Des méthodes numériques sont présentées qui permettent la simulation de jets chauds débouchants dans un écoulement transverse aux grands nombres de Reynolds et aux rapports des vitesses faibles. Différentes approches pour la modélisation de turbulence, c'est-à-dire URANS, SAS, DDES et ELES, sont validées par comparaison à des données expérimentales pour une configuration générique, soulignant la nécessité de résoudre les différentes échelles turbulentes pour une prévision correcte du mélange thermique. L'analyse de la solution instationnaire permet l'identification de processus dynamiques intrinsèques ainsi que des phénomènes de mélange et l'application de l'analyse en composantes principales révèle l'ondulation latérale du sillage de jet. Du fait du caractère multi-échelles qui se manifeste dans la simulation d'un jet débouchant sur une configuration avion, l'approche séquentielle basée sur le modèle SAS est mise en place. Comme les résultats pour la sortie d'un système de dégivrage de nacelle sont en bon accord avec les données d'essai en vol, cette approche est finalement appliquée à la sortie complexe d'un système de pre-cooler, mettant en valeur sa capacité à être appliquée dans un processus industriel. / Numerical methods for the simulation of hot jets in cross flow at high Reynolds numbers and small momentum ratios are presented. Different turbulence modeling strategies, i.e. URANS, SAS, DDES and ELES, are validated against experimental data on a generic configuration, highlighting the necessity of scale-resolution for a correct prediction ofthermal mixing. The analysis of transient flow simulations allows the identification of inherent flow dynamics as well as mixing phenomena and the application of the Proper Orthogonal Decomposition revealed the lateral wake meandering as being one of them. Due to the multi-scale problem which arises when simulating jets in cross flow on real aircraft configurations, the sequential approach based on the SAS turbulence model is introduced. As results for the exhaust of a nacelle anti-icing system comprising multiple jets in cross flow agree well with flight test data, the approach is applied in a last step to the complex exhaust of a pre-cooling system, emphasizing the capabilities of this methodology in an industrial environment.

In the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on hypersonic waverider are carried out at design Mach 6. The hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis.

Laminar turbulent transition is known for a long time as a critical phenomenon influencing the thermal load encountered by hypersonic vehicle during their planetary re-entry trajectory. Despite the efforts made by several research laboratories all over the world, the prediction of transition remains inaccurate, leading to oversized thermal protection system and dramatic limitations of hypersonic vehicles performances. One of the reasons explaining the difficulties encountered in predicting transition is the wide variety of parameters playing a role in the phenomenon. Among these parameters, surface roughness is known to play a major role and has been investigated in the present thesis.<p><p>A wide bibliographic review describing the main parameters affecting transition and their coupling is proposed. The most popular roughness-induced transition predictions correlations are presented, insisting on the lack of physics included in these methods and the difficulties encountered in performing ground hypersonic transition experiments representative of real flight characteristics. This bibliographic review shows the importance of a better understanding of the physical phenomenon and of a wider experimental database, including real flight data, for the development of accurate prediction methods.<p><p>Based on the above conclusions, a hypersonic experimental test campaign is realized for the characterization of the flow field structure in the vicinity and in the wake of 3D roughness elements. This fundamental flat plate study is associated with numerical simulations for supporting the interpretation of experimental results and thus a better understanding of transition physics. Finally, a model is proposed in agreement with the wind tunnel observations and the bibliographic survey.<p><p>The second principal axis of the present study is the development of a hypersonic in-flight roughness-induced transition experiment in the frame of the European EXPERT program. These flight data, together with various wind tunnel measurements are very important for the development of a wide experimental database supporting the elaboration of future transition prediction methods. / Doctorat en Sciences de l'ingénieur / info:eu-repo/semantics/nonPublished

A series of experiments were carried out in Purdue University’s Boeing/AFOSR Mach6 Quiet Tunnel, to understand the amplification of streamwise vortices across a separated region in a quiet flow regime. Streamwise vortices were induced on the upstream end of an axisymmetric model consisting of a 7-degree half-angle cone, a cylinder, and a 10-degree flare. The instabilities were seeded using a pre-existing set of roughness inserts, with small, discrete roughness elements. The elements varied in spacing, height, and number of elements. The model was aligned to near 0.0 degree angle of attack. <div><br></div><div>The streamwise, Gortler-like instabilities travelled across the separated region onto the flare, where they were measured with pressure transducers and infrared thermography. The amplification of the instabilities was measured at a variety of Reynolds numbers, under both quiet and conventional noise flow. The results were compared to those of a smooth insert. Heat transfer results showed a streaking pattern, with a peak in heating visible in the streak. Heat flux increased linearly with Reynolds number. If transition was induced, the heat flux would begin to decrease. Power spectral density measurements of the pressure fluctuations indicated that the region within the streak contained two notable instabilities, one between 70 and 150 kHz, and one between 200 and 250 kHz. Transition was only measured in the spectral content in the region on the flare where a ”filling in” of streaks was visible in heat transfer results. Heat flux increased in an nonlinear manner with increasing roughness height. </div><div><br></div><div>The streak positioning and peak heat flux showed a high sensitivity to small, uncontrollable changes in run conditions throughout. Heat transfer results were largely repeatable for small angles of attack, less than 0.1 degrees. The streaks shifted slightly in width and position for angles of attack near 0.1 degrees. Small changes in the streak positioning and heat transfer magnitude were seen in repeatability runs; this is mostly attributable to small changes in initial run conditions. </div>

