The reconstruction of entry, descent and landing (EDL) trajectories is significantly affected by the knowledge of the atmospheric conditions during flight. Away from Earth, this knowledge is generally characterized by a high degree of uncertainty, which drives the accuracy of many important atmosphere-relative states. One method of obtaining the in-flight atmospheric properties during EDL is to utilize the known vehicle aerodynamics in deriving the trajectory parameters. This is the approach taken by this research in developing a methodology for accurate estimation of ambient atmospheric conditions and atmosphere-relative states. The method, referred to as the aerodynamic database (ADB) reconstruction, performs reconstruction by leveraging data from flight measurements and pre-flight models. In addition to the estimation algorithm, an uncertainty assessment for the ADB reconstruction method is developed. This uncertainty assessment is a unique application of a fundamental analysis technique that applies linear covariance mapping to transform input variances into output uncertainties. The ADB reconstruction is applied to a previous mission in order to demonstrate its capability and accuracy. Flight data from the Mars Science Laboratory (MSL) EDL, having successfully completed on August 5th 2012, is used for this purpose. Comparisons of the estimated states are made against alternate reconstruction approaches to understand the advantages and limitations of the ADB reconstruction. This thesis presents a method of reconstruction for EDL systems that can be used as a valuable tool for planetary entry analysis.

During atmospheric entry, a spacecraft's aeroshell uses a thermal protection system (TPS) to withstand severe thermal loads. Heating to the vehicle surface arises as convective, catalytic and radiative heat flux due to the high temperature of the shockwave compressed gases surrounding the aeroshell. The problem for the TPS designer is that the heat load estimates are based on phenomenological models which have questionable validity and, thus, large uncertainty. As an example, recent analyses of heat loads for a proposed aerocapture vehicle designed for Titan differ by up to an order of magnitude. This uncertainty stems from the complexity of the blunt body flow field and the associated physical effects: thermochemical nonequilibrium; ablation and vehicle surface catalycity; and radiating flow. The motivation for this thesis is to develop computational tools that give accurate estimates of vehicle heat transfer as an input for design calculations. With that goal in mind, this thesis work has focussed on one aspect of this problem and that is the modelling of thermochemical nonequilibrium. The longer term goal is to produce tools which can be used to compute the high-temperature, radiating flow fields about aeroshell configurations; the modelling work presented here on thermochemical nonequilibrium effects is a foundation for tackling the radiating flow problem. The modelling work was implemented in an existing flow solver which solves the compressible Navier-Stokes equations with a finite volume method. As part of this work, the flow solver was verified by two methods: the Method of Manufactured Solutions to verify the spatial accuracy for purely supersonic flow; and the Method of Exact Solutions --- the flow problem being an oblique detonation wave --- to verify the spatial accuracy for flows with embedded shocks. Validation of the flow solver, without any of the complexity of thermochemical nonequilibrium, was performed by comparing numerical simulation results to experiments which measured shock detachment on spheres fired into noble gases. A model for chemical nonequilibrium based on the Law of Mass Action and using finite-rate kinetics was coupled with the flow solver. The implementation was verified on two test problems. The first treated a closed-vessel reactor of a hydrogen-iodine mixture, and the second computed the chemically relaxing flow behind a normal shock in air. For validation, the implementation was tested by computing ignition delay times in hydrogen-air mixtures and comparing to experimental results. It was found that the selection of a chemical kinetics scheme can complicate validation, that is, a poor choice of reaction scheme leads to poor computational results yet the implementation is correct. As further validation, a series of experiments on the shock detachment distance on spheres fired into air was compared against numerical simulations based on the present work. Two models for species diffusion were also implemented: Fick's first law approximation and the Stefan-Maxwell equations. These models were verified by comparison to an exact solution for binary diffusion of two semi-infinite slabs. The more general problem of thermochemical nonequilibrium was also pursued. A multi-temperature model, one translational/rotational temperature and multiple vibrational temperatures, was developed as appropriate for hypersonic flows. The model uses the Landau-Teller expression to compute the rate of vibrational-translational energy exchange and the Schwartz-Slawsky-Herzfeld expression for vibrational-vibrational energy exchange. The time constants for the rate expressions are estimated by a number of methods such as the use of SSH theory and the Millikan-White correlation. The coupling of vibrational nonequilibrium effects with the fluid dynamics was tested by computing the flow of nitrogen over an infinite cylinder. The simplified problem of a vibrationally relaxing flow behind a shock, without reactions, was compared to other calculations in the literature. This case tested the multi-temperature formulation, with oxygen and nitrogen each being ascribed their own vibrational temperatures. The coupling of chemistry and vibrational nonequilibrium uses the model by Knab, Fruehauf and Messerschmid. The complete model for thermochemical nonequilibrium was verified by calculating the relaxation of oxygen behind a strong shock. The models developed provide a basis for computing radiating flow fields, however the radiating flow problem cannot be attempted based on this work alone. Instead, a more immediate application of the modelling work was the simulation of expansion tube operation. It is desirable to simulate an impulse facility to give the experimenters access to aspects of experiment that are not directly attainable by experiment; especially a complete characterisation of the test flow properties. The modelling work and code development, as part of this thesis, addresses this need of experimenters. Two large-scale simulations are presented as a demonstration of the modelling work: (a) a simulation of an expansion tube in expansion mode; and (b) a simulation of an expansion tube in nonreflected shock tube mode.

