The NASA Langley Research Center hypersonic wind tunnels serve a vital role in the field of hypersonics in both helping validate CFD predictions and producing experimental results. These tunnels have been heavily utilized for decades by numerous planetary missions, such as MSL and Orion, commercial and academic partnerships, such as Sierra Space and the University of Maryland, and flight projects such as Artemis and LOFTID. The data acquisition method used in these tunnels is thermography, primarily phosphor and infrared. Image data are not collected during model injection, resulting in a data gap in the time-history of temperature. Historically, an approximate method has been used to obtain heating data with the data gap, but a new, higher-fidelity method has been developed that patches the data gap and performs integral heat transfer analysis on the temperature data, directly solving the heat equation and avoiding unnecessary assumptions. This method has been shown to model the surface heating much more accurately, agree with computational predictions better than the current method, and be an overall more robust method that collapses to a constant film coefficient value much more quickly. The culmination of these aspects results in a method that is a significant improvement over the approximate method and increases the fidelity of the heating results obtained from the NASA Langley Research Center hypersonic wind tunnels. / Master of Science / A new heat transfer analysis method that reduces image data taken of wind tunnel models in the NASA Langley Research Center hypersonic wind tunnels has been developed as a higher-fidelity successor to the existing approximate method. This new method patches the data gap that occurs during injection due to the fact that no images are taken of the model until it reaches the centerline of the tunnel. Then, this method performs integral heat transfer analysis on the image temperature data, directly solving the heat equation and avoiding unnecessary assumptions. This method has been shown to model the surface heating better, agree with computational predictions better, and be a more robust method that obtains a film coefficient value more quickly. This method is shown to be a significant improvement over the approximate method.

The experiments on this project proceeded on the premise that adding an array of auxiliary jets behind a main jet injector would alleviate the large region of low pressure typically found downstream of a normal, sonic injector in supersonic flow and also possibly increase in intensity of the upstream high-pressure region. The secondary jet would, in theory, "push" the primary jet further into the flow, increasing the size of the obstacle as seen by the flow. The resulting increased high pressure upstream of the flow would increase the force on the body. Also, the presence of secondary jets would reduce the intensity of the primary jet's low-pressure region. These results would be beneficial to increase the force and decrease the nose-down moment associated with sonic, normal injection into a supersonic crossflow. Therefore, in application to hypersonic, high-altitude missile maneuvering, the firing of a thruster with such an array would result in both added force and a reduction of the moment usually associated with the pressure field on the missile. Such an array could allow the missile to perform purely translational maneuvers with less fuel, all the while keeping the target in view. To accomplish this task, some modern missiles use a second injector far downstream from the primary injector. This second injector's primary function is to negate the nose-down moment, and it adds little to the overall jet effectiveness. To this end, two sets of experiments were conducted: one with low jet pressure ratio, Poj/P1 = 13.65, and low Mach number of 2.4 with Po,inf = 3.74 atm and To,inf = 293K for proof of concept and one at primary conditions Poj/P1 = 620, M1 = 4, Po,inf = 10.21 atm, To,inf = 293K. Spark Shadowgraphs were taken at both of these cases to study the structures present in the flow field and to qualitatively assess the effects of the secondary jet injectors. Placed under the Mach disk of the main jet, the secondary jets are hypothesized to push the plume of the main jet further up into the flow, increasing the force on the plate, and Shadowgraphs were used to test this hypothesis. Schlieren pictures were taken at the high M1, high-pressure ratio test case to further study the interaction of the secondary jets with the main jet. Pressure Sensitive Paint, PSP, was used in both cases to gain a greater understanding of the surface pressure near the injectors for different jet configurations. It was discovered that the addition of secondary jets could indeed both increase the force generated by the main jet and reduce the undesirable nose-down moment created by the main jet. In the low M1, low pressure ratio conditions, the addition of one pair of jets manipulated the surface pressure such that the force on the plate increased by 17% and the nose-down moment was increased by 9% over the main jet only case. The further addition of one more pair of injectors increased the surface pressure force on the plate by 34% and increased the nose-down moment on the plate by 3% when compared to the Main Jet Only case. It is important to note that, these increases are due solely to the manipulation of the surface pressure force field and not the thrust of the secondary jets. The added thrust would increase the force on the plate and their position would insure an increase of a nose-up moment. One pair of secondary jets increases the injectant mass flow by about 2.3%. Therefore, the effects reported above are seen to be disproportionate to the amount of added injectant. For the primary test conditions (M1 = 4, Poj/P1 = 620, Po = 10.21 atm, To = 293K) the addition of two pairs of secondary jets had a force increase