Modeling the atmospheric entry of spacecraft is challenging because of the large number of physical phenomena that occur during the process. In order to study thermal protection systems, engineers rely on high fidelity solvers to provide accurate predictions of both the thermochemical environment surrounding the heat shield, and its material response. Therefore, it is necessary to guarantee that the numerical models are correctly implemented and thoroughly validated. In recent years, a high-fidelity modeling tool has been developed at the University of Kentucky for the purpose of studying atmospheric entry. The objective of this work is to verify and validate this code. The verification consists of the development of an automated regression testing utility. It is intended to both aid code developers in the debugging process, as well as verify the correct implementation of the numerical models as these are developed. The validation process will be performed through comparison to relevant ablation experiments, namely arc-jet tests. Two modules of the code are used: fluid dynamics, and material response. First the fluid dynamics module is verified against both computational and experimental data on two distinct arc-jet tests. The material response module is then validated against arc-jet test data using PICA.

The objective of this thesis is to predict shape change due to ablation and to find temperature distribution of the thermal protection system of a supersonic vehicle under aerodynamic heating by using finite element method. A subliming ablative is used as thermal protection material. Required material properties for the ablation analyses are found by using DSC (Differential Scanning Calorimetry) and TGA (Thermogravimetric Analysis) thermal analysis techniques. DSC is a thermal analysis technique that looks at how a material&#039 / s specific heat capacity is changed by temperature and TGA is a technique in which the mass of a substance is monitored as a function of temperature. Moreover, oxyacetylene ablation tests are conducted for the subliming ablative specimens and measured recession values are compared with the analytically calculated values. Maximum difference between experimental results and analytical results is observed as 3% as seen in Table 7. For the finite element analyses, ANSYS Software is used. A numerical algorithm is developed by using programming language APDL (ANSYS Parametric Design Language) and element kill feature of ANSYS is used for simulation of ablation process. To see the effect of mesh size and time step on the solution of analyses, oxyacetylene test results are used. Numerical algorithm is also applied to the blunt-nosed section of a supersonic rocket which is made from subliming ablative material. Ablation analyses are performed for the nose section because nose recession is very important for a rocket to follow the desired trajectory and nose temperature is very important for the avionics in the inner side of the nose. By using the developed algorithm, under aerodynamic heating, shape change and temperature distribution of the nose section at the end of the flight are obtained. Moreover, effects of ablation on the trajectory of the rocket and on the flow around the rocket are examined by Missile DATCOM and CFD (computational fluid dynamics) analysis tools.

Heatshield design and analysis has traditionally been a decoupled process, the designer creates the geometry generally without knowledge about how the design variables affect the thermostructural response or how the system will perform under off nominal conditions. Heatshield thermal and structural response analyses are generally performed as separate tasks where the analysts size their respective components and feedback their results to the designer who is left to interpret them. The analysts are generally unable to provide guidance in terms of how the design variables can be modified to meet geometric constraints and not exceed the thermal or structural design specifications. In general, the thermal response analysis of ablative thermal protection systems has traditionally been performed using a one-dimensional finite difference calculation. The structural analyses are generally one, two, or three-dimensional finite element calculations. In this dissertation, the governing differential equations for ablative thermal response are solved in three-dimensions using the finite element method. Darcy' Law is used to model the flow of pyrolysis gas through the ablative material. The three-dimensional governing differential equations for Darcy flow are solved using the finite element method as well. Additionally, the equations for linear elasticity are solved by the finite element method for the thermal stress using temperatures directly from the thermal response calculations. This dissertation also links the analysis of thermal protection systems to their design. The link to design comes from understanding the variation in the thermostructural response over the range of the design variables. Material property sensitivities are performed and an optimum design is determined based on a deterministic analysis minimizing the design specification of bondline temperature subject to appropriate constraints. A Monte Carlo simulation is performed on the optimum design to determine the probability of exceeding the design specifications. The design methodology is demonstrated on the Orion Crew Exploration Vehicle's compression pad design.

