Ion Propulsion-Solar Wind Plasma Interactions
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Performance Characteristics of the NEXT Long-Duration Test After 16,550 h and 337 kg of Xenon Processed
The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg from mission analyses conducted utilizing the NEXT propulsion system. The LDT is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of June 25, 2008, the thruster has accumulated 16,550 h of operation: the first 13,042 h at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. Operation since 13,042 h, i.e., the most recent 3,508 h, has been at an input power of 4.7 kW with 3.52 A beam current and 1180 V beam power supply voltage. The thruster has processed 337 kg of xenon (Xe) surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster. The NEXT LDT has demonstrated a total impulse of 13.3 106 N s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. This paper presents the performance of the NEXT LDT to date with emphasis on performance variations following throttling of the thruster to the new operating condition and comparison of performance to the NSTAR extended life test.
Soulas, George C.↗
Parametric studies of electric propulsion systems for orbit transfer vehicles
The present parametric tradeoff study for OTV electric propulsion systems encompasses ammonia and hydrogen arcjets as well as Xe-ion propulsion systems with performance characteristics currently being projected for 1993 operation. In all cases, the power source is a nuclear-electric system with 30 kg/kW(e) specific mass, and the mission involves the movement of payloads from lower orbits to GEO. Attention is given to payload capabilities and associated propellant requirements. Mission trip time is identified as the key parameter for selection; while arcjets are preferable for shorter trip times, ion propulsion is more advantageous for longer trip times due to reduced propellant mass fraction.
Manvi, R.↗
Dawn: An Ion-Propelled Journey to the Beginning of the Solar System
The Dawn mission is designed to perform a scientific investigation of the two heaviest mainbelt asteroids Vesta and Ceres. These bodies are believed to preserve records of the physical and chemical conditions present during the formation of the solar system. The mission uses an ion propulsion system to enable the single Dawn spacecraft and its complement of scientific instruments to orbit both of these asteroids. Dawn's three science instruments - the gamma ray and neutron detector, the visible and infrared mapping spectrometer, and the primary framing camera - were successfully tested after launch and are functioning normally. The ion propulsion system includes three ion thrusters of the type flown previously on NASA's Deep Space 1 mission. A minimum of two ion thrusters is necessary to accomplish the Dawn mission. Checkout of two of the ion thrusters was completed as planned within 30 days after launch. This activity confirmed that the spacecraft has two healthy ion thrusters. While further checkout activities are still in progress, the activities completed as of the end of October indicate that the spacecraft is well on its way toward being ready for the start of the thrusting-cruise phase of the mission beginning December 15th.
Vesta↗
Dawn: an ion-propelled journey to the beginning of the Solar System
The Dawn mission is designed to perform a scientific investigation of the two heaviest mainbelt asteroids Vesta and Ceres. These bodies are believed to preserve records of the physical and chemical conditions present during the formation of the solar system. The mission uses an ion propulsion system to enable the single Dawn spacecraft and its complement of scientific instruments to orbit both of these asteroids. Dawn’s three science instruments – the gamma ray and neutron detector, the visible and infrared mapping spectrometer, and the primary framing camera – were successfully tested after launch and are functioning normally. The ion propulsion system includes three ion thrusters of the type flown previously on NASA’s Deep Space 1 mission. A minimum of two ion thrusters is necessary to accomplish the Dawn mission. Checkout of two of the ion thrusters was completed as planned within 30 days after launch. This activity confirmed that the spacecraft has two healthy ion thrusters. While further checkout activities are still in progress, the activities completed as of the end of October indicate that the spacecraft is well on its way toward being ready for the start of the thrusting-cruise phase of the mission beginning December 15th.
