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Spacecraft thermal design is an inverse problem that requires one to determine the choice of surface properties that yield a desired temperature distribution within a satellite. The current techniques for spacecraft thermal design are very much in the frame of trial and error. The goal of this work is to move away from that procedure, and have the thermal design solely dependent on heat transfer parameters. It will be shown that the only relevant parameters to attain this are ones which pertain to radiation. In particular, these parameters are absorptivity and emissivity. We intend to utilize an optimal/analytical approach, and obtain a solution via optimization. As mentioned in the motivation, having a purely passive thermal system will greatly reduce costs, and our optimization solution will enable that. This topic involves heat transfer (conduction and radiation), spacecraft thermal network models, numerical optimization, and materials selection.

The contribution of this thesis is on the temporal adjustment of the consensus weights, as applied to spacecraft formation control. Such an objective is attained by dynamically enforcing attitude synchronization via coupling terms included in each spacecraft controller. It is assumed that each spacecraft has identical dynamics but with unknown inertia parameters and external disturbances. By augmenting a standard adaptive controller that accounts for the unknown parameters, made feasible via an assumption on parameterization, with adaptation of the consensus weights, one opts to improve spacecraft synchronization. The coupling terms, responsible for enforcing synchronization amongst spacecraft, are weighted dynamically in proportion to the disagreement between the states of the spacecraft. The time adjustment of edge-dependent gains as well as the special cases of node-dependent and agent-independent constant gains are derived using Lyapunov redesign methods. The proposed adaptive control architectures which allow for adaptation of both parameter uncertainties and consensus penalty terms are demonstrated via extensive numerical studies of spacecraft networks with limited connectivity. By considering the sum of deviation-from-the-mean and rotational kinetic energy as appropriate metrics for synchronization and controller performance, the numerical studies also provide insights on the choice of optimal consensus gains.

Automated spacecraft docking is a concept of operations with several important potential applications. One application that has received a great deal of attention recently is that of an automated docking capable unmanned re-supply spacecraft. In addition to being useful for re-supplying orbiting space stations, automated shuttles would also greatly facilitate the manned exploration of nearby space objects, including the Moon, near-Earth asteroids, or Mars. These vehicles would allow for longer duration human missions than otherwise possible and could even accelerate human colonization of other worlds. This thesis develops an optimal docking controller for an automated docking capable spacecraft. An innovative vision-based relative navigation system called VisNav is used to provide real-time relative position and orientation estimates, while a Kalman post-filter generates relative velocity and angular rate estimates from the VisNav output. The controller's performance robustness is evaluated in a closed-loop automated spacecraft docking simulation of a scenario in circular lunar orbit. The simulation uses realistic dynamical models of the two vehicles, both based on the European Automated Transfer Vehicle. A high-fidelity model of the VisNav sensor adds realism to the simulated relative navigation measurements. The docking controller's performance is evaluated in the presence of measurement noise, with the cases of sensor noise only, vehicle mass errors plus sensor noise, errors in vehicle moments of inertia plus sensor noise, initial starting position errors plus sensor noise, and initial relative attitude errors plus sensor noise each being considered. It was found that for the chosen cases and docking scenario, the final controller was robust to both types of mass property modeling errors, as well as both types of initial condition modeling errors, even in the presence of sensor noise. The VisNav system was found to perform satisfactorily in all test cases, with excellent estimate error convergence characteristics for the scenario considered. These results demonstrate preliminary feasibility of the presented docking system, including VisNav, for space-based automated docking applications.

The Crew Exploration Vehicle (CEV) necessitates higher levels of automation than previous NASA vehicles due to program requirements for automation, including Automated Rendezvous and Docking (AR&D). Studies of spacecraft development often point to the locus of decision-making authority between humans and computers (i.e. automation) as a prime driver for cost, safety, and mission success. Therefore, a critical component in the CEV development is the determination of the correct level of automation. To identify the appropriate levels of automation and autonomy to design into a human space flight vehicle, NASA has created the Function-specific Level of Autonomy and Automation Tool (FLOAAT). This research develops a methodology for prototyping increased levels of automation for spacecraft rendezvous functions. This methodology was used to evaluate the accuracy of the FLOAAT-specified levels of automation, via prototyping. Two spacecraft rendezvous planning tasks were selected and then prototyped in Matlab using Fuzzy Logic (FL) techniques and existing Shuttle rendezvous trajectory algorithms. The prototyped functions are the determination of the maximum allowable Timeof- IGnition (TIG) slip for a rendezvous phasing burn and the evaluation of vehicle position relative to Transition initiation (Ti) position constraints. The methodology for prototyping rendezvous functions at higher levels of automation is judged to be a promising technique. The results of the prototype indicate that the FLOAAT recommended level of automation is reasonably accurate and that FL can be effectively used to model human decision-making used in spacecraft rendezvous. FL has many desirable attributes for modeling human decision-making, which makes it an excellent candidate for additional spaceflight automation applications. These conclusions are described in detail as well as recommendations for future improvements to the FLOAAT method and prototyped rendezvous functions.