The thermal response analysis of an ablative material on a two dimensional external surface is performed. The method is applied to both rectangular and cylindrical coordinate systems, where rectangular coordinate system is used for comparison with results available in literature. The current study solves the decomposition of the material at high temperatures by using the nth order Arrhenius equation but excludes the removal of char from the surface due to mechanical erosion or phase change and considers that the ablation process takes place in a finite zone. The method considers the whole domain as one computational domain, eliminating the necessity to check the positions of the start and end of decomposition zone. The decomposition of pyrolysis gases and/or char that may occur at high temperatures and the chemical reaction between pyrolysis gases and char is neglected while pyrolysis gases are assumed to behave as ideal gas. The pressure is taken as a constant value on a whole physical domain. The formulation for one-dimensional case is validated by experimental results obtained from literature. The two-dimensional case in a Cartesian geometry is formulated and an algebraic transformation is used to normalize the region in both directions and transformed at same time into a square computational domain in order to get a solution for the variable thickness domains. The formulation for two-dimensional case is revised for the cylindrical coordinates with finite length in the axial direction. To solve geometries where the outer surface deviates from cylindrical, the formulation is scaled and transformed into a non-dimensional square computational domain. The method is also applied to a two layer material problem in axisymmetric geometry. In all problems, the radiation and constant heat flux boundary conditions exist on the outer surface while whole domain is initially at a constant temperature. Case studies are performed to demonstrate the application of the solution method in optimizing the insulation material thickness.

The Space Shuttle Columbia accident and the recent excitement surrounding Discovery's return to space brought excessive media attention to the foam products used on the External Tank (ET). In both cases, videos showed chunks of foam or ablative material falling away from the ET during lift off. This led to several years of investigation and research into the exact cause of the accident and potential solutions to avoid the problem in the future. Several design changes were made prior to the return to flight this year, but the ET still shed foam during lift off. Since the Columbia accident, the loss of foam on ETs has been a significant area of interest for NASA, United Space Alliance, and Lockheed Martin. The Columbia Accident Investigation Board did not evaluate alternative materials but certainly highlighted the need for change. The majority of the research previously concentrated on improving the design and/or the application process of the current materials. Within recent years, some research and testing has been done to determine if a glass microsphere composite foam would be an acceptable alternative, but this work was overcome by the need for immediate change to return the shuttle to flight in time to deliver supplies to the International Space Station. Through a better understanding of the foam products currently used on the ET, other products can be evaluated for future space shuttle flights and potential applications on new space vehicles. The material properties and the required functionality of alternative materials can be compared to the current materials to determine if suitable replacement products exist. This research also lends itself to the development of future space flight and unmanned launch vehicles. In this paper, the feasibility of alternative material for the space shuttle's external tank will be investigated. Research on what products are used on the ET and a set of functional requirements driving the selection of those materials will be presented. The material properties of the current ET foam products will be collected and an evaluation of how those materials' properties meet the functional requirements will be accomplished. Then significant research on polymeric foams and ablative materials will be completed to learn how these various products can be applied in this industry. With this research and analysis, the knowledge gained will be used to select and evaluate the effectiveness of an alternate product and to determine feasibility of a product change with the current ET and the importance of maintaining the shuttle launch schedule. This research will also be used to evaluate the potential application of the alternative product on future platforms. There are several possible outcomes to this research. This research could result in a recommended change to the ET foam material or a perfectly acceptable alternative material that could result in a cost or schedule impact if implemented. It is also possible that there exists no suitable alternative material given the existing functional requirements. In any case, the alternative material could have future applications on new space vehicles. A set of results from the research and analysis will be provided along with a recommendation on a future material for use on space vehicles.

This study researched the history of Space Shuttle Reusable Surface Insulation which was designed and developed for use on the United States Orbiter fleet to protect from the high heating experienced during reentry through Earth's atmosphere. Specifically the tile system which is attached to the structure by the means of an RTV adhesive has experienced situations where the bonds are identified as subnominal. The history of these subnominal conditions is presented along with a recent identification of a subnominal bond between the Strain Isolation Pad and the tile substrate itself. Tests were run to identify the cause of these subnominal conditions and also to show how these conditions were proved to be acceptable for flight. The study also goes into cases that could be used to identify subnominal conditions on tile as a non-destructive test prior to flight. Several options of non-destructive testing were identified and recommendations are given for future research into this topic. A recent topic is also discussed in the instance where gap fillers were identified during the STS-114 mission that did not properly adhere to the substrate. The gap fillers were found protruding past the Outer Mold Line of the vehicle which required an unprecedented spacewalk to remove them to allow for a safe reentry through the atmosphere.

