For the past four decades, Mars missions have relied on Viking heritage technology for supersonic descent. Extending the use of propulsion, which is required for Mars subsonic deceleration, into the supersonic regime allows the ability to land larger payload masses. Wind tunnel and computational experiments on subscale supersonic retropropulsion models have shown a complex aerodynamic flow field characterized by the interaction of underexpanded jet plumes exhausting from nozzles on the vehicle with the supersonic freestream. Understanding the impact of vehicle and nozzle configuration on this interaction is critical for analyzing the performance of a supersonic retropropulsion system, as deceleration will have components provided by both the aerodynamic drag of the vehicle and thrust from the nozzles. This investigation focuses on the validity of steady state computational approaches to analyze supersonic retropropulsion flowfield structures and their effect on vehicle aerodynamics. Wind tunnel data for a single nozzle and a multiple nozzle configuration are used to validate a steady state, turbulent computational fluid dynamics approach to modeling supersonic retropropulsion. An analytic approximation to determine plume and bow shock structure in the flow field is also developed, enabling rapid assessment of flowfield structure for use in improved grid generation and as a configuration screening tool. Results for both the computational fluid dynamics and analytic approaches show good agreement with the experimental datasets. Potential limitations of the two methods are identified based on the comparisons with available data. Six additional geometries are defined to investigate the extensibility of the analytical model and determine the variation of supersonic retropropulsion performance with configuration. These validation geometries are split into two categories: three geometries with nozzles located on the vehicle forebody at varying nozzle cant angles, and three geometries with nozzles located on the vehicle aftbody at varying nozzle cant angles and number of nozzles. The forebody nozzle configurations show that nozzle cant angle is a significant driver in performance of a vehicle employing supersonic retropropulsion. Aerodynamic drag preservation for a given thrust level increases with increasing cant angle. However, increasing the cant angle reduces the contribution of thrust to deceleration. The tradeoff between these two contributions to the deceleration force is examined, noting that performance improvements are possible with modest nozzle cant angles. Static pitch stability characteristics are investigated for the lowest and highest cant angle configurations. The aftbody nozzle configuration results show that removing the plume flow from the region forward of the vehicle results in less interaction with the bow shock structure. This impacts aerodynamic performance, as the surface pressure remains relatively undisturbed for all thrust values examined. Static pitch stability characteristics for each of the aftbody nozzle configurations are investigated; noting that supersonic retropropulsion for these configurations exhibits a transition point from static stability to instability as a function of this center of mass location along the axis.

This dissertation is concerned with the problem of slewing large flexible structures in space and simultaneously suppressing any vibrations. The equations of motion for a three-dimensional spacecraft undergoing large rigid-body maneuvers are derived. The elastic motions are assumed to remain in the linear range. A method of substructure synthesis is presented which spatially discretizes the equations of motion. A perturbation approach is used to solve the equations of motion. The zero-order equations describing the rigid-body maneuver are independent of the first-order vibration problem which includes small rigid-body motions. The vibration problem is described by linear nonself-adjoint equations with time-dependent coefficients. Minimum-time, single-axis rotational maneuvers are considered. The axis of rotation is not necessarily a principal axis. The optimal maneuver force distribution is proportional to the corresponding rigid-body modes with the mass acting as the control gain. The premaneuver eigenvectors are used as admissible vectors to reduce the degrees of freedom describing the vibration of the spacecraft during the maneuver. Natural control and uniform damping control are used to suppress the vibrations during the maneuver. Actuator dynamics cause a degradation of control performance. The inclusion of the actuator dynamics in the control formulation partially offsets this effect. The performance of these control techniques is adversely affected by actuator saturation but they remain effective. Numerical results are presented for a spacecraft in orbit and in an earth-based laboratory. / Ph. D.

