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A Low Power Cylindrical Hall Thruster for Next Generation Microsatellites

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A Low Power Cylindrical Hall Thruster for Next Generation Microsatellites SSC15-P36 A Low Power Cylindrical Hall Thruster for Next Generation Microsatellites Carl E. Pigeon, Nathan G. Orr, Benoit P. Larouche, Vincent Tarantini, Grant Bonin, Robert E. Zee Space Flight Laboratory Microsatellite Science and Technology Center University of Toronto Institute for Aerospace Studies 4925 Dufferin Street, Toronto, Ontario, Canada, M3H 5T6; 416-667-7400 [email protected] ABSTRACT As the demand for highly capable microsatellite missions continues to grow, so too does the need for small, low power satellite technologies. One area which needs to be addressed is advanced propulsion systems capable of performing on-orbit maneuvers, station keeping and deorbit impulses with minimal propellant. A relatively low power Cylindrical Hall Thruster (CHT) is being developed at the Space Flight Laboratory (SFL) specifically to meet the low power and size requirements of microsatellite missions. The development of SFL’s sub 200 Watt, 26 mm ionization chamber diameter Hall Thruster will permit more capable microsatellite missions while minimizing the required propellant and thruster mass. The aim is to produce a CHT qualified to Technology Readiness Level (TRL) 6 by the end of 2015. The first development phase focused on a highly configurable prototype design which facilitated optimization of performance parameters such as thrust, specific impulse and power consumption. The second phase incorporated lessons learned from prototype testing into the development of a proto-flight model. Subsequent work will involve packaging and qualifying a standalone flight unit including all electrical interfaces and the propellant feed system. Evaluations of alternative propellants such as Krypton and Argon against the baseline Xenon propellant will also be performed. This paper presents the test results from the low power Cylindrical Hall Thruster development campaign and discusses the status of the program and future plans. INTRODUCTION monopropellant thruster. In accordance with The Space Flight Laboratory (SFL) was awarded a requirements, the monopropellant thruster uses contract under the Canadian Space Agency’s (CSA) alternative non-toxic propellants as oppose to Space Technology Development Program (STDP) to hydrazine. develop an on-board propulsion system for micro and small satellite station keeping and de-orbiting. The The second development initiative, which is the subject mandate of the STDP was to encourage research and of this paper, focuses on developing a low thrust, high development by awarding a contract to develop and demonstrate technologies that would meet the current and future needs of the Canadian Space Programs. One among the twelve highlighted key technologies was microsatellite propulsion systems. As part of the development initiative, it was required that the key technologies be demonstrated to Technology Readiness Level 6 (TRL6), which by definition means developing a representative model or prototype system that has been tested in a representative operating environment. SFL chose to address the STDP with two technology development initiatives that would be pursued in parallel. The first leverages the existing Canadian Nanosatellite Advance Propulsion System (CNAPS, see Figure 1) cold gas thruster technology of the CanX-4 & CanX-5 mission which successfully demonstrated microsatellite formation flying using this technology1. The CNAPS cold gas propulsion system is being upgraded into a low specific impulse, and high thrust Figure 1: CanX-4 and CanX-5 Cold Gas Thruster Propulsion System Carl E. Pigeon 1 29th Annual AIAA/USU Conference on Small Satellites specific impulse propulsion system capable of 1) Magnetic field characterization efficiently operating for long durations of time in order 2) Cathode testing to enable station keeping, orbit changes and end of life 3) Experimental optimization of mass flow rate, de-orbit impulse maneuvers. As a result, the initiative magnetic field, and thruster ionization chamber focuses on developing a low-power electric propulsion length system that meets the requirements of future Canadian 4) Thrust and specific impulse measurements microsatellite and small satellite missions. This is to be 5) Lifetime testing to observe thruster degradation Canada’s first flight qualified electric propulsion system. In order to ensure the propulsion system being Phase Two developed is relevant to future missions being planned The second phase incorporates lessons learned from the within Canada, SFL collaborated with Canada’s leading prototype model to design a proto-flight model that is industrial space companies including COM DEV Ltd., non-reconfigurable and representative of the flight MacDonald Detwiller and Associates Ltd., and model. The result is a design that reduced the mass by Magellan Aerospace-Winnipeg. 