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.

Most present day spacecrafts have large interconnected solar panels. The dynamic behavior of the spacecraft in orbit can be modeled as a free rigid mass with flexible elements attached to it. The natural frequencies of such spacecrafts with deployed solar panels are very low. The low values of the natural frequencies pose difficulties for maneuvering the spacecraft. The control torque required to maneuver the spacecraft is influenced by the flexibility of the solar arrays. The control torque sets up transient oscillations in the flexible solar panels which in turn induces disturbances in the rigid satellite body and the payload within. Therefore the payload operations can be carried out only after the disturbances die out. For any reduction of the above disturbances it is necessary to understand the dynamic behavior of such systems to an applied torque. The present work first studies the nature of the disturbances. The influence of structural parameters on these disturbances is then investigated. Finally, the use of passive damping treatment using viscoelastic material is investigated for the reduction of the disturbances. In order to understand the nature of vibrations induced in the flexible appendages of a satellite during maneuvers, we model the maneuver loads in terms of applied angular acceleration as well as varying torque. The transient decay of the disturbance of the rigid element is characterized by the dynamic characteristics of the flexible panels or appendages. It is shown that by changing the stiffness of the panel the response of the rigid element can be modified. A simple model consisting of an Euler-Bernoulli beam attached to a free mass is next considered. The influence of various parameters of the EulerBernoulli beam in mitigating vibration and thereby the disturbance in the rigid mass is investigated. As the response of the rigid system mounted with the large flexible panels are influenced by the dynamics of the flexible panels, reduction of these disturbances can be achieved by reducing the vibration in the flexible panels. Therefore application of viscoelastic materials for passive damping treatment is investigated. The loss factor of a structure is significantly improved by using constrained viscoelastic layer damping treatment. However providing a constrained layer damping treatment on the entire structure is very inefficient in terms of the additional mass involved. Therefore damping material is applied at suitable optimal locations. In previous studies reported in literature, modal strain energy distribution in the viscoelastic material as well as the base structure is used as a tool to arrive at the optimum location for the damping treatment. It is shown in this study that such locations selected are not the optimum. A new approach is proposed in this study by which both the above shortcomings are overcome. It is shown that use of a parameter that is the ratio of the strain in the viscoelastic material to the angle of flexure is a more reliable measure in arriving at optimal locations for the application of constrained viscoelastic layers. The method considers the deformations in the viscoelastic material and it is shown that significant values of loss factors are achieved by providing material in a small region alone. We also show that loss factor can be improved by providing damping material near the interface region. The loss factor can be further improved by incorporating spacers by using spacer material having higher extensional modulus. Also shown is the fact that loss factor is unaffected by the shear modulus of the spacer material. Experiments have been conducted to validate these results. In a related study we consider honeycomb type flexible structures since in most of the spacecraft applications honeycomb sandwich constructions are employed. But loss factors of sandwich panels with constrained layer damping treatment are seldom discussed in the literature. Use of viscoelastic layers to improve the loss factors of the honeycomb sandwich beams is explored. The results show that the loss factors are enhanced by increasing the inplane stiffness of the constraining layer. These conclusions too are validated by experimental results. Finally a typical satellite with flexible solar panels is considered, and the use of the viscoelastic material for improving the damping is demonstrated.

