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The current study aims to validate and extend RANS approaches, using linear and non-linear eddy viscosity models, by considering the modelling of hypersonic flows, especially those with SWBLIs. The main objective is to assess the performance of three EVMs, widely used in industrial applications, in predicting flow separation and aerodynamic heating in hypersonic flows, and to explore adaptations that can improve their performance, by accounting for influences such as strong compressibility and discontinuity in flow variables. The turbulence models employed are the linear k-ε model of Launder and Sharma (1974) (LSY model), the linear k-ω shear stress transport model of Menter et al. (2003) (SST model) and the non-linear k-ε model of Craft et al. (2000) (CLS model). To give an objective, uniform and reliable evaluation of the performance of the turbulence models selected, the formulation of RANS equations for high-speed flows has first been investigated. A new density-based compressible solver, with various options of including, or excluding, several terms that were typically omitted in low speed (and sometimes high speed) flow simulations has developed within the framework of OpenFOAM. The influence of partially, or fully, ignoring these terms in the RANS equations has been analysed with comparisons of spatial quantities and wall quantities in supersonic and hypersonic cases. The numerical results show that while omitting these terms is typically fine for low-speed applications, they can have a significant effect on predictions of SWBLI cases in the hypersonic regime. There-fore, the form of RANS equations with all these terms fully included is selected to run all the simulations in this research. The non-linear CLS model has been implemented in the OpenFOAM framework in a similar way as the linear models tested in the current study and then examined, together with the other two linear models, by comparing results over a range of typical hypersonic SWBLIs benchmark cases, 2D, axisymmetric or 3D, over a wide range of Mach numbers and Reynolds numbers. The three models tested do generally return reasonable predictions of wall pressure in most cases, while the non-linear CLS model presented the best capability among the three in predicting the flow separations. Nevertheless, the wall heat flux in the interaction region is overpredicted in most cases by all three models (sometimes showing quite a dramatic overprediction). To improve the accuracy of predictions of wall heat flux, a new source term in the modelled dissipation rate equation, aimed at restricting the turbulent length scale in the shock wave boundary layer interaction region, is proposed. This term will cause no impact on incompressible flows and only has a limited influence on the predic-tions of wall pressure in hypersonic flows. The introduction of this new source term to the LSY and CLS models is shown to largely eliminate their over predictions of the peaks of wall heat flux in the interaction region (in some test cases, the peak value was reduced by around 80%), while the extra cost of the numerical simulation is less than 2%.

Two sub-systems that present a significant challenge in the development of highperformanceair vehicle exceeding speeds of Mach 5 are the power generation and thermalmanagement sub-systems. The air friction experienced at high speeds, particularly aroundthe engine, generates large thermal loads that need to be managed. In addition, traditionaljet engines do not operate at speeds greater than Mach 3, therefore eliminating thepossibility of a rotating power generator. A multi-mode water-based Rankine cycle is aninnovative method to address both of these constraints of generating power and providingcooling. Implementing a Rankine cycle-based system allows for the waste heat from thevehicle to be used to meet the onboard power requirements. This application of a Rankinecycle differs from standard power plant applications because the transient system dynamicsbecome important due to rapid changes in thermal loads and electrical power requirements.Both an experimental and computational investigation is presented. An experimentalsteady state energy balance was used to determine a 5.1% and 11.5% thermal and SecondLaw efficiency, respectively. Transient testing showed an increase in power generation of283% in 30.5 seconds when starting from idle, with a steady state power generation of 230W. In addition to the power generation, the experimental system removed 10.7 kW fromthe hot oil loop which emulates a typical aircraft cooling fluid. Experimental results were used in the development of dynamic computational models using OpenModelica, an opensourcemodeling tool. Deviation between model and experimental results was within 5%for component models and 3.5% when analyzing the system energy balance. Testing of thevehicle level model included steady state, transient, and simulated mission, which was usedto characterize performance and develop the system controls. During transient testing, thesystem controls demonstrated the ability to meet both the cooling and power requirementsof the system through rapid response times and minimal temperature overshoot (2.72%).The development and testing of this model provides an opportunity for scaling andoptimization of a combined power and thermal management system across a wide range ofvehicle sizes and operating conditions.

