Langmuir probes are extensively employed to study the plasmas in space and laboratory environments. Successful measurements require a comprehensive modeling of both the plasma environment and the probe conditions in the form of current collection models. In this thesis, the surface morphology implications on the probe current collection are investigated. This problem is applied and solved in the context of a CubeSat regime. The first problem that is investigated is the consequence of surface structural variability on the current measurements. A new model for dealing with non-uniformity of the probe surface structure is developed in this paper. This model is applied to analyze the Langmuir probe data from a sounding rocket mission that was subjected to surface structural non-homogeneities. This model would be particularly useful for CubeSat platforms where elaborate probe design procedures are not feasible. The second problem that is investigated is the surface area implications on Langmuir probe measurements. It has been established that surface area ratio of the spacecraft to that of the probe needs to be sufficiently large to make successful plasma measurements. CubeSats would therefore pose a challenge for employing Langmuir-type instruments to study the space plasma. We inspect the feasibility of making plasma measurements using Langmuir probes subjected to CubeSat area constraints. This analysis is done for a forthcoming Utah State University (USU)/Space Dynamics Lab (SDL) CubeSat mission.

Pour mener à bien leur mission, les satellites de télécommunications doivent rester à la verticale d'un même point de la Terre, sur une orbite dite géostationnaire, pour laquelle la période de révolution des satellites sur leur orbite est identique à la période de rotation de la Terre sur elle-même. Cependant, à cause des perturbations orbitales, les satellites tendent à s'en éloigner, et il est alors nécessaire de concevoir des stratégies de commande pour les maintenir dans un voisinage de cette position de référence. Du fait de leur grande valeur de poussée, les systèmes à propulsion chimique ont largement été utilisés, mais aujourd'hui les systèmes à propulsion électrique avec leur grande impulsion spécifique sont des alternatives viables pour réduire la masse d'ergols du satellite, et ainsi le coût au lancement, ou allonger la durée de vie du satellite, ce qui permettrait de limiter l'encombrement dans l'espace. Cependant, l'utilisation d'un tel système propulsif induit des contraintes opérationnelles issues en partie du caractère limité de la puissance électrique disponible à bord. Ces contraintes sont difficiles à prendre en compte dans la transcription du problème de maintien à poste en un problème de contrôle optimal à consommation minimale avec contraintes sur l'état et le contrôle. Ce manuscrit propose deux approches pour résoudre ce problème de commande optimale. La première, basée sur le développement et l'exploitation de conditions nécessaires d'optimalité, consiste à découper le problème initial en trois sous-problèmes pour former une méthode de résolution à trois étapes. La première étape permet de résoudre un problème de maintien à poste expurgé des contraintes opérationnelles, tandis que la deuxième, initialisée par le résultat de la première, produit une solution assurant le respect de ces dernières contraintes. La troisième étape permet d'optimiser la valeur des instants d'allumage et d'extinction des propulseurs dans le cadre du formalisme des systèmes à commutation. La seconde approche, dite " directe ", consiste à paramétrer le profil de commande par une fonction binaire et à le discrétiser sur l'horizon temporel de résolution. Les contraintes opérationnelles sont ainsi facilement transcrites en contraintes linéaires en nombres entiers. Après l'intégration numérique de la dynamique, le problème de contrôle optimal se résume à un problème linéaire en nombres entiers. Après la résolution du problème de maintien à poste sur un horizon court d'une semaine, le problème est résolu sur un horizon long d'un an par résolutions successives sur des horizons courts d'une durée de l'ordre de la semaine. Des contraintes de fin d'horizon court doivent alors être ajoutées afin d'assurer la faisabilité de l'enchaînement des problèmes sur l'horizon court constituant le problème sur l'horizon long. / Geostationary spacecraft have to stay above a fixed point of the Earth, on a so-called geostationary Earth orbit. For this orbit, the orbital period of the spacecraft is equal to the rotation period of the Earth. Because of orbital disturbances, spacecraft drift away their station keeping position. It is therefore mandatory to create control strategies in order to make the spacecraft stay in the vicinity of the station keeping position. Due to their high thrust capabilities, chemical thrusters have been widely used. However nowadays electric propulsion based thrusters with their high specific impulse are viable alternative in order to decrease the spacecraft mass or increase its longevity. The use of such a system induce the necessity to handle operational constraints because of the limited on-board power. These operational constraints are difficult to take into account in the mathematical transcription of the station keeping problem in an optimal control problem with control and state constraints. This thesis proposed two techniques in order to solve this optimal control problem. The first one is based on the computation of first order necessary conditions and consists in decomposing the overall problem in three sub-problems, leading to a three-step decomposition method. The first step solves an optimal control problem without the operational constraints. The second steps enforces these operational constraints thanks to dedicated equivalence schemes and the third one optimises the switching times of the control profile thanks to a method borrowed from the switched systems theory. The second proposed method consists in parametrising the on-off control profile with binary functions. After a time discretisation of the station keeping horizons, the operational constraints are easily recast as linear constraints on integer variables, the dynamics is numerically integrated and the station keeping problem is recast as a mixed integer linear programming problem. After the resolution of the problem over a short time horizon of one week, the station keeping problem is solved over a long time horizon of one year. To this end, the long time horizon is split in shorter horizons over which the problem is successively solved. End-of-cycle constraints have been set up in order to ensure the feasibility of the solution one short horizon after another.

