Permanent Magnet Hall Thruster for satellite orbit raising with low power energy consumption

Electric propulsion is now a succeful method for primary propulsion of deep space long duration missions and for geosyncronous satellite attitude control. Closed Drift Thruster, so called Hall Thruster or SPT (Stationary Plasma Thruster), was primarily conceived in USSR (the ancient Soviet Union) and, since then, it has been developed by space agencies, space research institutes and industries in several countries such as France, USA, Israel, Russian Federation and Brazil. In this work we present the main features of the Permanent Magnet Hall Thruster (PMHT) developed at the Plasma Laboratory of the University of Brasilia. The idea of using an array of permanent magnets, instead of an electromagnet, to produce a radial magnetic field inside the plasma channel of the thruster is very significant. It allows the development of a Hall Thruster with power consumption low enough to be used in small and medium size satellites. Description of a new vacuum chamber used to test the second prototype of the PMHT (PHALL II) will be given. PHALL II has an aluminum plasma chamber and is smaller with 15 cm diameter and will contain rare earth magnets. We will show plasma density and temperature space profiles inside and outside the thruster channel. Ion temperature measurements based on Doppler broadening of spectral lines and ion energy measurements are also shown. Based on the measured plasma parameters we constructed an aptitude figure of the PMHT. It contains the specific impulse, total thrust, propellant flow rate and power consumption necessary for orbit raising of satellites. Based on previous studies of geosyncronous satellite orbit positioning we perform numerical simulations of satellite orbit raising from an altitude of 700 km to 36000 km using a PMHT operating in the 100 mN - 500 mN thrust range. In order to perform these calculations integration techniques were used. The main simulation paraters were orbit raising time, fuel mass, total satellite mass, thrust and exaust velocity. We conclude comparing our results with results obtainned with known space missions performed with Hall Thrusters. © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

Pulsed plasma thruster ignition system: investigation, test design and results

Fra i sistemi di propulsione elettrica per satelliti, il Pulsed Plasma Thruster, PPT, è quello dal design più semplice. È anche il primo sistema di propulsione elettrica utilizzato in un satellite artificiale, ossia ZOND-2 lanciato nel 1964 dall’Unione Sovietica. Tuttavia, dopo circa 50 anni di ricerca, la comprensione teorica e sperimentale di questo dispositivo rimane limitata. Questo elaborato di tesi magistrale indaga sul sottosistema di accensione del PPT, cercando di mettere in luce alcuni aspetti legati al lifetime della spark plug, SP. Tale SP, o candela, è l’attuatore del sottosistema di accensione. Questa produce una scintilla sulla sua superficie, la quale permette la realizzazione della scarica elettrica principale fra i due elettrodi del motore. Questa scarica crea una sottile parete di plasma che, per mezzo della forza elettromagnetica di Lorentz, produce la spinta del PPT. Poiché la SP si trova all’interno del catodo del motore e si affaccia nella camera di scarica, questa soffre di fenomeni di corrosione e di deposizione carbonacea proveniente dal propellente. Questi fenomeni possono limitare notevolmente il lifetime della SP. I parametri connessi alla vita operativa della SP sono numerosi. In questo elaborato si è analizzata la possibilità di utilizzare una elettronica di accensione della candela alternativa alla classica soluzione che utilizza un trasformatore. Il sottosistema di accensione classico e quello nuovo sono stati realizzati e testati, per metterne in luce le differenze ed i possibili vantaggi/svantaggi.

Hall-effect thruster-cathode coupling : the effect of cathode position and magnetic field topology

