Catalisadores de Ir/Al2O3, Ir-Ru/ Al2O3 e Ru/ Al2O3 com teores metálicos próximos a 30% em peso, foram preparados em vinte etapas de impregnação utilizando-se uma alumina sintetizada no LCP/INPE como suporte. Os catalisadores de Ir e Ir-Ru foram preparados a partir de soluções contendo precursores metálicos clorados pelo método de impregnação incipiente. Os catalisadores de Ru foram preparados a partir de dois precursores metálicos: um clorado e um precursor orgânico. Neste caso, o catalisador originado do precursor clorado foi preparado por impregnação incipiente, enquanto que o catalisador originado do precursor orgânico foi preparado pelo método de impregnação por excesso de volume. Todos os catalisadores foram caracterizados antes e depois dos testes em micropropulsor pelas técnicas: absorção atômica, para a determinação do teor metálico; fisissorção de nitrogênio, para determinações de área específica e distribuição do volume de mesoporos; quimissorção de hidrogênio e MET, para determinações da dispersão e do diâmetro médio das partículas metálicas (dQH e dMET). Os catalisadores foram testados na reação de decomposição de hidrazina em micropropulsor e comparados com o catalisador comercial Shell 405. Os resultados mostraram que os catalisadores contendo Ir apresentaram desempenho similar ao catalisador comercial e que os catalisadores de Ru não devem ser usados em partidas frias. / Ir/Al2O3, Ir-Ru/ Al2O3 and Ru/ Al2O3 catalysts with metallic loading of c. a. 30 %wt., were prepared in twenty impregnation steps using an alumina synthesized at LCP/INPE as support. The Ir and Ir-Ru catalysts were prepared from metallic chloride precursors solutions by incipient impregnation method. The Ru catalysts were prepared from two metallic different precursors: a chloride precursor and an organic precursor. In this case, the catalyst originated from the chloride precursor was prepared by the incipient impregnation method, while the catalyst originated from the organic precursor was prepared by volume excess impregnation method. All the catalysts were characterized before and after the microthruster tests by the following techniques: atomic absorption, for metallic content determination; nitrogen physiosorption, for specific area and mesoporous volume distribution; hydrogen chemisorption and TEM, for dispersion and metallic particles average diameter (dQH and dMET ). The catalysts were tested by the hydrazine decomposition reaction in microthruster and compared with commercial catalyst Shell 405. The results showed that the performance of Ir catalysts are similar to the commercial ones and the Ru catalysts should not be used in cold startups.

Modern gridded ion thrusters for CubeSats operate by generating high power and canpose challenging problems with Electromagnetic Interference (EMI). In order to verifycompatibility with neighbouring equipment, strict standards such as the militarystandard MIL-STD-461G, are required to be followed to achieve ElectromagneticCompatibility (EMC). To avoid abrupt and cataclysmic delays in production time, incase the product fails to comply with the requirements, companies integrate in-housepre-compliance tests into their development phase. The objective is to implementin-house measurement methods on an electric propulsion model NPT30 developedby ThrustMe. This document explains the process and methods to perform conductedemission test on power lines and radiated emission tests in the magneticfield. A custom measurement system integrity verification was developed for theradiated emission test. The presented results provide the engineers at ThrustMe aninsight on the electromagnetic behaviour on the ion thruster NPT30 and whethermodifications need to be included in the next development iteration to mitigate forthe detected excessive emission levels. When EMC methods are implemented earlyon in the development process, there are more pre-emptive mitigation options withless costs in time and money. By performing in-house pre-compliance tests andtaking measures to prepare for the tests at a certified EMC test house, the companycan be more confident in their product at passing the EMC tests. Based on the twoperformed in-house tests, the engineers at ThrustMe began to include mitigationmethods in the following circuit design iterations.

Es wird gezeigt, dass die angestrebte Steuerung der Lastverteilung auf Pilotmeißel und Erweiterungswerkzeug durch den Einsatz eines hydraulischen Thrusters im Pilotstrang erreicht werden kann. Das zur Steuerung bestehende operative Fenster einer solchen Bohrgarnitur ist jedoch im Allgemeinen klein und bedarf für jeden Einsatzfall einer detaillierten Vorausplanung. Ein entsprechender Berechnungsansatz zur Simulation der Vorgänge in einer komplexen Bohrgarnitur mit schaltbarem Erweiterungswerkzeug wurde entwickelt und getestet. Er stützt sich auf die Bingham-Gleichung für die Bohrgeschwindigkeit, die für die speziellen Belange der vorliegenden Untersuchung modifiziert wurde. Die Einflüsse individueller operativer und konstruktiver Parameter auf das Steuerverhalten der komplexen Bohrgarnitur wurden eingehend untersucht. Die Auswertung der Ergebnisse führte zu allgemeinen Planungs- und Einsatzempfehlungen.