No description available.

<p>Accurate and reliable prediction of laminar-turbulent boundary layer transition at hypersonic velocities is important for the development of a variety of practical high-speed flight systems currently under development. Boundary layer transition can cause up to an order of magnitude increase in skin friction and heat flux on a flight vehicle, meaning that understanding boundary layer behavior is critical to the design of weight-efficient thermal protection systems. Despite the importance of the topic, significant gaps remain in the community's current understanding of boundary layer transition and control. </p> <p>One of the biggest areas of concern in the field of high-speed boundary layer transition is the effect of facility noise on wind tunnel measurements. Conventional hypersonic wind tunnels are contaminated by freestream fluctuations which can be as much as two orders of magnitude higher than free-flight atmospheric conditions. These disturbances are typically produced by turbulent boundary layers on the tunnel walls; they are acoustic in nature and consist of pressure waves which radiate into the test section. This facility noise plays a leading role in high-speed transition phenomena in conventional hypersonic tunnels.</p> <p><br></p> <p>The current work studies the effects of facility noise on hypersonic transition using both linear stability theory and direct numerical simulation. A model for the freestream disturbance environment of the von Karman Facility's Tunnel B based on experimental measurements of the disturbance spectra present in the tunnel is created and used to study a past experiment performed in the same wind tunnel using a sharp cone and hollow cylinder. The results show that while linear stability theory accurately captures the behavior of second-mode instability growth, it fails to predict the growth of low-frequency instabilities recorded in the experiments. The stability theory analysis also suggests that very fine scale variation in nose tip geometry can play an outsize role in the development of boundary layer instabilities significantly farther downstream.</p> <p><br></p> <p>The direct numerical simulation demonstrates that, using an artificial body forcing term to implement the constructed tunnel noise model, the experimental effects of facility noise can be adequately captured with a sufficiently dense computational grid. For the conical geometry used in the experiments, calculations of surface heat flux indicate good experimental agreement with in prediction of transition location, and total temperature spectra extracted from the flow compare favorably with the experimental data as well. Visualizations of the flowfield confirm the onset of turbulence as a result of the freestream forcing. The computations also suggest that nonlinear interactions may be present in the turbulent breakdown region, leading to the production of streamwise streaks along the cone's surface. Transition on the hollow cylinder was not achieved due to suspected resolution issues, so detailed physical comparison of the two cases was not possible.</p> <p><br></p> <p>Overall, the results of this work suggest that direct numerical simulation is a capable tool for studying the effects of facility noise on hypersonic transition for simple geometries, albeit one which can be difficult to practically realize considering the required computational cost. Computational results indicate that two phenomena may play a role in the development of boundary layer instabilities for a sharp cone --- the fine-scale shape of the tip, which may change the behavior of the entropy layer near the nose; and the interactions between low- and high-frequency waveforms, which seems to cause nonlinear breakdown in line with current experimental understanding.</p>

<p dir="ltr">Pressure transducer, infrared heat transfer, and schlieren imaging data for a flared cone and cone-cylinder-flare in Mach 6 quiet flow are presented. Flared cone pressure transducer results show second-mode RMS values comparable to that found in prior experimental work. Second-mode frequency is found to linearly increase with increasing freestream unit Reynolds number, and frequency varies little between sensors for a given freestream unit Reynolds number. Turbulent intermittency begins to increase at a freestream unit Reynolds number 2x10⁶/m greater than the unit Reynolds number corresponding to peak second-mode RMS. peak RMS. High-speed schlieren imaging on the downstream section of the flared cone shows the second-mode disturbance following trends in power which correlate with PCB RMS. Infrared heat transfer results contain the azimuthal heating streak pattern observed for the flared cone in prior research, but the hot-cold-hot streak pattern is not seen due to limited model length. Streak heating occurs downstream of second-mode peak RMS over the freestream unit Reynolds number range of 6.4x10⁶/m to 10.4x10⁶/m. The heat transfer of streaks is found to vary significantly from streak to streak, while mean streak heating variation with freestream unit Reynolds number is small.</p><p dir="ltr">PCB results of the cone-cylinder-flare show intermittent turbulence at a freestream unit Reynolds number of 16.0x10⁶/m. Examination of shear-layer and second-mode instabilities show significant increases in RMS moving downstream along the flare and with increasing freestream unit Reynolds number. High-speed schlieren imaging of the shear-layer reattachment region on the flare show the presence of the shear-layer and second-mode instabilities when the model is configured with a sharp nose tip. The instabilities are not present with a blunt 5 mm radius nose tip. Heat transfer is observed to increase along the downstream portion of the flare. The sharp nose tip configuration has higher heat transfer rates than the 5 mm radius nose tip configuration.</p>

