20 - Electric propulsion systems

The Microsatellite Electric Propulsion System (MEPS) program has been originated by the increasing need to provide a low cost and low power Electric Propulsion System (EPS) for small satellites (<300Kg). The propulsion system of MEPS consists of three main parts, the Thruster Units (TUs) composed of a Hall Effect Thruster (Rafael's CAM-200 or SITAEL's HT100) and a Cathode (Rafael's Heaterless Hollow Cathode RHHC), the Propellant Management and Tank Assembly (PMA/PTA) which performs the regulation of xenon mass flow rate from the tank to the TU inlet, and the Power Processing Unit (PPU) which provides the necessary power to drive the TUs and the PMA/PTA and implements the system control logic. The objective of this paper is to present the development status of MEPS PPU highlighting the design optimizations and simplifications that have been performed. Particular emphasis is aimed at the most indicative improvements and how all modifications contributed to the PPU budget in terms of number of components, mass, size and cost, without affecting system efficiency, performance and reliability. Aggregated results for PPU budget are thoroughly illustrated, by taking into account the overall optimizations and simplifications. The results are very promising and point to a reduction on overall number of components, mass, area and cost in the range of 15% - 33%, all very significant especially in space market.

The application of fully electric propulsion satellites will increase the risk of low-pressure discharge in electric propulsion devices, especially ion thrusters, whose unexpected breakdown will directly affect the reliability of satellite applications. Existing on-orbit cases show that the unexpected breakdown occurred in the power processing unit (PPU) of the ion thruster during orbit change, which led to the failure of thruster. Therefore, this paper analyses the high-voltage components inside the PPU, and systematically discusses the PPU and its working environment, the control and protection technology of unexpected electrical breakdown of the electric propulsion system and its aerospace engineering. In this paper, we discuss the electric field intensity and discharge margin of components in the PPU under different conditions, and further propose a method for insulation evaluation based on partial discharge (PD) detection. The research results show that the discharge margin in the PPU is close to the lowest discharge voltage level of uniform electric field, so reasonable insulation protection measures should be taken against discharge under high voltage condition. Furthermore, we propose a feasibility analysis method based on PD insulation defect detection, which provides a basis for quantitative analysis and evaluation of component insulation design in PPU systems.

The objective of this paper is to present the design of the Power Processing Unit (PPU) of the Microsatellite Electric Propulsion System (MEPS) program currently being developed at Elegant Bread Board (EBB). High efficiency, small size and weight and high reliability are the main parameters that special focus is given. The propulsion system of MEPS consists of two small size and low power (up to 300W) Thruster Units (TUs) each one composed of a Hall Effect Thruster and a Cathode, a single Propellant Management and Tank Assembly (PMA/PTA) and a single PPU. The PPU designed for the EBB phase has no redundancy, is fully flexible and capable of driving two different TU types (Rafael’s CAM-200 coupled with Rafael’s Heaterless Hollow Cathode (RHHC) and SITAEL’s HT100 coupled with SITAEL’s Heated Hollow Cathode (HC1)). This paper illustrates the core design requirements and important aspects on which the design was based. Preliminary results obtained from the coupling tests on the critical parts developed at Bread Board (BB), which contributed to the current design are also briefly depicted. Last but not least, the future development steps of the program that can become the driving factor for the successful implementation of an easily adjustable PPU compatible with a wide range of low power Electric Propulsion Systems (EPSs) are demonstrated.

Power processing units (PPUs) in an electric propulsion system provide many challenging integration issues. The PPU must provide power to the electric thruster while maintaining compatibility with all of the spacecraft power and data systems. Inefficiencies in the power processor produce heat, which must be radiated to the environment in order to ensure reliable operation. Although PPU efficiencies are generally greater than 0.9, heat loads are often substantial. This heat must be rejected by thermal control systems which generally have specific masses of 15-30 kg/kW. PPUs also represent a large fraction of the electric propulsion system dry mass. Simplification or elimination of power processing in a propulsion system would reduce the electric propulsion system specific mass and improve the overall reliability and performance. A direct drive system would eliminate all or some of the power supplies required to operate a thruster by directly connecting the various thruster loads to the solar array. The development of concentrator solar arrays has enabled power bus voltages in excess of 300 V which is high enough for direct drive applications for Hall thrusters such as the Stationary Plasma Thruster (SPT). The option of solar array direct drive for SPTs is explored to provide a comparison between conventional and direct drive system mass.