A mission architecture consisting of two NASA LONESTAR-2 satellites in Low Earth Orbit is considered. The craft are equipped with cross-communication radios and GPS units. Analyses are conducted for ejection, thruster and attitude maneuvers to achieve objectives of the mission, including sustained communications between the craft. Simulations are conducted to determine the duration of the communication window following the initial separation of the two craft. Recommendations are made to maximize this window while accounting for attitude constraints and the effects of atmospheric drag. Orbital mechanics and control theory are employed to form an algorithm for filtering GPS position fixes. The orbit-determination algorithm accounts for the effects of drag and Earth’s oblateness. Procedures are formed for verifying the initial separation velocities of two spacecraft and for measuring the velocity imparted by impulsive thruster maneuvers. An algorithm is also created to plan the timing and magnitude of corrective thruster maneuvers to align the orbital planes of the two craft. When the craft pass out of communication range, a ground station is used to relay data and commands to conduct state rendezvous procedures. A plan for coordinated attitude maneuvers is developed to strategically utilize the cumulative effects of drag and orbit decay to align the craft over long time periods. The methodologies developed here extend prior research into close proximity operations, forming the foundation for autonomous on-orbit rendezvous under a broader set of initial conditions. / text

Current state-of-the-art plasma thrusters are limited in power density and thrust density by power losses to plasma-facing walls and electrodes. In the case of Hall effect thrusters, power deposition to the discharge channel walls and anode negatively impact the efficiency of the thruster and limit the attainable power density and thrust density. The current work aims to recreate thruster-relevant wall-interaction physics in a quiescent plasma and investigate them using electrostatic probes, in order to inform the development of the next generation of high-power-density / high-thrust-density propulsion devices. Thruster plasma-wall interactions are complicated by the occurrence of the plasma sheath, a thin boundary layer that forms between a plasma and its bounding wall where electrostatic forces dominate. Sheaths have been recognized since the seminal work of Langmuir in the early 1900’s, and the theory of sheaths has been greatly developed to the present day. The theories are scalable across a wide range of plasma parameters, but due to the difficulty of obtaining experimental measurements of plasma properties in the sheath region, there is little experimental data available to directly support the theoretical development. Sheaths are difficult to measure in situ in thrusters due to the small physical length scale of the sheath (order of micrometers in thruster plasmas) and the harsh plasma environment of the thruster. Any sufficiently small probe will melt, and available optical plasma diagnostics do not have the sensitivity and/or spatial resolution to resolve the sheath region. The goal of the current work is to experimentally characterize plasma sheaths xxvi in a low-density plasma that yields centimeter-thick sheath layers. By generating thick sheaths, spatially-resolved data can obtained using electrostatic probes. The investigation focuses on the effects of electron emission from the wall and several factors that influence it, including wall material, wall temperature, wall surface roughness and topology, as well as the scaling of sheaths from the low-density plasma environment towards thruster conditions. The effects of electron emission and wall material are found to agree with classical fluid and kinetic theory extended from literature. In conditions of very strong emission from the wall, evidence is found for a full transition in sheath polarities rather than a non-monotonic structure. Wall temperature is observed to have no effect on the sheath over boron nitride walls independent of outgassing on initial heat-up, for sub-thermionic temperatures. Wall roughness is observed to postpone the effects of electron emission to higher plasma temperatures, indicating that the rough wall impairs the wall’s overall capacity to emit electrons. Reductions in electron yield are not inconsistent with a diffuse-emission geometric trapping model. Collectively, the experimental data provide an improved grounding for thruster modeling and design.Current state-of-the-art plasma thrusters are limited in power density and thrust density by power losses to plasma-facing walls and electrodes. In the case of Hall effect thrusters, power deposition to the discharge channel walls and anode negatively impact the efficiency of the thruster and limit the attainable power density and thrust density. The current work aims to recreate thruster-relevant wall-interaction physics in a quiescent plasma and investigate them using electrostatic probes, in order to inform the development of the next generation of high-power-density / high-thrust-density propulsion devices. Thruster plasma-wall interactions are complicated by the occurrence of the plasma sheath, a thin boundary layer that forms between a plasma and its bounding wall where electrostatic forces dominate. Sheaths have been recognized since the seminal work of Langmuir in the early 1900’s, and the theory of sheaths has been greatly developed to the present day. The theories are scalable across a wide range of plasma parameters, but due to the difficulty of obtaining experimental measurements of plasma properties in the sheath region, there is little experimental data available to directly support the theoretical development. Sheaths are difficult to measure in situ in thrusters due to the small physical length scale of the sheath (order of micrometers in thruster plasmas) and the harsh plasma environment of the thruster. Any sufficiently small probe will melt, and available optical plasma diagnostics do not have the sensitivity and/or spatial resolution to resolve the sheath region. The goal of the current work is to experimentally characterize plasma sheaths xxvi in a low-density plasma that yields centimeter-thick sheath layers. By generating thick sheaths, spatially-resolved data can obtained using electrostatic probes. The investigation focuses on the effects of electron emission from the wall and several factors that influence it, including wall material, wall temperature, wall surface roughness and topology, as well as the scaling of sheaths from the low-density plasma environment towards thruster conditions. The effects of electron emission and wall material are found to agree with classical fluid and kinetic theory extended from literature. In conditions of very strong emission from the wall, evidence is found for a full transition in sheath polarities rather than a non-monotonic structure. Wall temperature is observed to have no effect on the sheath over boron nitride walls independent of outgassing on initial heat-up, for sub-thermionic temperatures. Wall roughness is observed to postpone the effects of electron emission to higher plasma temperatures, indicating that the rough wall impairs the wall’s overall capacity to emit electrons. Reductions in electron yield are not inconsistent with a diffuse-emission geometric trapping model. Collectively, the experimental data provide an improved grounding for thruster modeling and design.

