Direct drive options for 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.

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 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.

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 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.

The paper gives examples of orbital maneuver execution of spacecraft (SC) using orbit correction propulsion system (OCPS) based on electric propulsion (EP). The high velocity of the propellant flow achieved in EP allows for orbital maneuvers with significantly lower propellant flow rate than in conventional propulsion systems. The time required to perform the orbital maneuvers using EP is connected with the available on-board power. It is typically much greater than such time needed for conventional jet propulsions. The advantages of EP are realized if mission allows a long-time operation of the propulsion system. Since the early 1970s stationary plasma thrusters (SPTs) developed on the base on the concept proposed by A.I. Morozov are used on the SC Meteor. With the help of OCPS based on SPT EOL-1, SC Meteor was installed on the conventionally synchronous orbit, which provides a fixed grid of tracks with a period of T = 102.31 min. In this case, a complete overview of all daily Earth’s surface with the bandwidth of 2900 km is obtained. In recent decades, there has been a steady trend toward miniaturization of space equipment, which requires the development of acceptable thrusters to meet new requirements. Typical total impulses of thrust required for OCPS are reduced several times. At present, a number of space constellations are being developed in Russia based on small SC with a mass from 60 to 500 kg. VNIIEM Corporation creates a space constellation IONOZOND, designed to monitor the geophysical conditions. The description of the IONOZOND constellation is given and the options for using of various OCPS, in particular, based on SPT, ion and pulsed plasma thrusters are considered. It is shown that their use in small SC can significantly increase the economic efficiency of remote sensing orbital constellations.

Miniaturization perspectives of electrostatic propulsion for small spacecraft platforms

Most of marine electric propulsion ships are propelled by large electric motors. The speed control of the propulsion motor on the electric propulsion vessel is operated by adjusting the frequency. The power system has the inverter installed semiconductor element, such as IGBT or IGCT, for high-speed switching. However, the inverter for power conversion unit of the electric power propulsion on large vessels has fundamentally thermal problems of installed semiconductor, such problems could influence to considerable loss by switching and conducting the inverter in the power system transferring large current with high voltage. In order to calculate power loss and thermal loss of the inverter, this paper sets actual parameters of the electric propulsion system of the large ship in the simulation program considering condition of ship operation. The simulation proves that it is possible to analyze both power loss and thermal loss in the diodes and transistor during switching of the inverter in comparison with field oriented control and direct torque control. As a result, large power loss and thermal loss of switching and conducting the inverter by field oriented control considerably occurs and can also impact damages toward the power electronic devices by the thermal loss generated.

Three aspects of the proposed Asteroid Redirect Mission (ARM) could be extended to provide greater capability for future NASA missions: higher-power versions of the baseline asteroid redirect vehicle, in-space resource utilization, and planetary defense. The baseline ARM vehicle assumes the use of a 50 kW beginning-of-life solar array which provides a maximum of 40 kW to the electric propulsion system. Launch dates in mid to late 2020 could provide the opportunity for the development and implementation of higher-power solar arrays and electric propulsion systems that are farther along the path to the 100-kW-class systems that could be used to support human missions to Mars. The ARM robotic vehicle conceptual design provides a straightforward approach to increasing the solar array power to ∼100 kW for the first asteroid redirect mission. Transportation is also a major challenge for harvesting asteroids for the use of their material resources in space. ARM addresses this issue by selecting an asteroid that naturally returns close to Earth and then redirecting it into lunar orbit. Deriving propellants from asteroids is essential to a robust utilization of asteroid material resources. Two elements, magnesium and sulfur, abundantly available in common chondrite asteroids could be used as propellants in Hall thrusters and may be the key to asteroid mining. Finally, ARM has the potential to demonstrate two different planetary defense techniques: an enhanced gravity tractor, or an ion beam deflector. High-power solar electric propulsion (SEP) is needed for both techniques. Simple analyses highlight a clear choice between these options. To obtain the same force and total impulse applied to a potentially hazardous asteroid you can either develop higher power SEP systems for ion beam deflection or you can develop the capability to acquire hundreds of tons of mass from the asteroid for use with a lower power SEP system in an enhanced gravity tractor approach.