The primary source of electric propulsion development throughout NASA is managed by the In-Space Propulsion Technology Project at the NASA Glenn Research Center for the Science Mission Directorate. The objective of the Electric Propulsion project area is to develop near-term electric propulsion technology to enhance or enable science missions while minimizing risk and cost to the end user. Major hardware tasks include developing NASA's Evolutionary Xenon Thruster (NEXT), developing a long-life High Voltage Hall Accelerator (HIVHAC), developing an advanced feed system, and developing cross-platform components. The objective of the NEXT task is to advance next generation ion propulsion technology readiness. The baseline NEXT system consists of a high-performance, 7-kW ion thruster; a high-efficiency, 7-kW power processor unit (PPU); a highly flexible advanced xenon propellant management system (PMS); a lightweight engine gimbal; and key elements of a digital control interface unit (DCIU) including software algorithms. This design approach was selected to provide future NASA science missions with the greatest value in mission performance benefit at a low total development cost. The objective of the HIVHAC task is to advance the Hall thruster technology readiness for science mission applications. The task seeks to increase specific impulse, throttle-ability and lifetime to make Hall propulsion systems applicable to deep space science missions. The primary application focus for the resulting Hall propulsion system would be cost-capped missions, such as competitively-selected, Discovery-class missions. The objective of the advanced xenon feed system task is to demonstrate novel manufacturing techniques that will significantly reduce mass, volume, and footprint size of xenon feed systems over conventional feed systems. This task has focused on the development of a flow control module, which consists of a three-channel flow system based on a piezo-electrically actuated valve concept, as well as a pressure control module, which will regulate pressure from the propellant tank. Cross-platform component standardization and simplification are being investigated through the Standard Architecture task to reduce first user costs for implementing electric propulsion systems. Progress on current hardware development, recent test activities and future plans are discussed.

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

The piloted deep space exploration missions envisioned by the NASA Human Exploration and Development of Space initiative will require the development of advanced electric propulsion systems capable of providing high specific impulse for extended periods of operation. Current electric propulsion thrusters are well suited for orbit maneuvering and robotic exploration, but at present they cannot provide the combination of specific impulse, lifetime, and efficiency required for piloted deep space missions. The theta-pinch thruster concept is a high power plasma rocket that can potentially meet these future deep space propulsion requirements. Efficient, partial preionization of a gas propellant followed by rapid adiabatic magnetic compression is used to generate, heat, and expel a high velocity, high density plasma to provide thrust. The concept is electrodeless, and radial compression of the plasma by the magnetic field of the discharge coil mitigates material erosion to ensure long thruster life. Because the heated plasma is free to flow along axial magnetic field lines during compression, a magnetic mirror located at the entrance to the discharge chamber is used to direct the plasma flow out of the thruster. The thrust and specific impulse of the engine can be tailored for a given mission scenario through the selection of propellant species, mass flow rate, compression coil discharge current, and/or the compression coil repetition rate, making this a unique and versatile electric propulsion system.

Active development of electric thrustors began 10 years ago. Today, several kinds of thrustors have achieved efficiencies above 90 % and lifetimes of several thousand hours. The following article derives the basic theory of electric thrust production at constant exhaust velocity, and at variable exhaust velocity programmed for optimum vehicle performance. Electrothermal or arcjet; electrostatic or ion; and electrodynamic or plasma thrustors are described. At the present time, ion thrustors of the electron bombardment and of the surface ionization types are the most promising systems. Electric power in space may be generated by solar cells or nuclear-electric generators. It is expected that the incore thermionic converter will eventually be the preferred system. A variety of missions with electric propulsion systems appear feasible and highly desirable, among them orbital station keeping, attitude control, planetary probes, solar and out-of-the-ecliptic probes, deep-space probes, and manned Mars and Venus exploration. For each mission, a careful systems-design study must be made, which will provide the optimum selection of thrustor type, thrust level, exhaust velocity, thrust program, power source, trajectory, and flight plan.

Solar sails have long been known to be an attractive method of propulsion in the inner solar system if the areal density of the overall spacecraft (S/C) could be reduced to ∼10 g/m2. It has also long been recognized that the figure (precise shape) of useful solar sails needs to be reasonably good, so that the reflected light goes mostly in the desired direction. If one could make large reflective surfaces with reasonable figure at an areal density of ∼10 g/m2, then several other attractive options emerge. One is to use such sails as solar concentrators for solar-electric propulsion. Current flight solar arrays have a specific output of ∼100W/kg at 1 Astronomical Unit (AU) from the sun, and near-term advances promise to significantly increase this figure. A S/C with an areal density of 10 g/m2 could accelerate up to 29 km/s per year as a solar sail at 1 AU. Using the same sail as a concentrator at 30 AU, the same spacecraft could have up to ∼45 W of electric power per kg of total S/C mass available for electric propulsion (EP). With an EP system that is 50% power-efficient, exhausting 10% of the initial S/C mass per year as propellant, the exhaust velocity is ∼119 km/s and the acceleration is ∼12 km/s per year. This hybrid thus opens attractive options for missions to the outer solar system, including sample-return missions. If solar-thermal propulsion were perfected, it would offer an attractive intermediate between solar sailing in the inner solar system and solar electric propulsion for the outer solar system. In the example above, both the solar sail and solar electric systems don't have a specific impulse that is near-optimal for the mission. Solar thermal propulsion, with an exhaust velocity of the order of 10 km/s, is better matched to many solar system exploration missions. This paper derives the basic relationships between these three propulsion options and gives examples of missions that might be enabled by such hybrids.