Aujourd’hui les propulseurs à effet Hall ont gagné une position dominante dans le marché des propulseurs électriques spatiales. Ce grand succès est du surtout à leur simplicité de réalisation (par rapport aux autres typologies des propulseurs) et à leur efficacité (par rapport aux propulseurs chimiques traditionnels). Les propulseurs à effet Hall sont aujourd’hui utilisés sur un très grand nombre des plateformes satellitaires (surtout pour les télécommunications). Les composants principales d’un propulseur à effet Hall sont : le circuit magnétique, le canal plasma, l’anode (placé au fond du canal plasma avec injecteur du gaz) et la cathode (placée à l’extérieur du canal plasma). Le fonctionnement d’un propulseur à effet Hall est basé sur la génération d’un champ électrique axial (généré entre l’anode et la cathode) et d’un champ magnétique radial (perpendiculaires entre eux). Le champ magnétique a le rôle de former une zone de très forte concentration électronique (il emprisonne les électrons générés par la cathode) pour permettre aux atomes neutres du gaz de se ioniser. Le champ électrique a le rôle d’accélérer les ions vers l’extérieur du canal. Cette accélération génère la poussée. Le champ magnétique joue un rôle crucial dans le fonctionnement d’un propulseur à effet Hall. La forme du champ magnétique impacte sur les performances propulsifs et sur l’érosion du propulseurs. La topologie magnétique classique des propulseurs à effet Hall n’a subi presque pas des changements depuis les années de développement de cette technologie parce qu’elle garanti des performances propulsifs assez satisfaisantes. Aujourd’hui, avec les nouvelles exigences propulsifs, il y a une très forte nécessité des moteurs avec une durée de vie plus longue. Des nouvelles topologie magnétique innovante sont proposés aujourd’hui comme par exemple le "Magnetique-Shielding" ou le "Wall-Less" . Ces topologies magnétique bouleverse complètement la topologie magnétique classique (en gardant des performances propulsif satisfaisantes) pour protéger le moteur de l’érosion du plasma. Dans cette thèse une autre approche a été adopté. Nous avons pensé d’utiliser une topologie magnétique classique et de déplacer les parties du circuit magnétique attaquées par l’érosion vers des zones moins dangereuses. Nous avons agit sur la forme du circuit magnétique et pas sur la forme de la topologie magnétique pour garder les même performances propulsifs de la topologie magnétique classique. L’objectif de la thèse était de créer des outils pour le design et l’optimisation des circuits magnétiques des propulseurs à effet Hall. Un algorithme nommé ATOP a été créé dans l’équipe de recherche GREM3 du laboratoire LAPLACE de Toulouse. Cette thèse a contribué à la création de la section d’optimisation paramétrique (ATOPPO) et d’une section d’optimisation topologique basée sur les algorithmes génétiques (ATOPTOga) de l’algorithme ATOP. Les algorithme conçues dans cette thèse permettent d’optimiser des propulseurs existants (en terme de forme, masse et courant) ou de concevoir des nouveaux propulseurs (nécessité de concevoir un nouveau propulseur capable de reproduire une topologie magnétique précise). Les algorithmes développées ont démontrés leur efficacité à travers leur application sur un propulseur réel, le PPS-1350-E® de SAFRAN. Ce propulseur a été optimisé en terme de masse et de courant bobines (minimisation de la masse et du courant bobines). Les algorithmes développés ont démontré donc leur efficacité comme instrument d’optimisation et de design. Ces deux algorithmes ont été utilisé pour le design d’un circuit magnétique innovant qui a comme