of 62% and a nose-down moment decrease of 38% over that of the main jet only case. Three pairs increased the force 71% and the decreased the nose-down moment by 26% and four pairs increased the force 91% but increased the nose-down moment by 33%. These values do not account for the thrust of the secondary jets. Accounting for the beneficial effects of the thrust of the secondary jets, the force on the plate for two pairs of secondary jets increased the force 70% and decreased the moment 42%. Three pairs increased the force 83% and decreased the moment 35%. The increase of force for four pairs of secondary jets was 106% and the increase in nose-down moment was only 21%. A point of diminishing returns was reached. As more pairs of injectors are added further and further from the main injector, the beneficial force effects are offset by a growing moment penalty. By considering the locations of the secondary injectors to the main injector for both the low Mach number, low-pressure ratio tests and the main tests conditions, it can be surmised that the greatest benefit from the secondary jets can be extracted when the jets are placed within the main injector's downstream low-pressure region. / Master of Science

The abilities of Scramjets and Ramjets, in their respective operating ranges, are partially bridged by dual-mode Scramjets. The limitations of operation are due to making a static motor that is designed to function in both modes resulting in low and high speed restrictions. This study covers the analysis into the ability of morphing the combustor in a Scramjet to allow for expanded operational capacities through simple mechanisms. Through the restriction and expansion of combustor cavity volume, operational capabilities of the engine can, therefore, be modified to best match scenario requirements. Due to the engine's ability to match a wide variety of scenarios the limitations seen in that of the dual-mode Scramjet are avoided through the usage of a morphing combustor. From initial findings using the quasi-1D Canonical REactor Scramjet Simulation (CReSS) solver, progress was made to confirm results through the usage of Computational Fluid Dynamics (CFD). Prior analysis of the momentum balance between stages two and four of the simulated Scramjet engines, the results showed that the variable geometry matched or outperformed the baseline HiFiRE geometry. The analysis revealed points of Mach and altitude where certain combustor volumes demonstrated greater performance. This greater performance is only gained by the ability to tune the engine in flight to react to external factors as there is no dominant geometry for a given range of Machs and altitudes. This tuning allows for the usage of performance mapping to extract the greatest performance possible over a variety of conditions. Further, it allows for the project to be continuously expanded into mapping appropriate reactions to other initial conditions and stimuli. Using CFD modeling to perform a parametric study on the prior work allows for finer control and analysis of said initial conditions and the resulting flow paths in the variety of tested combustor volumes. From this a discussion is made in regards to the effectiveness of the prior CReSS based analysis of the novel approach. / Master of Science / The abilities of Scramjets and Ramjets (engines which contain no moving parts as the compression of the incoming air is accomplished by the speed at which they operate with the separating factor being that the scramjets internal flow does not go below supersonic speeds), in their respective operating ranges, are partially bridged by dual-mode Scramjets. Dual-mode Scramjets are scramjets which can function with both sub- and super-sonic internal flow speeds. This being below or above Mach 1 (343 m/s, 767.3 mph) respectively. The limitations of operation are due to making a static motor where the geometry does not change that is designed to function in both modes resulting in low and high speed restrictions. This study continues the analysis into the ability of morphing the combustor, the volume in which the air fuel mixture combusts, in a Scramjet to allow for expanded operational capacities through simple mechanisms. Through the restriction and expansion of combustor volume, operational capabilities of the engine can, therefore, be modified to best match scenario requirements. Due to the engine's ability to match a wide variety of scenarios the limitations seen in that of the dual-mode Scramjet are avoided through the usage of a morphing combustor where morphing in this case is a simple volume change equivalent to that of a slide whistle. From initial findings using the quasi-1D Canonical REactor Scramjet Simulation (CReSS) solver, progress was made to confirm results through the usage of Computational Fluid Dynamics (CFD). Prior analysis of the momentum balance between stages two and four of the simulated Scramjet engines, the results showed that the variable geometry matched or outperformed the baseline HiFiRE geometry. The analysis revealed points of Mach and altitude where certain combustor volumes demonstrated greater performance. This greater performance is only gained by the ability to tune the engine in flight to react to external factors as there is no dominant geometry for a given range of Machs and altitudes. This tuning allows for the usage of performance mapping to extract the greatest performance possible over a variety of conditions. Further, it allows for the project to be continuously expanded into mapping appropriate reactions to other initial conditions and stimuli. Using CFD modeling to perform a parametric study on the prior work allows for finer control and analysis of said initial conditions and the resulting flow paths in the variety of tested combustor volumes. From this a discussion is made in regards to the effectiveness of the prior CReSS based analysis of the novel approach.

