The next generation air breathing engine that would enable an alternative to conventional air travel or suborbital flight is the supersonic ramjet engine (SCRAM) which requires a highly accelerated rate of fuel mixing and combustion. For this type of engine, one of the improvement opportunities lies in the injection of the fuel into the supersonic flow inside the combustor. In order to determine the flow field that develops when a sonic jet of secondary gas is injected into a supersonic free stream, a dedicated test section was developed which was attached to the existing supersonic wind tunnel at the University of Central Florida. A Schlieren system was developed to visualize the associated flow fields and to provide a tool for future high-speed aerodynamic research. Mixing of a cross flow jet with a supersonic free stream is a topic of interest in the field of hypersonic air travel. In addition to a description and analysis of the resultant data, this thesis also serves as a reference for future research and work done with the UCF supersonic wind tunnel and the Schlieren system. The results agree with the literature, and the Schlieren images obtained show the associated shock and flow structures that are expected with a jet in cross flow.

The effect of superheat on the disruption of liquid droplets in a compressible gas flow was investigated experimentally in a small-scale, supersonic wind tunnel. Aerodynamically generated ethanol droplets of an average diameter of 0.1 mm were injected via a normal sonic jet into a Mach 1.8 freestream. Both nonsuperheated and superheated droplets were injected with initial Weber numbers of approximately 700. The droplets and flow structure were photographed using the shadowgraph method. The relatively high momentum of the liquid droplets typically caused them to pass out of the sonic jet structure. Nonsuperheated droplets showed no signs of disrupting after traveling over 200 mm downstream from the injection point. Only droplets with injection temperatures above the predicted boiling point at tunnel freestream static pressure (48°C) showed signs of disruption, typically after they left the sonic jet structure (30 to 100 mm downstream of the injection point). Droplets in this range of temperatures appeared to begin to boil from the downstream side of the droplet, shedding a vapor cloud before disrupting completely in the chaotic mode. Droplets with temperatures above the boiling point at the exit plane of the sonic jet began to disrupt in the chaotic mode almost instantly (within 1 exit nozzle diameter).

Minimizing numerical dissipation without compromising the robust shock-capturing attributes remains an outstanding challenge in the design of numerical methods for high-speed compressible flows. The conflicting requirements of low and high numerical dissipation for accurate resolution of discontinuous and smooth flow features, respectively, are the principal reason behind this challenge. In this work we pursue a recently proposed novel strategy of combining adaptive mesh redistribution with conservative high-order shock-capturing finite-volume discretization methodology to overcome this challenge. In essence, we perform high-order finite-volume WENO (weighted essentially non oscillatory) reconstruction on a continuously moving grid the nodes of which are repositioned adaptively in such a way that maximum spatial resolution is achieved in regions associated with sharpest flow gradients. Moreover, to reduce computational expense, the finite-volume WENO discretization strategy is combined with the midpoint quadrature so that only one reconstruction along each intercool location is necessary. To estimate a monotone upwind flux, a rotated HLLC (Harten-Lax-vanLeer-contact resolving) Riemann solver is employed at each intercool location with the state variables estimated from the high-order WENO reconstruction procedure. The effectiveness of this adaptive high-order discretization methodology is assessed on the well-known double Mach reflection test case for reconstruction orders ranging from five to eleven. We find that the resolution of the intricate flow features such as the wall-jet improves progressively with the reconstruction order, which is indicative of the reduced dissipation level of the adaptive high-order WENO discretization. The adaptive discretization methodology is applied to simulate a flow configuration consisting of a Mach 3 supersonic jet injected in a Mach 2 supersonic crossflow of similar ideal gas. It is found that the flow characteristics and especially features that are formed as a result of the Kelvin-Helmholtz instability are strongly influenced by the reconstruction order. The influence of the jet inclination angle on the overall flow features is analyzed.

Lors d’un vol hypersonique (Mach 6, 20 km d’altitude) la couche limite se développant sur l’avant-corps d’un véhicule hypersonique est laminaire. Cet état cause un désamorçage du moteur (statoréacteur) assurant la propulsion du véhicule. Pour pallier ce problème, il faut forcer la transition de la couche limite á l’aide d’un dispositif de contrôle dont l’effet est permanent (passif) ou modulable (actif) pendant le vol. Dans ce travail, nous analysons l’efficacité d’un dispositif actif d’injection d’air á la paroi pour forcer la transition de la couche limite sur un avant-corps générique. L’interaction jet d’air/couche limite est simulée numériquement avec une approche aux grandes échelles (LES). Une étude paramétrique sur la pression d’injection permet de quantifier l’efficacité du jet á déstabiliser la couche limite. L’influence des conditions de vol (altitude, Mach) sur la transition est également étudiée. Une analyse des résultats de simulation par Décomposition en Modes Dynamiques (DMD) est menée pour comprendre quels sont les modes dynamiques responsables de la transition et les mécanismes sous-jacents. Des essais dans la soufflerie silencieuse de l’université de Purdue (BAM6QT) ont été effectués pour tester expérimentalement l’efficacité des dispositifs passifs (rugosité isolée en forme de losange) et actifs (mono-injection d’air) pour faire transitionner la couche limite. Une peinture thermo-sensible et des capteurs de pression (PCB, Kulite) ont été utilisés pour déterminer la nature de la couche limite. Les résultats de ce travail montrent qu’une injection sonique suffit pour forcer la couche limite. On observe des essais, que pour une même hauteur de pénétration, les rugosités isolées sont moins efficaces que les jets (mono injection) pour déstabiliser la couche limite. / During a hypersonic flight (Mach 6, 20 km altitude), the boundary layer developing on the forebody of a vehicle is laminar. This state may destabilize the scramjet engine propelling the vehicle. To overcome this problem during the flight, the boundary layer transition has to be forced using a control device whose effect is fixed (passive) or adjustable (active). In this work, we analyze the efficiency of a jet in crossflow in forcing the boundary layer transition on a generic forebody. The flow is computed with a Large Eddy Simulations (LES) approach. A parametric study of the injection pressure allows the efficiency of the jet in tripping the boundary layer to be quantified. The influence of flight conditions (Mach, altitude) on the transition is also studied. Dynamic Mode Decomposition (DMD) is applied to the simulation results to determine the transition leading to dynamic modes and to understand underlying transition mechanisms. Experiments in the Purdue University quiet wind tunnel (BAM6QT) were performed to quantify the efficiency of a passive transition device (diamond roughnesses) and an active transition device (single air jet) in tripping the boundary layer. A thermo-sensitive paint and pressure transducers (Kulite, PCB) were used to determine the state of the boundary layer on the generic forebody. Experimental and numerical results show a sonic injection is sufficient to induce transition. We observe from the experiments that for the same penetration height, a single roughness is less efficient than a single air jet in destabilizing the boundary layer.