New Horizons, Voyager 1 & 2, and Pioneer 10 & 11 are the only spacecraft to ever venture past Pluto and provide information about space at those large distances. These spacecraft were very expensive and primarily designed to study planets during gravitational assist maneuvers. They were not designed to explore space past Pluto and their study of this environment is at best a secondary mission. These spacecraft rely on radioisotope thermoelectric generators (RTGs) to provide power, an expensive yet necessary approach to generating sufficient power. With Cubesats graduating to interplanetary capabilities, such as the Mars-bound MarCO spacecraft, matching the modest payload requirements to study the outer Solar System (OSS) with the capabilities of low-power nano-satellites may enable much more affordable access to deep space. This paper explores a design concept for a low-cost, small spacecraft, designed to study the OSS and satisfy mission requirements with solar power. The general spacecraft design incorporates a parabolic reflector that acts as both a solar concentrator and a high gain antenna. This paper explores a working design concept for a small spacecraft to operate up to 100 astronomical units (AU) from the sun. Deployable reflector designs, thermal and radiation environments, communications and power requirements, solar system escape trajectory options, and scientific payload requirements are detailed, and a working system is proposed that can fulfill mission requirements with expected near-future innovations in a few key technologies.

This study used images taken from the OSIRIS-REx spacecraft sample-return mission and projected onto a three-dimensional shape model to determine surface properties of the asteroid (101955) Bennu. Two major projects were completed. For the first project, images of the pre-sampled Nightingale site and post-sampled Nightingale were compared to determine how the TAGSAM sampling maneuver effected the surface of the asteroid directly at the sampling site and in the surrounding area. This analysis demonstrated how spacecraft can potentially affect a small body during this and future sample-return missions. For the second project, several craters on Bennu’s surface were selected and all of the boulders within their rims and up to one crater radius outward from their rims were counted and measured. The interior and exterior of all craters were compared which determined that there is a pattern that supports the existence of a subsurface layer of finer-grained material which could be responsible for cohesion on Bennu and potentially other rubble-pile asteroids

Mass spectrometers have proven to be vital to understanding the Solar System and the planets within it. Spacecraft containing mass spectrometers have been sent to numerous remote places and have determined important information about the atmospheric composition of Venus, Earth, Mars, Jupiter, and Saturn, along with other celestial bodies. Such results have shown a variety of small neutral molecules, such as CH4 NH3, H2O, CO2, and CO, neutral radicals such as atomic O, H, and N, and a host of small ions, such as H+, N+, and NH4+. Closed ion source inlets, which allow for the detection of these small neutral molecules, contain a spherical antechamber that allows the neutrals to thermalize with the walls of the chamber through many successive collisions before they are introduced into the ionization region of the spacecraft mass spectrometer. These collisions, however, energetically excite neutral molecules and lead to many chemical changes, such as racemization, ionization, or even dissociation. When these changes occur, smaller neutrals can be produced, even if they were not in the original sample from the atmosphere or surface. As a result, the determination of the true composition of an atmosphere or a surface is cast into doubt. Herein is given a brief description of mass spectrometry in space research and how the closed ion source has greatly assisted this process. Dissociation and other chemical changes caused by the high velocity impacts that occur in closed source antechambers is also addressed. A theoretical approach to understanding such dissociative processes that occur after high energy collisions in closed source antechambers is described and undertaken. Chapter 2 describes a proof-of-concept study using hexane as a representative molecule and determines the velocity at which widespread dissociation of hexane molecules is likely to occur in closed source antechambers. This same theoretical process is then utilized in Chapter 3 with many more members of the n-alkane family to probe what effect molecular weight has on the amount of dissociation. Alkanes of both higher and lower molecular weight than hexane (C6H14) are used to show the effect as a function of molecular weight. In all cases, it was found that the velocity at which half of the incoming neutral n-alkane molecules dissociate is roughly the same for all molecular weights studied. This result is then applied to current and future space research through a proposed hardware solution, which will reduce the amount of dissociation and a discussion of how this effect may be seen in the results obtained from future mission instruments. Lastly, future work with different molecular weights and with successive collisions (the second, third, fourth, etc.) is described. This future work will further expand the present study to show how different functional groups, which may be partly responsible for higher-than-expected levels of NH3 and CO2, are affected after a high velocity, high energy impact.