Airframes operating in the hypersonic regime are subjected to complex structural and thermal loads. Structural loads are a result of aggressive high G maneuvers, rapid vehicle acceleration and deceleration, and dynamic pressure, while thermal loads are a result of aerodynamic heating. For such airframes, structural members are typically constructed from steel, titanium and nickel alloys. However, with most materials, rapid elevations in temperature lead to undesirable changes in material properties. In particular, reductions in strength and stiffness are observed, along with an increase in thermal conductivity, specific heat and thermal expansion. Thus, hypersonic airframes are typically designed with external insulation, active cooling or a thermal protection system (TPS) added to the structure to protect the underling material from the effects of temperature. Such thermal protection may consist of adhesively bonded, pinned, and bolted thermal protection layers over exterior panels. These types of attachments create abrupt changes in thermal expansion and stiffness that make the structure susceptible to cracking and debonding as well as adding mass to the airframe. One of the promising materials concepts for extreme environments that was introduced in the past is the so-called Spatially Tailored Advanced Thermal Structures (STATS). The concept of STATS is rooted in functionally graded materials (FGMs), in which a directional variation of material properties exists. These materials are essentially composites and consist of two or more phases of distinct materials in which the volume fractions of each phase continuously change in space. Here, the graded material will serve a dual-purpose role as both the structural/skin member and thermal management with the goal of reducing the weight of the structure while maintaining structural soundness. This is achieved through the ability to tailor material properties to create a desired or enhanced thermomechanical response through spatial variation (e.g. grading). The objective of this study is to present a computational framework for modeling and evaluating the thermomechanical response of STATS and FGMs for highly maneuverable hypersonic (Mach > 5) airframes. To meet the objective of this study, four key steps have been defined to study the thermomechanical response of such materials in extreme environments. They involve: (1) modeling of graded microstructures; (2) validation of analytical and numerical modeling techniques for graded microstructures; (3) determination of effective properties of variable composition composites; (4) parametric studies to evaluate the performance of FGMs for use in the hypersonic operating environment; (5) optimization of the material spatial grading in hypersonic panels aiming to improve the thermomechanical performance. Modeling of graded microstructures, representing particulate reinforced FGMs, has been accomplished using power law distribution functions to specify the spatial variation of the constituents. Artificial microstructures consisting of disks and spheres have been generated using developed algorithms. These algorithms allow for the creation of dense packing fractions up to 0.61 and 0.91 for 2D and 3D geometry, respectively. Effective properties of FGMs are obtained using micromechanics models and finite element analysis of representative volume elements (RVEs). Two approaches have been adopted and compared to determine the proper RVE for materials with graded microstructures. In the first approach, RVEs are generated by considering regions that have a uniform to slow variation in material composition (i.e., constant volume fraction), resulting in statistically homogenous piecewise RVEs of the graded microstructure neglecting interactions from neighboring cells. In the second approach, continuous RVEs are generated by considering the entire FGM. Here it is presumed that modeling of the complete variation in a microstructure may influence the surrounding layers due to the interactions of varying material composition, particularly when there is a steep variation in material composition along the grading direction. To determine these effects of interlayer interactions, FGM microstructures were generated using three different types of material grading functions, linear, quadratic and square root, providing uniform, gradual and steep variations, respectively. Two- and three-dimensional finite element analysis was performed to determine the effective temperature-dependent material properties of the composite over a wide temperature range. The outcome of the computational analysis show that the similar effective properties are obtained by each of the modeling approaches. Furthermore, the obtained computational results for effective elastic, thermal, and thermal expansion properties are consistent with the known analytical bounds. Resulting effective temperature-dependent material properties were used to evaluate the time-dependent thermostructural response and effectiveness of FGM structural panels. Structural panels are