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Over the last decade, miniature ion thruster development has remained an active area of research do to its low power, low thrust, and high efficiency, however, due to several technical issues; a flight level miniature ion thruster has proved elusive. This thesis covers the design, fabrication, assembly, and test of an altered version of the Miniature Xenon Ion thruster (MiXI), originally developed by lead engineer Dr. Richard Wirz, at the California Institute of Technology (Caltech). In collaboration with Dr. Wirz, MiXI-CP-V3 was developed at Cal Poly San Luis Obispo with the goal of implementing of a ¼ inch hollow cathode and 3mmx3mm plasma confinement magnets in order to improve the plasma confinement characteristics, reliability, and performance of the MiXI design. Operational testing revealed a mass utilization efficiency of 35-75% and a discharge loss of 550-1200 eV/ion over plasma discharge currents of 0.5-1.5A and propellant flow rates of 0.8-1.3 SCCM. Testing revealed that the MiXI thruster can be operated with a hollow cathode and observations and data gained from this study have led to a greater understanding of the operational parameters of the MiXI thruster, and will contribute to the development and advancement of the MiXI baseline design, with the goal of creating an efficient and reliable flight level miniature ion thruster.

In order to support UCLA’s development of the 3 cm Miniature Xenon Ion (MiXI) thruster, Cal Poly has a 3 cm thruster under development. This version, called MiXI Cal Poly Version 1 (MiXI-CPv1), is complete and has been utilized in vacuum chamber thermal validation testing. Testing on this version was used to check the validity of heat transfer simulations modeled in SolidWorks. Investigations of the 3 cm ion thruster configuration were intended to discover the driving factors affecting the thermal behavior of the discharge chamber and surrounding design space. Numerical simulations indicate that the heating of the samarium cobalt permanent magnets can be mitigated through the implementation of two proposed modifications. The first modification is to implement a 2% thoriated tungsten filament cathode. This design exhibited maximum permanent magnet temperatures of 325°C, twenty-five degrees below the maximum upper temperature of 350°C. Since some magnetic degaussing effects have been observed at temperatures above 300°C, the aforementioned solution can be combined with a thruster design modification to achieve a reduced permanent magnet temperature of 298°C. This modification would involve increase the anode wall thickness from approximately 0.7 mm to 2 mm below the permanent magnet ring, creating a stepped anode design. Additionally, less effective solutions were proposed and modeled and are presented for completeness.

This master's thesis was conducted in collaboration with ECAPS, where a conjugate heat transfer analysis on their High Performance Green Propulsion (HPGP) 22N thruster was done. ECAPS is a Swedish propulsion company specializing in green propulsion. They develop thrusters for spacecraft orbit and attitude control, utilizing the propellant LMP-103S. LMP-103S is a non-toxic propellant, in contrast to the hazardous monopropellant hydrazine commonly used in thrusters. A previous master's thesis modified the original design of the 22N thruster to make it compatible with additive manufacturing. Some concerns about heat transfer in the feed tube surfaced with the new design as it showed elevated temperatures. The feed tube is a component that works as a pathway where liquid propellant is transported from the flow control assembly to the reactor chamber assembly, where combustion begins. The goal of this master's thesis was to determine the temperatures the liquid propellant reached, and to assess if the liquid propellant was at risk of vaporization in the feed tube before reaching the reactor chamber assembly. Since the feed tube is a limited volume, vaporization of the liquid propellant in the feed tube could have devastating consequences of the structure. Ansys Fluent was used as the Computational Fluid Dynamics (CFD) software, along with the Computer Aided Design (CAD) software NX and Matlab for data handling.  Four extreme case scenarios were determined to be simulated, varying the liquid propellant inlet temperatures from highest to lowest operable temperatures, as well as the thruster's highest and lowest operable inlet pressures. A literature study on conjugate heat transfer in CFD was done, along with determination and calculations of necessary parameters for a correct simulation setup, and a grid independence study. Both steady-state and transient simulations were conducted. The results indicate that when the thruster operates with the highest inlet pressure, there is a risk of vaporization, but critical consequences are less likely to have time to develop. However, for the cases where the thruster operates with its lowest inlet pressure, a significant risk of vaporization in the feed tube is present. The simulated temperature results suggest that the liquid propellant will rapidly vaporize, where increased pressure at the feed tube outlet will be building up as a result of the expanding vapor, leading to a feed tube failure for the vapor to escape through. Therefore, the new design change of the feed tube will most likely not work for the thruster to be able to work under all necessary conditions. New modifications to the feed tube are necessary, or alternatively, the original design of the feed tube could be added afterward to the 3D-printed structure, though this may result in the loss of some benefits of manufacturing the entire structure in one piece.

