Effect of Cooling on Hypersonic Boundary-Layer Stability

dc.contributor.advisorLaurence, Stuart Jen_US
dc.contributor.authorPaquin, Lauraen_US
dc.contributor.departmentAerospace Engineeringen_US
dc.contributor.publisherDigital Repository at the University of Marylanden_US
dc.contributor.publisherUniversity of Maryland (College Park, Md.)en_US
dc.description.abstractThe prediction of boundary-layer transition on hypersonic vehicles has long been considered a primary design concern due to extreme levels of heating and dynamic pressure loading this transition induces. While it has been predicted that the temperature gradient between the vehicle and the local freestream can drastically alter boundary-layer stability, experimental research on the topic over the past fifty years has provided conflicting results. This study investigates the relationship between the wall-to-edge temperature ratio and boundary-layer stability on a slender cone. Campaigns in two wind-tunnel facilities were conducted: one set within the HyperTERP reflected-shock tunnel at the University of Maryland, and one set at the high-enthalpy T5 reflected-shock tunnel at the California Institute of Technology. Both sets of campaigns employed non-intrusive, optical diagnostics to analyze the structures and spectral content within the boundary layer. In the first part of the study, performed in HyperTERP, an experimental methodology was developed to vary the wall temperature of the model using active cooling and passive thermal management. This allowed the wall temperature ratio to be varied at the same nominal test condition (and thus freestream disturbance environment), and three thermal conditions were established for analysis. Simultaneous schlieren and temperature-sensitive-paint (TSP) imaging were performed. Calibrated schlieren images quantified the unsteady density gradients associated with second-mode instabilities, and TSP contours provided insight into the thermal footprint of mean boundary-layer structures. It was found that, overall, cooling shrunk the boundary-layer thickness, increased second-mode disturbance frequencies, and increased the amplification rate of these instabilities. At nonzero angles of attack, cooling appeared to increase the azimuthal extent of flow separation on the leeward side of the cone. In the second part of the study, performed in T5, the disturbance structures and spectral content of laminar and transitional boundary layers were characterized under high-enthalpy conditions. Schlieren images indicated that, at these extremely low wall-to-edge temperature ratios, second-mode waves were confined very close to the wall in the laminar case. During the breakdown to turbulence, structures radiating out of the boundary layer and into the freestream were discovered. A texture-based methodology was used to characterize the Mach angles associated with these structures, and a wall-normal spectral analysis indicated a potential mechanism by which energy was transferred from the near-wall region to the freestream. The study presents some of the first simultaneous imaging of the flow structures and associated thermal footprint of boundary-layer transition within an impulse facility. The work also presents the first time-resolved, full-field visualizations of the second-mode dominated breakdown to turbulence at high enthalpy. Thus, the study imparts significant insight into the mechanics of boundary-layer transition at conditions representative of true hypervelocity flight.en_US
dc.subject.pqcontrolledAerospace engineeringen_US
dc.subject.pquncontrolledboundary layeren_US
dc.subject.pquncontrolledfluid dynamicsen_US
dc.titleEffect of Cooling on Hypersonic Boundary-Layer Stabilityen_US


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