subjected to time- and spatial-dependent thermal and mechanical loads resulting from hypersonic flight over a representative trajectory. Mechanical loads are the by-product of aggressive maneuvering at high air speeds and angles of attack. Thermal loads as a result of aerodynamic heating are applied to the material systems as laminar, turbulent and transitional heat flux on the outer surface. Laminar and turbulent uniform heat fluxes are used to evaluate the effectiveness of FGM panels graded in the through-thickness direction only. Transitional heat fluxes are used to evaluate the effectiveness of FGMs graded in two principal directions, e.g., through-thickness and the surface parallel to flow. The computational results indicate that when subjected to uniform surface heat flux, the graded material system can eliminate through-thickness temperature gradients that are otherwise present in traditional thermal protection systems. Furthermore, two-dimensional graded material systems can also eliminate through-thickness temperature gradients and significantly reduce in-plane surface temperature gradients when subjected to non-uniform surface aerodynamic heating.

The ability to predict and control laminar-to-turbulent transition in high-speed flow has a substantial effect on heat transfer and skin friction, thus improving the design and operation of hypersonic vehicles. The control of transition on blunt bodies is essential to improve the performance of lifting and control surfaces. The objective of this Ph.D. research is to develop efficient and accurate algorithms for the detection of high-frequency heat flux fluctuations supported by hypersonic flow in transitional boundary layers. The focus of this research is on understanding the mathematical properties of the reconstruction such as regularity, sensitivity to noise, multi-resolution, and accuracy. This research is part of an effort to develop small-footprint heat flux sensors able to measure high-frequency fluctuations on test articles in a hypersonic wind tunnel with a small curvature radius. In the present theoretical/numerical study a multi-resolution formulation for the direct and inverse reconstruction of the heat flux from temperature sensors distributed over a multidimensional solid in a hypersonic flow was developed and validated. The solution method determines the thermal response by approximating the system Green's function with the Galerkin method and optimizes the heat flux distribution by fitting the distributed surface temperature data. Coating and glue layers are treated as separate domains for which the Green's function is obtained independently. Connection conditions for the system Green's function are derived by imposing continuity of heat flux and temperature concurrently at all interfaces. The solution heat flux is decomposed on a space-time basis with the temporal basis a multi-resolution wavelet with arbitrary scaling function. Quadrature formulas for the convolution of wavelets and the Green's function, a reconstruction approach based on isoparametric mapping of three-dimensional geometries, and a boundary wavelet approach for inverse problems were developed and verified. This approach was validated against turbulent conjugate heat transfer simulations at Mach 6 on a blunted wedge at 0 angle of attack and wind tunnel experiments of round impinging jet at Mach 0.7 It was found that multidimensional effects were important near the wedge shoulder in the short time scale, that the L-curve regularization needed to be locally corrected to analyze transitional flows and that proper regularization led to sub-cell resolution of the inverse problem. While the L2 regularization techniques are accurate they are also computationally inefficient and lack mathematical rigor. Optimal non-linear estimators were researched both as means to promote sparsity in the regularization and to pre-threshold the inverse heat conduction problem. A novel class of nonlinear estimators is presented and validated against wind tunnel experiments for a flat-faced cylinder also at Mach 6. The new approach to hypersonic heat flux reconstruction from discrete temperature data developed in this thesis is more efficient and accurate than existing techniques. / Doctor of Philosophy / The harsh environment supported by hypersonic flows is characterized by high-frequency turbulent bursts, acoustic noise, and vibrations that pollute the signals of the sensors that probe at high frequencies the state of the boundary layers developing on the walls. This research describes the search for optimal estimators of the noisy signal, i.e., those that lead to the maximum attenuation of the risk of error pollution by non-coherent scales. This research shows that linear estimators perform poorly at high-frequency and non-linear estimators can be optimized over a sparse projection of the signal in a discrete wavelet basis. Optimal non-linear estimators are developed and validated for wind tunnel experiments conducted at Mach 6 in the Advanced Propulsion and Power Laboratory at Virginia Tech.