This dissertation describes model reduction techniques for the computation of aerodynamic heat flux and pressure loads for multi-disciplinary analysis of hypersonic vehicles.NASA and the Department of Defense have expressed renewed interest in the development of responsive, reusable hypersonic cruise vehicles capable of sustained high-speed flight and access to space. However, an extensive set of technical challenges have obstructed the development of such vehicles. These technical challenges are partially due to both the inability to accurately test scaled vehicles in wind tunnels and to the time intensive nature of high-fidelity computational modeling, particularly for the fluid using Computational Fluid Dynamics (CFD).The aim of this dissertation is to develop efficient and accurate models for the aerodynamic heat flux and pressure loads to replace the need for computationally expensive, high-fidelity CFD during coupled analysis. Furthermore, aerodynamic heating and pressure loads are systematically evaluated for a number of different operating conditions, including: simple two-dimensional flow over flat surfaces up to three-dimensional flows over deformed surfaces with shock-shock interaction and shock-boundary layer interaction. An additional focus of this dissertation is on the implementation and computation of results using the developed aerodynamic heating and pressure models in complex fluid-thermal-structural simulations.Model reduction is achieved using a two-pronged approach. One prong focuses on developing analytical corrections to isothermal, steady-state CFD flow solutions in order to capture flow effects associated with transient spatially-varying surface temperatures and surface pressures (e.g., surface deformation, surface vibration, shock impingements, etc.). The second prong is focused on minimizing the computational expense of computing the steady-state CFD solutions by developing an efficient surrogate CFD model.The developed two-pronged approach is found to exhibit balanced performance in terms of accuracy and computational expense, relative to several existing approaches. This approach enables CFD-based loads to be implemented into long duration fluid-thermal-structural simulations.

Four new two-equation turbulence models are recently added to LAURA (NASA Langley’s aerothermodynamics CFD code). LAURA is usually used to calculate the aerodynamic heating for reentry vehicles. These new two-equation turbulence models are: Wilcox 2006 k-w, Wilcox 1998 k-w, Menter SST and Menter two-equation model. In this research these models are run for the first time on blunt bodies, which are cases with a stagnation point. It has been shown that for hypersonic flows with stagnation point, it is better to calculate the turbulence kinetic energy production term by the vorticity-based method, rather than the strain-based method that is in the original formulation of these models. The models are run on two test cases: an axisymmetric hemisphere and an 86 block full scale half-body MPCV. On the axisymmetric hemisphere, the models with strain-based production term, generate large amounts of extra k right after the shock, near the stagnation point. This extra k disrupts the flow field, and prevents the case from converging. With vorticity-based production term, all the models gave results close to the Cebeci-Smith algebraic model that has been present in the code for a long time and is trusted for these types of attached flows. On the half-body MPCV case, we run the Wilcox 2006 k-w model with the strain-based and vorticity-based production terms, and we got very close results by the two methods

The objective of this thesis was to determine the feasibility of using loop heat pipes to dissipate waste heat from power electronics to the skin of a fighter aircraft and examine the performance characteristics of a titanium-water loop heat pipe under stationary and elevated acceleration fields. In the past, it has been found that the boundary condition at the condenser can be a controlling factor in the overall performance of this type of thermal management scheme. Therefore, the heat transfer removed from the aircraft skin has been determined by modeling the wing as a flat plate at zero-incidence as a function of the following parameters: airspeed: 0.8 ≤ Ma∞ ≤ 1.4; altitude: 0 ≤ H ≤ 22 km; wall temperature: 105 ≤ Tw ≤ 135°C. In addition, the effects of the variable properties of air have been taken into account. Heat transfer due to thermal radiation has been neglected in this analysis due to the low skin temperatures and high airspeeds up to Ma∞ = 1.4. It was observed that flight speed and altitude have a significant effect on the heat transfer abilities from the skin to ambient, with heat rejection becoming more difficult with increasing Mach number or decreasing altitude. An experiment has been developed to examine operating characteristics of a titanium-water loop heat pipe (LHP) under stationary and elevated acceleration fields. The LHP was mounted on a 2.44 m diameter centrifuge table on edge with heat applied to the evaporator via a mica heater and heat rejected using a high-temperature polyalphaolefin coolant loop. The LHP was tested under the following parametric ranges: heat load at the evaporator: 100 ≤ Qin ≤ 600 W; heat load at the compensation chamber: 0 ≤ Qcc ≤ 50 W; radial acceleration: 0 ≤ ar ≤ 10 g. For stationary operation (az = 1.0 g, ar = 0 g), the LHP evaporative heat transfer coefficient decreased monotonically, thermal resistance decreased to a minimum then increased, and wall superheat increased monotonically. Heat input to the compensation chamber was found to increase the evaporative heat transfer coefficient and decrease thermal resistance for Qin = 500 W. Flow reversal in the LHP was found for some cases, which was likely due to vapor bubble formation in the primary wick. Operating the LHP in an elevated acceleration environment (az = 1.0 g, ar > 0 g) revealed dry-out conditions from Qin = 100 to 400 W and varying accelerations and the ability for the LHP to reprime after an acceleration event that induced dry-out. Evaporative heat transfer coefficient and thermal resistance was found not to be significantly dependent on radial acceleration. However, wall superheat was found to increase slightly with radial acceleration.