Electric thrusters generate high exhaust velocities and can achieve specific impulses in excess of 1000 s. The low thrust generation and high specific impulse make electric propulsion ideal for interplanetary missions, spacecraft station keeping, and orbit raising maneuvers. Consequently, these devices have been used on a variety of space missions including Deep Space 1, Dawn, and hundreds of commercial spacecraft in Earth orbit. In order to provide the required total impulses, thruster burn time can often exceed 10,000 hours, making thruster lifetime essential. One of the main life-limiting components on ion engines is the hollow cathode, which serves as the electron source for ionization of the xenon propellant gas. Reactive contaminants such as oxygen can modify the cathode surface morphology and degrade the electron emission properties. Hollow cathodes that operate with reactive impurities in the propellant will experience higher operating temperatures, which increase evaporation of the emission materials and reduce cathode life. A deeper understanding of the mechanisms initiating cathode failure will improve thruster operation, increase lifetime, and ultimately reduce cost. A significant amount of work has been done previously to understand the effects of oxygen poisoning on vacuum cathodes; however, the xenon plasma adds complexity, and its role during cathode poisoning is not completely understood. The work presented here represents the first attempt at understanding how oxygen impurities in the xenon discharge plasma alter the emitter surface and affect operation of a 4:1:1 BaO-CaO-Al2O3 hollow cathode. A combination of experimentation and modeling was used to investigate how oxygen impurities in the discharge plasma alter the emitter surface and reduce the electron emission capability. The experimental effort involved operating a 4:1:1 hollow cathode at various conditions with oxygen impurities in the xenon flow. Since direct measurements of the emitter surface state cannot be obtained because of the cathode geometry and high particles fluxes, measurements of the emitter temperature using a two-color pyrometer were used to determine the oxygen surface coverage and characterize the rate processes that occur during poisoning. A model describing the material transport in the plasma discharge was developed and is used to predict the barium and oxygen fluxes to the emitter surface during cathode operation by solving the species continuity and momentum equations. The dominant ionization process for molecular oxygen in the plasma gas is resonant charge exchange with xenon ions. Barium is effectively recycled in the plasma; however, BaO and O2 are not. The model shows that the oxygen flux to the surface is not diffusion limited. Experimental results indicate that the oxygen poisoning rate is slow and that the oxygen poisoning coverage on the emitter surface is less than 3%. A time-dependent model of the reaction kinetics of oxygen and barium at the tungsten surface was developed using the experimental results. The experiments and kinetics model indicate that the dominant processes at the emitter surface are dissociative adsorption of O2, sputtering of the O2 precursor, and desorption of O. Ion sputtering of the weakly bound O2 precursor state limits the poisoning rate and yields low oxygen coverage. Removal of chemisorbed atomic oxygen is dominated by thermal processes. Based on the low oxygen coverage and long poisoning transients, plasma cathodes appear to be able to withstand higher oxygen concentrations than vacuum cathodes.

The space nuclear electric propulsion (NEP) system is a revolutionary space propulsion technology that converts nuclear heat energy into electrical energy and drives high-power electric thrusters to generate thrust. Compared with traditional propulsion technology, NEP has the advantages of high specific impulse, high power and long life, which is very suitable for large-scale deep space exploration missions in the future. In this paper, the specific mass model of high power NEP system is established based on NEP system composition and the theory of small thrust orbit to obtain the maximum payload ratio. The complex coupling relation of orbit time, specific mass, power and payload ratio of NEP spacecraft can be decoupled by this model, which provides a computational basis for task optimization. The model was used to evaluate the technical specifications of the NEP system to complete the NASA Juno space mission. The calculation shows that when the specific mass of the NEP system reaches 4.8kg/kWe, it can the ground transfer time of the Juno space mission from 2266 days to 665 days, and the payload from 160kg to 1179kg, which greatly improves the spacecraft's detection capability and providing a useful reference for subsequent design.

