Transitional hypersonic flow over slender cone/flare geometries
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Experiments are performed in a Mach-6 shock tunnel to examine the laminar-to-turbulent transition process associated with a sudden increase in surface angle on a slender body. A cone/flare geometry with a 5◦ frustum and compression angles ranging from 5◦ to 15◦ allow a range of mean flow configurations, spanning an attached shock-wave/boundary-layer interaction to a fully separated one; the unit Reynolds number of the flow is also varied to modify the state of incoming second-mode boundary-layer disturbances. Ultra-high-speed schlieren visualizations provide a global picture of the flow development, supplemented by high-frequency surface pressure measurements. For the 5◦ compression, the unsteady flow field is dominated by the second-mode waves, whose breakdown to turbulence is generally accelerated (compared with the straight-cone configuration) by encountering the angle change. As the compression angle is increased to induce separation, lower-frequency disturbances appear along the separated shear layer that exhibit much larger amplification rates than the incoming second-mode waves; the latter effectively freeze in amplitude downstream of the separation point before rapidly breaking down upon reattachment. The shear-layer disturbances become dominant at the largest compression angle tested. Radiation of disturbance energy to the external flow is consistently observed: this generally occurs along mean flow features (flare, separation or reattachment shocks) for the second-mode disturbances and spontaneously for the shear-layer waves. The combined application of spectral proper orthogonal decomposition and a global bispectral analysis allows the identification of important unsteady flow structures and the association of these with prominent nonlinear interactions in the various configurations.