Additive Manufacturing (AM, 3D printing) has made significant advancements over the past decade and has become a viable alternative to traditional machining techniques. AM offers several advantages over traditional manufacturing techniques including improved geometric freedom, reduction in part lead time, cost savings, enhanced customization, mass reduction, part elimination, and remote production. There are many different AM processes with the most commonly used process being Fused Filament Fabrication (FFF). Small satellites have also made significant advancements over the past two decades with the number of missions launched annually increased by orders of magnitude over that time span. Small satellites offer several advantages compared to traditional spacecraft architectures including increased access to space, lower development costs, and disaggregated architectures. On-orbit manufacturing and assembly have become major research and development topics for government and commercial entities seeking the capability to build very large structures in space. AM is well suited on-orbit manufacturing since the process is highly automated, produces little material waste, and allows for a large degree of geometric freedom. This dissertation seeks to address three major research objectives regarding applications of additive manufacturing in space systems: demonstrate the feasibility of 3D printing an ESPA class satellite using FFF, develop a FFF 3D printer that is capable of operating in high vacuum and characterize its performance, and analyze the coupled dynamics between a satellite and a robot arm used for 3D printing in-space. This dissertation presents the design, finite element analysis, dynamic testing, and model correlation of AdditiveSat, an additively manufactured small satellite fabricated using FFF. This dissertation also presents the design, analysis, and test results for a passively cooled FFF 3D printer capable of manufacturing parts out of engineering grade thermoplastics in the vacuum of space. Finally, this dissertation presents a numerical model of a free-flying small satellite with an attached robotic arm assembly to simulate 3D printing structures on-orbit with analysis of the satellite controls required to control the dynamics of the highly coupled system. / Doctor of Philosophy / 3D printing has made significant advancements over the past decade and has become common place in offices, schools, and even the homes of hobbyist. 3D printing has become an alternative to traditional machining techniques, such as machining parts from blocks of material. 3D printing offers several advantages over traditional manufacturing techniques including improved geometry freedom, reduction in part lead time, cost savings, enhanced customization, mass reduction, part elimination, and remote production. There are many different types of 3D printing with the most commonly used process being Fused Filament Fabrication (FFF) in which a thermoplastic is melded by a hotend and then extruded through a nozzle to deposited material layer-by-layer onto a printed part. Small satellites have also made significant advancements over the past two decades with the number of missions launched annually greatly increased over that time span. Small satellites offer several advantages compared to traditional spacecraft including increased access to space and lower development costs. On-orbit manufacturing and assembly have become major research and development topics for government and commercial entities seeking the capability to build very large structures in space. This dissertation seeks to address three major research objectives regarding applications of additive manufacturing in space systems: demonstrate the feasibility of 3D printing an ESPA class satellite using FFF, develop a FFF 3D printer that is capable of operating in high vacuum and characterize its performance, and analyze the coupled dynamics between a satellite and a robot arm used for 3D printing in-space. This dissertation presents the design, analysis, and test results of AdditiveSat, a 3D printed small satellite made using FFF. This dissertation also presents the development of a FFF 3D printer capable of operating in the vacuum of space. Finally, this dissertation presents a numerical simulation that models 3D printing structures on-orbit with a small satellite equipped with a robot arm.