No description available.

No description available.

Recent advances in plasma-based propulsion systems have led to the development of electromagnetic Radio-Frequency (RF) plasma generation and acceleration systems, called Helicon Plasma Thrusters (HPT). One of the pioneer companies developing this new type of space propulsion is T4i Technology for Propulsion and Innovation s.p.a., with its cutting-edge project called REGULUS, among which this study has been performed. A crucial part of HPT systems is the acceleration region, where, by the means of a magnetic nozzle, the thermal energy of the plasma is converted into axial acceleration and, in turn, into thrust. This study is focused on the numerically simulation of the plasma dynamics in the acceleration stage, using Xenon gas. A three-dimensional full Particle-In-Cell (PIC) simulation strategy is used to simulate the plume in the magnetic nozzle. The code developed for the plasma simulation is based on the open-source software Spacecraft Plasma Interaction Software (SPIS). The code has been conveniently modified and improved, neutrals and collision processes were added to evaluate their impact on the plasma properties. The features added improved the validity of the results, now one step closer to the physical reality. The code has been proven to be an extremely versatile and powerful tool for optimization and adaptation to different mission scenarios. / De senaste framstegen i plasmaframdrivning har lett till utvecklingen Helicon Plasma Thruster (HPT) som kombinerar elektromagnetisk högfrekvent (RF) plasmakälla och ett accelerationssystem. En av företagen som är pionjärer i att utveckla denna nya framdrivningsteknik är T4i Technology for Propulsion and Innovation s.p.a., med dess banbrytande projekt REGULUS, som detta arbete bygger på. En viktig del av HPT-systemet är accelerationsområde där plasmats termiska energin omvandlas till axiell accelleration i en magnetisk dysa. Denna rapport fokuserar på numeriska modelleringen av plasmadynamiken accelerationsområdet vid användning av Xenongasen. En tredimensionell Particle-In-Cell (PIC) simulering används för att studera plasmautflödet i magnetiska dysan. Koden bygger på den öppna mjukvaran Spacecraft Plasma interaction Software (SPIS). Koden har modifierats och förbättrats, en neutral komponent samt kollisionsprocesser har lagts till och deras påverkan på plasmabeteende har studerats. Dessa nya element förbättrade giltigheten av simulerings-resultaten. Nu ett steg närmre den fysiska verkligheten. Koden är ett mångsidigt och kraftfullt verktyg för optimering och anpassning till olika användningsscenarier.

Thesis (MScEng (Mechanical and Mechatronic Engineering))--University of Stellenbosch, 2009. / This thesis describes the development of a vacuum arc thruster (VAT) to be used as a potential low mass (< 500 g), low power (< 5–10W) propulsion system for nanosatellites. The thruster uses a high voltage capacitive circuit to initiate and power the arc process with a 400 ns high current (150–800A) pulse. A one-dimensional steady state analyticalmodel describing the cathode region of the vacuum arc was developed. The model made use of mass and energy balances at the sheath region and cathode surface respectively to predict key quantities such as thrust, ion velocity, ion-to-arc current ratio and erosion rate. Predicted results were shown to be within the limits of reported literature (∼63 μN/A, 26.12 km/s, 0.077 and 110 μg/C respectively). A sensitivity analysis of the analytical model found that a high electric field in the cathode region impedes and decelerates ion flow, which is used for thrust. This was confirmed experimentally for thrust values at arc voltages greater than 2000 V. Both direct and indirect means of measuring thrust were achieved by using a deflecting cantilever beam and an ion collector system, respectively. The transient response of the cantilever beam to impulsive thrust was analytically modeled, whilst the ion current was found by measuring the current induced on a plate subject to ion bombardment. Knowledge of the ion current density distribution was successfully used to approximate the effective normal thrust vector. Direct and indirect thrust levels were roughly 140 and 82 μN/A of average arc current, respectively. Measured thrust was found to be higher than predicted thrust due to thrust contributions fromthe ablation of Teflon insulation. The discrepancy is also due to the uncertainty in quantifying free parameters in the analytical model such as the fraction of generated ions flowing away from the cathode region. The thrust-topower ratio, specific impulse and efficiency of the vacuum arc thruster at an average arc current of 200 A was measured to be 0.6 μN/W, 160 s and 0.05 %, respectively. A thruster performance analysis and specification showed that the VAT is capable of achieving specific orbital and slew manoeuvres within a constant 5–10 W average power. It was concluded that thruster performance could be improved by using a two-stage arc circuit consisting of a high voltage, low current, short pulse trigger and a low voltage, high current, long pulse driver.