Hall-effect thruster (HET) cathodes are responsible for the generation of the free electrons necessary to initiate and sustain the main plasma discharge and to neutralize the ion beam. The position of the cathode relative to the thruster strongly affects the efficiency of thrust generation. However, the mechanisms by which the position affects the efficiency are not well understood. This dissertation explores the effect of cathode position on HET efficiency. Magnetic field topology is shown to play an important role in the coupling between the cathode plasma and the main discharge plasma. The position of the cathode within the magnetic field affects the ion beam and the plasma properties of the near-field plume, which explains the changes in efficiency of the thruster. Several experiments were conducted which explored the changes of efficiency arising from changes in cathode coupling. In each experiment, the thrust, discharge current, and cathode coupling voltage were monitored while changes in the independent variables of cathode position, cathode mass flow and magnetic field topology were made. From the telemetry data, the efficiency of the HET thrust generation was calculated. Furthermore, several ion beam and plasma properties were measured including ion energy distribution, beam current density profile, near-field plasma potential, electron temperature, and electron density. The ion beam data show how the independent variables affected the quality of ion beam and therefore the efficiency of thrust generation. The measurements of near-field plasma properties partially explain how the changes in ion beam quality arise. The results of the experiments show that cathode position, mass flow, and field topology affect several aspects of the HET operation, especially beam divergence and voltage utilization efficiencies. Furthermore, the experiments show that magnetic field topology is important in the cathode coupling process. In particular, the magnetic field separatrix plays a critical role in impeding the coupling between cathode and HET. Suggested changes to HET thruster designs are provided including ways to improve the position of the separatrix to accommodate the cathode.

Effect of segmented anodes on the performance and plume of a hall thruster

Development of alternative propellants for Hall thruster operation is an active area of research. Xenon is the current propellant of choice for Hall thrusters, but can be costly in large thrusters and for extended test periods. Condensible propellants may offer an alternative to xenon, as they will not require costly active pumping to remove from a test facility, and may be less expensive to purchase. A method has been developed which uses segmented electrodes in the discharge channel of a Hall thruster to divert discharge current to and from the main anode and thus control the anode temperature. By placing a propellant reservoir in the anode, the evaporation rate, and hence, mass flow of propellant can be controlled. Segmented electrodes for thermal control of a Hall thruster represent a unique strategy of thruster design, and thus the performance of the thruster must be measured to determine the effect the electrodes have on the thruster. Furthermore, the source of any changes in thruster performance due to the adjustment of discharge current between the shims and the main anode must be characterized. A Hall thruster was designed and constructed with segmented electrodes. It was then tested at anode voltages between 300 and 400 V and mass flows between 4 and 6 mg/s, as well as 100%, 75%, 50%, 25%, and <5% of the discharge current on the shim electrodes. The level of current on the shims was adjusted by changing the shim voltage. At each operating point, the thruster performance, plume divergence, ion energy, and multiply charged ion fraction were measured performance exhibited a small change with the level of discharge current on the shim electrodes. Thrust and specific impulse increased by as much as 6% and 7.7%, respectively, as discharge current was shifted from the main anode to the shims at constant anode voltage. Thruster efficiency did not change. Plume divergence was reduced by approximately 4 degrees of half-angle at high levels of current on the shims and at all combinations of mass flow and anode voltage. The fraction of singly charged xenon in the thruster plume varied between approximately 80% and 95% as the anode voltage and mass flow were changed, but did not show a significant change with shim current. Doubly and triply charged xenon made up the remainder of the ions detected. Ion energy exhibited a mixed behavior. The highest voltage present in the thruster largely dictated the most probable energy; either shim or anode voltage, depending on which was higher. The overall change in most probable ion energy was 20-30 eV, the majority of which took place while the shim voltage was higher than the anode voltage. The thrust, specific impulse, plume divergence, and ion energy all indicate that the thruster is capable of a higher performance output at high levels of discharge current on the shims. The lack of a change in efficiency and fraction of multiply charged ions indicate that the thruster can be operated at any level of current on the shims without detrimental effect, and thus a condensible propellant thruster can control the anode temperature without a decrease in efficiency or a change in the multiply charged ion fraction.

