This paper focused on two significant space forces that can affect the success of a spacecraft: the radiation and micrometeoroid environments. Both are looked at in the context of the region of space between Earth and Mars. The goal was create reference environments, to provide context to results of environmental modeling, and to provide recommendations to assist in early design decisions of interplanetary spacecraft. The radiation section of this report used NASA's OLTARIS program to generate data for analysis. The area of focus was on the radiation effects for crewed missions, therefore effective dose equivalent was the metric used to compare different models of radiation and shielding. Test spheres with one, two, or three different materials layers were compared, along with modifiers such as alloys or weight vs. thickness emphasis. Results were compared to limits set by the European and Russian Space Agencies to provide context. The results hinged heavily on the intensity of the Solar Particle Events (SPEs), with testing using additional temporary radiation shielding proving to be a requirement for feasible shielding masses. Differences in shield material effectiveness were found to be negligible for thin Galactic Cosmic Rays (GCRs) and thick SPEs. Thick shields were found to perform better when the more efficient shield was on the outside of the test sphere. The micrometeoroid section used equations and programs from multiple sources to generate state vectors, flux, and finally impact models for four different case studies. Impacts v were generated with mass, velocity, and impact angle/location statistics. The mass and velocity results were run through statistical software to generate information such as mean and standard deviation with confidence intervals. Also looked at were higher mass impacts, limited to above 10-3 grams as opposed to above 10-6 for the regular case. The results of this show that very thin monolithic shields (0.1 cm-0.25 cm) could protect against the average 10-6 impact. The Ram, Nadir, and Anti-sun faces received the highest quantity of impacts and Wake received the least. When looking at the worst cases average mass and velocity for the high mass impacts significantly higher shielding was required to prevent penetration (up to 5 cm for some cases). However, the test cases had probabilities of no high mass impacts greater than 46% of the time, with shorter mission having greater chances of no high mass impacts.

A transformation between minimum dimension adjoint variables and redundant adjoint variables is derived in this dissertation. The transformation is then applied between the adjoint variables associated with Cartesian position and velocity vectors and a set of redundant adjoint variables associated with certain regularized variables (Schumacher variables). This transformation proves to be very beneficial when it is applied to minimum-fuel space rendezvous and intercept problems. It facilitates using attributes from the two systems simultaneously; a new necessary condition in Schumacher adjoints is derived in this dissertation, and this together with classical necessary conditions for fuel-optimal transfer (existing in the position and velocity space) leads to a numerical algorithm which seems to be quite robust in finding candidate optimal control solutions for space transfer problems. / Ph. D.

This thesis investigates the implementation of a passive isolator with a pressurized air cushion for spacecraft payloads in mission architectures implementing in-space assembly technologies. A pressurized air bed capable of briefly surviving the space environment for cargo delivery was prototyped and experimentally evaluated for launch vehicle vibration dynamics resulting in a 72%, 93%, and 88% reduction in experienced GRMS loads for the X-Axis, Y-Axis, and Z-Axis, respectively. A preliminary Total Mass Loss evaluation of the Low-Density Polyethylene Film utilized for the air bed resulted in a mass loss of 0.7%, indicating that commercial off-the-shelf films might require minimal modification for flight readiness. An analytical model of a planar rectangular payload experiencing free vibrations with a Winkler foundation is generated and compared to the experimental results, showing a potential way for characterizing and designing such a foundation to reduce experienced vibrations. These preliminary results show a potential path for a non-cost-prohibitive method for space payloads to reduce loads experienced during launch as inspired by the successful hosted payloads program aboard the International Space Station.

This research involves review of value analysis and value engineering techniques and application of these techniques to a service organization. Application of value analysis techniques to service organizations required some modification due to the amount of labor intensive, manual operations typical of many service industries. The modified value analysis methodology was applied to a fueling operation performed by Kennedy Space Center as part of their space shuttle operations. The successful application of this technique illustrated that value analysis methodology can be applied to service organizations with slight modifications.

The strength reduction of a graphite-epoxy laminate due to dropping plies is investigated experimentally and analytically. Laminates were tested under uniaxial tension and compression. All the laminates were flat on one side with the plies dropped from the middle of a [( ± 45/0/90)_s (N_d) ( ± 45/0/90)_s]_T laminate, where N_d denotes the number of dropped-plies and their orientations. A total of 54 specimens were tested consisting of eight dropped-ply configurations, or values of N_d. This geometry creates an eccentric load path which causes local bending moments in the region of the ply drop-off. The strength of a laminate with dropped-plies is less than the strength of its thin section, and the compression specimens exhibited a lower strength than a tension specimen of the same configuration and width. For the laminates in this study, the reduction in strength is directly related to the axial stiffness change between the thick and thin sections of the laminate. The three-dimensional state of stress in the laminate was evaluated by the finite element method. The magnitude of the interlaminar stresses at the ply drop-off for N_d = [0₄]_T are greater than for N_d = [90₄]_T . The initial failure event for N_d = [0₄]_T was a delamination between the dropped plies and upper sublaminate at the drop-off. The tensile interlaminar failure criterion predicts this as the critical location in the finite element model as well. However, the tensile interlaminar criterion underestimates the failure initiation load. The cause for this may be due to the inaccuracies in the modeling of the ply drop-off geometry in the finite element analysis. The magnitude of the stresses at the ply drop-off are sensitive to changes in the finite element mesh geometry. The N_d = [90₄]_T laminates were stronger than the N_d = [0₄]_T laminates and the N_d = [90₄]_T laminates failed in the thin section away from the drop-off. / M.S.

