays a more important role in the after body flow. The results of the experiments and the computations points to the necessity of other correlation parameters for the after body heat transfer in hypervelocity flows. Then the difference between the shock tunnel experiment and actual flight was examined. The most dominant factor is the difference of free stream temperature. One method to estimate the heat flux in actual flight from experimental data was proposed and this method compensates the difference of the temperature. The result shows very good agreement with numerical computational results