Spacecraft thermal design is an inverse problem that requires one to determine the choice of surface properties that yield a desired temperature distribution within a satellite. The current techniques for spacecraft thermal design are very much in the frame of trial and error. The goal of this work is to move away from that procedure, and have the thermal design solely dependent on heat transfer parameters. It will be shown that the only relevant parameters to attain this are ones which pertain to radiation. In particular, these parameters are absorptivity and emissivity. We intend to utilize an optimal/analytical approach, and obtain a solution via optimization. As mentioned in the motivation, having a purely passive thermal system will greatly reduce costs, and our optimization solution will enable that. This topic involves heat transfer (conduction and radiation), spacecraft thermal network models, numerical optimization, and materials selection.

Thesis (Ph. D.)--University of Washington, 1997. / Vita. Includes bibliographical references (leaves [213]-218).

Thesis (M.S.)--University of Wisconsin--Madison, 1963. / Typescript. eContent provider-neutral record in process. Description based on print version record. Includes bibliographical references (leaves [279]-284).

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1992. / Vita. Abstract. Includes bibliographical references (leaves 76-81). Also available via the Internet.

This research on Reusable Launchers was motivated by the need to reduce substantially the cost of space transportation. The specific objective was to explore the perception that launcher reusability is the key to achieving these major cost reductions. The exploration was achieved by undertaking a comparative system study on potentially feasible reusable launcher concepts, using a consistent set of design tools, a standard analysis methodology and a standard reference mission. To set the background for the research, the results of an extensive literature review are presented on the vehicle studies and technology developments that are engaged across the world on reusable launchers. Comprehensive vehicle studies appear to be engaged without justification for the choice of selected concepts in the absence of results from comparative system studies of reusable launchers. Technology developments also appear to be engaged without clear links to needs derived from vehicle system studies. The challenge of reusability is then addressed. Firstly, to set the performance and cost targets of reusable launchers, the capabilities of current expendable launchers are derived. Secondly, to establish the operational requirements for reusable launchers, the probable space transportation needs for the early 21st century are derived. Thirdly, the concepts and characteristics of reusable launchers are derived, allowing the selection, on a rationale basis, of a short-list of 13 potentially feasible reusable launcher concepts for analysis in the research. The performance equations of reusable launchers are then derived, leading to the preparation of the comparative analysis tools. The major work of the research, which comprises the performance analysis, technical feasibility assessment and cost analysis of each candidate vehicle, are then presented and compared. A set of acceptance requirements for performance, technical feasibility and operational costs of reusable launchers is then derived. The results of the comparative analysis for each candidate launcher are then measured against these requirements. The results of the comparative analysis show that only 2 of the 13 candidate reusable launcher concepts are able to meet all the acceptance requirements. These two acceptable vehicles are both rocket-propelled. They are, in order of preference: a single-stage-to-orbit, rocket-propelled, vertical launch and vertical landing vehicle; a two-stage-to-orbit, rocket-propelled, vertical launch and horizontal landing vehicle. The operational costs per launch for these two vehicles, based on a utilisation plan of 3 vehicles operating for 20 years at a launch rate of 12 launches per year, was calculated to be about 20% of the current costs of the European Ariane 44L expendable launcher. This warrants their further evaluation in a thorough feasibility study. The more complex, air-breathing propelled, horizontal launch and landing vehicles were found to be unable to meet the performance, technical feasibility and cost requirements; Several vehicles were found to be unable to deliver a positive payload mass to orbit; Several vehicles were found to have technology requirements that were deemed to be infeasible to achieve; Several vehicles were found to have operational costs ranging from equal to double that of the European Ariane 44L expendable launcher, which was adopted as a comparative reference vehicle. The contributions of this research to the advancement of knowledge on reusable launchers are: a clear identification of the performance capability limits of 13 plausible reusable launcher concepts; an analysis methodology for determining the performance capability limits for any reusable launcher concept; a clear identification of the reasons for the poor practical performance of air-breathing propulsion systems for Earth-to-orbit launchers, which results from their installed operational characteristics.