Experiments to generate multiple shock waves in an axisymmetric model at hypersonic speeds were conducted in a small reflected shock tunnel. Conical surfaces were used to generate shock waves inside a circular duct chosen to be representative of a scramjet combustor. These shock waves impinged on turbulent boundary layers to produce shock wave/boundary layer interactions (SWBLIs). In the process of observing this phenomenon, the commonly used empirical correlations of Korkegi were tested for accuracy, i.e. the combined pressure ratio across these shocks can be measured and compared to that predicted by these correlations. Korkegi correlates only with Mach number, and is independent of Reynolds number and on how the pressure is applied. A major contribution of this work is to examine how the details of the compression process effect separation. In this study, the history of applying the compression was varied. An analytical method was developed for theoretically estimating the onset of incipient separation using an integrated computation of the momentum flux contained in the boundary layer. By including the summed (negative) contribution of wall shear stress on the integrated momentum flux, the upstream history of the boundary layer was considered. The overall result has a form similar to the Korkegi correlations, plus an additional correction term relating to momentum loss through wall shear stress. The correction term was determined to be a second order effect, which explains why the Reynolds number independent Korkegi correlations work so well over such a large range of conditions. A hypersonic flow test condition conducive to the generation of high Reynolds number flows and turbulent boundary layer production was developed in a small reflected shock tunnel. The experimentally measured flow parameters were matched by numerical simulation using a number of in-house codes at The University of Queensland. This has allowed the unmeasured parameters which are numerically derived to be stated with greater confidence. An internal centre-body with a conical forebody was used to generate conditions of incipient separation. This provided benchmark data for comparison with subsequent experiments with multiple compressions. A semi-vertex angle of 15o was selected based on Large Eddy Simulation (LES) numerical results once the experimental and numerical static wall pressure and heat flux were matched. A two-cone experimental model, which provided for adjustment of the axial separation between the two shock systems, was tested at the same flow conditions as used in the single-cone experiments. A technique of incrementally moving the instrumentation (relative to the centre-body) and repeating the same condition to achieve high resolution in pressure and heat flux distributions with a limited number of transducers was successful. The results verified that it was possible to subject a hypersonic turbulent boundary layer to two quantified compression-expansion systems with an adjustable axial separation between them and capture the first reflected shock in a “shock trap” to remove it's influence from the second SWBLI. The data from this initial two-cone model provided non-separated pressure and heat flux data which was used as a reference to help interpret data from separated flows. The commercially available Reynolds Averaged Navier-Stokes (RANS) numerical code, CFD-Fastran, was used to help design an experimental model which produces boundary layer separation. Algebraic and two-equation turbulence models were applied to a modified two-cone model to show greater pressure rises which would produce boundary layer separation. A modified two-cone model was tested and demonstrated boundary layer separation. Three configurations with varying axial separation between SWBLIs were tested which all produced separation. The configuration that produced the largest pressure ratio and largest separation region at the second SWBLI may represent a geometry whereby the distance from the hollow cylinder inlet and the second cone may represent a critical value. The amount of viscous interaction, generated from the leading edge of the shock trap, and the proximity of the two interactions may be coupled to produce higher than expected values. It is postulated that the boundary layer momentum recovery for the configuration where the second SWBLI was furthest downstream (30 mm configuration), prevented severe separation from occurring. An in-house RANS code, elmer3, was used to simulate the flow of the modified two-cone model. An algebraic turbulence model was applied to this model and comparisons of experimentally measured static wall pressure and heat flux have given good agreement. The wall shear stress was investigated to provide further information concerning the position and size of flow reversal regions. The use of the numerical codes utilised in this study has reinforced their effectiveness for model design and comparison of experimental results.

The main objective of the thesis is to provide a detailed assessment of the performance of four types of Low Reynolds Number (LRN) Eddy Viscosity Models (EVM), widely used for industrial purposes, on flows featuring SWBLI, using experimental and direct numerical simulation data. Within this framework the two-equation linear k-ε of Launder and Sharma (1974) (LS), the two-equation linear k-ω SST, the four-equation linear φ-f of Laurence et al. (2004) (PHIF) and the non-linear k-ε scheme of Craft et al. (1996b,1999) (CLSa,b) have been selected for testing. As initial test cases supersonic 2D compression ramps and impinging shocks of different angles and Reynolds numbers of the incoming boundary layer have been selected. Additional test cases are then considered, including normal shock/isotropic turbulence interaction and an axisymmetric transonic bump, in order to examine the predictions of the selected models on a range of Mach numbers and shock structures. For the purposes of this study the PHIF and CLSa,b models have been implemented in the open source CFD package OpenFOAM. Some results from validation studies of these models are presented, and some explorations are reported of certain modelled source terms in the ε-equation of the PHIF and CLSb models in compressible flows. Finally, before considering the main applications of the study, an examination is made of the performance of different solvers and numerical methods available in OpenFOAM for handling compressible flows with shocks. The performance of the above models, is analysed with comparisons of wall-quantities (skin-friction and wall-pressure), velocity profiles and profiles of turbulent quantities (turbulent kinetic energy and Reynolds stresses) in locations throughout the SWBLI zones. All the selected models demonstrate a broadly consistent performance over the considered flow configurations, with the CLSb scheme generally giving some improvements in predictions over the other models. The role of Reynolds stress anisotropy in giving a better representation of the evolution of the boundary layer in these flows is discussed through the performance of the CLSb model. It is concluded that some of the main deficiencies of the selected models is the overestimation of the dissipation rate levels in the non-equilibrium regions of the flow and the underestimation of the amplification of Reynolds stress anisotropy, especially within the recirculation bubble of the flows. Additionally, the analysis of the performance of the considered EVM's in a normal shock/isotropic turbulence interaction illustrates some drawbacks of the EVM formulation similar to the ones observed in normally-strained incompressible flows. Finally, a hybrid Detached Eddy Simulation (DES) approach is incorporated for the prediction of the transonic buffet around a wing.

