Pulsed plasma jets (also called "SparkJets'") were investigated for use in controlling the mean separation location induced by shock wave-boundary layer interaction. These synthetic jet actuators are driven by electro-thermal heating from an electrical discharge in a small cavity, which forces the gas in the cavity to exit through a small hole as a high-speed jet. With this method of actuation, pulsed plasma jets can achieve pulsing frequencies on the order of kilohertz, which is on the order of the instability frequency of many lab-scale shock wave-boundary layer interactions (SWBLI). The interaction under investigation was generated by a 20° compression ramp in a Mach 3 flow. The undisturbed boundary layer is transitional with Re[subscript theta] of 5400. Surface oil streak visualization is used in a parametric study to determine the optimum pulsing frequency of the jet, the optimum distance of the jet from the compression corner, and the optimum injection angle of the jets. Three spanwise-oriented arrays of three plasma jets are tested, each with a different pitch and skew angle on the jet exit port. The three injection angles tested were 22° pitch and 45° skew, 20° pitch and 0° skew, and 45° pitch and 0° skew. Jet pulsing frequency is varied between 2 kHz and 4 kHz, corresponding to a Strouhal number based on separation length of 0.012 and 0.023. Particle image velocimetry is used to characterize the effect that the actuators have on the reattached boundary layer profile on the ramp surface. Results show that plasma jets pitched at 20° from the wall, and pulsed at a Strouhal number of 0.018, can reduce the size of an approximate measure of the separation region by up to 40% and increase the integrated momentum in the downstream reattached boundary layer, albeit with a concomitant increase in the shape factor. / text

Large heat transfer rates are a major source of possible failure in flight vehicles due to increases in temperature being linked to weakening material properties. Aircraft in high-Mach number flow generate excessive aerodynamic heat that may increase temperatures above limits of structural integrity. Even without reducing speed or changing material, it is possible to mitigate heat transfer by altering vehicle geometry. The purpose of this thesis is to study the extent of heat transfer in gap regions of various sizes by computationally simulating flow over an idealized missile-fin configuration. Maximum levels of heat transfer are analyzed as well as surface distributions that identify key design points. The Department of Defense software package with computational fluid dynamics capabilities, Kestrel, was employed to use the Reynolds-averaged Navier-Stokes equations to simulate turbulent Mach~6 flow over the missile model. Results are compared to data obtained by the Air Force Research Laboratory via wind tunnel tests of the same flow. Experiments and simulations both found an order of magnitude increase in heat transfer when an offset fin was attached, but this heating could be reduced by minimizing the offset distance. Simulated baseline properties agreed very well with experimental measurements and simulations of the gap region more precisely identified the locations of maximum heating.

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Une interaction entre une onde de choc oblique et une couche limite turbulente sur plaque plane, à un nombre de Mach de 2.3 a été étudiée expérimentalement.<br />Ces interactions, pour des ondes de choc assez fortes, engendrent le décollement et le recollement de la couche limite, et sont le siège d'instationnarités basses fréquences dont les origines sont mal connues. Ces instationnarités ont été caractérisées expérimentalement en partie dans des travaux précédents, et des similarités entre l'interaction étudiée ici et d'autres configurations d'interactions ainsi qu'avec les décollements de couche limite subsonique ont été mis en évidence, suggérant que les mécanismes responsables des instationnarités sont de même nature.<br />Pour ces travaux, la Vélocimétrie par Imagerie de Particules (PIV) a été utilisée afin de décrire spatialement l'organisation longitudinale et transversale de cette interaction. L'exploitation des mesures a mis en évidence un lien statistique fort entre les mouvements basses fréquences du choc réfléchi et les contractions/dilatations successives du bulbe décollé. L'interprétation proposée est que les grands mouvements du choc sont liés aux pulsations basses fréquences du décollement, associées à sa réalimentation intermittente en air frais.<br />Un modèle aérodynamique en a été déduit et permet de préciser les principaux paramètres contrôlant l'échelle de temps du phénomène. En particulier, il permet de déterminer la fréquence des battements du choc. Ce modèle a été appliqué aux interactions sur plaques planes ainsi que pour d'autres configurations expérimentales, pour un éventail de nombres de Mach allant de M=0 à 5, et montre un bon accord avec les mesures.

<p dir="ltr">This thesis concentrates on the improvement of the quality of shock-dominated flows in supersonic inlets by controlling shock wave / boundary layer interactions (SWBLIs). SWBLI flow control has been a major issue relevant to scramjet-associated endeavors for many years. The ultimate goal of this study is to numerically investigate SWBLI flow control through the application of steady-state thermal sources --- which were defined to replicate the Joule heating effect produced by Quasi-DC electric discharges --- and compare the results with data obtained from previous experiments.</p><p dir="ltr">Numerical solutions were obtained using both a three-dimensional, unsteady Reynolds-averaged Navier-Stokes (RANS) solver with a Spalart-Allmaras (SA) Detached Eddy Simulation (DES) turbulence modeling method and also a simple three-dimensional, compressible RANS solver with a SA turbulence model. Computations employed an ideal gas thermodynamic model. The numerical code is Stanford University Unstructured (SU2), an open-source, unstructured grid, computational fluid dynamics code. The SU2 code was modified to include volumetric thermal source terms to represent the Joule heating effect of electric current flowing through the gas. The computational domain, source term configuration, and flow conditions were defined in accordance with experiments carried out at the University of Notre Dame. Mach 2 flow enters the three-dimensional test domain with a stagnation pressure of 1.7 bar. The test domain is contained by four isothermal side walls maintained at room temperature, as well as an inlet and outlet. A shock wave (SW) generator, a symmetric 10 degree wedge, is positioned on the upper surface of the test domain. The overall length of the test sections is 910 mm and inlet length of the computational domain is increased prior to the location of shock wave generator in order to allow for adequate boundary layer growth. Volumetric heating source terms were positioned on the lower surface of the test domain in the reflected SW region.</p><p dir="ltr">Experimental results show that the thermal sources create a new shock train within the duct and do not initiate significant additional pressure losses. What remains to be explored is the overall characterization of the 3D flow features and dynamics of the thermally induced SW and the effect of gas heating on total pressure losses in the test section.</p><p dir="ltr">Numerical solutions validate what is observed experimentally, and offer the ability to gather more temporally and spatially-resolved measurements to better understand and characterize shock-dominated flow control in a supersonic inlet or duct. Although thermally driven SWBLI flow control requires additional research, this study alleviates the dependency on experimentally driven data and adds insight into the nature of the complex unsteady, three-dimensional flowfield.</p>