Thermal control in spacecraft will be increasingly important as the spacecraft grows smaller and more compact. Such spacecraft with low thermal mass will have to be designed to retain or reject heat more efficiently. The passive smart radiation device (SRD) is a new type of thermal control material for spacecraft. Current space thermal control systems require heaters with an additional power penalty to maintain spacecraft temperatures during cold swings. Because its emissivity can be changed without electrical instruments or mechanical part, the use of SRD decreases the request of spacecraft power budget. The (SRD) based on VO2 films is one of the most important structures of the functional thermal control surface, being lighter, more advanced and without a moving devices. A large portion of the heat exchange between an object in space and the environment is performed throughout radiation, which is in turn determined by the object surface properties. The modulation device is coated on the spacecraft surface and thus provides a thermal window that can adapt to the changing conditions in orbit. VO2 is well known to have a temperature driven metal to insulator transition ≈ 68ᴼC accompanying a transformation of crystallographic structure, from monoclinic (M-phase, semiconductor) at temperature below 68ᴼC to tetragonal (R-phase, metal) at temperature above 68ᴼC. This transition temperature is accompanied by an increase of infrared reflectivity and a decrease of infrared emissivity with increasing temperature. This flexibility makes VO2 potentially interesting for optical, electrical, and electro-optical switches devices, and as window for energy efficiency buildings applications. This study reports on effect of thickness on VO2 as well as the effect of proton irradiation on VO2 for active smart radiation device (SRD) application. VO2 was deposited on mica by Pulsed laser deposition techniques. The thickness of the film was varied by varying the deposition time. To characterize VO2 the following techniques were performed: XRD, AFM, SEM, TEM, XPS, RBS, RAMAN and transport measurements for optical properties. The effect of proton irradiation was observed using the SEM, where the change in structure, from crystal grains to rods, was observed. / Physics / M.Sc. (Physics)

Thermal control in spacecraft will be increasingly important as the spacecraft grows smaller and more compact. Such spacecraft with low thermal mass will have to be designed to retain or reject heat more efficiently. The passive smart radiation device (SRD) is a new type of thermal control material for spacecraft. Current space thermal control systems require heaters with an additional power penalty to maintain spacecraft temperatures during cold swings. Because its emissivity can be changed without electrical instruments or mechanical part, the use of SRD decreases the request of spacecraft power budget. The (SRD) based on VO2 films is one of the most important structures of the functional thermal control surface, being lighter, more advanced and without a moving devices. A large portion of the heat exchange between an object in space and the environment is performed throughout radiation, which is in turn determined by the object surface properties. The modulation device is coated on the spacecraft surface and thus provides a thermal window that can adapt to the changing conditions in orbit. VO2 is well known to have a temperature driven metal to insulator transition ≈ 68ᴼC accompanying a transformation of crystallographic structure, from monoclinic (M-phase, semiconductor) at temperature below 68ᴼC to tetragonal (R-phase, metal) at temperature above 68ᴼC. This transition temperature is accompanied by an increase of infrared reflectivity and a decrease of infrared emissivity with increasing temperature. This flexibility makes VO2 potentially interesting for optical, electrical, and electro-optical switches devices, and as window for energy efficiency buildings applications. This study reports on effect of thickness on VO2 as well as the effect of proton irradiation on VO2 for active smart radiation device (SRD) application. VO2 was deposited on mica by Pulsed laser deposition techniques. The thickness of the film was varied by varying the deposition time. To characterize VO2 the following techniques were performed: XRD, AFM, SEM, TEM, XPS, RBS, RAMAN and transport measurements for optical properties. The effect of proton irradiation was observed using the SEM, where the change in structure, from crystal grains to rods, was observed. / Physics / M.Sc. (Physics)

A Variable Speed Control Moment Gyro (VSCMG) is a recently introduced actuator for spacecraft attitude control. As its name implies, a VSCMG is essentially a single-gimbal control moment gyro (CMG) with a flywheel allowed to have variable spin speed. Thanks to its extra degrees of freedom, a VSCMGs cluster can be used to achieve additional objectives, such as power tracking and/or singularity avoidance, as well as attitude control. In this thesis, control laws for an integrated power/attitude control system (IPACS) for a satellite using VSCMGs are introduced. The power tracking objective is achieved by storing or releasing the kinetic energy in the wheels. The proposed control algorithms perform both the attitude and power tracking goals simultaneously. This thesis also provides a singularity analysis and avoidance method using CMGs/VSCMGs. This issue is studied for both the cases of attitude tracking with and without a power tracking requirement. A null motion method to avoid singularities is presented, and a criterion is developed to determine the momentum region over which this method will successfully avoid singularities. The spacecraft angular velocity and attitude control problem using a single VSCMG is also addressed. A body-fixed axis is chosen to be perpendicular to the gimbal axis, and it is controlled to aim at an arbitrarily given inertial direction, while the spacecraft angular velocity is stabilized. Finally, an adaptive control algorithm for the spacecraft attitude tracking in case when the actuator parameters, for instance the spin axis directions, are uncertain is developed. The equations of motion in this case are fully nonlinear and represent a Multi-Input-Multi-Output (MIMO) system. The smooth projection algorithm is applied to keep the parameter estimates inside a singularity-free region. The design procedure can also be easily applied to general MIMO dynamical systems.