Solutions to trajectory optimization problems are carried out by the direct and indirect methods. Under broad heading of these methods, numerous algorithms such as collocation, direct, indirect and multiple shooting methods have been developed and reported in the literature. Each of these algorithms has certain advantages and limitations. For example, direct shooting technique is not suitable when the number of nonlinear programming variables is large. Indirect shooting method requires analytical derivatives of the control and co-states function and a poorly guessed initial condition can result in numerical unstable values of the adjoint variable. Multiple shooting techniques can alleviate some of these difficulties by breaking down the trajectory into several segments that help in reducing the non-linearity effects of early control on later parts of the trajectory. However, multiple shooting methods then have to handle more number of variables and constraints to satisfy the defects at the segment joints. The sie of the nonlinear programming problem in the collocation method is also large and proper locations of grid points are necessary to satisfy all the path constraints. Stochastic methods such as Genetic algorithms, on the other hand, also require large number of function evaluations before convergence. To overcome some of the limitations of the conventional methods, improved solution techniques are developed. Three improved methods are proposed for the solution of trajectory optimization problems. They are • a genetic algorithm employing dominance and diploidy concept. • a collocation method using chebyshev polynomials , and • a hybrid method that combines collocation and direct shooting technique A conventional binary-coded genetic algorithm uses a haploid chromosome, where a single string contains all the variable information in the coded from. A diploid, as the name suggests, uses pair of chromosomes to store the same characteristic feature. The diploid genetic algorithm uses a dominant map for decoding genotype into a stable, consistent phenotype. In dominance, one allele takes precedence over another. Diploidy and dominance helps in retaining the previous best solution discovered and shields them from harmful selection in a changing environment. Hence, diploid and dominance affect a king of long-term memory in the genetic algorithm. They allow alternate solutions to co-exist. One solution is expressed and the other is held in abeyance. In the improved diploid genetic algorithm, dominant and recessive genes are defined based on the fitness evaluation of each string. The genotype of fittest string is declared as the dominant map. The dominant map is dynamic in nature as it is replaced with a better individual in future generations. The concept of diploidy and dominance in the improved method mimics closer to the principles used in human genetics as compared to any such algorithms reported in the literature. It is observed that the improved diploid genetic algorithm is able to locate the optima for a given trajectory optimization problem with 10% lower computational time as compared to the haploid genetic algorithm. A parameter optimization problem arising from an optimal control problem where states and control are approximated by piecewise Chebyshev polynomials is well known. These polynomials are more accurate than the interpolating segments involving equal spaced data. In the collocation method involving Chebyshev polynomials, derivatives of two neighboring polynomials are matched with the dynamics at the nodal points. This leads to a large number of equality constraints in the optimization problem. In the improved method, derivative of the polynomial is also matched with the dynamics at the center of segments. Though is appears the problem size is merely increased, the additional computations improve the accuracy of the polynomial for a larger segment. The implicit integration step size is enhanced and overall size of the problem is brought down to one-fourth of the problem size defined with a conventional collocation method using Chebyshev polynomials. Hybrid method uses both collocation and direct shooting techniques. Advantages of both the methods are combined to give more synergy. Collocation method is used in the starting phase of the hybrid method. The disadvantage of standalone collocation method is that tuning of grid points is required to satisfy the path constraints. Nevertheless, collocation method does give a good guess required for the terminal phase of the hybrid method, which uses a direct shooting approach. Results show nearly 30% reduction in computation time for the hybrid approach as compared to a method in which direct shooting alone is used, for the same initial guess of control. The solutions obtained from the three improved methods are compared with an indirect method. The indirect method requires derivations of the control and adjoint equations, which are difficult and problem specific. Due to sensitivity of the costate variables, it is often difficult to find a solution through the indirect method. Nevertheless, these methods do provide an accurate result, which defines a benchmark for comparing the solutions obtained through the improved methods. Trajectory design and optimization of a RLV(Reusable Launch Vehicle) Demonstrator mission is considered as a test problem for evaluating the performance of the improved methods. The optimization problem is difficult than a conventional launch vehicle trajectory optimization problem because of the following two reasons. • aerodynamic lift forces in the RLV add one more dimension to the already complex launch vehicle optimization problem. • as RLV performs a sub orbital flight, the ascent phase trajectory influences the re-entry trajectory. Both the ascent and re-entry optimization problem of the RLV mission is addressed. It is observed that the hybrid method gives accurate results with least computational effort, as compared with other improved techniques for the trajectory optimization problem of RLV during its ascent flight. Hybrid method is then successfully used during the re-entry phase and in designing the feasible optimal trajectories under the dispersion conditions. Analytical solutions obtained from literature are used to compare the optimized trajectory during the re-entry phase. Trajectory optimization studies are also carried out for the off-nominal performances. Being a thrusting phase, the ascent trajectory is subjected to significant deviations, mainly arising out of solid booster performance dispersions. The performance index during rhe ascent phase is modified in a novel way for handling dispersions. It minimizes the state errors in a least square sense, defined at the burnout conditions ensure possibilities of safe re-entry trajectories. The optimal trajectories under dispersion conditions serve as a benchmark for validating the closed-loop guidance algorithm that is developed for the ascent phase flight. Finally, an on-line trajectory command-reshaping algorithm is developed which meets the flight objectives under the dispersion conditions. The guidance algorithm uses a pre-computed trajectory database along with some real-time measured parameters in generating the optimal steering profiles. The flight objectives are met under the dispersion conditions and the guidance generated steering profiles matches closely with the optimal trajectories.