Development of a High-Propellant Throughput Small Spacecraft Electric Propulsion System to Enable Lower Cost NASA Science Missions

Abstract

This paper describes recent progress at the NASA Glenn Research Center (GRC) in the development and demonstration of an integrated high-propellant throughput small spacecraft electric propulsion (HT-SSEP) system based on a Hall-effect thruster. A center-mounted cathode and an innovative magnetic circuit topology were implemented in the design of the Hall-effect thruster to achieve high-propellant throughput, high performance, and efficient packaging. To minimize technical risk, the HT-SSEP development approach sought to limit design features and materials to those with a clear path-to-flight. A propellant throughput capability of greater than 100 kg at a minimum thruster efficiency of 50% was targeted. The proof-of-concept NASA-H64M laboratory model (LM) thruster was designed, fabricated, and tested at GRC in fiscal year 2018. The thruster development leveraged heritage Hall-effect thruster design and manufacturing processes wherever appropriate. Recent NASA advances in Hall-effect thruster technology were also leveraged. A scalable discharge power supply (DPS) capable of powering the H64M-LM was developed, then demonstrated as part of an integrated system test. The DPS uses commercial off-the-shelf components with spaceflight equivalents. A keeper supply with DC ignitor was breadboarded, then demonstrated with a laboratory cathode. Finally, feed system trade studies were performed to ascertain what feed system architecture might be appropriate for an HT-SSEP system. This paper details the motivations for the project, the development approach, the chosen sub-system architectures, design considerations, and test results.

Performance characterization of a low-power magnetically shielded Hall thruster with an internally-mounted hollow cathode

“…These values decreased with SN 2, which demonstrated 2.8%, 1.5%, and 1.3% at 0.5 mg s −1 , 1 mg s −1 , and 3 mg s −1 , respectively. While comparable data is limited in the open literature, these results represent a significant improvement over previously published results for Hall thrusters of all power levels [50,54] The MaSMi-DM was operated over a discharge voltage range of 200-600 V in increments of 100 V, a discharge current range of 0.5-4 A in increments of 0.5 A, and a discharge power range of 150-1000 W. At discharge voltages beyond 400 V, the minimum discharge current was increased to 1 A. Additionally, data was collected at 2.67 and 3.33 A for the 300 V condition. A summary plot of the operating points examined is presented in figure 8.…”

“…This is an important result as numerous Hall thrusters demonstrate significant (∼3%-15%) performance reductions within the first 1 kh of operation MaSMi-DM operation at 300 V-600 W (top row), 400 V-800 W (middle row), and 500 V-1000 W (bottom row) to thermal steady state, displaying thruster temperatures (left column) and performance (right column) versus time. [19,54,57]. This can generally be attributed to ionbombardment erosion changing the discharge channel geometry (specifically, creating or increasing the chamfers at the channel exit), yielding an associated increase in beam divergence and reduction in propellant utilization efficiency.…”

Polarization Re-Configurable Antenna with Increase Gain for Small Satellites

“…The small satellites are of tinier sizes and have less weight, and therefore, they are becoming the most appealing area of research in the aerospace industry. Small satellites usually weigh less than 500 Kg and are categorized accordingly as mini, micro, nano (cube satellite), pico, and femto-satellites [2]. These small satellites enable the growth of new space technologies by allowing companies, educational organizations, scientists, and engineers to have access to space at lower cost [3].…”