Interplanetary space exploration has thus far consisted of single, expensive spacecraft missions. Mission costs are particularly high on missions to the outer planets and while invaluable, finite budgets limit our ability to perform extensive and frequent investigations of the planets. Planetary systems such as Jupiter and Saturn provide extremely complex exploration environments with numerous targets of interest. Exploring these targets in addition to the main planet requires multiple fly-bys and long mission timelines. In LEO, CubeSats have changed the exploration paradigm, offering a fast and low cost alternative to traditional space vehicles. This new mission development philosophy has the potential to significantly change the economics of interplanetary exploration and a number of missions are being developed to utilize CubeSat class spacecraft beyond earth orbit (e.g., NEAScout, Lunar Ice Cube, Marco and BioSentinel). This paper takes the CubeSat philosophical approach one step further by investigating the potential for small satellite swarms to provide extensive studies of the Saturn system. To do this, an architecture was developed to best replicate the Cassini Primary Mission science objectives using swarms of CubeSats. Cassini was chosen because of its complexity and it defines a well-understood baseline to compare against. The paper outlines the overall mission architecture developed and provides a feasible initial design for the spacecraft in the architecture. The number of swarms needed, number of CubeSats per swarm, size of the CubeSats, overall science output and estimated mission cost are all presented. Additional science objectives beyond Cassini's capabilities are also proposed. Significant scientific returns can be achieved by the swarm based architecture and the risk tolerance afforded by the utilization of large numbers of low-cost sensor carriers. This study found a potential architecture that could reduce the cost of replicating Cassini by as much as 63%. The results of this investigation are not constrained to Saturn and can be easily translated to other targets such as Uranus, Neptune or the asteroid belt.

The space environment presents various challenges when designing systems and selecting materials for applications beyond Earth’s atmosphere. For mission success, these challenges must be considered. One of the detrimental aspects of the space en- vironment is Atomic Oxygen, AO. Only present in harmful quantities in Lower Earth Orbit, LEO, AO causes significant damage to materials by breaking molecular bonds. California Polytechnic State University’s, Cal Poly’s, space environments laboratory features an apparatus capable of simulating this environment. Very thin or short samples were tested to observe the mass loss due to erosion of the sample material. Recent modifications to the system allow it to expose surfaces of three dimensional objects to AO rather than only those two dimensional objects. Simulating this effect on taller samples makes available the opportunity to test coupons that are then used in additional testing to measure the effect of that erosion on other properties. Challenges in adapting the AO system are explored and addressed, as well as some possible use cases for future work. As a use case, bending moment specimens were exposed to AO prior to testing in four point bending. Multiple regression models were constructed to determine variables contributing to slope changes between specimen pairs’ linear-elastic regions of force-displacement graphs. Results show that AO exposed specimens had significantly gentler slopes in the linear elastic region of the force-displacement curve, meaning that AO exposure reduced structural rigidity of the coupons.

The space environment is a volatile and challenging place for satellites to survive in. For Low Earth Orbiting (LEO) satellites, atomic oxygen (AO) is a constant corrosive effect that degrades the outer surface of satellites over long durations. Atomic oxygen exists in the atmosphere between 180 and 675 km and has a relatively high energy at 4.5 eV, which allows AO to break molecular bonds in materials on the surfaces of spacecraft. As the number and complexity of CubeSat missions increase, there is an increased risk that AO degradation on commercial off the shelf parts (COTS), such as antenna, could degrade the satellite’s ability to communicate with ground systems. This thesis looks at how AO erosion affects the performance of patch antennas for CubeSat applications. Patch antennas are small, cheap, low-profile antennas that can be used on CubeSats to communicate with the ground or other satellites. Patch antennas are semi-directional, providing higher gain and higher available frequencies than omnidirectional antennas. An AO chamber in the California Polytechnic State University San Luis Obispo’s (Cal Poly) Spacecraft Environments Testing Lab was used to expose the patch antennas for 24-hour and 48-hour tests. The 24-hour exposure saw an average AO fluence of 8.757 ± 0.807•1020 atoms/cm2 which corresponds to roughly 3.5 months of on-orbit AO exposure on the Ram side when in a 28.5° inclined orbit with an altitude of 400 km. The 48-hour exposure saw an average AO fluence of 1.595 ± 0.076•1021 atoms/cm2 which corresponds to approximately 6.4 months of on-orbit AO exposure on the Ram side when in a 28.5° inclined orbit with an altitude of 400 km. To test the performance of the patch antenna before and after AO exposure, an anechoic chamber in the Microwave Lab at Cal Poly was used to measure boresight gain and radiation pattern in the E-plane and H-plane. From the testing in the anechoic chamber it was determined that there was no apparent difference in the patch antenna’s gain and radiation pattern before and after AO exposure. By using a Fourier Transform Infrared Spectrometer (FTIR) it was discovered that the outer surface of the patch antennas were forming a silicon dioxide layer, which did not affect the performance of the patch antenna. Since silicon dioxide is resistant to AO erosion, it may be beneficial for CubeSats to include silica additives to their exposed antenna surfaces to prevent erosion.