20 - Electric propulsion systems

The mission targets of the Chinese asteroid exploration are an earth co-orbiting celestial body 2016HO3 and the main-belt comet in the asteroid belt. Compared with Lunar Exploration and Mars Exploration missions, asteroid exploration mission is special. The light intensity varies widely during the mission, resulting in a large change in the output power range of the solar array, which seriously affects the power system design and the load design during the mission; the spacecraft uses a high-power electric propulsion system; the power system has strict weight requirements, the existing power supply system design cannot meet the requirements. Aiming at the design constraints of the large variation of solar wing light intensity, the large difference in power requirements between the electric propulsion and the platform, and the lightweight design of the spacecraft, a dual-bus power supply system is proposed. High-voltage bus power the electric propulsion system. The solar cell array directly converges to form a high-voltage bus, which supplies power for the electric propulsion system. The low-voltage bus adopts a semi-adjustable bus, and the solar cell array adopts a circular flexible solar wing. The main detector and return cabin are equipped with lithium-ion battery pack and lithium fluorocarbon battery pack respectively. Aiming at the requirement of long-term on-orbit operation of electric propulsion system, a solar wing output power measurement system is designed. Product test results show that the power system design can meet the working requirements of electric propulsion and platform loads. The power system design can provide reference for the other deep space exploration power system.

A data-driven approach for analyzing Hall thruster discharge instability leading to plasma blowoff

Abstract. The report presents the results of the development of the PPU-500, which is part of the SPS-500 electric propulsion system based on the ST-40 Hall thruster. PPU-500 contains all power supplies necessary for the operation of the propulsion system: discharge; electromagnet; cathode heater and keeper; power supplies for valves and heaters of the storage and feeding working substance, as well as a standby power supply that ensures the standby mode of PPU operation.The converter of the discharge power supply is built using a bridge topology with a phase-shifted output power control circuit. The remaining power supplies are built according to the topology of a step-down synchronous converter with current or voltage stabilization. The discharge power supply is a controlled power supply. The power supplies for the cathode heater and electromagnet current sources are. The accuracy of power supplies output parameters in not worse than 1.5%. The valves of the feeding system are controlled in order to stabilize the pressure in the receiver by periodically turning on/turning off the high- and low-pressure valves. For control the valves, signals taken from pressure sensors installed in the receiver are used. The PPU control unit ensures an accuracy of maintaining pressure in the receiver no worse than 2.5%. The control unit of the PPU is built using a microcontroller with an ARM core. It implements the functions of monitoring the condition of the program, restoring the program firmware if an integrity violation is detected, and the ability to remotely flash a new version. To communicate the SPS-500 with the on-board control system, a galvanically isolated CAN interface is used. It can be also the RS485/RS42 interface.During operation of the propulsion system, operating parameters are measured using multi-channel ADCs and DACs. The obtained data, on the one hand, is used to assess the current state of the propulsion system and control Hall thruster operating modes. On the other hand, they are transmitted as telemetry signals reflecting the current state of the SPS-500, operating mode and emerging emergency conditions of the equipment. PPU is possible to alternately control two thrusters. Solid-state switches are used to switch the corresponding power sources.Testing of the PPU-500 as part of the SPS-500 propulsion system with the ST-40 Hall thruster and a flight prototype of the working substance feeding system confirmed the technical solutions adopted. The operating efficiency of PPU-500 for the nominal operating mode of the propulsion system is 94%.

This paper examines the role of the TOPAZ II and the 40 kWe thermionic reactors in powering electric propulsion systems for orbital transfer and on‐orbit repositioning applications. Specifically, the feasibility of substituting a TOPAZ II power system and the Russian Stationary Plasma Thruster (SPT) for the conventional power and propulsion systems that might be used on typical Air Force navigation, surveillance, and communications satellites were examined. Additionally, two advanced high‐power missions were analyzed. Four major conclusions were drawn from this top‐level investigation: (1) Stepdown to smaller launch vehicles is possible for both TOPAZ II and 40 kWe systems even though the TOPAZ II at 6 kWe and 1100 kg, is not an extremely effective system for orbital transfer applications; (2) On‐orbit repositioning is an area that profits by the utilization of electric propulsion systems; (3) Advanced high‐power missions are arguably the greatest benefactor of nuclear electric propulsion; and (4) Placing minimum operational altitude requirements on reactor‐propelled systems will significantly diminish their payload capability to the destination orbit.

ABSTRACTThe increasing environmental requirements in the air transport sector pose great challenges to the aviation industry and are key drivers for innovation. Besides various approaches for increasing the efficiency of conventional gas turbine engines, electric propulsion systems have moved into the focus of aviation research. The first electric concepts are already in service in general aviation. This study analyses the potentials of electric and turbo hybrid propulsion systems for commercial aviation. Its purpose is to compare various architectures of electrical powertrains with a conventional turboprop on a regional aircraft, similar to the ATR 72, on engine and flight mission levels. The considered architectures include a turbo-electric (power controlled and direct driven), hybrid-electric (serial and parallel) and a pure electric concept. Their system weights are determined using today’s technology assumptions. With the help of performance models and flight mission calculations the impact on fuel consumption, CO ${}_{2}$ emissions and aircraft performance is evaluated.