This work develops a methodology to facilitate the optimum design of spacecraft power processing systems. Emphasis is placed on the battery charge and discharge systems. A comparison of several battery charge and discharge topologies is presented. Characteristics which effect the overall system performance are addressed including size, weight, efficiency, dynamic performance, electromagnetic interference, and reliability. A detailed comparison, using nonlinear design optimization techniques, is performed for three candidate topologies for application to the NASA Earth Observing System (EOS) satellite battery discharger. Experimental verification is provided. A novel zero-voltage-switching (ZVS) bidirectional converter topology is presented. A bidirectional battery charger / discharger has several advantages, including a potentially substantial system weight savings. However, it is shown that most of the advantages can not be realized unless zero-voltage-switching is incorporated. The design of a novel ZVS, multi-module, multi-phase, bidirectional converter with application to the NASA EOS satellite battery charger / discharger system is presented. The system not only yields high efficiency and light weight, but also possesses superior dynamic characteristics when compared to either a conventional bidirectional converter or separate charge and discharge converters. Design considerations for system control of the NASA EOS satellite are presented. A methodology is presented to insure system stability with an unknown complex load. / Ph. D.

Large-signal analysis of various spacecraft power systems is performed to predict the bus dynamics in various modes of operation. The large-signal trajectories of the system's operating point are analyzed employing qualitative graphical representation. The analyses are verified through simulation using EASY5 software. Small-signal dynamic characteristics of spacecraft battery charge converter systems are analyzed to facilitate the design of control loop for optimum performance and stability. Control-loop designs for the charge converters in bus voltage regulation mode, charge current regulation mode, and peak power tracking mode are discussed. / Ph. D.

Pontryagin's principle is applied to several significant problems associated with optimal large angle spacecraft rotational maneuvers. Both rigid and flexible body dynamical models for these vehicles are considered. Three relaxation/analytic continuation methods are developed for iteratively solving the two-point-boundary value problem which results in the treatment of these problems. The solutions obtained are required to rigorously satisfy the necessary conditions derived from Pontryagin's principle. These methods include: (1) boundary condition relaxation processes; (2) differential equation relaxation processes; and (3) hybrid relaxation processes, combining (1) and (2) above. In the literature these relaxation processes are closely related to a number of methods for solving nonlinear equations, known as Davidenko's method, imbedding, and homotopy chain methods. For rigid vehicles a general nonsingular optimal maneuver formulation is obtained, treating all kinematic and dynamical nonlinearities, for general orientation and angular velocity boundary conditions. For flexible vehicles restricted to single axis maneuvers and anti-symmetric elastic deflection modes, a general optimal maneuver formulation is obtained; treating all kinematic, dynamical, and first order structural nonlinearities. In the case of general motion for a flexible vehicle a general formulation is provided, though a solution is not obtained; due to a previously unidentified and as yet unresolved computational difficulty associated symmetry in the dynamical model for the spacecraft. / Ph. D.

Results of an experimental study of the failure and crippling of thin-walled open section prismatic compression members are presented. Twenty-four specimens were tested, 13 of which were channel sections and the other 11 were zee sections. Three specimens were made of 2024-T3 aluminum. The remaining 21 specimens were made of AS4-3502 graphite-epoxy. All specimens were tested to failure. Seventeen specimens exhibited local buckling of flanges and webs prior to failure, four exhibited global column buckling prior to failure, two exhibited material short column failure, and one exhibited unstable postbuckling behavior prior to failure. The buckling loads for each specimen were also calculated by a computer code and compared to experimental buckling loads. Good correlation was achieved for specimens that did not buckle as columns. The graphite-epoxy specimens which buckled locally had significant postbuckling response prior to failure at a maximum load (crippling). Differences in the crippling failure and compressive strength failure are discussed for the graphite-epoxy specimens. / Master of Science

A battery charger design for the NASA Earth Observing System (EOS) Space Platform has been developed and tested. This thesis discusses the design of the battery charger power stage and its current and voltage control loops. The charger was designed to minimize the mass and to maximize the efficiency. In addition to restoring energy to the batteries, this charger regulates the spacecraft bus voltage during the transition between eclipse and sunlight. The battery charger design and analysis was facilitated by use of the model for the pulse-width-modulated (PWM) switch and the new continuous-time model for current-mode control. Analyses of the battery charger small-signal behavior are compared to hardware measurements to verify modeling accuracy. / Master of Science