Space debris has become an increasingly hazardous obstacle to continued spaceflight operations. In an effort to mitigate this problem an investigation of the feasibility of a microsatellite active debris removal system was conducted. Through proposing a novel concept of operation, utilizing a grapple-and-tug system architecture, and by analyzing each resultant mission phase in the frame of a representative example, it was found that microsatellite scale systems are capable of fulfilling the active debris removal mission. Analysis of rendezvous, docking, control and deorbit mission requirements determined that the design of a grapple-and-tug system will be driven by sizing of the propellant required to deorbit the target vehicle. Further sensitivity analysis determined that target altitude and mass are critical factors in determining the capabilities of a microsatellite mission. Preliminary sizing demonstrated that hardware considerations for both satellite core and mission related activities do not impede microsatellite feasibility. Further investigation of microsatellite debris removal missions including detailed design analysis and engineering is suggested.

As space technologies continue to develop rapidly, there is a common desire to launch astronauts beyond the ISS to return to the Moon and put human footsteps on Mars. One of the largest hurdles that still needs to be addressed is the protection of astronauts from the radiation environment seen in deep space. The most effective way to defend against radiation is increasing the thickness of the shield, however this is limited by strict mass requirements. In order to increase the thickness of the shield, it is beneficial to make mission critical items double as shielding material. The human rated Orion spacecraft has procedures in place for astronauts to create an emergency bunker using food and water in the event of a forewarned radiation storm. This can provide substantial support to defend against radiation storms when there is an adequate amount of warning time, however, fails to protect against Galactic Cosmic Radiation (GCR) or Solar Particle Events (SPE) without sufficient warning. Utilizing these materials as a permanent shielding method throughout the mission could be a beneficial alternative to the Orion programs current protection plan to provide constant safety to the crew. This thesis analyzes the effect in the radiation dosage seen by astronauts in the Orion Crew Module through use of on-board water as a permanent shielding fixture. The primary method used to analyze radiation is NASA’s OLTARIS (On-Line Tool for the Assessment of Radiation In Space) program, which enables users to input thickness distributions to determine a mission dosage profile. In addition this thesis further develops a ray tracing code which enables users to import male and female models into the vehicle model to produce gender specific radiation dosage results. The data suggests the permanent inclusion of water as a shielding material provides added support for GCR as well as SPE radiation that can extend the mission lifetime of humans in space.

Today, Mars is one of the most interesting and important destinations for humankind and copious methods have been proposed to accomplish these future missions. One of the more fascinating methods is the Earth-Mars cycler trajectory which is a trajectory that accomplishes repeat access to Earth and Mars with little to no fuel-burning maneuvers. This would allow fast travel to and from Mars, as well as grant the possibility of multiple missions using the same main vehicle. Insertion from Earth-orbit onto the cycler trajectory has not been thoroughly ex- plored and the only existing method so far is a Hohmann-esque transfer via direct burn. The use of manifolds from gravitational equilibrium points has not been con- sidered for low energy transfer to the cycler trajectory. This work is primarily focused on closing this gap and analyzing the feasibility of this maneuver. To accomplish this, a study of the cycler trajectory – and the S1L1-B class specif- ically – was completed. The required gravity assist maneuvers at each planet was analyzed through V∞ matching and the entire trajectory was generated over the re- quired inertial period. This method allowed for the generation of 2 cycler trajectories of the inbound and outbound classes, which combine to allow for a reduction in the amount of time the astronauts spend in space. The Earth-Sun L2 point is analyzed as a potential hub for the maneuver and a halo orbit about this libration point is optimized for low energy transfer from and Earth parking orbit. The associated invariant manifold is then optimized for launch date and distance to the first trajectory on the cycler in order to burn from a trajectory on the manifold to the cycler trajectory. iv The comparisons of this work lie in the required ∆V to perform each maneuver compared to a direct burn onto the cycler trajectory. These values are compared and the practicality of this maneuver is drawn from these comparisons. It was found that the total required ∆V for the manifold method is larger than a direct burn from Earth orbit. However, this considers the trajectory from Earth to the halo orbit and if this is removed from consideration the ∆V is significantly reduced. It was shown that the feasibility of this method relies heavily on the starting position of the cycler vehicle. If the vehicle begins in Earth-orbit, a direct burn is preferred, however, if the vehicle began in a halo orbit (say it was assembled there) the manifold maneuver is largely preferable.