Fast sample return from the outer Solar System would open an entirely new avenue for space science, but the vast distances make this a daunting task. The achievable transit velocity and the need for extra propellant on the return trip limit the feasibility of returning extraterrestrial samples to Earth. To keep the mission duration short enough to be of interest, sample return from objects farther out in the Solar System requires increasingly higher velocities. High specific impulse, electric propulsion reduces the propellant required for the outbound and return trips, but decelerating the spacecraft at the inner Solar System from high velocity still involves a long, inward spiral trajectory. The use of solar sails to rapidly decelerate incoming sample capsules and eliminate propellant is explored in this paper. The sail is essentially a ''solar parachute'' used for braking at the end of the interplanetary return flight, permitting a higher transit speed and truncating the deceleration spiral. In this application the sail is relatively small and manageable since only the sample capsule and its sail are decelerated. A comparison is made between using all-electric propulsion versus combining electric propulsive acceleration with sail deceleration for sample return from the distances of Saturn, Uranus, and Pluto. Solar-sail braking dramatically reduces the return flight time by one-third or more compared to using electric rocket deceleration. To elucidate the technology requirements, wide ranges for both the loaded sail density and electric propulsion specific mass are considered in this initial parametric study.

The NASA s Evolutionary Xenon Thruster (NEXT) project is developing an advanced ion propulsion system for future NASA missions for solar system exploration. An engineering model (EM) power processing unit (PPU) for the NEXT project was designed and fabricated by L-3 Communications under contract with NASA Glenn Research Center (GRC). This modular PPU is capable of processing up from 0.5 to 7.0 kW of output power for the NEXT ion thruster. Its design includes many significant improvements for better performance over the state-of-the-art PPU. The most significant difference is the beam supply which is comprised of six modules and capable of very efficient operation through a wide voltage range because of innovative features like dual controls, module addressing, and a high current mode. The low voltage power supplies are based on elements of the previously validated NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) PPU. The highly modular construction of the PPU resulted in improved manufacturability, simpler scalability, and lower cost. This paper describes the design of the EM PPU and the results of the bench-top performance tests.

A novel concept of cost-effective active debris removal spacecraft system

Onboard the ESA SMART-1 spacecraft, (Small Mission for Advanced Research and Technology), the primary Electric Propulsion Subsystem (EPS) must operate at optimum performance under limited power available and with a variable power feature. As EPS Contractor for ESA-ESTEC, SNECMA present in detail the major's features of the complete electric propulsion system, as well as the development efforts of each main component. The development tests of the PPS-1350-G hall effect plasma thruster, including an extensive number of start-up tests in the whole useful operational domain have been performed. This paper describes the results of the PPS-1350 low-power characterization tests, and presents the new approach for having a limited inrush power when the thruster is switched on. The test campaign has been successful and the results obtained during the campaign have been fully taken into account in an optimization process for providing the foreseen performances of the SMART-1 mission. Those performances are the thrust, the total xenon flow, the specific impulse and the additional magnet trim current. Acceptance tests of each components of the EPS have been performed and finally an to end test has been successfully performed at satellite level. These new features of primary electric propulsion subsystem and especially the low-power start-up and variable power features can be also a significant added value for any commercial application using electric propulsion for station-keeping and/or orbit transfer. Introduction Although High Specific Impulse Thrusters have already propelled numerous satellites in the frame of station-keeping missions, their use for orbit transfer and deep space missions has only recently begun, bringing out a need for new thrust strategies. The Snecma PPS1350-G Hall-effect thruster is primarily intended for North South station keeping of Geo-stationary satellites. For that kind of application, the spacecraft available power is much higher than the PPU maximum power demand, so there is no risk to overload the spacecraft power bus. In the case of SMART-1 , on the contrary, the PPU power could exceed the total spacecraft power. In addition, the thruster shall be able to work with less than half its nominal power. To fulfil these requirements, it was necessary to modify the thruster/PPU interfaces and to validate these modifications by tests. SMART-1 Electric Propulsion Subsystem (EPS) The SMART-1 power is generated by two solar panels enabling thruster operation at a discharge power of maximum 1190 W at beginning of life . In some cases, (e.g. failure of one solar wing in addition to solar cells) the available power can be reduced to 700 W or less. This is therefore an essential requirement for thruster and PPU to work within a range of pre-set power levels. This capability is also required for deep-space missions involving variable sun distances. The whole EPS is designed for the three following main functions : u Xenon system u Electrical power system and thruster u Digital interface and communication system 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 20-23 July 2003, Huntsville, Alabama AIAA 2003-4545 Copyright © 2003 by Snecma Moteurs and ESA. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. 2 PPU/TSU Bus Input, Input Switch