京都大学 / 0048 / 新制・課程博士 / 博士(工学) / 甲第20375号 / 工博第4312号 / 新制||工||1668(附属図書館) / 京都大学大学院工学研究科電気工学専攻 / (主査)教授 山川 宏, 教授 引原 隆士, 准教授 海老原 祐輔 / 学位規則第4条第1項該当 / Doctor of Philosophy (Engineering) / Kyoto University / DFAM

京都大学 / 0048 / 新制・課程博士 / 博士(工学) / 甲第21069号 / 工博第4433号 / 新制||工||1689(附属図書館) / 京都大学大学院工学研究科電気工学専攻 / (主査)教授 山川 宏, 教授 松尾 哲司, 准教授 海老原 祐輔 / 学位規則第4条第1項該当 / Doctor of Philosophy (Engineering) / Kyoto University / DFAM

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The plasma environment of an active comet provides a unique setting for plasma physics research. The complex interaction of newly created cometary ions with the flowing plasma of the solar wind gives rise to a plethora of plasma physics phenomena, that can be studied over a large range of activity levels as the distance to the sun, and hence the influx of solar energy, varies. In this thesis, we have used measurements of the spacecraft potential by the Rosetta Langmuir probe instrument (LAP) to study the evolution of activity of comet 67P/Churyumov-Gerasimenko as it approached the sun from 3.6 AU in August 2014 to 2.1 AU in March 2015. The measurements are validated by cross-calibration to a fully independent measurement by an electrostatic analyzer, the Ion Composition Analyzer (ICA), also on board Rosetta. The spacecraft was found to be predominantly negatively charged during the time covered by our investigation, driven so by a rather high electron temperature of ~5 eV resulting from the low collision rate between electrons and the tenuous neutral gas. The spacecraft potential exhibited a clear covariation with the neutral density as measured by the ROSINA Comet Pressure Sensor (COPS) on board Rosetta. As the spacecraft potential depends on plasma density and electron temperature, this shows that the neutral gas and the plasma are closely coupled. The neutral density and negative spacecraft potential were higher in the northern hemisphere, which experienced summer conditions during the investigated period due to the nucleus spin axis being tilted toward the sun. In this hemisphere, we found a clear variation of spacecraft potential with comet longitude, exactly as seen for the neutral gas, with coincident peaks in neutral density and spacecraft potential magnitude roughly every 6 h, when sunlit parts of the neck region of the bi- lobed nucleus were in view of the spacecraft. The plasma density was estimated to have increased during the investigated time period by a factor of 8-12 in the northern hemisphere and possibly as much as a factor of 20-44 in the southern hemisphere, due to the combined effects of seasonal changes and decreasing heliocentric distance. The spacecraft potential measurements obtained by LAP generally exhibited good correlation with the estimates from ICA, confirming the accuracy of both of these instruments for measurements of the spacecraft potential. / <p>QC 20200602</p>

