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.

"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."

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>

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.

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.

<p>Prior to the mid-20th century, the idea of electric propulsion had been all but a foreign one that manifested itself in the topic of science fiction. It was around this time when companies and agencies like NASA began to take interest in the topic of space propulsion, as most famously seen in the landing of the Apollo 11 mission on the moon. It was not until the early-1960s where the idea of a pulsed plasma thruster was first realized, with its first test being in 1964 aboard the Russian Zond-2 satellite which contained 6 ablative Polytetrafluoroethylene (PTFE, or “Teflon”) pulsed plasma thrusters.</p> <p>In this paper, a new low-voltage, liquid-fed pulsed plasma thruster was developed, tested, and characterized. This project took influence from the previous low-voltage, liquid-fed pulsed plasma thruster in Purdue’s EPPL and desired to transition it from a current gas-fed system to its intended liquid-fed system. The two main objectives for this project included conducting direct studies of the cathode’s erosion rate using a simple weighing method after simulating a lifetime of discharging the thruster, and completing the initial design of the liquid-fed pulsed plasma thruster using AF-M315E as its propellant while gathering data on its required breakdown voltage, exhaust velocity, and specific impulse.</p> <p>Both objectives were successfully completed, with the following parameters being measured or calculated. The required breakdown voltage was seen to be less than 26kV to keep the ignition spark inside the chamber. For the subsequent results measured however, the breakdown voltage was kept between 10-16kV for all successive tests. The peak current measured for all discharges was an average of 11kA, far exceeding similar geometries such as MPD thrusters. The operational voltage was less than 200V, although an operational voltage closer to 100V is expected after further optimization of the system is completed. The erosion rate of the tungsten cathode at this operational setting was found to be 15.4046 +/- 0.592 microgram/Coulomb which is much less than the cathode spot erosion rate reported for tungsten in literature of about 60 microgram/Coulomb and is beneficial for extending system lifetime. The exhaust velocity was calculated to be 30.6 +/- 4.8km/s which is typical of state-of-the-art PPT electric propulsion devices. The specific impulse was also extrapolated from the ion’s exhaust velocity, calculating to be 3,119 +/- 489 seconds. Future work would require optimization of the propellant injection mechanism to minimize propellant loss.</p>

"The ablative Teflon pulsed plasma thruster (PPT) is an onboard electromagnetic propulsion enabling technology for small spacecraft missions. The integration of PPTs onboard spacecraft requires the understanding and evaluation of possible thruster/spacecraft interactions. To aid in this effort the work presented in this thesis is directed towards the development and application of Langmuir probe techniques for use in the plume of PPTs. Double and triple Langmuir probes were developed and used to measure electron temperature and density of the PPT plume. The PPT used in this thesis was a laboratory model parallel plate ablative Teflon® PPT similar in size to the Earth Observing (EO-1) PPT operating in discharge energies between 5 and 40 Joules. The triple Langmuir probe was operated in the current-mode technique that requires biasing all three electrodes and measuring the resulting probe currents. This new implementation differs from the traditional voltage-mode technique that keeps one probe floating and requires a voltage measurement that is often susceptible to noise in the fluctuating PPT plume environment. The triple Langmuir probe theory developed in this work incorporates Laframboise’s current collection model for Debye length to probe radius ratios less than 100 in order to account for sheath expansion effects on ion collection, and incorporates the thin-sheath current collection model for Debye length to probe radius ratios greater than 100. Error analysis of the non-linear system of current collection equations that describe the operation of the current-mode triple Langmuir probe is performed as well. Measurements were taken at three radial locations, 5, 10, and 15 cm from the Teflon® surface of the PPT and at angles of 20 and 40 degrees to either side of the thruster centerline as well as at the centerline. These measurements were taken on two orthogonal planes, parallel and perpendicular to the PPT electrodes. A data-processing software was developed and implements the current-mode triple Langmuir probe theory and associated error analysis. Results show the time evolution of the electron temperature and density. Characteristic to all the data is the presence of hot electrons of approximately 5 to 10 eV at the beginning of the pulse, occurring near the peak of the discharge current. The electron temperature quickly drops off from its peak values to 1-2 eV for the remainder of the pulse. Peak electron densities occur after the peak temperatures. The maximum electron density values on the centerline of the plume of a laboratory PPT 10 cm from the Teflon® surface are 6.6x10^19 +/- 1.3x10^19 m^-3 for the 5 J PPT, 7.2x10^20 +/- 1.4x10^20 m^-3 for the 20 J PPT, and 1.2x10^21 +/- 2.7x10^20 m^-3 for the 40 J PPT. Results from the double Langmuir probe taken at r=10 cm, theta perpendicular=70 degrees and 90 degrees of a laboratory PPT showed good agreement with the triple probe method."

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