Development of a direct evaporation bismuth Hall thruster

Hall thrusters have been under active development around the world since the 1960’s. Thrusters using traditional propellants such as xenon have been flown on a variety of satellite orbit raising and maintenance missions with an excellent record. To expand the mission envelope, it is necessary to lower the specific impulse of the thrusters but xenon and krypton are poor performers at specific impulses below 1,200 seconds. To enhance low specific impulse performance, this dissertation examines the development of a Hall-effect thruster which uses bismuth as a propellant. Bismuth, the heaviest non-radioactive element, holds many advantages over noble gas propellants from an energetics as well as a practical economic standpoint. Low ionization energy, large electron-impact crosssection and high atomic mass make bismuth ideal for low-specific impulse applications. The primary disadvantage lies in the high temperatures which are required to generate the bismuth vapors. Previous efforts carried out in the Soviet Union relied upon the complete bismuth vaporization and gas phase delivery to the anode. While this proved successful, the power required to vaporize and maintain gas phase throughout the mass flow system quickly removed many of the efficiency gains expected from using bismuth. To solve these problems, a unique method of delivering liquid bismuth to the anode has been developed. Bismuth is contained within a hollow anode reservoir that is capped by a porous metallic disc. By utilizing the inherent waste heat generated in a Hall thruster, liquid bismuth is evaporated and the vapors pass through the porous disc into the discharge chamber. Due to the high temperatures and material compatibility requirements, the anode was fabricated out of pure molybdenum. The porous vaporizer was not available commercially so a method of creating a refractory porous plate with 40-50% open porosity was developed. Molybdenum also does not respond well to most forms of welding so a diffusion bonding process was also developed to join the molybdenum porous disc to the molybdenum anode. Operation of the direct evaporation bismuth Hall thruster revealed interesting phenomenon. By utilizing constant current mode on a discharge power supply, the discharge voltage settles out to a stable operating point which is a function of discharge current, anode face area and average pore size on the vaporizer. Oscillations with a 40 second period were also observed. Preliminary performance data suggests that the direct evaporation bismuth Hall thruster performs similar to xenon and krypton Hall thrusters. Plume interrogation with a Retarding Potential Analyzer confirmed that bismuth ions were being efficiently accelerated while Faraday probe data gave a view of the ion density in the exhausted plume.

EVALUATION OF MAGNESIUM AS A HALL THRUSTER PROPELLANT

In this study, the use of magnesium as a Hall thruster propellant was evaluated. A xenon Hall thruster was modified such that magnesium propellant could be loaded into the anode and use waste heat from the thruster discharge to drive the propellant vaporization. A control scheme was developed, which allowed for precise control of the mass flow rate while still using plasma heating as the main mechanism for evaporation. The thruster anode, which also served as the propellant reservoir, was designed such that the open area was too low for sufficient vapor flow at normal operating temperatures (i.e. plasma heating alone). The remaining heat needed to achieve enough vapor flow to sustain thruster discharge came from a counter-wound resistive heater located behind the anode. The control system has the ability to arrest thermal runaway in a direct evaporation feed system and stabilize the discharge current during voltage-limited operation. A proportional-integral-derivative control algorithm was implemented to enable automated operation of the mass flow control system using the discharge current as the measured variable and the anode heater current as the controlled parameter. Steady-state operation at constant voltage with discharge current excursions less than 0.35 A was demonstrated for 70 min. Using this long-duration method, stable operation was achieved with heater powers as low as 6% of the total discharge power. Using the thermal mass flow control system the thruster operated stably enough and long enough that performance measurements could be obtained and compared to the performance of the thruster using xenon propellant. It was found that when operated with magnesium, the thruster has thrust ranging from 34 mN at 200 V to 39 mN at 300 V with 1.7 mg/s of propellant. It was found to have 27 mN of thrust at 300 V using 1.0 mg/s of propellant. The thrust-to-power ratio ranged from 24 mN/kW at 200 V to 18 mN/kW at 300 volts. The specific impulse was 2000 s at 200 V and upwards of 2700 s at 300 V. The anode efficiency was found to be ~23% using magnesium, which is substantially lower than the 40% anode efficiency of xenon at approximately equivalent molar flow rates. Measurements in the plasma plume of the thruster—operated using magnesium and xenon propellants—were obtained using a Faraday probe to measure off-axis current distribution, a retarding potential analyzer to measure ion energy, and a double Langmuir probe to measure plasma density, electron temperature, and plasma potential. Additionally, the off axis current distributions and ion energy distributions were compared to measurements made in krypton and bismuth plasmas obtained in previous studies of the same thruster. Comparisons showed that magnesium had the largest beam divergence of the four propellants while the others had similar divergence. The comparisons also showed that magnesium and krypton both had very low voltage utilization compared to xenon and bismuth. It is likely that the differences in plume structure are due to the atomic differences between the propellants; the ionization mean free path goes down with increasing atomic mass. Magnesium and krypton have long ionization mean free paths and therefore require physically larger thruster dimensions for efficient thruster operation and would benefit from magnetic shielding.