Hypersonic vehicles are exposed to severe heating loads during their flight in the atmosphere. In order to minimize the heating problem, a variety of cooling techniques are presently available for hypersonic blunt bodies. Introduction of a forward-facing cavity in the nose tip of a blunt body configuration of hypersonic vehicle is one of the most simple and attractive methods of reducing the convective heating rates on such a vehicle. In addition to aerodynamic heating, the overall drag force experienced by vehicles flying at hypersonic speeds is predominate due to formation of strong shock waves in the flow. Hence, the effective management of heat transfer rate and aerodynamic drag is a primary element to the success of any hypersonic vehicle design. So, precise information on both aerodynamic forces and heat transfer rates are essential in deciding the performance of the vehicle. In order to address the issue of both forces and heat transfer rates, right kind of measurement techniques must be incorporated in the ground-based testing facilities for such type of body configurations. Impulse facilities are the only devices that can simulate high altitude flight conditions. Uncertainties in test flow conditions of impulse facilities are some of the critical issues that essentially affect the final experimental results. Hence, more reliable and carefully designed experimental techniques/methodologies are needed in impulse facilities for generating experimental data, especially at hypersonic Mach numbers. In view of the above, an experimental program has been initiated to develop novel techniques of measuring both the aerodynamic forces and surface heat transfer rates. In the present investigation, both aerodynamic forces and surface heat transfer rates are measured over the test models at hypersonic Mach numbers in IISc hypersonic shock tunnel HST-2, having an effective test time of 800 s. The aerodynamic coefficients are measured with a miniature type accelerometer based balance system where as platinum thin film sensors are used to measure the convective heat transfer rates over the surface of the test model. An internally mountable accelerometer based balance system (three and six-component) is used for the measurement of aerodynamic forces and moment coefficients acting on the different test models (i.e., blunt cone with after body, blunt cone with after body and frustum, blunt cone with after body-frustum-triangular fins and sharp cone with after body-frustum-triangular fins), flying at free stream Mach numbers of 5.75 and 8 in hypersonic shock tunnel. The main principle of this design is that the model along with the internally mounted accelerometer balance system are supported by rubber bushes and there-by ensuring unrestrained free floating conditions of the model in the test section during the flow duration. In order to get a better performance from the accelerometer balance system, the location of accelerometers plays a vital role during the initial design of the balance. Hence, axi-symmetric finite element modeling (FEM) of the integrated model-balance system for the missile shaped model has been carried out at 0° angle of attack in a flow Mach number of 8. The drag force of a model was determined using commercial package of MSC/NASTRAN and MSC/PATRAN. For test flow duration of 800 s, the neoprene type rubber with Young’s modulus of 3 MPa and material combinations (aluminum and stainless steel material used as the model and balance) were chosen. The simulated drag acceleration (finite element) from the drag accelerometer is compared with recorded acceleration-time history from the accelerometer during the shock tunnel testing. The agreement between the acceleration-time history from finite-element simulation and measured response from the accelerometer is very good within the test flow domain. In order to verify the performance of the balance, tests were carried out on similar standard AGARD model configurations (blunt cone with cylinder and blunt cone with cylinder-frustum) and the results indicated that the measured values match very well with the AGARD model data and theoretically estimated values. Modified Newtonian theory is used to calculate the aerodynamic force coefficient analytically for various angles of attack. Convective surface heat transfer rate measurements are carried out by using vacuum sputtered platinum thin film sensors deposited on ceramic substrate (Macor) inserts which in turn are embedded on the metallic missile shaped body. Investigations are carried out on a model with and without fin configurations in HST-2 at flow Mach number of 5.75 and 8 with a stagnation enthalpy of 2 MJ/kg for zero degree angle of attack. The measured heating rates for the missile shaped body (i.e., with fin configuration) are lower than the predicted stagnation heating rates (Fay-Riddell expression) and the maximum difference is about 8%. These differences may be due to the theoretical values of velocity gradient used in the empirical relation. The experimentally measured values are expressed in terms of normalized heat transfer rates, Stanton numbers and correlated Stanton numbers, compared with the numerically estimated results. From the results, it is inferred that the location of maximum heating occurs at stagnation point which corresponds to zero velocity gradient. The heat-transfer ratio (q1/Qo)remains same in the stagnation zone of the model when the Mach number is increased from 5.75 to 8. At the corners of the blunt cone, the heat transfer rate doesn’t increase (or) fluctuate and the effects are negligible at two different Mach numbers (5.75 and 8). On the basis of equivalent total enthalpy, the heat-transfer rate with fin configuration (i.e., at junction of cylinder and fins) is slightly higher than that of the missile model without fin. Attempts have also been made to evaluate the feasibility of using forward facing cavity as probable technique to reduce the heat transfer rate and to study its effect on aerodynamic coefficients on a 41° apex angle missile shaped body, in hypersonic shock tunnel at a free stream Mach number of 8. The forward-facing circular cavities with two different diameters of 6 and 12 mm are chosen for the present investigations. Experiments are carried out at zero degree angle of attack for heat transfer measurements. About 10-25 % reduction in heat transfer rates is observed with cavity at gauge locations close to stagnation region, whereas the reduction in surface heat transfer rate is between 10-15 % for all other gauge locations (which is slightly downstream of the cavity) compared with the model without cavity. In order to understand the influence of forward facing cavities on force coefficients, measurement of aerodynamic forces and moment coefficients are also carried out on a missile shaped body at angles of attack. The same six component balance is also being used for subsequent investigation of force measurement on a missile shaped body with forward facing cavity. Overall drag reductions of up to 5 % is obtained for a cavity of 6 mm diameter, where as, for the 12 mm cavity an increase in aerodynamic drag is observed (up to about 10%). The addition of cavity resulted in a slight increase in the missile L/D ratio and did not significantly affect the missile lateral components. In summary, the designed balances are found to be suitable for force measurements on different test models in flows of duration less than a millisecond. In order to compliment the experimental results, axi-symmetric, Navier-Stokes CFD computations for the above-defined models are carried out for various angles of attack using a commercial package CFX-Ansys 5.7. The experimental free stream conditions obtained from the shock tunnel are used for the boundary conditions in the CFD simulation. The fundamental aerodynamic coefficients and heat transfer rates of experimental results are shown to be in good agreement with the predicted CFD. In order to have a feeling of the shock structure over test models, flow visualization experiments have been carried out by using the Schlieren technique at flow Mach numbers of 5.75 and 8. The visualized shock wave pattern around the test model consists of a strong bow shock which is spherical in shape and symmetrical over the forebody of the cone. Experimentally measured shock stand-off distance compare well with the computed value as well as the theoretically estimated value using Van Dyke’s theory. These flow visualization experiments have given a factual proof to the quality of flow in the tunnel test section.