The CalPoly Spacecraft Attitude Dynamics Simulator mimics the rotational dynamics of a spacecraft in orbit and acts as a testbed for spacecraft attitude control system development and demonstration. Prior to this thesis, the simulator platform and several subsystems had been designed and manufactured, but the total simulator system was not yet capable of closed-loop attitude control. Previous attempts to make the system controllable were primarily mired by data transport performance. Rather than exporting data to an external command computer, the strategy implemented in this thesis relies on a compact computer onboard the simulator platform to handle both attitude control processing and data acquisition responsibilities. Software drivers were created to interface the computer’s data acquisition boards with Matlab, and a Simulink library was developed to handle hardware interface functions and simplify the composition of attitude control schemes. To improve the usability of the system, a variety of actuator control, hardware testing, and data visualization utilities were also created. A closedloop attitude control strategy was adapted to facilitate future sensor installations, and was tested in numerical simulation. The control model was then updated to interface with the simulator hardware, and for the first time in the project history, attitude control was performed onboard the CalPoly spacecraft attitude dynamics simulator. The demonstration served to validate the numerical model and to verify the functionality of the entire simulator system.

The constant growth of space debris and the associated risks force the space community to find solutions to mitigate them. Today the most advanced solutions to dispose of satellites and rocket stages after the end of mission consists of moving them either into a graveyard orbit or towards an atmospheric re-entry ending in the demise of both spacecraft and its materials. Alternative solutions should be considered, such as providing a sustainable solution by reusing materials in space. However, it is crucial to understand better the ageing process of the materials present in currently active spacecraft and space debris. The space environment causes degradation and damage over time, making the state of those materials uncertain for potential re-use. Degradation effects have been studied as a source mechanism to result in paint flakes, ejecta particles, or delaminated insulation foils released into the space environment and sustaining a positive feedback loop through potential impacts into spacecraft. A better understanding of degradation effects would also help to better characterize the small debris environment and its evolution. The current materials databases used by the space industry could be useful tools to select materials for satellite missions with respect to their reusability, but they often do not include the evolution of material properties in space after the end of mission.This study will investigate the impact of the damage effects of radiation and meteoroid impact on solar panels. During this research, the methodology used to analyse these effects was explained. The results showed that radiation caused the most damage and could cause solar panels to lose more than a third of their performance over a period of 50 years. It was therefore possible to estimate the quantity of solar panels available for re-use. It was concluded that the results were valid, but that to obtain more accurate data, all the different types of deterioration would also have to be considered. / Creaternity

No description available.

<p dir="ltr">Design under uncertainty (DUU) for spacecraft is crucial in ensuring mission success, especially given the criticality of their failure. To obtain a more realistic understanding of space systems, it is beneficial to holistically couple the modeling of the spacecraft and its trajectory as a multidisciplinary analysis (MDA). In this work, a MDA model is developed for an Earth-Mars mission by employing the general mission analysis tool (GMAT) to model the mission trajectory and rocket propulsion analysis (RPA) to design the engines. By utilizing this direct MDA model, the deterministic optimization (DO) of the system is performed first and yields a design that completed the mission in 307 days while requiring 475 kg of fuel. The direct MDA model is also integrated into a Monte Carlo simulation (MCS) to investigate the uncertainty quantification (UQ) of the spacecraft and trajectory system. When considering the combined uncertainty in the launch date for a 20-day window and the specific impulses, the time of flight ranges from 275 to 330 days and the total fuel consumption ranges from 475 to 950 kg. The spacecraft velocity exhibits deviations ranging from 2 to 4 km/s at any given instance in the Earth inertial frame. The amount of fuel consumed during the TCM ranges from 1 to 250 kg, while during the MOI, the amount of fuel consumed ranges from 350 to 810 kg. The usage of the direct MDA model for optimization and uncertainty quantification of the system can be computationally prohibitive for DUU. To address this challenge, the effectiveness of utilizing surrogate-based approaches for performing UQ is demonstrated, resulting in significantly lower computational costs. Gaussian processes (GP) models trained on data from the MDA model were implemented into the UQ framework and their results were compared to those of the direct MDA method. When considering the combined uncertainty from both sources, the surrogate-based method had a mean error of 1.67% and required only 29% of the computational time. When compared to the direct MDA, the time of flight range matched well. While the TCM and MOI fuel consumption ranges were smaller by 5 kg. These GP models were integrated into the DUU framework to perform reliability-based design optimization (RBDO) feasibly for the spacecraft and trajectory system. For the combined uncertainty, the DO design yielded a poor reliability of 54%, underscoring the necessity for performing RBDO. The DUU framework obtained a design with a significantly improved reliability of 99%, which required an additional 39.19 kg of fuel and also resulted in a reduced time of flight by 0.55 days.</p>