A technical report on analytical-experimental correlation of radiation loss from an argon arcjet.

At head of title: Dept. of the Air Force. Office of Aerospace Research. Aeronautical Research Laboratory. Thermomechanics Branch. Contract AF 33(657)-9962.

Growth of ultrananocrystalline diamond film by DC Arcjet plasma enhanced chemical vapor deposition

Self-standing diamond films were grown by DC Arcjet plasma enhanced chemical vapor deposition (CVD). The feed gasses were Ar/H 2/CH 4, in which the flow ratio of CH 4 to H 2 (FCH4/FH2) was varied from 5% to 20%. Two distinct morphologies were observed by scanning electron microscope (SEM), i.e. the pineapple-like morphology and the cauliflower-like morphology. It was found that the morphologies of the as-grown films are strongly dependent on the flow ratio of CH 4 to H 2 in the feed gasses. High resolution transmission electron microscope (HRTEM) survey results revealed that there were nanocrystalline grains within the pineapple-like films whilst there were ultrananocrystalline grains within cauliflower-like films. X-ray diffraction (XRD) results suggested that (110) crystalline plane was the dominant surface in the cauliflower-like films whilst (100) crystalline plane was the dominant surface in the pineapple-like films. Raman spectroscopy revealed that nanostructured carbon features could be observed in both types of films. Plasma diagnosis was carried out in order to understand the morphology dependent growth mechanism. It could be concluded that the film morphology was strongly influenced by the density of gas phases. The gradient of C2 radical was found to be different along the growth direction under the different growth conditions. © 2012 Elsevier B.V. All rights reserved.

A Strategy to Characterize the LISA-Pathfinder Cold Gas Thruster System

The cold gas micro-propulsion system that will be used during the LISA-Pathfinder mission will be one of the most important component used to ensure the "free-fall" of the enclosed test masses. In this paper we present a possible strategy to characterize the effective direction and amplitude gain of each of the 6 thrusters of this system.

Development of numerical tools for the simulation, design and optimization of a Helicon Plasma Thruster

This work thesis focuses on the Helicon Plasma Thruster (HPT) as a candidate for generating thrust for small satellites and CubeSats. Two main topics are addressed: the development of a Global Model (GM) and a 3D self-consistent numerical tool. The GM is suitable for preliminary analysis of HPTs with noble gases such as argon, neon, krypton, and xenon, and alternative propellants such as air and iodine. A lumping methodology is developed to reduce the computational cost when modelling the excited species in the plasma chemistry. A 3D self-consistent numerical tool is also developed that can treat discharges with a generic 3D geometry and model the actual plasma-antenna coupling. The tool consists of two main modules, an EM module and a FLUID module, which run iteratively until a steady state solution is converged. A third module is available for solving the plume with a simplified semi-analytical approach, a PIC code, or directly by integration of the fluid equations. Results obtained from both the numerical tools are benchmarked against experimental measures of HPTs or Helicon reactors, obtaining very good qualitative agreement with the experimental trend for what concerns the GM, and an excellent agreement of the physical trends predicted against the measured data for the 3D numerical strategy.

ROAZ Autonomous Surface Vehicle Design and Implementation

The design of an Autonomous Surface Vehicle for operation in river and estuarine scenarios is presented. Multiple operations with autonomous underwater vehicles and support to AUV missions are one of the main design goals in the ROAZ system. The mechanical design issues are discussed. Hardware, software and implementation status are described along with the control and navigation system architecture. Some preliminary test results concerning a custom developed thruster are presented along with hydrodynamic drag calculations by the use of computer fluid dynamic methods.