Hall Effect Thrusters (HETs) are a form of electric propulsion device which uses external electrical energy to produce thrust. When compared to various other electric propulsion devices, HETs are excellent candidates for future orbit transfer and interplanetary missions due to their relatively simple configuration, moderate thrust capability, higher thrust to power ratio, and lower thruster mass to power ratio. Due to the short history of HETs, the current design process of a new HET is a largely empirical and experimental science, and this has resulted in previous designs being developed in a narrow design space based on experimental data without systematic investigations of parameter correlations. In addition, current preliminary low-thrust trajectory optimizations, due to inherent difficulties in solution procedure, often assume constant or linear performances with available power in their applications of electric thrusters. The main obstacles come from the complex physics involved in HET technology and relatively small amounts of experimental data. Although physical theories and numerical simulations can provide a valuable tool for design space exploration at the inception of a new HET design and preliminary low-thrust trajectory optimization, the complex physics makes theoretical and numerical solutions difficult to obtain. Numerical implementations have been quite extensively conducted in the last two decades. An investigation of current methodologies reveals that to date, none provide a proper methodology for a new HET design at the conceptual design stage and the coupled low-thrust trajectory optimization. Thus, in the first half of this work, an efficient, robust, and self-consistent numerical method for the analysis of HETs is developed with a new approach. The key idea is to divide the analysis region into two regions in terms of electron dynamics based on physical intuition. Intensive validations are conducted for existing HETs from 1 kW to 50 kW classes. The second half of this work aims to construct a simultaneous design optimization environment though collaboration with experts in low-thrust trajectory optimization where a new HET and associated optimal low-thrust trajectory can be designed simultaneously. A demonstration for an orbit raising mission shows that the constructed simultaneous design optimization environment can be used effectively and synergistically for space missions involving HETs. It is expected that the present work will aid and ease the current expensive experimental HET design process and reduce preliminary space mission design cycles involving HETs.

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.

Weight growth has been a significant factor in nearly every space and launch vehicle development program. In order to account for weight growth, program managers allocate a design margin. However, methods of estimating design margin are not well suited for the task of assigning a design margin for a novel concept. In order to address this problem, a hybrid method of estimating margin is developed. This hybrid method utilizes range estimating, a well-developed method for conducting a bottom-up weight analysis, and a new forecasting technique known as executable morphological analysis. Executable morphological analysis extends morphological analysis in order to extract quantitative information from the morphological field. Specifically, the morphological field is extended by adding attributes (probability and mass impact) to each condition. This extended morphological field is populated with alternate baseline options with corresponding probabilities of occurrence and impact. The overall impact of alternate baseline options can then be estimated by running a Monte Carlo analysis over the extended morphological field. This methodology was applied to two sample problems. First, the historical design changes of the Space Shuttle Orbiter were evaluated utilizing original mass estimates. Additionally, the FAST reference flight system F served as the basis for a complete sample problem; both range estimating and executable morphological analysis were performed utilizing the work breakdown structure created during the conceptual design of this vehicle.

Thesis (MEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2009. / Autonomous rendezvouz and docking is seen as an enabling technology. It allows, among others, the construction of larger space platforms in-orbit and also provides a means for the in-orbit servicing of space vehicles. In this thesis a docking sequence is proposed and tested in both simulation and practice. This therefore also requires the design and construction of a test platform. A model hovercraft is used to emulate the chaser satellite in a 2-dimensional plane as it moves relatively frictionlessly. The hovercraft is also equipped with a single camera (monocular vision) that is used as the main sensor to estimate the target’s pose (relative position and orientation). An imitation of a target satellite was made and equipped with light markers that are used by the chaser’s camera sensor. The position of the target’s lights in the image is used to determine the target’s pose using a modified version ofMalan’s Extended Kalman Filter [20]. This information is then used during the docking sequence. This thesis successfully demonstrated the autonomous and reliable identification of the target’s lights in the image, and the autonomous docking of a satellite pair using monocular camera vision in both simulation and emulation.

This thesis explores how an Asteroid Sample Return Mission might make use of solar electric propulsion to send a spacecraft on a journey to the asteroid 1989ML and back. It examines different trajectories that can be used to get an asteroid sample return or similar spacecraft to an interplanetary destination and back in the most fuel-efficient manner. While current plans call for keeping such a spacecraft on the asteroid performing science experiments for approximately 90 days, it is prudent to inquire how lengthening or shortening this time period may affect mission fuel requirements. Using optimal control methods, various mission scenarios have been modeled and simulated. The results suggest that the amount of time that the spacecraft may spend on the asteroid surface can be approximated as a linear function of the available fuel mass. Furthermore, It can be shown that as maximum available thrust is decreased, the radial component of the optimal thrust vector becomes more pronounced.