During atmospheric entry, a spacecraft's aeroshell uses a thermal protection system (TPS) to withstand severe thermal loads. Heating to the vehicle surface arises as convective, catalytic and radiative heat flux due to the high temperature of the shockwave compressed gases surrounding the aeroshell. The problem for the TPS designer is that the heat load estimates are based on phenomenological models which have questionable validity and, thus, large uncertainty. As an example, recent analyses of heat loads for a proposed aerocapture vehicle designed for Titan differ by up to an order of magnitude. This uncertainty stems from the complexity of the blunt body flow field and the associated physical effects: thermochemical nonequilibrium; ablation and vehicle surface catalycity; and radiating flow. The motivation for this thesis is to develop computational tools that give accurate estimates of vehicle heat transfer as an input for design calculations. With that goal in mind, this thesis work has focussed on one aspect of this problem and that is the modelling of thermochemical nonequilibrium. The longer term goal is to produce tools which can be used to compute the high-temperature, radiating flow fields about aeroshell configurations; the modelling work presented here on thermochemical nonequilibrium effects is a foundation for tackling the radiating flow problem. The modelling work was implemented in an existing flow solver which solves the compressible Navier-Stokes equations with a finite volume method. As part of this work, the flow solver was verified by two methods: the Method of Manufactured Solutions to verify the spatial accuracy for purely supersonic flow; and the Method of Exact Solutions --- the flow problem being an oblique detonation wave --- to verify the spatial accuracy for flows with embedded shocks. Validation of the flow solver, without any of the complexity of thermochemical nonequilibrium, was performed by comparing numerical simulation results to experiments which measured shock detachment on spheres fired into noble gases. A model for chemical nonequilibrium based on the Law of Mass Action and using finite-rate kinetics was coupled with the flow solver. The implementation was verified on two test problems. The first treated a closed-vessel reactor of a hydrogen-iodine mixture, and the second computed the chemically relaxing flow behind a normal shock in air. For validation, the implementation was tested by computing ignition delay times in hydrogen-air mixtures and comparing to experimental results. It was found that the selection of a chemical kinetics scheme can complicate validation, that is, a poor choice of reaction scheme leads to poor computational results yet the implementation is correct. As further validation, a series of experiments on the shock detachment distance on spheres fired into air was compared against numerical simulations based on the present work. Two models for species diffusion were also implemented: Fick's first law approximation and the Stefan-Maxwell equations. These models were verified by comparison to an exact solution for binary diffusion of two semi-infinite slabs. The more general problem of thermochemical nonequilibrium was also pursued. A multi-temperature model, one translational/rotational temperature and multiple vibrational temperatures, was developed as appropriate for hypersonic flows. The model uses the Landau-Teller expression to compute the rate of vibrational-translational energy exchange and the Schwartz-Slawsky-Herzfeld expression for vibrational-vibrational energy exchange. The time constants for the rate expressions are estimated by a number of methods such as the use of SSH theory and the Millikan-White correlation. The coupling of vibrational nonequilibrium effects with the fluid dynamics was tested by computing the flow of nitrogen over an infinite cylinder. The simplified problem of a vibrationally relaxing flow behind a shock, without reactions, was compared to other calculations in the literature. This case tested the multi-temperature formulation, with oxygen and nitrogen each being ascribed their own vibrational temperatures. The coupling of chemistry and vibrational nonequilibrium uses the model by Knab, Fruehauf and Messerschmid. The complete model for thermochemical nonequilibrium was verified by calculating the relaxation of oxygen behind a strong shock. The models developed provide a basis for computing radiating flow fields, however the radiating flow problem cannot be attempted based on this work alone. Instead, a more immediate application of the modelling work was the simulation of expansion tube operation. It is desirable to simulate an impulse facility to give the experimenters access to aspects of experiment that are not directly attainable by experiment; especially a complete characterisation of the test flow properties. The modelling work and code development, as part of this thesis, addresses this need of experimenters. Two large-scale simulations are presented as a demonstration of the modelling work: (a) a simulation of an expansion tube in expansion mode; and (b) a simulation of an expansion tube in nonreflected shock tube mode.

