Timothy R. Sarver-Verhey

Ion propulsion development activities at the NASA Glenn Research Center

The NASA Glenn Research Center (GRC) ion propulsion program addresses the need for high specific impulse ion propulsion systems and technology across a broad range of mission applications and power levels. Development areas include high-throughput NSTAR derivative engine and power processing technology, lightweight high-efficiency sub-kilowatt ion propulsion, micro-ion propulsion concepts, engine and component technologies for highpower (30 kW class) ion engines, and fundamentals. NASA GRC is also involved in two highly focussed activities: development of 5/10-kW class next-generation ion propulsion system technology, and development of high-specific impulse (> 10,000 seconds) ion propulsion technology applicable to deep-space and interstellar-precursor missions.

Solar Electric Propulsion Vehicle Design Study for Cargo Transfer to Earth-Moon L1

ABSTRACTA design study for a cargo transfer vehicle using solar electric propulsion was performed for NASA’s Revolu-tionary Aerospace Systems Conceptsprogram. Targetedfor 2016, the solar electric propulsion (SEP) transfervehicle is required to deliver a propellant supply module with a mass ofapproximately 36 metric tons fromLow Earth Orbit to the first Earth-Moon libration point (LL1) within 270 days. Following an examination ofpropulsion and power technology options, a SEP transfer vehicle design was selected that incorporated large-area (~2700 m 2 ) thin film solar arrays and a clustered engine configuration of eight 50 kW gridded ionthrusters mountedonanarticulatedboom. Refinement of the SEP vehicle designwasperformediteratively toproperly estimate the required xenon propellant load for the out-bound orbit transfer. The SEP vehicle per-formance, including the xenon propellant estimation, was verified via the SNAP trajectory code. Further ef-fortsare underway to extendthissystem model to otherorbit transfer missions.INTRODUCTION

Concurrent Mission and Systems Design at NASA Glenn Research Center: The Origins of the COMPASS Team

Abstract Established at the NASA Glenn Research Center (GRC) in 2006 to meet the need for rapid mission analysis and multi-disciplinary systems design for in-space and human missions, the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team is a multidisciplinary, concurrent engineering group whose primary purpose is to perform integrated systems analysis, but it is also capable of designing any system that involves one or more of the disciplines present in the team. The authors were involved in the development of the COMPASS team and its design process, and are continuously making refinements and enhancements. The team was unofficially started in the early 2000s as part of the distributed team known as Team JIMO (Jupiter Icy Moons Orbiter) in support of the multi-center collaborative JIMO spacecraft design during Project Prometheus. This paper documents the origins of a concurrent mission and systems design team at GRC and how it evolved into the COMPASS team, including defining the process, gathering the team and tools, building the facility, and performing studies.

High Current Cathode Development for 50 kW C lass Hall Thrusters

50 kW class Hall thrusters developed at NASA G lenn Research Center (GRC) requir ed holl ow cathodes to operate at emission currents up to 100 A. Future Hall thruster designs may operate at over 100 kilowatts , requiring cathodes capable of operating at emission currents over 200 A. A test program was undertaken to det ermine the hollow cathode geometr y needed to meet the se high emission current requirements while provid ing long life. To enable this, a laboratory model hollow cathode assembly with a reconfigurable geometry was designed and characteriz ed at emission curr ents up to 100 A . This cathode design incorporated some characteristics of the NASA plasma contactor design but was modified for integration along the central axis of a Hall thruster. The reconfigurable geometry allowed for variation of the following par ameters: cathode orifice size, emitter & cathode diameter, cathode -keeper spacing , and propellant flow rate. Cathode temperature s were measured to determine the optimal geometry for an emission current of 100 A.

Hollow Cathode Assembly Development for the HERMeS Hall Thruster

To support the operation of the HERMeS 12.5 kW Hall Thruster for NASAs Asteroid Redirect Robotic Mission, hollow cathodes using emitters based on barium oxide impregnate and lanthanum hexaboride are being evaluated through wear-testing, performance characterization, plasma modeling, and review of integration requirements. This presentation will present the development approach used to assess the cathode emitter options. A 2,000-hour wear-test of development model Barium Oxide (BaO) hollow cathode is being performed as part of the development plan. Specifically this test is to identify potential impacts cathode emitter life during operation in the HERMeS thruster. The cathode was operated with a magnetic field-equipped anode that simulates the HERMeS hall thruster operating environment. Cathode discharge performance has been stable with the device accumulating 743 hours at the time of this report. Observed voltage changes are attributed to keeper surface condition changes during testing. Cathode behavior during characterization sweeps exhibited stable behavior, including cathode temperature. The details of the cathode assembly operation of the wear-test will be presented.