Pavri, Betina↗
NEXT Long-Duration Test After 11,570 h and 237 kg of Xenon Processed
The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to validate and qualify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the mission-derived throughput requirement of 300 kg. This wear test is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of September 1, 2007, the thruster has accumulated 11,570 h of operation primarily at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. The thruster has processed 237 kg of xenon surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 (DS1) flight spare. The NEXT LDT has demonstrated a total impulse of 9.78 10(exp 6) N(dot)s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. Lifetime-limiting component erosion rates have been consistent with the NEXT service life assessment, which predicts the earliest failure sometime after 750 kg of xenon propellant throughput; well beyond the mission-derived lifetime requirement. The NEXT wear test data confirm that the erosion of the discharge keeper orifice, enlarging of nominal-current-density accelerator grid aperture cusps at full-power, and the decrease in cold grid-gap observed during NSTAR wear testing have been mitigated in the NEXT design. NEXT grid-gap data indicate a hot grid-gap at full-power that is 60 percent of the nominal cold grid-gap. This paper presents the status of the NEXT LDT to date with emphasis on comparison to the NSTAR extended life test results.
Soulas, George C.↗
Status of the NEXT Long-Duration Test After 23,300 Hours of Operation
The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated in June 2005, to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg per thruster from mission analyses. The LDT is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of July 2009, the thruster has accumulated 23,300 h of operation with extensive durations at the following input powers: 6.9, 4.7, 1.1, and 0.5 kW. The thruster has processed 427 kg of xenon surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster and approaching the NEXT development qualification throughput goal. The NEXT LDT has demonstrated a total impulse of 16.0 10(exp 6) N/s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. The NSTAR first-failure mode, accelerator aperture erosion leading to electron backstreaming, has been mitigated in the NEXT design. The severe NSTAR discharge cathode assembly erosion has been mitigated by a graphite keeper in the NEXT thruster. Tracking of the NEXT first failure mode, charge-exchange ion impingement on the accelerator grid causing hexagonal groove erosion, is consistent with model predictions and indicates thruster life greater than or equal to 750 kg throughput. This paper presents the status, performance data, and wear characteristics of the NEXT LDT to date.
Herman, Daniel A.↗
Planned flight test of a mercury ion auxiliary propulsion system. I - Objectives, systems descriptions, and mission operations
A planned flight test of an 8-cm diameter, electron-bombardment mercury ion thruster system is described. The primary objective of the test is to flight qualify the 5 mN thruster system for auxiliary propulsion applications. A seven year north-south stationkeeping mission was selected as the basis for the flight test operating profile. The flight test, which will employ two thruster systems, will also generate thruster system space performance data, measure thruster-spacecraft interactions, and demonstrate thruster operation in a number of operating modes. The flight test is designated as SAMSO-601 and will be flown aboard the Shuttle-launched Air Force Space Test Program P80-1 satellite in 1981. The spacecraft will be 3-axis stabilized in its final 740 km circular orbit, which will have an inclination of at least 73 degrees. The spacecraft design lifetime is three years.
Power, J. L.↗
Planned flight test of a mercury ion auxiliary propulsion system. 1: Objectives, systems descriptions, and mission operations
A planned flight test of an 8 cm diameter, electron-bombardment mercury ion thruster system is described. The primary objective of the test is to flight qualify the 5 mN (1 mlb.) thruster system for auxiliary propulsion applications. A seven year north-south stationkeeping mission was selected as the basis for the flight test operating profile. The flight test, which will employ two thruster systems, will also generate thruster system space performance data, measure thruster-spacecraft interactions, and demonstrate thruster operation in a number of operating modes. The flight test is designated as SAMSO-601 and will be flown aboard the shuttle-launched Air Force space test program P80-1 satellite in 1981. The spacecraft will be 3- axis stabilized in its final 740 km circular orbit, which will have an inclination of approximately greater than 73 degrees. The spacecraft design lifetime is three years.
Power, J. C.↗
NEXT Long-Duration Test Plume and Wear Characteristics after 16,550 h of Operation and 337 kg of Xenon Processed
The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art. The NEXT ion propulsion system provides improved mission capabilities for future NASA science missions to enhance and enable Discovery, New Frontiers, and Flagship-type NASA missions. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to validate and qualify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the mission-derived throughput requirement of 300 kg. This wear test is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of June 25, 2008, the thruster has accumulated 16,550 h of operation: the first 13,042 h at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. Operation since 13,042 h, i.e., the most recent 3,508 h, has been at an input power of 4.7 kW with 3.52 A beam current and 1180 V beam power supply voltage. The thruster has processed 337 kg of xenon (Xe) surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare. The NEXT LDT has demonstrated a total impulse of 13.3 106 N s; the highest total impulse ever demonstrated by an ion thruster. Thruster plume diagnostics and erosion measurements are obtained periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Observed thruster component erosion rates are consistent with predictions and the thruster service life assessment. There have not been any observed anomalous erosion and all erosion estimates indicate a thruster throughput capability that exceeds ~750 kg of Xe, an equivalent of 36,500 h of continuous operation at the full-power operating condition. This paper presents the erosion measurements and plume diagnostic results for the NEXT LDT to date with emphasis on the change in thruster operating condition and resulting impact on wear characteristics. Ion optics grid-gap data, both cold and operating, are presented. Performance and wear predictions for the LDT throttle profile are presented.
