As conventional chemical rockets reach the outer limits of their abilities, significant research is going into alternative thruster technologies, some of which decouple the maximum thrust and efficiency from the propellant's internal chemical energy by supplying energy to the propellant as needed. Of particular interest and potential is the electrically powered thruster, which promises very high specific thrust using relatively inexpensive and stable propellant gasses. Some such thrusters, specifically ion thrusters, have achieved significant popularity for various applications. However, there exist other classes of electrical thrusters which promise even higher levels of efficiency and performance. This thesis will focus on one such thruster type - the magnetoplasmadynamic thruster - which uses an ionized propellant flow and large currents to accelerate the propellant gas by electrical and magnetic force interactions. The necessary background will be presented in order to understand and characterize the operation of such devices, and a theoretical model will be developed in order to estimate the levels of performance which can be expected. Simulations will be performed and analyzed in order to better understand the principles on which these devices are designed. Finally, a thruster package will be designed and built in order to test the performance of the device and accuracy of the model. This will include a high-current power supply, ignition circuit, gas delivery system, and nozzle. Finally, the measured performance of this thruster package will be measured and compared to the theoretical predictions in order to validate the models constructed for this type of thruster.

There has been a sharp increase in the market for electric propulsion systems for small satellites in the recent years. Electric propulsion systems have become smaller, more efficient and cheaper, which made them ideal for small satellites because they have a low thrust requirement and benefit significantly from the high specific impulse (Isp) that is characteristic to electric thrusters. These thrusters are generally fabricated and tested manually and there is a low degree of automation in the process. As the demand for the thrusters increases, there is a need to improve the speed of the fabrication and testing process. The Test Platform for Electronics Modules in Electric Thrusters (TESPEMET) project at ThrustMe is an attempt to design a system that addresses this issue. The vision is to have a test platform that facilitates the testing of ThrustMe's Electric Thrusters by applying various source and load conditions, emulating events while performing instrumentation during the test process and generating a test report at the end of the test procedure. The development of such a test platform would enable and accelerate the test and qualification process of the thrusters significantly. This thesis presents the technical design of this test platform along with the results obtained, encountered problems and solutions. Future work and design changes have also been proposed based on the knowledge gained during the Research and Development process.