Ion Thruster Development at NASA Lewis Research Center

Recent ion propulsion technology efforts at NASA’s Lewis Research Center including development of kW‐class xenon ion thrusters, high power xenon and krypton ion thrusters, and power processors are reviewed. Thruster physical characteristics, performance data, life projections, and power processor component technology are summarized. The ion propulsion technology program is structured to address a broad set of mission applications from satellite stationkeeping and repositioning to primary propulsion using solar or nuclear power systems.

Discharge ignition behavior of the Space Station plasma contactor

Ignition testing of hollow cathode assemblies being developed for the Space Station plasma contactor system has been initiated to validate reliable multiple restart capability. An ignition approach was implemented that was derived from an earlier arcjet program that successfully demonstrated over 11,600 ignitions. For this, a test profile was developed to allow accelerated cyclic testing at expected operating conditions. To date, one hollow cathode assembly has been used to demonstrate multiple ignitions. A prototype hollow cathode assembly has achieved 3,615 successful ignitions at a nominal anode voltage of 18.0 V. During the ignition testing several parameters were investigated, of which the heater power and pre-heat time were the only parameters found to significantly impact ignition rate.

Functional Testing of the Space Station Plasma Contactor

A plasma contactor system has been baselined for the International Space Station Alpha (ISSA) to control the electrical potentials of surfaces to eliminate/mitigate damaging interactions with the space environment. The system represents a dual-use technology which is a direct outgrowth of the NASA electric propulsion program and, in particular, the technology development effort on ion thruster systems. The plasma contactor subsystems include a hollow cathode assembly, a power electronics unit, and an expellant management unit. Under a pre-flight development program these subsystems are being developed to the level of maturity appropriate for transfer to U.S. industry for final development. Development efforts for the hollow cathode assembly include design selection and refinement, validating its required lifetime, and quantifying the cathode performance and interface specifications. To date, cathode components have demonstrated over 10,000 hours lifetime, and a hollow cathode assembly has demonstrated over 3,000 ignitions. Additionally, preliminary integration testing of a hollow cathode assembly with a breadboard power electronics unit has been completed. This paper discusses test results and the development status of the plasma contactor subsystems for ISSA, and in particular, the hollow cathode assembly.

International Space Station Cathode Life Testing Status

To demonstrate adequate lifetime and performance capabilities of a hollow cathode for use on the International Space Station (ISS) plasma contactor system, life tests of multiple hollow cathode assemblies (HCAs) were initiated at operating conditions simulating on-orbit operation. Three HCAs are presently being tested. These HCAs are operated with a continuous 6 sccm xenon flow rate and 3 A anode current. Emission current requirements are simulated with a square waveform consisting of 50 minutes at a 2.5 A emission current and 40 minutes with no emission current. As of July 1998, these HCAs have accumulated between 1 1,700 and 14,200 hours. While there have been changes in operatin, behavior the three HCAs continue to operate stably within ISS specifications and are expected to demonstrate the required lifetime.

Preparation for Hollow Cathode Testing for the Advanced Electric Propulsion System at NASA Glenn Research Center

NASA Glenn Research Center is performing activities to support the unique needs of hollow cathode development and testing for the Advanced Electric Propulsion System (AEPS). Three existing vacuum facilities have been outfitted as cathode test facilities, and each will serve a different role in upcoming testing. Vacuum Facility 67 is being developed to serve as a longduration test facility for the Engineering Development Unit cathode, which is to be delivered by the AEPS contractor. It will feature a thruster-like magnetic field simulator and cold-cycle capability via a liquid nitrogen-cooled cold plate. Vacuum Facility 17 is being developed as a test facility for shortto medium-duration experiments in order to provide auxiliary support for the long-duration testing. It will feature a magnetic field simulator but not cold-cycling. Finally, Vacuum Facility 1 will be a high-pumping speed cathode development environment, and will feature an array of plasma and temperature diagnostics. In addition to the facility preparation work, a new cathode, referred to as the Mark II, has been designed. The Mark II is an evolution of the Technology Demonstration Unit cathodes that better evokes the geometry, fabrication, and construction of the forthcoming Engineering Development Unit. This cathode serves as a transition between the Technology Demonstration Unit cathodes used during early thruster development and the forthcoming Engineering Development Unit cathodes. It will be used as a means of verifying the new test facilities prior to arrival of Engineering Development Unit hardware. Details of the Mark II design and key features are presented, as well as details of future work to be performed.