A New ROV Design: Issues on Low Drag and Mechanical Symmetry

This paper reports the design of a new remotely operated underwater vehicle (ROV), which has been developed at the Underwater Systems and Technology Laboratory (USTL) - University of Porto. This design is contextualized on the KOS project (Kits for underwater operations). The main issues addressed here concern directional drag minimization, symmetry, optimized thruster positioning, stability and layout of ROV components. This design is aimed at optimizing ROV performance for a set of different operational scenarios. This is achieved through modular configurations which are optimized for each different scenario.

Azimuth thrusterin rungon lujuustekninen tarkastelu

Työssä perehdyttiin azimuth potkurilaitteen rakenteelliseen kestävyyteen vaikuttaviin seikkoihin. Työn tavoitteena oli tutkia eri rakenneratkaisuja ja yksityiskohtia sekä niiden vaikutusta koko potkurilaitteen toimintaan. Potkurilaitteiden lujuustekninen tarkastelu sisälsi rakenteen eheyden tarkistuksen mekaanisten kuormitusten alla kuin myös värähtelykäyttäytymisen tutkimisen. Tarkastelun pääsääntöinen kohde oli potkurilaitteen alarunko, eli laivan ulkopuolelle jäävä osa. Diplomityö oli osa kahden uuden potkurilaitteen kehitysprojektia. Projektissa esiin tulleita rakennemuutoksia vertailtiin vanhoihin, jo toteutettuihin rakenneyksityiskohtiin. Lisäksi tutkittiin olemassa olevien rakenneratkaisujen hyödyntämistä uusissa laitteissa. Huomiota kiinnitettiin myös hitsien muotoiluun ja näiden mitoituskäytäntöihin. Rungon ehyyttä koskevat laskelman ovat hyvin pitkälti suoritettu elementtimenetelmää apuna käyttäen. Elementtimenetelmää varten luotiin useita malleja kattaen kokonaisia potkurilaittee runkoja kuin myös sen yksittäisiä osia. Elementtimenetelmän avulla saatuja tuloksia vertailtiin käytettävissä oleviin mittaustuloksiin. Työssä havaittiin potkurilaitteen rungon mitoittavaksi tekijäksi staattisessa tilanteessa muodostuvan voimansiirtolinja ja siinä sijaitsevien hammaspyörien siirtymät. Tämän seurauksena rungon taipumat ovat määräävämpiä geometrian kannalta kuin materiaalien sallitut jännitykset.

Potkurilaitteen ohjausputken laserpinnoitus

Laserpinnoitus on additiivinen prosessi, jossa lasertehon avulla sulatetaan pinnoitusainetta pinnoitettavan kappaleen pintaan. Laserpinnoituksella on mahdollista aikaansaada tiivis sekä kova, kulutusta ja korroosiota kestävä pinnoite kappaleen pintaan. Tässä työssä kohteena oli potkurilaitteen ohjausputken laserpinnoitus. Kyseisessä kohteessa pinnoitteelta vaaditaan hyvää korroosion- ja kulumisen kestoa. Työn tavoitteena oli lisätä tietoisuutta laserpinnoitusprosessista ja siinä vaikuttavista tekijöistä. Myös materiaaliominaisuuksien vaikutuksia sekä pinnoitus- että pinnoitettavien aineiden osalta selvitettiin. Työhön sisältyi myös kokeellinen osuus, jossa tavoitteena oli löytää kustannustehokkain pinnoiteaine tutkittuun kohteeseen. Kokeissa ohjausputki- sekä levyaihioon tehtiin koepinnoituksia. Kokeiden ja edelleen materiaalitutkimuksien avulla tutkittiin neljää eri pinnoitusainetta, jotka olivat Stellitti 21, Inconel 625, AISI 316L ja AISI 431. Kovuus- ja taivutuskokeiden perusteella kyseisistä aineista vain Stelliitti 21 täytti ohjausputken pinnoitteelle asetetut vaatimukset. Näin ollen Stelliitti 21 valikoitui kustannustehokkaimmaksi pinnoitusaineeksi ohjausputkien laserpinnoitukseen.