An accurate satellite attitude and orbit control is a key factor for a successful mission. It guarantees for example sun acquisition on solar panels, fine pointing for optimal telescope usage or satellite lifting to reach higher orbits, when required. Furthermore attitude and orbit control is applied to compensate any occurring disturbances within the space environment. The problem tackled in the present thesis is the optimization of thruster commanding to perform spacecraft orbital manoeuvres. The main objective is to develop different algorithms that are suitable for on-board implementation and to compare their performance. For an optimal thruster command distribution the algorithms shall solve linear programming (or optimization) problems, more exact they shall compute thruster on-times to generate desired torques and/or forces, which are requested by the on-board software. In total three different algorithms are developed of which the first one is based on the pseudoinverse of a matrix, the second one is a variation of the Simplex method and the third one is based on Karmarkar’s algorithm, which belongs to the interior-point methods. The last two methods are well known procedures to solve linear programming problems and in theory they have been analyzed before. However this paper proves their practical application and industrial feasibility for orbital manoeuvres of the weather satellites of ESA’s MTG project and their scalability to any number of thrusters on a generic satellite for 6 degrees of freedom manoeuvres. There are 6 MTG satellites and each has 16 one-sided reaction control thrusters, placed at specific positions and pointed towards defined directions. Physical mechanisms limit the thrusters output to minimum on- and off-times. The focus of this thesis will be on the orbital transfer mode, due to the high disturbances that arise during four motor firing sessions at the apogee, executed to reach higher orbits and finally GEO. The firing sessions are performed by a liquid apogee engine and while this engine is in boost mode, the thrusters shall be used for attitude control only. The technique (nominal case) developed by OHB for this maneuver and currently operational uses 4 thrusters only, which are all pointing in the engine’s direction. They are also used to settle the fuel before the engine is turned on. For control the Pseudoinverse method is applied. If one of the 4 thrusters fails, the backup scenario takes place, which includes using 4 totally different thrusters and no fuel settling, due to their unfavorable position with respect to the engine. The initial idea of this work was to develop a controller for 6 thrusters, using only 2 of the 4 nominal case thrusters, to have a better control performance in the backup case. The Pseudoinverse method was developed by OHB before, thus only small changes needed to be applied to work with 6 thrusters. The two other algorithms, based on the Simplex and Karmarkar method, were completely developed from scratch. To analyze their performance several tests were executed. This includes unit tests on a simple computer hardware with different input, Monte Carlo simulations on a cluster to test if the algorithms are suitable for MTG orbital manoeuvres and the application to 12 thrusters, mounted on a generic satellite to generate torques and forces at the same time for 6 degrees of freedom manoeuvres. For each thruster configuration the worst case outputs are shown in so called minimum control authority plots. The performance analysis consists of the maximum and average deviation between requested and generated torque/force, the average computed thruster on-times, the algorithms computation(running) time and iteration steps. For MTG the test results clearly confirm that the usage of 6 thrusters leads to more accurate generated torques and better control authority, than using only 4 thrusters. The Simplex method stands out here in particular, showing excellence performance regarding torque precision. Nevertheless the accuracy goes at the expense of computation effort. While the Pseudoinverse method is very fast and needs only one iteration step, the Simplex is half a magnitude, the Karmarkar one magnitude slower. But the latter lead to lower thruster on-times in terms of firing duration and thus fuel consumption is reduced. Also it is shown that Simplex and Karmarkar can control 12 thrusters at the same time to generate torques and forces, which proves their scalability to any thruster distribution. In the end it comes to the question whether generating a more accurate torque/force or the computational effort, which is strongly hardware dependent, is more important. A decision which depends on the mission’s objective. This paper shows that all three implemented algorithms are able to handle attitude control in the MTG backup scenario and beyond.