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The mean flowfield and time-dependent characteristics of a Mach 5 cylinder-induced transitional shock-wave/boundary-layer interaction have been studied experimentally. The objectives of the study were to: (i) provide a detailed description of the mean flow structure of the interaction, and (ii) characterize the unsteadiness of the interaction based on fluctuating pressure measurements. / text

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

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<p>The compressible turbulent boundary layer developing over atwo-dimensional bump which leads to a supersonic pocket with aterminating shock wave has been studied. The measurements havebeen made with hot-wire anemometry and Particle ImageVelocimetry (PIV).</p><p>A method to calibrate hot-wire probes in compressible ow hasbeen developed which take into account not only the ow velocitybut also the inuence of the Mach number, stagnation temperatureand uid density. The calibration unit consists of a small jetow facility, where the temperature can be varied. The hot wiresare calibrated in the potential core of the free jet. The jetemanates in a container where the static pressure can becontrolled, and thereby the gas density. The calibration methodwas verfied in the at plate zero pressure gradient turbulentboundary layer in front of the bump at three different Machnumbers, namely 0.3, 0.5 and 0.7. The profiles were alsomeasured at different static pressures in order to see theinuence of varying density. Good agreement between the profilesmeasured at different pressures, as well as with the standardlogarithmic profile was obtained.</p><p>The PIV measurements of the boundary layer ow in front ofthe 2D bump showed good agreement with the velocity profilesmeasured with hotwire anemometry. The shock wave boundary layerinteraction was investigated for an inlet Mach number of 0.69.A lambda shock wave was seen on the downstream side of thebump. The velocity on both sides of the shock wave as measuredwith the PIV was in good agreement with theory. The shock wavewas found to cause boundary layer separation, which was seen asa rapid growth of the boundary layer thickness downstream theshock. However, no back ow was seen in the PIV-data, probablybecause the seeding did not give enough particles in theseparated region. The PIV data also showed that the shock wavewas oscillating, i.e. it was moving approximately 5 mm back andforth. This distance corresponds to about five boundary layerthicknesses in terms of the boundary layer upstream theshock.</p><p><b>Descriptors:</b>Fluid mechanics, compressible ow,turbulence, boundary layer, hot-wire anemometry, PIV, shockwave boundary layer interaction, shape factor.</p>

The compressible turbulent boundary layer developing over atwo-dimensional bump which leads to a supersonic pocket with aterminating shock wave has been studied. The measurements havebeen made with hot-wire anemometry and Particle ImageVelocimetry (PIV). A method to calibrate hot-wire probes in compressible ow hasbeen developed which take into account not only the ow velocitybut also the inuence of the Mach number, stagnation temperatureand uid density. The calibration unit consists of a small jetow facility, where the temperature can be varied. The hot wiresare calibrated in the potential core of the free jet. The jetemanates in a container where the static pressure can becontrolled, and thereby the gas density. The calibration methodwas verfied in the at plate zero pressure gradient turbulentboundary layer in front of the bump at three different Machnumbers, namely 0.3, 0.5 and 0.7. The profiles were alsomeasured at different static pressures in order to see theinuence of varying density. Good agreement between the profilesmeasured at different pressures, as well as with the standardlogarithmic profile was obtained. The PIV measurements of the boundary layer ow in front ofthe 2D bump showed good agreement with the velocity profilesmeasured with hotwire anemometry. The shock wave boundary layerinteraction was investigated for an inlet Mach number of 0.69.A lambda shock wave was seen on the downstream side of thebump. The velocity on both sides of the shock wave as measuredwith the PIV was in good agreement with theory. The shock wavewas found to cause boundary layer separation, which was seen asa rapid growth of the boundary layer thickness downstream theshock. However, no back ow was seen in the PIV-data, probablybecause the seeding did not give enough particles in theseparated region. The PIV data also showed that the shock wavewas oscillating, i.e. it was moving approximately 5 mm back andforth. This distance corresponds to about five boundary layerthicknesses in terms of the boundary layer upstream theshock. <b>Descriptors:</b>Fluid mechanics, compressible ow,turbulence, boundary layer, hot-wire anemometry, PIV, shockwave boundary layer interaction, shape factor. / NR 20140805