A three-rotational degree of freedom attitude control system has been developed for the NASA Jet Propulsion Laboratory’s INSPIRE Project by the Texas Spacecraft Laboratory at The University of Texas at Austin. Using 3D plastic printing manufacturing techniques, a cold gas thruster system was created in order to detumble and maintain the attitude of two 3U CubeSats traveling through interplanetary space. A total of four thruster units were produced, including two engineering designs and two flight units. The units feature embedded sensors and millisecond level thrust control while using an inert, commercially-available refrigerant as a propellant. The thrust, minimum impulse bit, and specific impulse performance of the cold gas units was characterized using a ballistic pendulum test stand within a microtorr vacuum chamber. A heating element was used to change the temperature conditions of the propellant and determine the relationship between temperature and performance. The flight units were delivered in January of 2014 and the INSPIRE satellites are expected to launch in the upcoming year. / text

The present thesis is the result of a research period at the Institute of Space and Astronautical Science of the Japanese Aerospace Exploration Agency, ISAS/JAXA within Funaki Laboratory of the Department of Space Flight Systems that followed the path of plume plasma diagnostics for space electric propulsion drives. During the experimental studies two high-current hollow cathodes and an innovative prototype of a low-voltage fully external discharge plasma thruster (XPT) had their plasma plumes diagnosed using electrostatic probes and retarding potential analyser (RPA). A Hall thruster and hollow cathode plume is defined as an unmagnetised quasi-neutral plasma which is mainly formed of neutral particles, electrons, singly and doubly charged ions. Plasma diagnostic techniques provide information through practical observations in order to fully understand the dynamics of the aforementioned plume components, the physical processes taking place within the plume and their effects on the spacecraft, for instance. Mastering these aspects of the plasma plume of space electric propulsion drives bolster the design processes, leading to highly efficient devices. Firstly, the introduction provides insights on the fundamental principles of hollow cathodes and Hall thrusters and a brief presentation of the plasma diagnostic techniques used during the research: single and double Langmuir probes, emissive probes and retarding potential analyser. Then, the fundamental plume diagnostics principles are depicted in an exhaustive way, departing from classical plasma kinetic theory, energy distribution functions and ending with an overview on the theory of charge collection by cylindrical probes. Subsequently, peculiarities of various analysis techniques are exposed for the Langmuir probes, emissive probes and RPA, with an emphasis on their strengths and demerits. The experimental setups for the cathodes and XPT plume diagnostic procedures are then outlined. The experimental logic, setup and electrical diagrams as well as a presentation of each probe design and manufacturing details are extensively discussed. The hollow cathodes experimental results are exposed with a discourse that aims of overviewing the difference between the various data analysis methods applied for the raw data. A discussion ensued based on the results in order to effectively identify mechanisms that produced the observed plasma parameters distributions. For the first time, the plume of a fully external discharge plasma thruster was diagnosed utilising double Langmuir probes.  