This thesis examines the modeling of the shock-layer radiative heating associated with hypersonic vehicles entering the atmospheres of Earth and Titan. For Earth entry, flight conditions characteristic of lunar-return are considered, while for Titan entry, the Huygens probe trajectory is considered. For both cases, the stagnation region flowfield is modeled using a two-temperature chemical nonequilibrium viscous shock layer (VSL) approach. This model is shown to provide results that are in agreement with the more computationally expensive Navier-Stokes solutions. A new radiation model is developed that applies the most up-to-date atomic and molecular data for both the spectrum and non-Boltzmann modeling. This model includes a new set of atomic-lines, which are shown to provide a significant increase in the radiation (relative to previous models) resulting from the 1 - 2 eV spectral range. A new set of electronic-impact excitation rates was compiled for the non-Boltzmann modeling of the atomic and molecular electronic states. Based on these new rates, a novel approach of curve-fitting the non-Boltzmann population of the radiating atomic and molecular states was developed. This new approach provides a simple and accurate method for calculating the atomic and molecular non-Boltzmann populations. The newly-developed nonequilibrium VSL flowfield and nonequilibrium radiation models were applied to the Fire II and Apollo 4 cases, and the resulting radiation predictions were compared with the flight data. For the Fire II case, the present radiation-coupled flowfield model provides intensity values at the wall that predicted the flight data better than any other previous study, on average, throughout the trajectory for the both the 0.2 - 6.0 eV and 2.2 - 4.1 eV spectral ranges. The present results over-predicted the calorimeter measurements of total heat flux over most of the trajectory. This was shown to possibly be a result of the super-catalytic assumption for the wall boundary condition, which caused the predicted convective heating to be too high. For the Apollo 4 case, over most of the trajectory the present model over-predicted the flight data for the wall radiative intensity values between 0.2 - 6.2 eV. For the analysis of Huygens entry into Titan, the focus of the radiation model was the CN violet band. An efficient and accurate method of modeling the radiation from this band system was developed based on a simple modification to the smeared rotational band (SRB) model. This modified approach, labeled herein as SRBC, was compared with a detailed line-by-line (LBL) calculation and shown to compare within 5% in all cases. The SRBC method requires many orders-of-magnitude less computational time than the LBL method, which makes it ideal for coupling to the flowfield. The non-Boltzmann modeling of the CN electronic states, which govern the radiation for Huygens entry, is discussed and applied. The radiation prediction resulting from the non-Boltzmann model is up to 70% lower than the Boltzmann result. A new method for treating the escape factor in detail, rather than assuming a value equal to one, was developed. This treatment is shown to increase the radiation from the non-Boltzmann model by about 10%. / Ph. D.

No description available.

The focus of this dissertation is aimed at advancing autonomous structural health monitoring. All the research is based on developing the impedance method for monitoring structural health. The impedance technique utilizes piezoelectric patches to interrogate structures of interested with high frequency excitations. These patches are bonded directly to the structure, so information about the health of the structure can be seen in the electrical impedance of the piezoelectric patch. However, traditional impedance techniques require the use of a bulky and expensive impedance analyzer. Research presented here describes efforts to miniaturize the hardware necessary for damage detection. A prototype impedance-based structural health monitoring system, incorporating wireless based communications, is fabricated and validated with experimental testing. The first steps towards a completely autonomous structural health monitoring sensor are also presented. Power harvesting from ambient energy allows a prototype to be operable from a rechargeable power source. Aerospace vehicles are equipped with thermal protection systems to isolate internal components from harsh reentry conditions. While the thermal protection systems are critical to the safety of the vehicle, finding damage in these structures presents a unique challenge. Impedance techniques will be used to detect the standard damage mechanism for one type of thermal protection system. The sensitivity of the impedance method at elevated temperatures is also investigated. Sensors are often affixed to structures as a means of identifying structural defects. However, these sensors are susceptible to damage themselves. Sensor diagnostics is a field of study directed at identifying faulty sensors. The influence of temperature on these techniques is largely unstudied. In this dissertation, a model is generated to identify damaged sensors at any temperature. A sensor diagnostics method is also adapted for use in developed hardware. The prototype used is completely digital, so standard sensor diagnostics techniques are inapplicable. A new method is developed to work with the digital hardware. / Ph. D.