A numerical method has been developed to predict the three-dimensional nonequilibrium flowfield past the space shuttle orbiter at high angles-of-attack (up to 50-deg). An existing viscous shock-layer method for perfect gas flows has been extended to include finite-rate chemical reactions of multi-component ionizing air. A general nonorthogonal computational grid system was introduced to treat the nonaxisymmetric geometry. At shuttle reentry flight conditions, nonequilibrium real gas effects on the surface-measurable quantities are significant. Computational solutions have been obtained for chemically reacting flowfields over the entire windward surface of the space shuttle orbiter at high angles-of-attack. Boundary conditions studied include noncatalytic wall, finite-catalytic wall, fully-catalytic wall, and nonequilibrium slip conditions at the wall and/or shock. The nonequilibrium solutions with a finite-catalytic wall are compared to both fully-catalytic and noncatalytic wall solutions. The present solutions are also compared to chemical equilibrium air solutions, perfect gas solutions, and the shuttle flight heating and pressure data. The comparisons show good agreement and correlations with flight-derived surface heat-transfer and pressure distributions. Three-dimensional effects are clearly shown in the flight-derived data for the first time based upon the results of this study. / Ph. D.

The present study was undertaken to determine the effects of the aerodynamics on a close earth orbit and reentry trajectory. The aerodynamic influence is compared to the effects of the earth’s oblateness, the sun, and the moon. In order to obtain maximum accuracy and computational speed Encke’s perturbative procedures were used during orbital periods and Cowell’s integration procedures during thrust and reentry periods. / M.S.

A computer program for predicting the three-dimensional nonequilibrium viscous shock-layer flows over blunt spherecones, straight. and bent mul ticonics at angle-of-attack has been developed. The method used is the viscous shock-layer approach- for nonequilibrium, multi-component ionizing air. A seven species chemical reaction model with single ionizing species and an eleven species chemical reaction model with five ionizing species are used to represent the chemistry. The seven species model considers 7 reactions whereas the eleven species model considers 26 reactions and the results obtained using these models are compared with perfect gas and equilibrium air results. This code is capable of analyzing shock-slip or no-shock-slip boundary conditions and equilibrium or non-catalytic wall boundary conditions. In this study the diffusion model is limited to binary diffusion. A sphere-cone-cylinder-flare with moderate flare angle, a straight biconic, and a bent biconic with seven deg. bend angle and a sphere-cone at various flight conditions are analyzed using this method. The bent biconic has been analyzed up to an angle-of-attack of 20 deg. with respect to the aft-cone axis and sample results are compared with inviscid and viscous results. The surface pressure distribution computed by this code compares well with that from a parabolized Navier-Stokes method. The diffusion heat transfer is about 15% of the total heat transfer for most cases. The aerodynamic forces and moments at the base of the body and computing time required for all cases are presented. The shock layer profiles at a streamwi se location of 8. 8 nose radii for one case computed using seven and eleven species models compare very well with each other. / Ph. D.