A computational study aimed at improving the understanding of rotating instability in the LP steam turbine last stage working under low flow rate conditions is described in this thesis. A numerical simulation framework has been developed to investigate into the instability flow field. Two LP model turbine stages are studied under various flow rate conditions. By using the 2D simulations as reference and comparing the results to those of the 3D simulations, the basic physical mechanism of rotating instability is analysed. The pressure ratio characteristics across the rotor row tip are found to be crucial to the inception of rotating instability. The captured instability demonstrates a 2D mechanism based on the circumferential variation of unsteady separation flow in the rotor row. The 3D tip clearance flow is found not a necessary cause of the instability onset. Several influential parameters on the instability flow are also investigated by a set of detailed studies on different turbine configurations. The results show that the instability flow pattern and characteristics can be altered by the gap distance between the stator and rotor row, the rotor blading and the stator row stagger angle. Some flow control approaches are proposed based on the observations, which may also serve as design reference. The tip region 3D vortex flow upstream to the rotor row is also captured by the simulations under low flow rate conditions. Its appearance is found to be able to suppress the inception of rotating instability by disrupting the interaction between the rotor separation flow and the incoming flow. Finally, some recommendations for further work are proposed.

This thesis presents the derivation of the theory leading to the determination of the governing equations of chemically reacting flows under local thermodynamic equilibrium, which rigorously takes into account effects of elemental (de)mixing. As a result, new transport coefficients appear in the equations allowing a quantitative predictions and helping to gain deeper insight into the physics of chemically reacting flows at and near local equilibrium. These transport coefficients have been computed for both air and carbon dioxide mixtures allowing the application of this theory to both Earth and Mars entry problems in the framework of the methodology for the determination of the catalytic activity of Thermal Protections Systems (TPS) materials. Firstly, we analyze the influence of elemental fraction variations on the computation of thermochemical equilibrium flows for both air and carbon dioxide mixtures. To this end, the equilibrium computations are compared with several chemical regimes to better analyze the influence of chemistry on wall heat flux and to observe the elemental fractions behavior along a stagnation line. The results of several computations are presented to highlight the effects of elemental demixing on the stagnation point heat flux and chemical equilibrium composition for air and carbon dioxide mixtures. Moreover, in the chemical nonequilibrium computations, the characteristic time of chemistry is artificially decreased and in the limit the chemical equilibrium regime, with variable elemental fractions, is achieved. Then, we apply the closed form of the equations governing the behavior of local thermodynamic equilibrium flows, accounting for the variation in local elemental concentrations in a rigorous manner, to simulate heat and mass transfer in CO2/N2 mixtures. This allows for the analysis of the boundary layer near the stagnation point of a hypersonic vehicle entering the true Martian atmosphere. The results obtained using this formulation are compared with those obtained using a previous form of the equations where the diffusive fluxes of elements are computed as a linear combination of the species diffusive fluxes. This not only validates the new formulation but also highlights its advantages with respect to the previous one : by using and analyzing the full set of equilibrium transport coefficients we arrive at a deep understanding of the mass and heat transfer for a CO2/N2 mixture. Secondly, we present and analyze detailed numerical simulations of high-pressure inductively coupled air plasma flows both in the torch and in the test chamber using two different mathematical formulations: an extended chemical non-equilibrium formalism including finite rate chemistry and a form of the equations valid in the limit of local thermodynamic equilibrium and accounting for the demixing of chemical elements. Simulations at various operating pressures indicate that significant demixing of oxygen and nitrogen occurs, regardless of the degree of nonequilibrium in the plasma. As the operating pressure is increased, chemistry becomes increasingly fast and the nonequilibrium results correctly approach the results obtained assuming local thermodynamic equilibrium, supporting the validity of the proposed local equilibrium formulation. A similar analysis is conducted for CO2 plasma flows, showing the importance of elemental diffusion on the plasma behavior in the VKI plasmatron torch. Thirdly, the extension of numerical tools developed at the von Karman Institute, required within the methodology for the determination of catalycity properties for thermal protection system materials, has been completed for CO2 flows. Non equilibrium stagnation line computations have been performed for several outer edge conditions in order to analyze the influence of the chemical models for bulk reactions. Moreover, wall surface reactions have been examined, and the importance of several recombination processes has been discussed. This analysis has revealed the limits of the model currently used, leading to the proposal of an alternative approach for the description of the flow-surface interaction. Finally the effects of outer edge elemental fractions on the heat flux map is analyzed, showing the need to add them to the list of parameters of the methodology currently used to determine catalycity properties of thermal protection materials.