In this report we present a graphical tool, Behavioral Analysis of Spacecraft Systems (BASS), that can be used by spacecraft designers to perform system-level behavioral analysis of small satellites. The domain-specific spacecraft meta-model is created in the visual modeling tool Generic Modeling Environment (GME) such that spacecraft designs created using the meta-model appear familiar to the spacecraft designers. Users can model scenarios that are to be verified for the design in BASS. The graphical models are assigned formal semantics facilitating the creation of formally verifiable spacecraft models. The C++ application that translates the modeling objects to equivalent mathematical representation of interest is called BASS Interpreter and is bound to the meta-model. BASS Interpreter that generates Communicating Sequential Processes (CSP) semantics for the visual spacecraft models is supported in the current work. The model-checker for CSP called Failures Divergences and Refinement (FDR) is run to explore the state-space of the spacecraft process model to comment on the design. We demonstrate the feasibilty and advantage of incorporating BASS into initial design phases of small satellite development by successfully verifying the design of Tomographic Remote Observer of Ionospheric Disturbances (TOROID).

Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1980. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO. / Includes bibliographical references. / by Edward Gregory Taylor. / Ph.D.

This research aims to investigate the influence of 3D through-thickness stitching on the gas permeability and transverse microcracking of cryogenically cycled carbon/epoxy composites. 3D through-thickness stitching can be used to improve the interlaminar properties of polymer matrix composites (PMCs) and produce lightweight, unitized structures for cryogenic storage tanks. To fully utilize stitched composite structures for these applications, their inherent gas permeability challenges must be understood. Therefore, in this study, the stitched composites' damage evolution and gas permeability was experimentally characterized under a) pure thermal stress, b) thermal and uniaxial mechanical stress, and c) thermal and biaxial mechanical stress. Helium gas permeability was measured for each specimen at room or cryogenic temperatures under a mechanically strained state following the thermo-mechanical cycles. Optical microscopy was used to measure microcrack densities and monitor their evolution through the thickness of the composite specimens. Thin plies, graphene nanoplatelets (GNP) modified resin, and a hybrid barrier layer comprising of both were incorporated in the stitched specimens as barrier layers to reduce their gas permeability. The dependence of gas permeability of stitched composites on the mechanical strain, test temperature, and load history was evaluated and correlated to microcrack density. A significant reduction in permeability and damage evolution (transverse microcracks and delaminations) was obtained for all thermo-mechanical cases using the hybrid barrier layer laminate. Additionally, the permeability was several orders of magnitude lower than the allowable. Overall, the hybrid barrier layer shows tremendous promise as a viable barrier layer for stitched/unstitched composites undergoing thermo-mechanical fatigue involving a cryogenic environment.

This thesis outlines an orbit determination and 3-axis attitude determination system for use on orbit as applicable to 1U CubeSats and other picosatellites. The constraints imposed by the CubeSat form factor led to the need for a simple configuration and relaxed accuracy requirements. To design a system within the tight mass, volume, and power constraints inherent to CubeSats, a balance between hardware complexity, software complexity and accuracy is sought. The proposed solution consists of a simple orbit propagator, magnetometers with a magnetic field look-up table, Sun sensors with an analytic Sun direction model, and the TRIAD method to combine vector observations into attitude information. The orbit propagator is a simple model of a circular trajectory with several frequently updated parameters and can provide orbital position data with average and maximum errors—when compared to SGP4—of less than 3.7km and 10.7km for 14 days. The magnetic field look up table provides useful information from a small memory footprint; only 480 data points provide a mean error of approximately 0.2° and a maximum error of approximately 2°—when compared to the IGRF model. The Sun’s direction is modeled, and as expected, can be modeled simply and accurately. Combining the magnetic field and Sun direction models with inaccurate sensors and the TRIAD method results in useful attitude information from a very simple system. A system with Sun sensor error standard deviation of 1° and magnetometer error standard deviation of 5° yields results with average error of only 2.74°, and 99% of the errors in this case are less than approximately 13°. The system outlined provides crude attitude determination with software and hardware requirements that are well within the capabilities of current 1U CubeSats—something that many other systems, such as Kalman filters or star trackers, cannot do. It also provides an excellent starting point for future ADCS systems, which will significantly increase the ability of CubeSats.