<p>The first force and moment measurements were conducted in the BAM6QT. Three 7-degree half-angle sharp cones were tested, one with base radius of 4.5 in. and two with base radius of 3.5 in. made out of different materials. Models were tested at 0 and 2 degrees angle of attack. Models were tested over a range of burst pressures and Reynolds numbers. Models were fitted onto a strain gauge, 6 component, internal, moment balance. Multiple assemblies were tested that mounted the balance in the BAM6QT. High-speed schlieren video was used to monitor flow conditions and track the movement of the tunnel and model. Three entries were performed in the BAM6QT. The improvement in data quality with each new entry is shown and the startup and running loads from entry 3 are analyzed.</p> <p>Startup loads were measured and are of importance in determining the load range needed to operate in the BAM6QT. Large startup loads up to 40X the running load were identified. Tunnel movement was measured and was used to approximate the inertial loading during startup and the run. The inertial loading was not found to be the cause of the large startup loads. Schlieren video was used to qualitatively review the startup flow. It was found the large startup loads in axial force were plausibly from the high-pressure subsonic flow evacuating the nozzle. For normal force and pitching moment, the startup loads peak at a different time than axial force and appear to be from a shock-shock interaction nearby the model. Trends in startup load with changing model geometry, AoA, and burst pressure were put together to form an empirical estimation for startup loads sharp cones. </p> <p>Running loads were profiled and found to be trending with burst pressure and model geometry similarly to Newtonian flow theory predictions. However, due to the lack of a base pressure measurement, the results are uncorrected for sting effects and differ from Newtonian flow theory by a scalar. A 5.3 Hz oscillation in axial force was identified. The frequency of the oscillation is the same as the frequency of the quasi-steady flow periods caused by the reflection of the expansion fan in the driver tube. Normal force during the running load was found to be measuring positive loads when at 0 degrees angle of attack. Both the axial and normal force phenomena were unexpected and were investigated but both require further research. </p> <p><br></p> <p><br></p> <p><br></p> <p><br></p>

<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>