Electromagnetic compatibility (EMC) is a very important and increasingly relevant technology in use and is closely related to many technologies such as automobiles and aerospace industry technology including aircraft and spacecraft. Achieving electromagnetic compatibility between equipment inside the device is more than necessary to avoid the problem of interference leading to serious problems. More than that, achieving electromagnetic compatibility for the device as a complete system is quite challenging. The device as a whole system is necessary to be compatible with its electromagnetic environment in order to avoid the problem of interference with other devices which also leads to safety issues. The objective of the thesis is to measure EMC radiated emission from the aircraft as a complete system and to know its compatibility with its electromagnetic environment, by building knowledge of the challenges that arise when conducting electromagnetic compatibility measurements of the aircraft as a whole system outside a protected environment. The challenge here is that there are no standards for EMC radiated emission of the aircraft as a complete system in an Open Area Test Site (OATS). This required us to research and try to relate what was done in this field to try to plan to build an OATS to conduct EMC radiated emission on aircraft as a whole system. We have come up with, that the measurement of EMC radiated emissions performed on the aircraft as a complete system in an OATS at Skellefteå Airport, shows similar results to those obtained on Aircraft SIMulator (ASIM) tested in a Fully Anechoic Chamber (FAC). In addition, we performed a civil airport of Skellefteå site validation measurement and obtained the result that the site complies with the OATS requirements in CISPR 16-1-4: 2019 standard. This work is considered a building block for other EMC studies in the field of space technology, which calls for the need to think about achieving electromagnetic compatibility to avoid all that results from electromagnetic interference (EMI) on the safety of pilots, astronauts, and all human life.

Low-thrust interplanetary spacecraft trajectory optimization poses a uniquely difficult problem to solve because of the inherent nonlinearities of the dynamics and constraints as well as the large size of the search space of possible solutions. Tools currently exist that optimize low-thrust interplanetary trajectories, but these tools are rarely openly available to the public, and when they are available they require multiple interfaces between multiple different packages. The goal of this work is to present a new piece of low-thrust interplanetary spacecraft trajectory optimization software that is open-source and entirely self-contained so that more people can have access to the ability to design interplanetary trajectories. To achieve this goal, a gradient-descent based nonlinear programming method, called the interior point method, was used. The nonlinear programming method was chosen so that results from this work could be compared and contrasted with results from Spacecraft Trajectory Optimization Suite (STOpS), which uses heuristics to iterate towards a solution. Interior point methods are popular because of their ability to handle large amounts of equality and inequality constraints, which is a characteristic that is valuable for low-thrust interplanetary spacecraft trajectories. The software developed, Interior Point Optimizer (IP Optimizer), was then validated against test cases with known solutions to ensure that the software delivered the intended results. Lastly, a constraint satisfaction, a minimum-time, and a maximum-final-mass optimization problem were solved and compared with literature to illustrate the advantages of IP Optimizer and the methods it employs. For the constraint satisfaction problem, IP Optimizer was able to find a solution that exactly satisfied the desired terminal constraints whereas STOpS had an error of 2.29 percent. In this case, IP Optimizer had a reduced runtime of 15 percent compared to STOpS as well. When minimizing time for a spacecraft transfer, IP Optimizer improved upon the solution found by STOpS by 5.3 percent. The speed of convergence for IP Optimizer was almost twice as fast as STOpS for this case. These results show that IP Optimizer is faster than STOpS at converging on a solution and the solution it converges to has a better objective value and more accurately satisfies the terminal constraints than STOpS. Lastly, the maximum-final-mass problem resulted in an objective value that was only 0.5 percent lower than the value found in literature.

Hardware configurations, a control law, and a steering law are developed for a mixed hardware spacecraft that uses both control moment gyros and reaction wheels. Replacing one or more gyros in a spacecraft with a reaction wheel has potential for cost savings while still achieving much greater performance than using reaction wheels alone. Several simulated tests are run to compare the performance to a traditional all reaction wheel or all control moment gyro spacecraft, including analysis of failure modes and singular configurations. The mixed system performed similarly to all gyro systems, responding within 6% of the gyro system’s time for all nominal cases. It far exceeds the performance of reaction wheel systems, taking only a fourth of the time. It also handles failures better than reduced size gyro systems. As such, it can be an effective cost saving measure for certain satellite missions.