Herman, Daniel A.↗
The High Power Electric Propulsion (HiPEP) Ion Thruster
Practical implementation of the proposed Jupiter Icy Moon Orbiter (JIMO) mission, which would require a total delta V of approximately 38 km/s, will require the development of a high power, high specific impulse propulsion system. Initial analyses show that high power gridded ion thrusters could satisfy JIMO mission requirements. A NASA GRC-led team is developing a large area, high specific impulse, nominally 25 kW ion thruster to satisfy both the performance and the lifetime requirements for this proposed mission. The design philosophy and development status as well as a thruster performance assessment are presented.
Foster, John E.↗
A Status of the Advanced Space Transportation Program from Planning to Action
A Technology Plan for Enabling Commercial Space Business was presented at the 48th International Astronautical Congress in Turin, Italy. This paper presents a status of the program's accomplishments. Technology demonstrations have progressed in each of the four elements of the program; (1) Low Cost Technology, (2) Advanced Reusable Technology, (3) Space Transfer Technology and (4) Space Transportation Research. The Low Cost Technology program element is primarily focused at reducing development and acquisition costs of aerospace hardware using a "design to cost" philosophy with robust margins, adapting commercial manufacturing processes and commercial off-the-shelf hardware. The attributes of this philosophy for small payload launch are being demonstrated at the component, sub-system, and system level. The X-34 "Fastrac" engine has progressed through major component and subsystem demonstrations. A propulsion system test bed has been implemented for system-level demonstration of component and subsystem technologies; including propellant tankage and feedlines, controls, pressurization, and engine systems. Low cost turbopump designs, commercial valves and a controller are demonstrating the potential for a ten-fold reduction in engine and propulsion system costs. The Advanced Reusable Technology program element is focused on increasing life through high strength-to-weight structures and propulsion components, highly integrated propellant tanks, automated checkout and health management and increased propulsion system performance. The validation of rocket based combined cycle (RBCC) propulsion is pro,-,ressing through component and subsystem testing. RBCC propulsion has the potential to provide performance margin over an all rocket system that could result in lower gross liftoff weight, a lower propellant mass fraction or a higher payload mass fraction. The Space Transfer Technology element of the program is pursuing technology that can improve performance and dramatically reduce the propellant and structural mass of orbit transfer and deep space systems. Flight demonstration of ion propulsion is progressing towards launch. Ion propulsion is the primary propulsion for Deep Space 1; a flyby of comet West-kohoutek-lkemura and asteroid 3352 McAuliffe. Testing of critical solar-thermal propulsion subsystems have been accomplished and planning is continuing for the flight demonstration of an electrodynamic tether orbit transfer system. The forth and final element of the program, Space Transportation Research, has progressed in several areas of propulsion research. This element of the program is focused at long-term (25 years) breakthrough concepts that could bring launch costs to a factor of one hundred below today's cost or dramatically expand planetary travel and enable interstellar travel.
Lyles, Garry↗
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.
Sovey, James S.↗
Status of a Power Processor for the Prometheus-1 Electric Propulsion System
NASA is developing technologies for nuclear electric propulsion for proposed deep space missions in support of the Exploration initiative under Project Prometheus. Electrical power produced by the combination of a fission-based power source and a Brayton power conversion and distribution system is used by a high specific impulse ion propulsion system to propel the spaceship. The ion propulsion system include the thruster, power processor and propellant feed system. A power processor technology development effort was initiated under Project Prometheus to develop high performance and lightweight power-processing technologies suitable for the application. This effort faces multiple challenges including developing radiation hardened power modules and converters with very high power capability and efficiency to minimize the impact on the power conversion and distribution system as well as the heat rejection system. This paper documents the design and test results of the first version of the beam supply, the design of a second version of the beam supply and the design and test results of the ancillary supplies.
