The following thesis regards the use of small aperture telescopes for space domain awareness efforts. The rapidly populating space domain was motivation for the development of a new operation scheme to conduct space domain awareness feasibility studies using small telescopes. Two 14-inch Schmidt-Cassegrain Telescopes at the California Polytechnic State University and the Air Force Research Lab in Kirtland AFB, NM, in conjunction with a dedicated CCD camera and a commercial DSLR camera, were utilized to conduct optical observations on satellites in Earth orbit. Satellites were imaged during August 2019, and from January 2020 to March 2020, resulting in the collection of 77 valid images of 16 unique satellites. These images were used to obtain celestial spherical coordinates, which were used in Gauss and Double-R angles-only initial orbit determination methods. Initial orbit determination methods successfully produced valid results, reaffirming the feasibility of using small aperture telescopes for such methods. These orbit determinations were used to propagate orbit states forward in time to determine the feasibility of future imaging of the targets with the same apparatus. Propagation results demonstrated that initial orbit determinations rapidly decayed in accuracy over distant times and are most accurate for immediate satellite passes. In addition, an attempt to combine multiple initial orbit determinations using Lambert’s problem solutions was made. Combination of these multiple initial orbit determinations resulted in either no orbit state accuracy improvement compared to individual initial orbit determinations, or a decrease in accuracy compared to these methods. Ultimately, efforts demonstrated that small telescope usage is feasible for orbit determination operations, however there may be a need for hardware and operational revisions to improve the ability of the apparatus.

Over the last decade, miniature ion thruster development has remained an active area of research do to its low power, low thrust, and high efficiency, however, due to several technical issues; a flight level miniature ion thruster has proved elusive. This thesis covers the design, fabrication, assembly, and test of an altered version of the Miniature Xenon Ion thruster (MiXI), originally developed by lead engineer Dr. Richard Wirz, at the California Institute of Technology (Caltech). In collaboration with Dr. Wirz, MiXI-CP-V3 was developed at Cal Poly San Luis Obispo with the goal of implementing of a ¼ inch hollow cathode and 3mmx3mm plasma confinement magnets in order to improve the plasma confinement characteristics, reliability, and performance of the MiXI design. Operational testing revealed a mass utilization efficiency of 35-75% and a discharge loss of 550-1200 eV/ion over plasma discharge currents of 0.5-1.5A and propellant flow rates of 0.8-1.3 SCCM. Testing revealed that the MiXI thruster can be operated with a hollow cathode and observations and data gained from this study have led to a greater understanding of the operational parameters of the MiXI thruster, and will contribute to the development and advancement of the MiXI baseline design, with the goal of creating an efficient and reliable flight level miniature ion thruster.

The Miniaturized Ultraviolet Imager (MUVI), is a compact wide field UV imaging instrument in development at UC Berkeley Space Sciences Laboratory and Cal Poly, San Luis Obispo. MUVI is designed to fit in a 2U CubeSat form factor and provide wide field, high resolution images of the ionosphere at far ultraviolet wavelengths. This thesis details the design and analyses of MUVI’s deployable cover mirror mounting flexures. Three different flexure geometries were evaluated, an optimal candidate was determined based on a number of criteria including isolation of vibration and stress to the mirrors, manufacturability and cost. The design of the flexure system includes the flexure blades themselves, Invar pads bonded to the mirror to mitigate the difference in CTEs of the different material, mounting of flexure blades to the deployable cover and ground support equipment for assembly and testing. During the design of the flexures, various materials were studied, and Titanium was concluded as the optimum material due to its combination of high strength and flexibility compared to stainless steel, aluminum and other metals. Utilizing titanium, several flexure designs were proposed, and three candidates were selected to be manufactured and tested. Throughout the design phase, all flexures went through several rounds of analysis utilizing finite element analysis to simulate quasi-static loads, modal analysis of the systems natural frequency as well as random vibration simulations to simulate testing environments. Once the front-runner designs were selected and manufactured, several tests were conducted. Testing included adhesive bond coupon testing of the adhesive in tension and bending to experimentally validate the bonding size of the invar pads would be sufficient. The adhesive bond testing conducted tension and three-point bend tests to characterize the epoxy adhesive used in the flexure assembly. Testing also consisted of sine sweep and random vibration environment in accordance with the NASA General Environmental Verification Standard to qualify the hardware for spaceflight. Throughout the vibration testing, an autocollimator was used pre and post-test to measure shifts in the optical alignment of the mirror after it underwent vibration qualification testing. Experimental and analytical models were compared once all testing was completed. The Curved Blade showed to test in the real world very close to that predicted by the finite element model, however, the Bent Blade and Z Blade showed a larger difference between analysis and test. Discussion into the reasoning for this difference and lessons learned is included.

The low size and budget of typical nanosatellite missions limit the available sensors for attitude estimation. Relatively high noise MEMS gyroscopes often must be employed when accurate knowledge of the spacecraft’s angular velocity is necessary for attitude determination and control. This thesis derived and tested in simulation the “Virtual Gyroscope” algorithm, which replaced a standard gyroscope with an array of spatially distributed accelerometers for a 1U CubeSat mission. A MEMS accelerometer model was developed and validated using Root Allan Variance, and the Virtual Gyroscope was tested both in the open loop configuration and as a replacement for a gyroscope in a Multiplicative Extended Kalman Filter. It was found that the quality of the Virtual Gyroscope’s rate measurement improved with a larger and higher quality array, but the error in the estimate was very large. The low signal-to-noise ratio and the unknown bias in the accelerometers caused the angular velocity estimate from the accelerometer array to be too poor for use in the propagation step of the Kalman filter. The Kalman filter performed better with attitude measurements alone than with the Virtual Gyroscope, even when the attitude were delivered at a low rate with added noise. Overall, the current Virtual Gyroscope algorithm that is presented in this thesis is not suitable to replace a MEMS gyroscope in a nanosatellite mission, although there is room for future improvements using bias prediction for the individual accelerometers in the array.