Cette thèse est centrée sur le couplage entre les phénomènes d’écoulement, d’ablation et de rayonnement au voisinage du point d’arrêt de véhicules d’entrée atmosphérique pourvus d’un système de protection thermique de type carbonephénolique. La recherche est divisée en trois parties : 1) le développement de méthodes numériques et d’outils pour la simulation d’écoulements hypersoniques hors équilibre autour de corps émoussés, 2) la mise en oeuvre d’un nouveau modèle de transport du rayonnement hors équilibre dans ces écoulements, y compris dans les couches limites contaminées par les produits d’ablation, et 3) l’application de ces outils à des conditions réelles de vol.Les effets du couplage entre l’ablation et le rayonnement sont étudiés pour les rentrées terrestres. Il est démontré que les produits d’ablation dans la couche limite peuvent augmenter le blocage radiatif à la surface du véhicule. Pour les conditions de flux maximum d’Apollo 4, les effets de couplage entre le rayonnement et l’ablation réduisent le flux conductif de 35%. L’accord avec les données radiométriques est excellent, ce qui valide partiellement la méthode de couplage et la base de données radiatives. L’importance d’une modélisation précise du soufflage du carbone dans la couche limite est également établie. / This thesis focuses on the coupling between flow, ablation, and radiation phenomena encountered in the stagnation region of atmospheric entry vehicles with carbon-phenolic thermal protection systems. The research is divided into three parts : 1) development of numerical methods and tools for the simulation of hypersonic, non equilibrium flows over blunt bodies, 2) implementation of a new radiation transport model for calculating nonequilibrium radiative heat transfer in atmospheric entry flows, including ablation contominated boundary layers, and 3) application of these tools to study real flight conditions.The effects of coupled ablation and radiation are studied for Earth entries. It’s shown that ablation products in the boundary layer can increase the radiation blockage to the surface of the vehicle. An analysis of the Apollo 4 peak heating condition shows coupled radiation and ablation effects reduce the conducted heat flux by as much as 35% for a fixed wall temperature of 2500 K. Comparison with the radiometer data shows excellent agreement, partially validating the coupling methodology and radiation database. The importance of accurately modeling the amount of carbon blown into the boundary layer is demonstrated by contrasting the results of other researchers.

During hypersonic atmospheric entry, spacecraft are exposed to enormous aerodynamic heat. To prevent the payload from overheating, charring ablative materials are favored to be applied as the heat shield at the exposing surface of the vehicle. Accurate modeling not only prevents mission failures, but also helps reduce cost. Existing models were mostly limited to one-dimensional and discrepancies were shown against measured experiments and flight-data. To help improve the models and analyze the charring ablation problems, a multidimensional material response module is developed, based on a finite volume method framework. The developed computer program is verified through a series of test-cases, and through code-to-code comparisons with a validated code. Several novel models are proposed, including a three-dimensional pyrolysis gas transport model and an orthotropic material model. The effects of these models are numerically studied and demonstrated to be significant.

Space vehicles, while entering the planetary atmosphere, experience high loads of heat. Ablative materials are commonly used for a thermal protection system, which undergo mass removal mechanisms to counter the heat rates. Spallation is one of the ablative processes, which is characterized by the ejection of solid particles from the material into the flow. Numerical codes that are used in designing the heat shields ignore this phenomenon. Hence, to evaluate the effectiveness of spallation phenomenon, a numerical model is developed to compute the dynamics and chemistry of the particles. The code is one-way coupled to a CFD code that models high enthalpy flow field around a lightweight ablative material. A parametric study is carried out to examine the variations in trajectories with respect to ejection parameters. Numerical results are presented for argon and air flow fields, and their effect on the particle behavior is studied. The spallation code is loosely coupled with the CFD code to evaluate the impact of a particle on the flow field, and a numerical study is conducted.

Since the 1970's, NASA has relied on the use of rigid aeroshells and supersonic parachutes to enable robotic mission to Mars. These technologies are constrained by size and deployment condition limitations that limit the payload they can deliver to the surface of Mars. One candidate technology envisioned to replace the supersonic parachute is the supersonic inflatable aerodynamic decelerator (IAD). This dissertation presents an overview of work performed in maturing a particular type of IAD, the tension cone. The tension cone concept consists of a flexible shell of revolution that is shaped so as to remain under tension and resist deformation. Systems analyses that evaluated trajectory impacts of a supersonic IAD demonstrated several key advantages including increases in delivered payload capability of over 40%, significant gains in landing site surface elevation, and the ability to accommodate growth in the entry mass of a spacecraft. A series of supersonic wind tunnel tests conducted at the NASA Glenn and Langley Research Centers tested both rigid and flexible tension cone models. Testing of rigid force and moment models and pressure models demonstrated the new design to have favorable performance including drag coefficients between 1.4 and 1.5 and static stability at angles of attack from 0º to 20º. A separate round of tests conducted on flexible tension cone models showed the system to be free of aeroelastic instability. Deployment tests conducted on an inflatable model demonstrated rapid, stable inflation in a supersonic environment. Structural modifications incorporated on the models were seen to reduce inflation pressure requirements by a factor of nearly two. Through this test program, this new tension cone IAD design was shown to be a credible option for a future flight system. Validation of CFD analyses for predicting aerodynamic IAD performance was also completed and the results are presented. Inviscid CFD analyses are seen to provide drag predictions accurate to within 6%. Viscous analyses performed show excellent agreement with measured pressure distributions and flow field characteristics. Comparisons between laminar and turbulent solutions indicate the likelihood of a turbulent boundary layer at high supersonic Mach numbers and large angles of attack.