Pinero, Luis R.↗
Development of a 50,000-s, Lithium-fueled, Gridded Ion Thruster
The ion propulsion system on NASA’s Dawn mission provided over 11 km/s delta-V to the spacecraft. There is potential interest in missions that have delta-V’s an order of magnitude greater than this, i.e., 100 km/s to 200 km/s. To perform such missions would require a thruster that can produce a specific impulse roughly an order of magnitude greater than the 3100 s of the Dawn ion propulsion system. A 50-kW gridded ion thruster is being developed for operation with lithium propellant to produce a specific impulse of 50,000 s. The resulting thruster design requires a net accelerating voltage of 9 kV and a beam current of 5.5 A. Discharge chamber modeling is used to design a 35-cm diameter ring-cusp discharge chamber with six magnet rings. The discharge chamber is masked down to produce an active grid area that is ~25 cm diameter. Modeling suggests that the unique ionization characteristics of lithium may enable discharge chamber operation at a propellant efficiency of 99%. Operation at such a high propellant efficiency could significantly reduce the production of charge-exchange ions and thereby significantly reduce erosion of the accelerator grid.
Goebel, Dan M.↗
Performance of the NASA 30 cm Ion Thruster
A 30 cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for missions of national interest, and is being proposed for use on the USAF/TRW Space Surveillance, Tracking and Autonomous Repositioning (SSTAR) platform to validate ion propulsion. The thruster incorporates innovations in design, materials, and fabrication techniques compared to those employed in conventional ion thrusters. Specific development efforts include thruster design optimizations, component life testing and validation, vibration testing, and performance characterizations. Under this test program, the ion thruster will be brought to engineering model development status. This paper discusses the performance and power throttling test data for the NASA 30 cm diameter xenon ion thruster over an input power envelope of 0.7 to 4.9 kW, and corresponding thruster lifetime expectations.
Patterson, Michael J.↗
Initial assessment of the effects of energetic ion injections in the magnetosphere due to the transport of satellite power system components from low earth orbit to geosynchronous earth orbit
Potentially serious environmental effects exist when cargo orbital transfer vehicle (COTV) ion propulsion is used on the scale proposed in the preliminary definition studies of the Satellite Power System. These effects of the large scale injections of ion propulsion exhaust in the plasmasphere and in the outer magnetosphere were shown to be highly model dependent with major differences existing in the predicted effects of two models, the ion cloud model and the ion sheath model. The expected total number density deposition of the propellant Ar(+) in the plasmasphere, the energy spectra of the deposited Ar(+) and time dependent behavior of the Ar(+) injected into the plasmasphere by a fleet of COTV vehicles differ drastically between the two models. The ion sheath model was demonstrated to be applicable to the proposed Ar(+) beam physics if the beam was divergent and turbulent whereas the ion cloud model was not a realistic approximation for such a beam because the "frozen-field" assumption on which it is based is not valid.
Curtis, S. A.↗
Thermal Development Test of the NEXT PM1 Ion Engine
NASA's Evolutionary Xenon Thruster (NEXT) is a next-generation high-power ion propulsion system under development by NASA as a part of the In-Space Propulsion Technology Program. NEXT is designed for use on robotic exploration missions of the solar system using solar electric power. Potential mission destinations that could benefit from a NEXT Solar Electric Propulsion (SEP) system include inner planets, small bodies, and outer planets and their moons. This range of robotic exploration missions generally calls for ion propulsion systems with deep throttling capability and system input power ranging from 0.6 to 25 kW, as referenced to solar array output at 1 Astronomical Unit (AU). Thermal development testing of the NEXT prototype model 1 (PM1) was conducted at JPL to assist in developing and validating a thruster thermal model and assessing the thermal design margins. NEXT PM1 performance prior to, during and subsequent to thermal testing are presented. Test results are compared to the predicted hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.
Anderson, John R.↗