In recent years, the Helicon Plasma Thruster (HPT) has become one of the most promising technologies of in-space electric propulsion. T4i Technology for Propulsion and Innovation S.P.A. is one of the leading companies working with this new type of systems, and their thruster, REGULUS, is the first HPT ever to be operated in orbit. To better assess the performance of the motor, the company has developed, in conjunction with the University of Padova and the University of Bologna, a numerical tool called 3DVIRTUS (3Dimensional adVanced fluId dRifT diffUsion plaSma solver), which simulates the plasma dynamics in the production stage of the thruster. The model describes the species present in the plasma (electrons, ions, excited and neutrals) by means of a fluid approach, as the plasma density in this part of the motor is in the order of 1017-1018 m−3. Particularly, the tool considers the Drift-Diffusion (DD) approximation instead of the full set of fluid momentum equations. Unfortunately, for typical discharges applied to HPTs, this assumption is accurate only for the electrons species, but not for the heavy species in the plasma, i.e. ions, excited and neutrals. The thesis project presented in this report, executed in collaboration with T4i S.P.A, proposes an updated numerical tool which solves the fully coupled continuity and momentum equations for the neutrals species in the plasma. The new solver is implemented with OpenFOAM®, a finite volume library written in C++, and the Semi-Implicit Method for Pressure Linked Equations (SIMPLE) is utilised to resolve the pressure-velocity coupling in the continuity and momentum equations. Four different test cases are considered: a one-dimensional typical discharge, a cylindrical discharge, the Schwabedissen GECICP reactor experiment and the Piglet helicon reactor of Lafleur. The obtained results have been compared against the original drift-diffusion solver, and when available, with experimental data. The new tool produced similar results to the older one, even though the neutrals density computed with the former generally presented stronger gradients. Additionally, in the case of the GECICP and Piglet reactors, the agreement in terms of electrons density computed with the new solver was satisfactory compared to the empirical data. Nevertheless, all the analysis performed during the thesis project revealed that the keys to obtain physically realistic results are the boundary conditions for the neutrals’ pressure and velocity, which greatly affects the outcome of the simulations. Overall, the new solver has shown to provide accurate results with reasonable computational time. / Under de senaste åren har Helicon Plasma Thruster (HPT) blivit en av de mest lovande teknikerna för elektrisk framdrift i rymden. T4i Technology for Propulsion and Innovation S.P.A. är ett av de ledande företagen som arbetar med denna nya typ av system, och deras motor, REGULUS, är den första HPT som har demonstrerats fungera i omloppsbana. För att bättre kunna bedöma motorns prestanda har företaget tillsammans med universitetet i Padova och universitetet i Bologna utvecklat ett numeriskt verktyg som kallas 3DVIRTUS (3Dimensional adVanced fluId dRifT diffUsion plaSma solver), som simulerar plasmadynamiken i thrusterns produktionsstadium. Modellen beskriver de typer av partikler som finns i plasma (elektroner, joner, exciterade och neutrala) med hjälp av en vätskeapproximation, eftersom plasmatätheten i denna del av motorn är i storleksordningen 10171018 m−3. Särskilt överväger verktyget approximationen Drift-Diffusion (DD) istället för hela uppsättningen vätska ekvationer. Dessvärre, för typiska urladdningar som appliceras på HPT, är detta antagande korrekt endast för elektroner, men inte för de tunga partiklarna i plasma, dvs joner, exciterade och neutrala partiklar. Avhandlingsprojektet som presenteras i denna rapport, utfört i samarbete med T4i S.P.A, föreslår ett uppdaterat numeriskt verktyg som löser de fullständigt kopplade kontinuitets och rörelseekvationerna för neutrala partiklar i plasma. Den nya lösaren implementeras med OpenFOAM®, ett begränsat volymbibliotek skrivet i C++, och Semi-Implicit Method for Pressure Linked Equations (SIMPLE) används för att lösa tryck hastighetskopplingen i kontinuitets och rörelseekvationer. Fyra olika testfall övervägs: en endimensionell typisk urladdning, en cylindrisk urladdning, Schwabedissen GECICP reaktorförsöket och Piglet helicon reaktorn i Lafleur. De erhållna resultaten har jämförts med det ursprungliga driftdiffusions antagandet, och när möjligt, med experimentella data. Det nya verktyget gav liknande resultat som det äldre, även om densiteten av neutrala partiklar beräknad med den tidigare generellt visade starkare gradienter. Dessutom, när det gäller GECICP och Piglet reaktorerna, var överenskommelsen i termer av elektrontäthet beräknad med den nya lösaren tillfredsställande jämfört med empiriska data. Ändå avslöjade all analys som gjordes under avhandlingsprojektet att nycklarna för att få fysiskt realistiska resultat är randvillkoren för de neutrala partiklarnas tryck och hastighet, vilket i hög grad påverkar resultatet av simuleringarna. Sammantaget har den nya lösaren visat sig ge noggranna resultat med rimlig beräkningstid.

In order to enable comparison of Helicon ion thruster performance across different vacuum test facilities, an understanding of the effect of operating pressure on plasma plume properties is required. Plasma property measurements are compared for thruster operation at two separate vacuum facility operating pressures to determine the effect of neutral ingestion on Helicon ion thruster operation. The ion energy distribution function (IEDF), electron temperature, ion number density, and plasma potential are measured along the thruster main axis for a replica of the Madison Helicon eXperiment. Plasma property values recorded at the ‘high-pressure condition’ (3.0×10^(-4) Torr corrected for argon) are compared to values recorded at the ‘low-pressure condition’ (1.2×10^(-5) Torr corrected for argon) for thruster operation at 100 - 500 watts radio frequency forward power, 340 – 700 gauss source region magnetic field strength, and 1.3 - 60 sccm argon volumetric flow rate (0.039-1.782 mg/s). Differences in plasma behavior at the ‘high-pressure condition’ result from two primary neutral-plume interactions: collisions between accelerated beam ions and ingested neutrals leading to a reduction of ion energy and neutral ionization downstream of the thruster exit due to electron-neutral collisions. Electron temperature at higher operating pressures is lowered due to an electron cooling effect resulting from repeated collisions with neutral atoms. Results suggest that Helicon ion thruster plasma properties are greatly influenced when subjected to neutral ingestion.