objectif de réduire l’érosion du moteur. Les résultats de ce processus de design ont amené à la réalisation et à la construction d’un prototype qui possède la même topologie magnétique du propulseur PPS- 1350-E® commercialisé par SAFRAN mais avec une circuit magnétique de forme différente. / Nowadays, two types of space propulsion engines exist: the most common ones are the chemical propulsion engines which provide high thrust impulses allowing fast orbit transfers. But this technology requires a large amount of propellant and is not suitable for interplanetary displacements, whose propellant mass requirements are too high. The second type of propulsion engine is based on electrical propulsion that provide very low but continuous thrust, resulting in huge propellant savings at the cost of longer spacecraft transfer time. The main advantage of electric thrusters lies in their highly efficient utilization of propellant mass. The corresponding reduction in necessary propellant supply makes it possible to board a greater portion of useful payload possible. Hall thruster belongs to the electric spacecraft engines typology and it is constituted of a cylindrical plasma channel, an interior anode, an external cathode and a magnetic circuit that generates a primarily radial magnetic field across the plasma channel. The magnetic circuit of a Hall effect thruster must generate a specific electromagnetic distribution inside and near the outlet of the plasma channel. In a Hall Effect thruster the magnetic circuit constitutes more than half of the whole thruster. Consequently the design of this magnetic circuit must be optimized in order to minimize the embedded mass. The main components of this circuit are the coils which produce the magnetic and ferromagnetic parts which guide the and to shape the density. Usually the magnetic circuit includes four (or more) external coils located around the exterior radius of the plasma channel and one internal coil around the interior radius of the plasma channel. All the coils are supplied by the same DC. Two coils located at the rear of the plasma channel can be also used to perform the magnetic topography. The first objective of the design process of this type of structure is to obtain a specific magnetic topography in the thruster channel with given magnetic field radial component values and a certain inclination of the corresponding field lines. By considering nowadays the requirements in terms of lifetime new specifications concerning in particular the erosion of ceramic wall have to be taken into account.This weakness has its origins in plasma-surface interaction inside the discharge chamber. Thus, to solve this problem it has been proposed to move the ionization zone outside the thruster channel in order to avoid contact between the ions and ceramic material. Thanks to new studies carried on the impact of magnetic topology, new magnetic configurations have been highlighted to improve the efficiency and reduce the erosion of the ceramic walls. The aim of this work is to develop tools for solving this inverse magnetostatic problem and to find new magnetic structures that are able to produce these new magnetic cartographies. Methods based on topological optimization have already been developed for these structures. The algorithm ATOPTO (Algorithm To Optimize Propulsion with Topology Optimization) has already demonstrated its efficiency. In this work we try to extend the scope of the algorithm ATOP by adding a new parametric optimization section called ATOPPO. The ATOP algorithm becomes a very versatile optimization tool for Hall Effect thruster magnetic circuits.

