No description available.

No description available.

Not available / text

For direct reentry from a lunar or deep space mission, considerable variation in reentry plane, reentry point, and reentry angle must be anticipated. The returning vehicle must therefore, possess the ability to control its range after reentry in order to touch down in the desired recovery area. Recent studies have indicated that considerable ranging capability is available with even low lift-drag ratio vehicles operating wholly within the atmosphere if aerodynamic maneuvering is initiated while the vehicle still possesses greater than satellite velocity. In these studies, maneuvering was initiated shortly after the initial pull-up. Range control is also available during the initial pull-up, but such control results in little gain in longitudinal ranging capability in most cases. It is the purpose of the present thesis to investigate the feasibility of increasing lateral ranging capability by banking during the initial pull-up. Low lift-drag ratio vehicles reentering the atmosphere in a banked attitude are considered and the effects of such reentries on allowable corridor width, and lateral range capability are studied. Equations are developed for the motion of a vehicle reentering the atmosphere of a spherical, non-rotating earth, and some permissible approximations applicable for the present problems are discussed. Numerical results obtained for the developed system of equations through use of an IBM 7090 high-speed computer are used throughout the investigation to furnish accurate evaluations of the effects being studied and to check the validity of some of the approximations used. Particular emphasis is placed on reentry at escape velocity, but the effects determined apply in character to reentry at other supercircular velocities. / Master of Science

The terminal guidance problem for a fixed-trim re-entry body is formulated with the objective of synthesizing a closed-loop steering law. A transformation of variables is made that reduces the order of the state system for the guidance problem, and a subsequent linearization with motion along the sight line to the target as a reference produces a further order reduction. The final, reduced-order system, although nonlinear and time-varying, is simple enough to lend itself to synthesis of a class of guidance laws. A generalization of the feedforward device of classical control theory is successfully employed for compensation of roll-autopilot lags. A comparison with existing fixed trim guidance laws is carried out computationally with a simulation model idealizing the navigation and control systems as error-free. The proposed guidance law exhibits superior miss-distance performance in the comparison. / Ph. D.

Supersonic deceleration has been identified as a critical deficiency in extending heritage technologies to the high-mass systems required to achieve long-term exploration goals at Mars. Supersonic retropropulsion (SRP), or the use of retropropulsive thrust while an entry vehicle is traveling at supersonic conditions, is an approach addressing this deficiency. The focus of this dissertation is aerodynamic and performance evaluation of SRP as a decelerator technology for high-mass Mars entry systems. This evaluation was completed through a detailed SRP performance analysis, establishment of the relationship between vehicle performance and the aerodynamic-propulsive interaction, and an assessment of the required fidelity and computational cost in simulating SRP flowfields, with emphasis on the effort required in conceptual design. Trajectory optimization, high-fidelity computational aerodynamic analysis, and analytical modeling of the SRP aerodynamic-propulsive interaction were used to define the fidelity and effort required to evaluate individual SRP concepts across multiple mission scales.

Thermal Protection System (TPS) is required to shield an atmospheric entry vehicle against the high surface heating environment experienced during hypersonic flight. There are significant uncertainties in the tools and models currently used for the prediction of entry aeroheating and TPS material thermal response. These uncertainties can be reduced using experimental data. Analysis of TPS ground and flight data has been traditionally performed in a direct fashion. Direct analyses center upon comparison of the computational model predictions to data. Qualitative conclusions about model validity may be drawn based on this comparison and a limited number of model parameters may be iteratively adjusted to obtain a better match between predictions and data. The goal of this thesis is to develop a more rigorous methodology for the estimation of surface heating and TPS material response using inverse estimation theory. Built on theoretical developments made in related fields, this methodology enables the estimation of uncertainties in both the aeroheating environment and material properties from experimental temperature data. Unlike direct methods, the methodology developed here is capable of estimating a large number of independent parameters simultaneously and reconstructing the time-dependent surface heating profile in an automated fashion. This methodology is applied to flight data obtained from thermocouples embedded in the Mars Pathfinder and Mars Science Laboratory entry vehicle heatshields.

In order to investigate the thermal stress and expansion as well as the associated strain effect on material properties caused by high temperature and large temperature gradient, a two-way thermo-mechanical coupling solver is developed. This solver integrates a new structural response module to the Kentucky Aerothermodynamics and Thermal response System (KATS) framework. The structural solver uses a finite volume approach to solve either hyperbolic equations for transient solid mechanics, or elliptic equations for static solid mechanics. Then, based on the same framework, a quasi-static approach is used to couple the structural response and thermal response to estimate the thermal expansion and stress within Thermal Protection System (TPS) materials. To better capture the thermal expansion and study its impacts on material properties such as conductivity and porosity, a moving mesh scheme is also developed and incorporated into the solver. Grid deformation is transferred among different modules in the form of variations of geometric parameters and strain effects. By doing so, a bi-direction information loop is formed to accomplish the two-way strong thermo-mechanical coupling. Results revealed that the thermal stress experienced during atmospheric re-entry concentrates in a banded area at the edge of the pyrolysis zone and its magnitude can be large enough to cause the failure of the TPS. In addition, thermal expansion causes the whole structure to deform and the changes in material properties. Results also indicated that the impacts coming from structural response should not be ignored in thermal response.

The present thesis provides a description of historical and current modeling methods with recent discoveries within the ablation community. Several historical assumptions are challenged, namely the presence of water in Thermal Protection System (TPS) materials, presence of coking in TPS materials, non-uniform elemental production during pyrolysis reactions, and boundary layer gases, more specifically oxygen, interactions with the charred carbon interface. The first topic assess the potential effect that water has when present within the ablator by examining the temperature prole histories of the recent flight case Mars Science Laboratory. The next topic uses experimental data to consider the instantaneous gas species produced as the ablator pyrolyzes. In this study, key gas species are identified and assumed to be unstable within the gas phase; thus, equilibrating to the solid phase. This topic investigates the potential effects due to the these process. The finial topic uses a simplified configuration to study the role of carbon oxidation, from diatomic oxygen, on the ablation modes of a TPS, surface versus volumetric ablation. Although each of these topics differ in their own right, a common theme is found by understanding the role that common pyrolysis and boundary layer gases species have as they interacts with the porous TPS structure. The main objective of the present thesis is to investigate these questions.

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.