The PEGASES thruster (Plasma Propulsion with Electronegative Gases) is a novel type of electric thruster for space propulsion. It uses negative and positive ions produced by an inductively coupled radio frequency discharge to create the thrust by electrostatically accelerating the ions through a set of grids. A magnetic filter is used to increase the amount of negative ions in the cavity of the thruster. The PEGASES thruster is not only a source to create a strongly negative ion plasma or even an ion-ion plasma but it can also be used as a classical ion thruster. This means that a plasma is created and only the positive ions are extracted and accelerated making it necessary to neutralize the plasma behind the acceleration stage like in other ion thrusters. The performances of the PEGASES thruster have been investigated mainly in xenon in order to compare the obtained results with RIT-type ion thrusters. The thruster has been investigated with the help of a variety of probes such as a Langmuir probe, a planar probe, a capacitive probe and a RPA (Retarding Potential Analyzer). In addition, an ExB probe has been developed to measure the velocity of the ions leaving the thruster and to differentiate between the ion species present in the plasma.

The most common type of electric propulsion in space exploration is the Hall Effect Thruster (HET), mainly due to its high specific impulse and high thrust to power ratio. However, uncertainties about the thruster's lifetime prediction have prevented widespread integration of HETs. Among these limitations, wall erosion of acceleration channel is of greatest concern. The experimental methods of erosion are time consuming and costly, and they are often limited to one single configuration. Hence, developing a computational model not only decreases the costs but also shortens the design time of a HET. This thesis investigates the feasibility of a uid erosion modelling with a multi-physics software (COMSOL) to further decrease the time and the development cost. First of all, this thesis provides an overview of available plasma modelling techniques and the physics behind the erosion phenomenon. Moreover, the effective parameters and available modules in the multiphysics software as well as their theoretical background were studied and discussed in detail. The Electron Anomalous phenomenon and pressure instability are determined as the main limiting factors for such a model. A non-magnetized model is included to find an optimal value for pressure and to reduce the probability of pressure instability occurrence in magnetized model. To fulfill this task, several simulations for various pressure values (0.005 Torr, 0.05 Torr, and 0.5 Torr) were conducted. Next, the simulation of magnetized/full model has been carried out with addition of magnetic coils in non-magnetized model. To avoid the Electron Anomalous phenomenon, the Bohm diffusion approach was implemented. In addition, a full Particle-In-Cell (PIC) simulation of a typical HET (SPT-100) with the similar input parameters as in fluid model was conducted, and the results were compared and validated using experimental data. The PIC model was intended to be utilized to investigate the accuracy of erosion model in multiphysics software. The results of this thesis indicate that current application of erosion model in COMSOL is not possible whilst high accuracy of the erosion model based on PIC approach can be achieved. Finally, the application of semi-empirical method through direct input of magnetic field data can allow short time simulation of a HET in COMSOL to gain insight about the preliminary behaviour of plasma, however, the simulation of an erosion model requires either a built-in PIC algorithm in COMSOL or a PIC based code.

"The pulsed plasma thruster (PPT) is an onboard electromagnetic propulsion device currently being considered for use in various small satellite missions. The work presented in this thesis is directed toward improving PPT performance using a control engineering approach along with externally applied magnetic fields. An improved one dimensional electromechanical model for PPT operation is developed. This slug model represents the PPT as an LRC circuit with a dynamics equation for the ablated plasma. The improved model includes detailed derivation for the induced magnetic field and a model for the plasma resistance. A modified electromechanical model for the case of externally applied magnetic fields is also derived for the parallel plate geometry. A software package with a graphical user interface (GUI) is developed for the simulation of various PPT types, geometric configurations, and parameters The simulations show excellent agreement with data from the Lincoln Experimental Satellite (LES)-6, the LES-8/9 PPT and the Univ. of Tokyo PPT. The control objective employed in this thesis involves the maximization of the specific impulse and thrust efficiency of the PPT, which are each directly related with the exhaust velocity of the thruster. This objective is achieved through the use of an externally applied magnetic field as a system actuator. To simulate an open-loop constant-input controller the modified electromechanical PPT model is applied to the various PPT configurations. In this controller the external magnetic field was applied as constant throughout or portions of the PPT channel. For the Univ. of Tokyo PPT a magnetic field applied over the entire 6-cm long channel increases the specific impulse and thrust efficiency by 10% over the case that the filed is applied in the first 1.75 cm of the PPT channel. The magnitude of these increases compare well with the results of the UOT applied B-field experiments. For the LES-6 and LES-9 PPTs, the simulations predicts significant performance enhancements with approximately linear increases for the specific impulse, thrust efficiency and impulse bit. "

