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.

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

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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

<p>This research focuses on experiments on compression-induced shock wave/boundary-layer interactions conducted in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) at Purdue University. The BAM6QT facilitates a low-freestream-noise hypersonic test environment more similar to that experienced in flight than a conventional wind tunnel. Measurements were captured on two sliced 7° half-angle cones with finite-span compression ramps. On the first, the slice was cut parallel to the axis of the cone to build upon previous measurements in hypersonic flow. While similar geometries have been analyzed for over 30 years in experiment and computation, there are significant gaps in understanding of the underlying mechanisms leading to instability and transition on the ramp. Further, in low-noise Mach 6 flow, the boundary layer separated at the leading edge of the slice, which is unlikely to occur on a real flight vehicle. Thus, on the second model, the slice was cut at a 4° incline to the</p> <p>cone axis to facilitate the growth of an attached laminar boundary layer on the slice. Using this configuration, the ramp-induced boundary-layer thickening initiated between the slice leading edge and the ramp leading edge, allowing the investigation of a ‘naturally’ formed separated region. </p> <p><br></p> <p>Data were captured at angles of attack ranging from 0° to 6°, on compression ramp angles ranging from 10° to 20°, and for freestream Reynolds numbers of 2.5×10^6/m to 12×10^6/m. To analyze the mean-flow behavior of the separation bubble as it changes with the above parametrics, time-averaged schlieren visualization was used to provide off-surface visualization of the flowfield, allowing estimates of reattachment position and separation bubble size. In all cases, reattachment position was shown to move upstream with an increase in angle of attack, an increase in ramp angle, and an increase in Reynolds number. However, on the model with the inclined slice, the Reynolds number impacted reattachment location to a much lesser extent. </p> <p><br></p> <p>Heat transfer measurements on the ramp revealed regions with the most significant aerothermal loading. Streamwise streaks of high heating originating at the ramp edges and centerline were observed to increase in magnitude with an increase in Reynolds number, angle of attack, and ramp angle. On the model with the inclined slice, many streaks of high heating were observed that increased in quantity and magnitude with angle of attack and ramp angle. Root mean squared pressure fluctuations computed from surface pressure measurements were shown to follow similar trends to centerline heat transfer results for both models. Angle of attack, ramp angle, and slice angle are shown to play a dominant role in transition. Finally, the importance of quiet tunnels is made remarkably clear, as the BAM6QT operating in its conventional-noise configuration resulted in drastically different results.</p> <p><br></p> <p>For measurement of shock wave/boundary-layer instabilities, schlieren frames were captured at 100,000 fps to allow measurement of low-to-mid-frequency fluctuations of the recirculation zone edge. Shear layer flapping frequencies were found to occur at around 1100–1200 Hz, which increased with angle of attack to up to 1600 Hz. It is likely that this is an inherent instability in the separation bubble itself, rather than a function of freestream disturbances, and may be indicative of an ‘expansion and relaxation’ effect known as bubble breathing. Additional measurements using low-frequency-capable pressure sensors must be captured to determine whether this breathing effect manifests on the model slice or ramp. </p>

<p>This dissertation covers a computational investigation into hypersonic flight vehicle geometric imperfections, with a focus on wing-elevon-cove configurations. The primary region of focus for the overall research was the cove region at the juncture of the main wing element and the elevon. This region is associated with the shock-wave/boundary-layer interaction produced by the control surface deflection. There also exists a centrifugal instability at the cove, due to streamline curvature, which is associated with the production of Görtler vortices. The content includes three projects revolving around hypersonic wing-elevon-cove flows. These flows were computed with improved delayed detached-eddy simulation.</p> <p><br></p> <p>The first project was a computational investigation simulating the NASA experimental study done by W.D. Deveikis and W. Bartlett in 1978. This experiment consisted of hypersonic high Reynolds number wind tunnel tests for a shuttle-type reentry vehicle. The computational aerothermodynamic surface loadings for this project were compared to the experimental published data. Grounded with the agreement with mean surface data, this project expanded on the topics explored in the experimental study to include topics such as flow visualization and statistical analysis. The second and third project are extensions of this work and were done in collaboration with Purdue University and the University of Tennessee Space Institute (UTSI). A swept wing-elevon-cove model was designed by Carson Lay, of Purdue University, and is currently being employed in ongoing experiments in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) and at the Tennessee Aerothermodynamics Laboratory (TALon). A computational investigation on hypersonic high Reynolds number wing-elevon-cove flows was conducted with this model, where both corresponding experimental facility conditions were employed. At this time, the experimental data are limited; however, future experimental and computational collaboration is expected.</p> <p><br></p> <p>The motivation behind this research was to expand the knowledge on hypersonic wing-elevon-cove flows, gap heating, and the low-frequency unsteadiness in shock-wave/boundary-layer interactions. Therefore, the intended goal of this work was to provide an accurate characterization of the three hypersonic wing-elevon-cove flows. This was accomplished by using computational data to produce flowfield visualizations, analyze aerothermodynamic loadings, and conduct statistical flow analyses. The results on the three hypersonic wing-elevon-cove computations are presented, analyzed, and discussed throughout this dissertation.</p>