Les satellites en orbite terrestre se déplacent dans le plasma ionosphérique, un mélange de particules chargées, et éventuellement de particules neutres. Des électrons et des ions issus de ce plasma, ainsi que les émissions Ultra-Violets(UV) en provenance du soleil, interagissent avec les surfaces du satellite et modifient sa charge électrostatique. Cette chargement peut induire elle-même des décharges électrostatiques aux conséquences allant de perturbations électromagnétiques (fausses commandes par exemple) à la perte du satellite. En orbites de basse altitude (LEO) l'énergie cinétique et thermique du plasma est généralement faible et donc, les satellites vont rarement présenter des décharges importantes. Néanmoins, les missions scientifiques qui embarquent des instruments très performants et précis peuvent être affectées par cette interaction satellite-plasma-émissions UV. Cette thèse s'intéresse particulièrement à ces phénomènes de charge des structures externes du satellite et à l'impact de ce chargement sur les mesures scientifiques effectuées à bord, i.e. mesures du champ électrique et de la densité du plasma thermique. / Earth-orbiting satellites travel in ionospheric plasma, a mixture of charged particles, and possibly neutral particles. Electrons and ions from this plasma, as well as Ultra-Violet (UV) emissions from the sun, interact with the surfaces of the satellite and modify its electrostatic charge. This loading can itself induce electrostatic discharges to the consequences ranging from electromagnetic disturbances (false commands for example) to the loss of the satellite. In low-Earth orbits (LEO), the kinetic and thermal energy of the plasma is generally low and therefore satellites rarely exhibit large discharges. Nevertheless, scientific missions that carry high-performance and accurate instruments can be affected by this satellite-plasma-UV-emissions interaction. This thesis is particularly interested in these phenomena of charge of the external structures of the satellite and the impact of this load on the scientific measurements carried out on board, i.e. measures of the electric field and the density of the thermal plasma.

The contemporary terrestrial atmosphere loses matter at a rate of around 100,000 tons per year. A major fraction of the net mass loss is constituted by ions, mainly H+ and O+, which escape from the Earth’s ionosphere in the polar regions. Previously, the outflow has only been measured at low altitudes, but to understand what fraction actually escapes and does not return, the measurements should be conducted far from the Earth. However, at large geocentric distances the outflowing ions are difficult to detect with conventional ion instruments on spacecraft, since the spacecraft electrostatic potential normally exceeds the equivalent energy of the ions. This also means that little is known about the ion outflow properties and distribution in space far from the Earth. In this thesis, we present a new method to measure the outflowing low-energy ions in those regions where they previously have been invisible. The method is based on the detection by electric field instruments of the large wake created behind a spacecraft in a flowing, low-energy plasma. Since ions with low energy will create a larger wake, the method is more sensitive to light ions, and our measured outflow is essentially the proton outflow. Applying this new method on data from the Cluster spacecraft, we have been able to make an extensive statistical study of ion outflows from 5 to 19 Earth radii in the magnetotail lobes. We show that cold proton outflows dominate in these large regions of the magnetosphere in both flux and density. Our outflow values of low-energy protons are close to those measured at low altitudes, which confirms that the ionospheric outflows continue far back in the tail and contribute significantly to the magnetospheric content. We also conclude that most of the ions are escaping and not returning, which improves previous estimates of the global outflow. The total loss of protons due to high-latitude escape is found to be on the order of 1026 protons/s.