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Thesis (Ph.D.) - University of Queensland, 2003. / Includes bibliography.

Thesis (Ph. D.)--Rutgers University, 2010. / "Graduate Program in Mechanical and Aerospace Engineering." Includes bibliographical references (p. 86-90).

A study is conducted to analyze the performance of different turbulence models when applied to flow through a Mach 7.4 hypersonic inlet. The analysis, which is two-dimensional, is done by comparing computational results from a Parabolized Navier Stokes code, with experimental data. The McDonald Camarata (MC) and Baldwin Lomax (BL) models were the two zero-equation models used in the study. The Turbulent Kinetic Energy (TKE) model was chosen as a representative higher order model. The MC model, when run with transition of flow, provides a solution which compares excellently with the data. Transition has a first order effect on the overall solution provided by the code. The BL model predicts separation of flow in the inlet, which contradicts experimental findings. The TKE model does not perform any better than the MC and BL models, despite the fact that it is a higher order turbulence model. The BL and TKE models predict transition in the inlet at a location which is much earlier than observed in the experiment. This may be attributed to the empirical constants used to determine the point of transition. / M.S.

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

One of the characteristics of the high speed ows over blunt bodies is the detached shock formed in front of the body. The distance of the shock from the stagnation point measured along the stagnation streamline is termed as the shock stand o distance or the shock detachment distance. It is one of the most basic parameters in such ows. The need to know the shock stand o distance arises due to the high temperatures faced in these cases. The biggest challenge faced in high enthalpy ows is the high amounts of heat transfer to the body. The position of the shock is relevant in knowing the temperatures that the body being subjected to such ows will have to face and thus building an efficient system to reduce the heat transfer. Despite being a basic parameter, there is no theoretical means to determine the shock stand o distance which is accepted universally. Deduction of this quantity depends more or less on experimental or computational means until a successful theoretical model for its predictions is developed. The experimental data available in open literature for spherical bodies in high speed ows mostly lies beyond the 2 km/s regime. Experiments were conducted to determine the shock stand o distance in the velocity range of 1-2 km/s. Three different hemispherical bodies of radii 25, 40 and 50 mm were taken as test models. Since the shock stand o distance is known to depend on the density ratio across the shock and hence gamma (ratio of specific heats), two different test gases, air and carbon dioxide were used for the experiments here. Five different test cases were studied with air as the test gas; Mach 5.56 with Reynolds number of 5.71 million/m and enthalpy of 1.08 MJ/kg, Mach 5.39 with Reynolds number of 3.04 million/m and enthalpy of 1.42 MJ/kg Mach 8.42 with Reynolds number of 1.72 million/m and enthalpy of 1.21 MJ/kg, Mach 11.8 with Reynolds number of 1.09 million/m and enthalpy of 2.03 MJ/kg and Mach 11.25 with Reynolds number of 0.90 million/m and enthalpy of 2.88 MJ/kg. For the experiments conducted with carbon dioxide as test gas, typical freestream conditions were: Mach 6.66 with Reynolds number of 1.46 million/m and enthalpy of 1.23 MJ/kg. The shock stand o distance was determined from the images that were obtained through schlieren photography, the ow visualization technique employed here. The results obtained were found to follow the same trend as the existing experimental data in the higher velocity range. The experimental data obtained was compared with two different theoretical models given by Lobb and Olivier and was found to match. Simulations were carried out in HiFUN, an in-house CFD package for Euler and laminar own conditions for Mach 8 own over 50 mm body with air as the test gas. The computational data was found to match well with the experimental and theoretical data

The hypersonic laminar to turbulent transition problem above Mach 10 is addressed experimentally in the short duration VKI Longshot gun tunnel. Reentry conditions are partially duplicated in terms of Mach and Reynolds numbers. Pure nitrogen is used as a test gas with flow enthalpies sufficiently low to avoid its dissociation, thus approaching a perfect gas behavior. The stabilizing effects of Mach number and nosetip bluntness on the development of natural boundary layer disturbances are evaluated over a 7 degrees half-angle conical geometry without angle of attack. <p><p>Emphasis is initially placed on the flow characterization of the Longshot wind tunnel where these experiments are performed. Free-stream static pressure diagnostics are implemented in order to complete existing stagnation point pressure and heat flux measurements on a hemispherical probe. An alternative method used to determine accurate free-stream flow conditions is then derived following a rigorous theoretical approach coupled to the VKI Mutation thermo-chemical library. Resulting sensitivities of free-stream quantities to the experimental inputs are determined and the corresponding uncertainties are quantified and discussed. The benefits of this different approach are underlined, revealing the severe weaknesses of traditional methods based on the measurement of reservoir conditions and the following assumptions of an isentropic and adiabatic flow through the nozzle. The operational map of the Longshot wind tunnel is redefined accordingly. The practical limits associated with the onset of nitrogen flow condensation under non-equilibrium conditions are also accounted for. <p><p>Boundary layer transition experiments are then performed in this environment with free-stream Mach numbers ranging between 10-12. Instrumentation along the 800mm long conical model includes flush-mounted thermocouples and fast-response pressure sensors. Transition locations on sharp cones compare favorably with engineering correlations. A strong stabilizing effect of nosetip bluntness is reported and no transition reversal regime is observed for Re_RN<120000. Wavelet analysis of wall pressure traces denote the presence of inviscid instabilities belonging to Mack's second mode. An excellent agreement with Linear Stability Theory results is obtained from which the N-factor of the Longshot wind tunnel in these conditions is inferred. A novel Schlieren technique using a short duration laser light source is developed, allowing for high-quality flow visualization of the boundary layer disturbances. Comparisons of these measurement techniques between each other are finally reported, providing a detailed view of the transition process above Mach 10. / Doctorat en Sciences de l'ingénieur / info:eu-repo/semantics/nonPublished