Em função das limitações apresentadas pelos relés eletromecânicos, a proteção térmica de motores foi historicamente tratada como um problema de coordenação de sobrecorrente, sem levar em consideração a dinâmica e o histórico térmico envolvido na operação contínua do motor. Os atuais relés microprocessados para proteção de motores implementam equações diferenciais de primeira ordem, cujos algoritmos, processados em tempo real, possibilitam uma nova abordagem para uma adequada proteção térmica, utilizando modelos matemáticos. Especialmente para os motores industriais de grande porte e de maior importância operacional, somente os recentes relés de proteção microprocessados e seus algoritmos digitais tem sido efetivamente capazes de fornecer proteção adequada, baseados em modelos térmicos que realisticamente estimam, continuamente e em tempo real, o nível térmico atual do motor. A proteção térmica de motores de indução trifásicos tem sido uma das maiores áreas onde a proteção numérica, baseado em microprocessadores, tem proporcionado um aprimoramento do nível básico das funções de proteção de motores. O método da proteção térmica tem sido aperfeiçoado, de forma a implementar modelos que levam em consideração o aquecimento do motor devido às correntes de seqüência positiva e negativa e as características térmicas de um motor de indução. A capacidade do processamento digital de sinais tem possibilitado a implementação de novas soluções para as deficiências de proteção de motores industriais trifásicos apresentadas pelas tecnologias convencionais de proteção, até então fundamentadas em proteção de sobrecorrente. As principais funções de proteção aplicáveis para motores trifásicos industriais, bem como os aspectos do estado da arte de hardware, software e filtros digitais implementados nos atuais relés de proteção microprocessados são discutidos neste trabalho. O equacionamento de um sistema térmico de primeira ordem e os requisitos de modelo para a implementação da proteção térmica de motores são também aqui analisados. São discutidas as dinâmicas de dois modelos térmicos, um baseado em proteção por sobrecorrente e outro baseado em um sistema térmico de primeira ordem. São simulados e comparados os desempenhos destes dois diferentes algoritmos de proteção térmica de motores, quando submetidos às correntes de carga e de sobrecarga, tanto constantes como cíclicas. / On account of the limitations presented for the electromechanical relays, the motor thermal protection was historically treated as an overcurrent coordination issue, without taking into account the dynamics and the thermal historical involved in the process. The modern microprocessor-based relays for motor protection implement discrete time first-order differential equations, whose algorithms, based on the power of the real time signal processing, make possible a new approach for a proper thermal protection, applying mathematical models. Especially for large and critical operational significance industrial motors, only the recent numerical relays for motor protection and its digital algorithms has been efficiently suitable to provide an adequate protection, based in thermal models that realistically take into account, continuously and in real time, the actual motor thermal level. The thermal protection of three-phase induction motors has been one of the biggest areas where the numerical protection, based in microprocessor-based relays, has provide an improvement of the basic level of the motor protection functions. The method of the thermal protection has been improved, in such wise as to implement models that take into account the motor heating due to both positive and negative sequence currents, and the thermal characteristics of an induction motor. The capacity of the digital signal processing has made possible the implementation of new solutions for the deficiencies of three-phase industrial motors protection, established on the conventional protection technologies, till then based on overcurrent protection. The main applicable protection functions for industrial three-phase motors, as well as the aspects of the state of the art of the hardware, software and digital filters implemented in the actual microprocessor-based protection relays are discussed in this work. The derivation of a first-order thermal system and the requirements of model for the implementation of the motor thermal protection also are studied in this work. The dynamics of two thermal models, one based in overcurrent protection and another based on a first-order thermal system are analyzed. The performances of these two different algorithms of motor thermal protection are simulated and compared, when subjected to both constants and cyclic, load and overload currents.