The purpose of the following research is to explore molten regolith electrolysis (MRE) methodology for in-situ resource utilization (ISRU) of Highlands lunar regolith, to be explored during the initial Artemis missions. An analysis of potential commercial launch providers for MRE-equipment based on technology-readiness level (TRL), payload mass support, and $ USD/kg payload price is provided. SpaceX is ultimately proposed as a launch provider of MRE equipment following multi-factorial analysis, with the SpaceX Starship human landing system (HLS) variant proposed for supporting MRE payload. Finally, customers of regolith-derived oxygen, aluminum, and silicon are distinguished to form the business case for operating MRE equipment on the lunar surface.

The exploration of celestial bodies has recently advanced from rovers to rotorcraft. This includes the recent flights of Mars Ingenuity and the upcoming Dragonfly mission to explore the terrain of Saturn’s moon Titan as part of NASA’s New Frontiers Program. Flight-based landers can travel quickly to sites kilometers apart and land in complex terrain. Although cruise conditions for these rotorcrafts are well understood, studies are necessary to understand take-off and landing. In ground effect conditions, a rotor wake impinges and reflects off the ground, creating changes in aerodynamics such as increased lift. Additionally, operating over loose surfaces, the rotors can create clouds of dust obscuring the vehicle’s sensors, a hazard termed “brownout” from rotorcraft landing in sandy and snowy conditions on Earth. Take-off and landing events involve interactions between the rotor wake, fuselage, and ground, and lead to a multi-phase interface between the fluid atmosphere and the dispersed dust particles. The objective of this study is to computationally model and evaluate ground effect aerodynamic forces on the Dragonfly rotorcraft lander. A calculation of sediment distribution across the surface of the vehicle will provide insight to which components might be most affected by brownout.

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Earth and Planetary Science, 1980. / Microfiche copy available in Archives and Science. / Includes bibliographical references. / by Thomas Marshall Eubanks. / M.S.

The compression strength of dropped-ply, graphite-epoxy laminated plates for the delamination mode of failure is studied by analysis and corroborated with experiments. The nonlinear response of the test specimens is modeled by a geometrically nonlinear finite element analysis. The methodology for predicting delamination is based on a quadratic interlaminar stress criterion evaluated at a characteristic distance from the ply drop-off. The details of the complex state of stress in the region of the thickness discontinuity are studied using three-dimensional solid elements, while the uniform sections of the plate are modeled with quadrilateral shell elements. A geometrically nonlinear transition element was developed to couple the shell elements to the solid elements. The analysis was performed using the COmputational MEchanics Testbed (COMET), an advanced structural analysis software environment developed at the NASA Langley Research Center to provide a framework for research in structural analysis methods. Uniaxial compression testing of dropped-ply, graphite-epoxy laminated plates has confirmed that delamination along the interfaces above and/or below the dropped plies is a common mode of failure initiation. The compression strength of specimens exhibiting a linear response is greater than the compression strength of specimens with the same layup exhibiting geometrically nonlinear response. Experimental and analytical results also show a decrease in laminate strength with increasing number of dropped plies. For linear response there is a large decrease in compression strength with increasing number of dropped plies. For nonlinear response there is less of a reduction in compression strength with increasing number of dropped plies because the nonlinear response causes a redistribution and concentration of interlaminar stresses toward the unloaded edges of the laminate. / Ph. D.

The integration of two Zero-Voltage-Switched Bidirectional Battery Charger Discharger (ZVS-BBCD) units into a space power system is shown. A robust design featuring: four interleaved phases, commandable charge rates, overcurrent protection, overvoltage protection, soft starting, reliable gate drive circuitry, high efficiency, and good dynamics is demonstrated. The ZVS-BBCD is compared to separate hard-switched multimodule charge and discharge units in weight and efficiency and is found to be significantly lighter with comparable losses. The ZVS-BBCD has similar characteristics in discharge and in charge bus regulation modes and allows the use of a common control design for both modes. The two ZVS-BBCDs are integrated into the NASA power system testbed built at the Virginia Power Electronics Center (VPEC) to test their dynamics. The result shows good characteristics including low bus impedance and fast transient response. / Master of Science

An open loop control law is applied to a flexible spacecraft that admits translational, as well as rotational and flexural motion. The translational degrees of freedom are coupled to the rotation and deformation through the active controls applied to the structure. The objective of any maneuver is to control the attitude of the craft as well as to dissipate any vibrations of the structure. Depending on the type of maneuver specified, the equations of motion may be divided into two distinct optimal control problems. Single-axis rotational maneuvers (with small flexural deformations) constitute a strictly linear problem. The solution of the resulting two Q point boundary value problem is accomplished through the use of matrix exponential functions. Maneuvers which involve the translational degrees of freedom, are described by nonlinear equations. The solution method presented is a algorithm that generates successive approximations similar to quasi-linearization. A perturbed linear optimal control problem is solved for each approximation. Examples are presented which illustrate the effectiveness of the solution methods for both types of maneuvers. / Master of Science