Les écoulements transsoniques rencontrés dans le cadre de la propulsion aéronautique et spatiale sont associés à l'apparition d'ondes de choc. En impactant la couche limite se développant sur une paroi, un gradient de pression adverse est généré qui conduit à l'épaississement ou au décollement de la couche limite. Lors de la vibration de la structure, l'onde de choc oscille et interagit avec la couche limite, générant une fluctuation de la pression statique à la paroi. Il s'ensuit alors un échange d'énergie entre le fluide et la structure qui peut être stabilisant ou au contraire conduire à une instabilité aéroélastique (flottement). La modélisation de la réponse instationnaire de l'interaction onde de choc / couche limite pour l'étude des interactions fluide-structure est l'objet de ce travail de recherche. Il s'appuie sur la résolution des équations de Navier-Stokes moyennées (RANS) et la modélisation de la turbulence. Les méthodes et modèles utilisés ont été validés à partir de résultats expérimentaux issus d'une tuyère transsonique dédiée à l'étude des interactions fluide-structure. Ces travaux sont ensuite appliqués à l'amélioration de la prédiction du flottement en turbomachine. Une méthode linéarisée en temps permettant la résolution des équations RANS dans le domaine fréquentiel est utilisée. Nous confirmons l'importance de la dérivation du modèle de turbulence lors de la prédiction d'une interaction forte entre une onde de choc et une couche limite décollée. Une méthode de régularisation est présentée puis appliquée aux opérateurs non dérivables du modèle de turbulence k-! de Wilcox (2006). La prédiction de la réponse instationnaire de l'interaction onde de choc / couche limite dans une tuyère est évaluée à partir de simulations bidimensionnelles et présente un bon accord avec les données expérimentales. En évaluant l'influence de la fréquence réduite, une instabilité aéroélastique de type flottement transsonique est identifiée. Un dispositif de contrôle, reposant sur la génération d'ondes de pression rétrogrades à l'aval de la tuyère, est proposé puis validé numériquement. Enfin, une méthodologie est proposée pour comprendre les mécanismes aérodynamiques conduisant au flottement. Pour cela, il a été réalisé un dessin provisoire d'une soufflante transsonique à fort taux de dilution. Cette soufflante, l'ECL5, est destinée à l'étude expérimentale des instabilités aérodynamiques et aéroélastiques. La méthodologie proposée repose sur la simulation 2D d'une coupe de tête et met à profit la linéarisation pour analyser la contribution de sources locales en fonction de la fréquence réduite, du diamètre nodal et de la déformée modale / Transonic flows, which are common in aeronautical and spatial propulsion systems, produce shock-waves over solid boundaries. When a shock-wave impacts the boundary layer, an adverse pressure gradient is generated and a thickening or even a separation of the boundary layer is induced. If the solid boundary vibrates, the shock-wave oscillates, interacts with the boundary layer and produce a fluctuation of the static pressure at the wall. This induces an exchange of energy between the fluid and the structure which can be stabilising or lead to an aeroelastic instability (flutter).The main objective of this PhD thesis is the modelling of the unsteady behaviour the simulation of the shock-wave/boundary layer interaction for fluid-structure interaction studies. To this end, simulations have been carried out to solve Reynolds-Averaged Navier-Stokes equations using two equations turbulence model. The method is validated thanks to experimental data obtained on a transonic nozzle dedicated to aeroelastic studies. This method is then use to increase the predictability of flutter events in turbomachinery.A time linearised frequency-domain method is applied to RANS equations. It is shown that the unsteady behaviour of the turbulent boundary-layer contributes to the fluctuating static pressure when the shock-wave boundary layer interaction is strong. Hence, the frozen turbulence assumption is not valid and the turbulence model must be derivated. Thus, the regularisation of the non derivable operators is proposed and applied on k-? Wilcox (2006) turbulence model.The unsteady behaviour of the shock-wave/boundary-layer interaction in a transonic nozzle is evaluated thanks to 2D numerical simulations and shows good agreement with experimental data. When varying the reduced frequency an aeroelastic instability is found, known as transonic flutter. An active control device generating backward travelling pressure waves is then designed and numerically validated.Finally, a methodology is proposed to understand the aerodynamic onsets of transonic flutter. To this end, a preliminary design of a high bypass ratio transonic fan has been carried out. This fan, named ECL5, is dedicated to experimental aerodynamic and aeroelastic studies. The methodology relies on 2D simulations of a tip blade passage and uses linearisation to analyse the contribution of local sources as a function of reduced frequency, nodal diameter and mode shape