<div>A sharp, circular 7° half-angle cone was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel</div><div>at 6° angle of attack, extending several previous experiments on the growth and breakdown of</div><div>stationary crossflow instabilities in the boundary layer. </div><div><br></div><div>Measurements were made using infrared</div><div>imaging and surface pressure sensors. Detailed measurements of the stationary and traveling</div><div>crossflow vortices, as well as various secondary instability modes, were collected over a large</div><div>region of the cone.</div><div><br></div><div>The Rod Insertion Method (RIM) roughness, first developed for use on a flared cone, was</div><div>adapted for application to crossflow work. It was demonstrated that the roughness elements were</div><div>the primary factor responsible for the appearance of the specific pattern of stationary streaks</div><div>downstream, which are the footprints of the stationary crossflow vortices. In addition, a roughness</div><div>insert was created with a high RMS level of normally-distributed roughness to excite the naturally</div><div>most-amplified stationary mode.</div><div><br></div><div>The nonlinear breakdown mechanism induced by each type of roughness appears to be</div><div>different. When using the discrete RIM roughness, the dominant mechanism seems to be the</div><div>modulated second mode, which is significantly destabilized by the large stationary vortices. This</div><div>is consistent with recent computations. There is no evidence of the presence of traveling crossflow</div><div>when using the RIM roughness, though surface measurements cannot provide a complete picture.</div><div>The modulated second mode shows strong nonlinearity and harmonic development just prior</div><div>to breakdown. In addition, pairs of hot streaks merge together within a constant azimuthal</div><div>band, leading to a peak in the heating simultaneously with the peak amplitude of the measured</div><div>secondary instability. The heating then decays before rising again to turbulent levels. This nonmonotonic</div><div>heating pattern is reminiscent of experiments on a flared cone and earlier computations</div><div>of crossflow on an elliptic cone.</div><div><br></div><div>When using the distributed roughness there are several differences in the nonlinear breakdown</div><div>behavior. The hot streaks appear to be much more uniform and form at a higher wavenumber,</div><div>which is expected given computational results. Furthermore, the traveling crossflow waves become</div><div>very prominent in the surface pressure fluctuations and weakly nonlinear. In addition there</div><div>appears in the spectra a higher-frequency peak which is hypothesized to be a type-I secondary instability</div><div>under the upwelling of the stationary vortices. The traveling crossflow and the secondary</div><div>instability interact nonlinearly prior to breakdown.</div>

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. 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. 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. 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.

abstract: This dissertation considers an integrated approach to system design and controller design based on analyzing limits of system performance. Historically, plant design methodologies have not incorporated control relevant considerations. Such an approach could result in a system that might not meet its specifications (or one that requires a complex control architecture to do so). System and controller designers often go through several iterations in order to converge to an acceptable plant and controller design. The focus of this dissertation is on the design and control an air-breathing hypersonic vehicle using such an integrated system-control design framework. The goal is to reduce the number of system-control design iterations (by explicitly incorporate control considerations in the system design process), as well as to in&#64258;uence the guidance/trajectory specifications for the system. Due to the high computational costs associated with obtaining a dynamic model for each plant con&#64257;guration considered, approximations to the system dynamics are used in the control design process. By formulating the control design problem using bilinear and polynomial matrix inequalities, several common control and system design constraints can be simultaneously incorporated into a vehicle design optimization. Several design problems are examined to illustrate the effectiveness of this approach (and to compare the computational burden of this methodology against more traditional approaches). / Dissertation/Thesis / Ph.D. Electrical Engineering 2014