abstract: Two methods of improving the life and efficiency of the Pulsed Inductive Thruster (PIT) have been investigated. The first is a trade study of available switches to determine the best device to implement in the PIT design. The second is the design of a coil to improve coupling between the accelerator coil and the plasma. Experiments were done with both permanent and electromagnets to investigate the feasibility of implementing a modified Halbach array within the PIT to promote better plasma coupling and decrease the unused space within the thruster. This array proved to promote more complete coupling on the edges of the coil where it had been weak in previous studies. Numerical analysis was done to predict the performance of a PIT that utilized each suggested switch type. This model utilized the Alfven velocity to determine the critical mass and energy of these theoretical thrusters. / Dissertation/Thesis / Masters Thesis Aerospace Engineering 2018

Since 2000, a relatively new electric thruster concept has been in research, development, and production at Thales Electron Devices in Germany. This High Efficiency Multistage Plasma Thruster, or HEMPT, has promising lifetime capabilities due to its plasma confinement system. However, the permanent magnet system that offers this and other benefits also creates a hollow plume, where ions are accelerated at angles rather than up the thruster centerline, causing a dip in ion current along the centerline. A laboratory model, built at JPL, was run at Cal Poly to characterize this plume shape and implement a shield to restore a conical shape to the plume. A similar solution was used on a different type of thruster, a cylindrical hall thruster, at Princeton with excellent results. A shield was designed to shunt the magnetic field outside the thruster, where the Princeton experiments have identified a radial magnetic field as the cause for this hollow plume. The thruster was run with and without the shield, taking measurements of the ion current in the plume using a linear probe drive. The shield fixed the plume shape, increasing centerline current by 48%, however it also had detrimental effects on thruster performance, causing a decrease in thrust, specific impulse, and cut the total efficiency in half. The shield design was reexamined and a new design has been suggested for future testing of the HEMPT to restore performance while still fixing the plume shape.

<p>The commercial CubeSat industry has experienced explosive growth recently, and with falling  costs  and  growing  numbers  of  launch  providers,  the  trend  is  likely  to  continue.  The scientific missions CubeSats could complete are expanding, and this has resulted in a demand for reliable  high  specific  impulse  nanosatellite  propulsion  systems.  Interest  in  liquid-fed  pulsed plasma thrusters (LF-PPTs) to fulfill this role has grown lately. Prior work on a nanosatellite LF-PPT was done in the Purdue Electric Propulsion and Plasma Laboratory, but its high operational voltage and electrode size would be disadvantageous for integration on a CubeSat, which have strict volume limitations and provide only tens of Watts in power at low voltages. This work aims to address those disadvantages and further advance the development of a nanosatellite LF-PPT by reducing the operating voltage and removing long plate electrodes to prevent energy losses on components other than the expelled plasma sheet. Two major objectives are pursued: to construct a  coaxial  pulsed  plasma  thruster  operating  with  10s  to  100s  of  volts  and  to  characterize  the temporal evolution of the discharge parameters in this low-voltage operation scenario. </p> <p>It  took  three  experimental  design  iterations,  all  of  which  used  a  260  <em>uF</em> ,  400 <em>V</em> film capacitor, to arrive at a functional coaxial pulsed plasma thruster. First, a button gun was tested. It produced  a  peak  current  of ~16<em> kA</em>,  which  serves  as  the  expected  maximum  for  the  later experiments. Due to the presence of parasitic arcing, it revealed that electrical lines needed to be removed from vacuum chamber to enable testing at a wide range of pressures. Second, a coaxial PPT was designed, built, and tested. This design confirmed operation at discharge voltages <100 <em>V</em> across the plasma, achieving one of the project’s aims, and produced a peak current of 7.4 <em>kA</em>. However,  necessity  to  better  align  the  cathode and  provide  an  unobstructed  camera  view  for observation of the discharge column attachment to the cathode surface forced additional system redesign. Third, a revised coaxial PPT was built and tested. Using air as a propellant, the discharge generated a peak current of 10.4 <em>kA</em> at a mass flow rate of 2 mgs. The PPT cathode was imaged with an ICCD camera over a wide range of pressures, and the photos indicated “spotless” diffuse arc attachment to the cathode, which serves as evidence to expect low erosion rates. The direct measurements of the cathode erosion rate are planned for future. </p>

Electric propulsion systems utilize electrical energy to produce thrust for spacecraft propulsion. These systems have multiple applications ranging from Earth orbit North-South station keeping to solar system exploratory missions such as NASA’s Discovery, New Frontiers, and Flagship class missions that focus on exploring scientifically interesting targets. In an electromagnetic thruster, a magnetic field interacting with current in an ionized gas (plasma) accelerates the propellant to produce thrust. Pulsed inductive thrusters rely on an electrodeless discharge where both the magnetic field in the plasma and the plasma current are induced by a time-varying current in an external circuit. The multi-dimensional acceleration model for a pulsed inductive plasma thruster consists of a set of circuit equations describing the electrical behavior of the thruster coupled to a one-dimensional momentum equation that allow for estimating thruster performance. Current models lack a method to account for the time-varying energy distribution in an inductive plasma accelerator.