"As new, smaller satellites are built, the need for improved on-board propulsion systems has grown. The pulsed plasma thruster has received attention due to its low power requirements, its simple propellant management, and the success of initial flight tests. Successful integration of PPTs on spacecraft requires the comprehensive evaluation of possible plume-spacecraft interactions. The PPT plume consists of neutrals and ions from the decomposition of the Teflon propellant, material from electrode erosion, as well as electromagnetic fields and optical emissions. To investigate the PPT plume, an on-going program is underway at WPI that combines experimental and computational investigations. Experimental investigation of the PPT plume is challenging due to the unsteady, pulsed as well as the partially ionized character of the plume. In this thesis, a triple Langmuir probe apparatus was designed and used to obtain electron temperature and density measurements in the plume of a PPT. This experimental investigation provides further characterization of the plume, much needed validation data for computational models, and is useful in thruster optimization studies. The pulsed plasma thruster used in this study is a rectangular geometry laboratory model built at NASA Lewis Research Center for component lifetime tests and plume studies. It is almost identical in size and performance to the LES 8/9 thruster, ablating 26.6 ug of Teflon, producing an impulse bit of 256 uN-s and a specific impulse of 986 s at 20 J. All experiments were carried out at NASA LeRC Electric Propulsion Laboratory. The experimental setup included triple Langmuir probes mounted on a moveable probe stand, to collect data over a wide range of locations and operating conditions. Triple probes have the ability to instantaneously measure electron temperature and density, and have the benefit of being relatively simple to use, compared to other methods used to measure these same properties. The implementation of this measuring technique is discussed in detail, to aid future work that utilizes these devices. Electron temperature and density was measured from up to 45 degrees from the centerline on planes parallel and perpendicular to the thruster electrodes, for thruster energy levels of 5, 20 and 40 J. Radial distances extend from 6 to 20 cm downstream from the Teflon surface. These locations cover the core of the PPT plume, over a range of energy levels that corresponds to proposed mission operating conditions. Data analysis shows the spatial and temporal variation of the plume. Maximum electron density near the exit of the thruster is 1.6 x 1020, 1.6 x 1021, and 1.8 x 1021 m-3 for the 5, 20 and 40 J discharges, respectively. At 20 cm downstream from the Teflon surface, densities are 1 x 1019, 1.5 x 1020 and 4.2 x 1020 for the 5, 20 and 40 J discharges, respectively. The average electron temperature at maximum density was found to vary between 3.75 and 4.0 eV for the above density measurements at the thruster exit, and 20 cm from the Teflon surface the temperatures are 0.5, 2.5, and 3 eV for the 5, 20 and 40 J discharges. Plume properties show a great degree of angular variation in the perpendicular plane and very little in the parallel plane, most likely due to the rectangular geometry of the PPT electrodes. Simultaneous electron temperature and density traces for a single thruster discharge show that the hottest electrons populate the leading edge of the plume. Analysis between pulses shows a 50% variation in density and a 25% variation in electron temperature. Error analysis estimates that maximum uncertainty in the temperature measurements to be approximately +/- 0.75 eV due to noise smoothing, and the maximum uncertainty in electron density to be +/- 60%, due to assumptions related to the triple probe theory. In addition, analysis of previously observed slow and fast ion components in the PPT plume was performed. The analysis shows that there is approximately a 3 us difference in creation time between the fast and slow ions, and that this correlates almost exactly with the half period of the oscillations in the thruster discharge current."

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.