The thesis highlights the main results obtained for the XPT far-field plume plasma diagnostics. The experimental findings for both thruster centreline positions and 2D plume maps for several axial distances away from the anode plate offer a ground basis for future measurements, a comparison term and a database to support ongoing computational codes. The results are discussed and related to the thruster performances data obtained during previous experiments. The thesis includes consistency analyses between the experimental data and the numerical simulation results and the uncertainties in measured plasma parameters associated with each data analysis procedure are evaluated for each data set. Last, the conclusions underline the main aspects of the research and further work on the previously mentioned plasma diagnostic techniques for hollow cathodes and XPT is suggested. / La présente thèse est le résultat d'une période de recherche à l'Institut des Sciences Spatiales et Astronautiques de l'Agence Spatiale Japonaise, ISAS / JAXA qui a suivi la voie des diagnostics du plasma de la plume de propulseurs électriques spatiaux. Au cours des études expérimentales, deux cathodes creuses à fort courant et un prototype innovant d'un propulseur basse tension à décharge externe de plasma (XPT) avaient leurs faisceaux de plasma diagnostiqués en utilisant des sondes électrostatiques et un analyseur à potentiel retardé. La plume d’un propulseur à effet Hall et d’une cathode creuse est définie comme un plasma quasi-neutre non-magnétisé qui est principalement formé de particules neutres, d’électrons, d’ions monovalents et bivalents. Les techniques de diagnostic du plasma fournissent des informations, via des observations pratiques, afin de bien comprendre la dynamique des composants de la plume mentionnés ci-dessus, les processus physiques qui se déroulent dans la plume et leurs effets sur une sonde spatiale, par exemple. La maîtrise de ces aspects du plasma de la plume généré par les propulseurs électriques spatiaux renforce les processus de conception de ce type de propulsion, ce qui conduit à des dispositifs hautement efficaces. Tout d'abord, l'introduction donne un aperçu sur les principes fondamentaux de cathodes creuses et de propulseurs à effet Hall, et une brève présentation des techniques de diagnostic du plasma utilisées lors de la recherche : sondes de Langmuir simples et doubles, des sondes émissives et d’analyseur à potentiel retardé. Ensuite, les principes fondamentaux de diagnostic de la plume sont représentés de manière exhaustive, d’abord la théorie cinétique classique du plasma, les fonctions de distribution en énergie et pour terminer une vue d'ensemble de la théorie de la collecte de charge par des sondes cylindriques. Par la suite, les particularités des diverses techniques d'analyse sont exposées pour les sondes de Langmuir, les sondes émissives et RPA, en mettant l'accent sur leurs avantages et leurs inconvénients. Les montages expérimentaux pour les procédures de diagnostic de la plume-plasma de cathodes et du XPT sont ensuite décrits. La logique expérimentale, les schémas électriques ainsi qu'une présentation de la conception et de la fabrication de chaque sonde sont largement discutés. Les résultats expérimentaux pour les cathodes creuses sont exposés de façon à présenter la différence entre plusieurs méthodes d'analyse de données appliquées aux données brutes. Une discussion s’ensuit, basée sur les résultats afin d'identifier efficacement les mécanismes qui ont produits les propriétés électroniques observées. Pour la première fois, la plume d'un propulseur à décharge externe de plasma a été diagnostiquée en utilisant des sondes de Langmuir doubles. La thèse met en évidence les principaux résultats obtenus pour le diagnostic en champ lointain de la plume-plasma du XPT. Les résultats expérimentaux pour les positions sur l'axe du propulseur et le cartes 2D de la plume pour plusieurs distances axiales loin de l’anode offrent une base pour de futures mesures, un terme de comparaison et une base de données pour appuyer les codes numériques. Les résultats sont discutés et sont rapportés aux données de performances du propulseur obtenus lors des essais précédents. La thèse comprend des analyses de la cohérence entre les données expérimentales et les résultats de simulation numérique, et les incertitudes des paramètres mesurés du plasma associées à chaque procédure d'analyse des données sont évaluées pour chaque ensemble de données. Enfin, les conclusions soulignent les principaux aspects de la recherche et une poursuite des travaux sur les techniques de diagnostic de plasma pour les cathodes creuses et le XPT est suggérée.