The purpose of this paper is to report the results of studies on the aerobic, mesophilic intestinal flora of guinea pigs subjected to conditions similar to those encountered by man in spacecraft.

A hybrid thruster system utilizes propellants in two different stages, traditionally a solid fuel and a gaseous or liquid oxidizer. Recently hybrid thrusters have become a popular topic of research due to the high demand of a ”green” replacement for hydrazine. Not only are hybrid thruster systems typically much safer than hydrazine, but they are also a low-cost system with a high reliability in performance. The Propulsion Research Laboratory (PRL) at Utah State University (USU) has developed a hybrid thruster system using 3-D printed acrylonitrile butadiene styrene (ABS) as the fuel and gaseous oxygen (GOX) as the oxidizer. This system has been spaceflight flown and tested in a hard vacuum environment with success. However, GOX has a low density and must be stored at high pressures to be considered viable. This thesis investigates the use of N2O/O2 mixtures, ”Nytrox”, and more commonly known as ”laughing gas”, as a higher density replacement oxidizer for GOX. Ina manner directly analogous to the creation of soda-water using dissolved carbon dioxide, Nytrox is created by bubbling gaseous oxygen under high pressure into nitrous oxide until the solution reaches saturation level. Oxygen in the mixture ullage dilutes the nitrous oxide vapor, and increases the required decomposition activation energy of the fluid by several orders of magnitude. Data from tests using each oxidizer are analyzed and presented for performance comparisons. Comparisons include, ignition reliability, ignition energy, thrust coefficient, characteristic velocity, specific impulse, and regression rate. Nytrox is shown to work effectively as a “drop in” replacement for gaseous oxygen, exhibiting slightly reduced specific impulse and regression rate, but with the trade of a significantly higher volumetric efficiency.

Research has been completed in order to rapidly manufacture spacecraft thermal control technologies embedded in spacecraft structural panels using ultrasonic consolidation. This rapid manufacturing process enables custom thermal control designs in the time frame necessary for responsive space. Successfully embedded components include temperature sensors, heaters, wire harnessing, pre-manufactured heat pipes, and custom integral heat pipes. High conductivity inserts and custom integral pulsating heat pipes were unsuccessfully attempted. This research shows the viability of rapid manufacturing of spacecraft structures with embedded thermal control using ultrasonic consolidation.

The number of objects occupying orbital regimes beyond Geosynchronous Earth Orbit and cislunar space are expected to grow in the coming years; Especially with the Moon reemerging as latest frontier in the race for space exploration and technological superiority. In order to support this growth, new methods of autonomously navigating in cislunar space are necessary to reduce demand and reliance on ground based tracking infrastructure. Periodic orbits about the first libration point offer favorable vantage points for scientific or military spacecraft missions involving the Earth or Moon. This thesis develops a new autonomous spacecraft navigation method for cislunar space and analyzes its performance applied to Lyapunov and halo orbits around $L_1$. This method uses existing lunar ranging retroreflectors (LRRR) installed on the Moon's surface in the 1960s and 1970s. A spacecraft can make laser ranging measurements to the LRRR to estimate its orbit states. A simulation platform was created to test this concept in the circular restricted three body problem and evaluate its performance. This navigation method was found to be successful for a subset of Lyapunov and halo orbits when cycling the five measurement targets. Simulation data showed that sub-kilometer position estimation and sub 2 centimeter per second velocity accuracies are achievable without receiving any state updates from external sources. / Master of Science / The number of objects occupying the space between the Earth and Moon (cislunar space) is expected to grow in the coming years as the Moon regains popularity in the latest race for space exploration and technological superiority. In order to support this growth, new methods of determining a spacecraft's position and velocity while in this region of space are necessary to reduce demand and dependence on Earth based methods, which have historically relied upon. Repeating orbits around the equilibrium point between the Earth and Moon provide valuable observation points for scientific and military spacecraft missions. This thesis develops a new spacecraft navigation method for cislunar space and analyzes how well it performs in two different types of orbits, Lyapunov and halo orbits. This method uses existing laser reflector panels that were installed on the Moon's surface in the 1960s and 70s. A spacecraft can use these panels to make range or distance measurements in order to estimate its position and velocity. Software was written to simulate the motion of a spacecraft as it is acted on by gravity from the Earth and Moon. Different scenarios were then simulated and used to test this concept and evaluate its performance. Lunar laser ranging was found to be successful for a some Lyapunov and halo orbits when switching between the five different reflector panels on the Moon. Data generated from the simulations show that position can be estimated with errors less than SI{1}{kilo meter}, and velocity error on the order of a few centimeters per second, all without receiving any additional information from Earth based systems.

A deep space concept vehicle created from a core stage barrel section of the Space Launch System rocket has been designed by the National Aeronautics and Space Administration for use in future manned Mars missions. The spacecraft, known as the Space Launch System-Derived Habitat, features a dedicated space for exercise equipment. A human factors assessment was performed to determine whether or not the exercise area has adequate volume for multiple microgravity exercise machines to be used by multiple crew members simultaneously. It was determined that in its current design the exercise area does not have adequate volume to house the machines required for bone and muscle maintenance as required for long-duration spaceflight missions. It was recommended that the volume either be vastly expanded or dissolved entirely in favor of multiple, smaller exercise volumes that could each house one machine.