76% and includes the proper mechanical and electrical interfaces in preparation for the next phase. Similar to REQUIREMENTS Phase One, the end of this phase includes performance The requirements for the propulsion program are evaluation of the thruster. derived from the needs of industrial partners and SFL to Phase Three ensure that the thruster being developed is able to In Phase Three, drive electronics including power satisfy Canada’s future microsatellite missions. A brief regulators and control logic will be integrated in order summary of the key requirements that drove the design to make a complete stand-alone system. At the end of of the Electric Propulsion System are found in Table 1. the program the proto-flight model will undergo full Table 1: Propulsion System Key Requirements flight qualification testing in order to bring the design to TRL6 status. Environmental testing will include: Number Requirement 1) Vibration testing including Sine Sweep, Shock 1 The propulsion system shall impart a minimum ΔV of and Random at specified space qualification 100 m/s to the reference spacecraft levels 2 The reference spacecraft shall have a 150 kg dry 2) Thermal vacuum (TVAC) testing mass, not including the mass of the thruster assembly 3 The propulsion system shall not require more than 300 W DC power while in operation DESIGN JUSTIFICATION 4 The propulsion system shall have a specific impulse After a comprehensive review it was determined that a greater than 1000 s at the design point high specific impulse electric propulsion (EP) system 5 The targeted thrust range should be between 5 mN capable of enabling significant on-orbit and 10 mN maneuverability and orbit changes would be beneficial 6 The propellant shall be green for Canada’s future microsatellite missions. Several EP MICRO-PROPULSION PROGRAM STRUCTURE technologies were evaluated for compatibility with the requirements along with their development risk. The The program has been organized into three phases, each outcome of the trade study determined that a having key development goals and test milestones to Cylindrical Hall Thruster (CHT) was the best candidate complete. Organized in this fashion, SFL mitigates as it is suitable for operation at power levels technical risk while achieving a rapid development between 50 W to 300 W and has a mass and volume schedule. As of the time of writing, the program is in suitable for microsatellites. Phase Two. Traditional annular Hall Thrusters (HT) which are Phase One characterized by an annular ionization chamber are The first phase focused on designing and manufacturing mechanically simpler and produce higher thrust to a thruster prototype that would operate on laboratory power ratios than their counterpart – the Gridded Ion power supplies with manually controlled propellant Thruster. Since microsatellites are generally power flow regulators. The thruster was built to be fully limited, a high thrust/power ratio is desirable. However, reconfigurable in order to allow for experimental conventional HTs have drawbacks when miniaturized. optimization. The objective of this phase was to have a When reducing the size of the ionization chamber for working prototype to perform basic functional testing optimal low power operation, their annular geometry which included the following: produces a high channel surface area to channel volume Carl E. Pigeon 2 29th Annual AIAA/USU Conference on Small Satellites Figure 2: Schematic of a Cylindrical Hall Thruster Figure 3: SFL’s Prototype Cylindrical Hall Thruster on its Test Stand 2 ratio. This increases the interaction of plasma with the parameters that would maximize specific impulse and channel walls resulting in heating and erosion of thrust between 50 and 200 W of discharge power. thruster parts, most prominently the critical inner parts of the coaxial channel and the magnetic circuit. Thus, compared to larger HTs, low power HTs are inefficient A photograph of the 26 mm diameter SFL prototype at 6-30 % efficiency at 0.1-0.2 kW compared to greater CHT is shown in Figure 3. Weighing 1.6 kg (excluding than 50 % efficiency at 1 kW.3,4 cathode) the thruster is oversized for what is needed. Much of the mass is attributed to the use of To mitigate these issues, Princeton Plasma Physics Lab electromagnets which tend to be heavier and bulkier (PPPL) proposed and developed a Hall Thruster with a than permanent magnets. Within the thruster, a steel cylindrical ionization chamber5 as shown in Figure 2. magnetic core strengthens and shapes the magnetic This configuration reduces the surface area to volume field inside the ionization chamber. Several materials ratio thus limiting electron transport and ion losses.6 A were evaluated for the magnetic
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