Hypersonic flight at sea-level conditions induces severe thermal loads not seen by any other type of current hypersonic system. Appropriate design of the hypersonic round requires a solid understanding of the thermal environment. Numerous codes were obtained and assessed for their applicability to the problem under study, and outside of the GASP Conjugate Heat Transfer module, Navier-Stokes code from Aerosoft, Inc., no efficient codes are available that can model the aerodynamic heating response for a fully detailed projectile, including all subassemblies, over an entire trajectory. Although the codes obtained were not applicable to a fully detailed thermal soak analyses they were useful in providing insight into ablation effects. These initial trade studies indicated that ablation of up to 1.25 inches could be expected for a Carbon-Carbon nosetip in this flight environment. In order to capture the thermal soak effects a new methodology (BMA) was required. This methodology couples the Sandia aerodynamic heating codes with a full thermal finite element model of the desired projectile, using the finite element code ANSYS from ANSYS, Inc. Since ablation can be treated elsewhere it was not included in the BMA methodology. Various trajectories of quadrant elevations of 0.5, 10, 30, 50, and 80 degrees were analyzed to determine thermal time histories and maximum operating temperatures. All of the trajectories have the same launch condition, Mach 8 sea-level, and therefore will undergo the same initial thermal spike in temperature at the nose-tip of approximately 3,100 K (5600R). Of the five trajectories analyzed the maximum internal temperatures experienced occurred for the 50 degree quadrant elevation trajectory. This trajectory experienced temperatures in excess of 1,000 K (1800R) for more than 80% of its flight time. The BMA methodology was validated by comparisons with experiment and computational fluid solutions with an uncertainty of 10% at a cost savings of over three orders of magnitude.

Due to renewed interest in the sub orbital rocket program of the Florida Space Authority and a surplus of Super Loki Sounding Rockets, an effort to improve the usefulness of this surplus is herein undertaken. Currently, the capacity of the payload compartment in the upper stage of the Super Loki system is very limited. A redesign of the upper stage will allow larger and more versatile payloads to be carried, assuming the appropriate design parameters are met. It has therefore been undertaken to create a design procedure that is comprehensive in scope in order to affect this redesign. This procedure includes five major components. These are the separation of the upper and lower stages, the stability of the vehicle, the altitude and velocity of the rocket, the mechanical loading and finally the aerodynamic heating. Semi-empirical methods were used whenever possible to allow comparison with experimental data. This procedure revealed that larger diameter upper stages might be used up to a reasonable maximum of four inches. The four-inch modification is found to be stable as were the smaller modifications considered. The altitude and velocity of the rocket were found via a simple Eulerian time stepping scheme resulting in an estimate of approximately 148,000ft for the four-inch dart. The mechanical loading analysis allowed for the material selection for the rocket components. Reinforced steel fins and carbon fiber tubing, for the payload section, are adequate to meet expected mechanical loads, those being, 16000psi for the fin section due to launcher forces, 22800psi for compressive plus torsion forces on the composite section and 18000psi for the ejection stresses. An ablative coating is considered necessary to counteract the 760ºF temperatures along the composite tube.

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An analysis is given which treats experimental heat-transfer data obtained during flights of three V-2 rockets and two Aerobee rockets. This work was conducted largely at Boston University under Air Force Contracts W28-099-ac-395 and AF 33(615)-382. The instrumentation and rocket-flight phases extended from 1947 to 1952, while the analysis was carried out in the period from 1953 to 1955. A total of nine V-2 rockets were instrumented to some extent for the purposes of this experiment. The missiles for which data have been analyzed were chosen on the basis of best performance on the part of both instrumentation and rocket. Experimental equipment of special design was produced and installed by the Boston University group in the rockets which were launched in the White Sands area in New Mexico. Extensive field trips were required in this connection, both for the experiment on heat0transer and others in which the same group participated at the same time. The general program was one of research in physics of the upper atmosphere. Since the conclusion of the supporting contracts in October of 1954, additional consideration of the experimental results has continued on a private basis. The data obtained during the course of the present experiment were in the form of temperature variations as experienced by heat meters which were placed at selected locations over the body and fin of the rockets. The heat meters were small circular diaphragms of stainless steel which were placed flush with the rocket skin. Each diaphragm had on its central inner surface a pair of resistance thermometers made of nickel wire, with a diameter of 0.002". The thermometers were switched in sequence into a resistance bridge. One thermometer had always a higher resistance than the other, an arrangement which served two purposes. Firstly, the two values of voltage for a given temperature permited unambiguous determination of the direction in which the resistance bridge was unbalanced. Secondly, these two values of voltage were handled independently by our own instrumentation, and by the recording or telemetry devices. When two such different values of voltage were finally translated back into their equivalent temperatures, close agreement of the temperature values lent considerable support to their accuracy. When the two thermometers reported temperatures which were significantly different from each other, restraint was indicated in trusting their accuracy. Fortunately, close agreement of temperatures reported by the dual thermometers was the more usual experience. [TRUNCATED]