No description available.

Translational and rotational motion are coupled for spacecraft performing formation flying missions. This motion is coupled because orbital control is dependent on the spacecraft attitude for vectored thrust. Formation flying spacecraft have a limited mass and volume for propulsion systems. We want to maximize the efficiency of the spacecraft, which leads to minimizing the error introduced by thrusting in the wrong direction. This thrust direction error leads to the need for a coupled attitude and orbital control strategy. In this thesis a coupled control system is developed using a nonlinear Lyapunov attitude controller and a nonlinear Lyapunov-based orbital controller. A nonlinear Lyapunov attitude controller is presented for a spacecraft with three-axis momentum wheel control. The nonlinear Lyapunov-based orbital controller is combined with a mean motion control strategy to provide a globally asymptotically stable controller. The attitude and orbit control laws are verified separately using numerical simulation, and then are integrated into a coupled control strategy. The coupled system simulations verify that the coupled control strategy is able to correct for an initial relative position error between two spacecraft. / Master of Science

The COLISEUM framework is a tool designed for electric propulsion plume interactions. Virginia Tech has been developing a module for COLISEUM called DRACO, a particle-in-cell based code capable of plume modeling for geometrically complex spacecraft. This work integrates a charging module into DRACO. Charge is collected via particle impingement on the spacecraft surface and converted to potential. Charge can be stored in the surface, or added to a local ground potential. Current can flow through the surface and is governed by the internal electric field in the spacecraft. Several test cases were run to demonstrate the code's capabilities. The first was a plume impingement of a composite spherical probe by a xenon thruster. It was shown that the majority of current conducted will reach the interior of the spacecraft, not other surface elements. A conductive interior will therefore result in a uniform surface potential, even for low surface conductivities. A second test case showed a composite spacecraft exposed to a solar wind. This test showed that when a potential gradient is applied to a conductive body, the ground potential of the spacecraft will lower significantly to compensate and maintain a zero net current. The case that used the semiconductive material showed that the effect of the potential gradient can be lowered in cases such as this, where some charge will always be stuck in the surface. If a dielectric material is used, the gradient will disappear altogether. The final test case showed the effect of charge exchange ion backflow on the potential of a spacecraft similar to the DAWN spacecraft. This case showed that CEX ion distribution is not even along the spacecraft and will result in a transverse potential gradient along the panel. / Master of Science

The problem of reorienting a spacecraft to acquire a moving target is investigated. The spacecraft is modeled as a rigid body with N axisymmetric wheels controlled by axial torques, and the kinematics are represented by Modified Rodriques Parameters. The trajectory, denoted the reference trajectory, is one generated by a virtual spacecraft that is identical to the actual spacecraft. The open-loop reference attitude, angular velocity, and angular acceleration tracking commands are constructed so that the solar panel vector is perpendicular to the sun vector during the tracking maneuver. We develop a nonlinear feedback tracking control law, derived from Lyapunov stability and control theory, to provide the control torques for target tracking. The controller makes the body frame asymptotically track the reference motion when there are initial errors in the attitude and angular velocity. A spacecraft model, based on the X-ray Timing Explorer spacecraft, is used to demonstrate the effectiveness of the Lyapunov controller in tracking a given target. / Master of Science

Neutral gas pressure measurements in low Earth orbit (LEO) can facilitate the monitoring of atmospheric gravity waves, which can trigger instabilities that severely disrupt radio frequency communication signals. The Space Neutral Pressure Instrument (SNeuPI) is a low-power instrument detecting neutral gas density in order to determine neutral gas pressure. SNeuPI consists of an ionization chamber and a logarithmic electrometer circuit. The Rev. 1 SNeuPI electrometer prototype does not function as designed. A Rev. 2 electrometer circuit must be designed and its performance characterized across specified operating temperature and input current ranges. This document presents a design topology for the Rev. 2 electrometer and a derivation of the theoretical circuit transfer function. Component selection and layout are discussed. A range of predicted operating input currents is calculated using modeled neutral density data for a range of local times, altitudes, and latitudes corresponding to the conditions expected for the Lower Atmosphere/Ionosphere Coupling Experiment (LAICE) CubeSat mission. Laboratory test setups for measurements performed both under vacuum and at atmospheric pressure are documented in detail. Test procedures are presented to characterize the performance of the Rev. 2 electrometer at a range of controlled operating temperatures. The results of these tests are then extrapolated in order to predict the operation of the circuit at specified temperatures outside of the range controllable under laboratory test conditions. The logarithmic conformance, accuracy, sensitivity, power consumption, and deviations from expected response of the circuit are characterized. The results validate the electrometer for use under its expected flight conditions. / Master of Science