Jonathan L. Van Noord

Lifetime Assessment of the NEXT Ion Thruster

Ion thrusters are low thrust, high specific impulse devices with required operational lifetimes on the order of 10,000-100,000 hours. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hours of operation at the highest throttling point. Currently, a NEXT engineering model ion thruster with prototype model ion optics is undergoing a long duration test to determine wear characteristics and establish propellant throughput capability. The NEXT thruster includes many improvements over previous generations of ion thrusters, but two of its component improvements have a larger effect on thruster lifetime. These include the ion optics with tighter tolerances, a masked region and better gap control, and the discharge cathode keeper material change to graphite. Data from the NEXT 2000 hour wear test, the NEXT long duration test, and further analysis is used to determine the expected lifetime of the NEXT ion thruster. This paper will review the predictions for all of the anticipated failure mechanisms. The mechanisms will include wear of the ion optics and cathode’s orifice plate and keeper from the plasma, depletion of low work function material in each cathode’s insert, and spalling of material in the discharge chamber leading to arcing. Based on the analysis of the NEXT ion thruster, the first failure mode for operation above a specific impulse of 2000 s is expected to be the structural failure of the ion optics at 750 kg of propellant throughput, 1.7 times the qualification requirement. An assessment based on mission analyses for operation below a specific impulse of 2000 s indicates that the NEXT thruster is capable of double the propellant throughput required by these missions.

Application of the NEXT Ion Thruster Lifetime Assessment to Thruster Throttling

Abstract Ion thrusters are low thrust, high specific impulse devices with typical operational lifetimes of 10,000 to 30,000 hr over a range of throttling conditions. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hr of operation at the highest input power throttling point. This paper will provide a brief review the previous life assessment predictions for various throttling conditions. A further assessment will be presented examining the anticipated accelerator grid hole wall erosion and related electron backstreaming limit. The continued assessment of the NEXT ion thruster indicates that the first failure mode across the throttling range is expected to be in excess of 36,000 hr of operation from charge exchange induced groove erosion. It is at this duration that the groove is predicted to penetrate the accelerator grid possibly resulting in structural failure. Based on these lifetime and mission assessments, a throttling approach is presented for the Long Duration Test to demonstrate NEXT thruster lifetime and validate modeling .

NEXT multi-thruster array test -Engineering demonstration

A multi-thruster array test was executed at NASA Glenn Research Center, focusing on the characterization of individual thruster, and array, performance and behavior – as affected by the simultaneous operation of multiple ion thrusters; a key step in development of the NEXT ion propulsion system. The subject of this characterization effort was a four engineering model NEXT thruster array in a 3+1 flight-representative configuration where one thruster was dormant (a spare). This test was executed concurrent with detailed plasma environments and plume measurements documented elsewhere. The array was operated over a broad range of conditions including the simultaneous firing of 3 thrusters at 20.6 kW total input power, yielding a total thrust of about 710 mN, at 4190 seconds specific impulse and approximately 71 percent efficiency. Major findings from a series of tests include: the performance observed for a thruster during operation in an array configuration appears to be consistent with that measured during singular thruster operation with no apparent deleterious interactions; and, operation of 1 neutralizer to neutralize 2-or-more thruster beams appears to be a potentially viable fault-recovery mode, and viable system architecture with significant system performance advantages. Overall, the results indicating single thruster operations are generally independent of array configuration have potentially significant implications with respect to testing requirements and architectural flexibility for multi-thruster systems.

Environmental Testing of the NEXT PM1R Ion Engine

Abstract The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, power processing unit, feed system components, and thruster gimbal. The Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 G

NEXT Ion Thruster Thermal Model

As the NEXT ion thruster progresses towards higher technology readiness, it is necessary to develop the tools that will support its implementation into flight programs. An ion thruster thermal model has been developed for the latest prototype model design to aid in predicting thruster temperatures for various missions. This model is comprised of two parts. The first part predicts the heating from the discharge plasma for various throttling points based on a discharge chamber plasma model. This model shows, as expected, that the internal heating is strongly correlated with the discharge power. Typically, the internal plasma heating increases with beam current and decreases slightly with beam voltage. The second is a model based on a finite difference thermal code used to predict the thruster temperatures. Both parts of the model will be described in this paper. This model has been correlated with a thermal development test on the NEXT Prototype Model 1 thruster with most predicted component temperatures within 5-10 °C of test temperatures. The model indicates that heating, and hence current collection, is not based purely on the footprint of the magnet rings, but follows a 0.1:1:2:1 ratio for the cathode-to-conical-to-cylindrical-tofront magnet rings. This thermal model has also been used to predict the temperatures during the worst case mission profile that is anticipated for the thruster. The model predicts ample thermal margin for all of its components except the external cable harness under the hottest anticipated mission scenario. The external cable harness will be re-rated or replaced to meet the predicted environment.

Results of the 2000 hr Wear Test of the HiPEP Ion Thruster with Pyrolytic Graphite Ion Optics

A 25 kW, long-life ion thruster was developed and wear tested at the NASA Glenn Research Center in support of Project Prometheus. The 2000 hr wear test was undertaken to quantify known erosion phenomena such as ion optics erosion due to charge exchange ion impingement and discharge cathode keeper erosion and to identify unknown wear mechanisms associated with such high-specific impulse, high-power thrusters. The discussion provides a comparison between predicted wear and deposition rates and an analysis of the impact of the various phenomena observed. Trends in observed erosion of the ion optics were consistent with expectations and the negligible wear of the discharge keeper and neutralizer keeper was less than expected. The HiPEP thruster was designed and developed at the NASA Glenn Research Center (GRC) during the ongoing development of the NASA Evolutionary Xenon Thruster (NEXT) 2 and following the successful demonstration of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) ion thruster on the Deep Space 1 spacecraft. 3 Both of these programs incorporated relatively short duration wear tests (~2000 hr) at GRC early in their development efforts. 4,5 These tests were conducted to validate design approaches, identify unknown wear mechanisms, and quantify wear rates before the thrusters were developed to higher fidelity. It is in this context that the HiPEP thruster was incorporated into a wear test very early in its development. While this test was ongoing, a higher fidelity development model thruster (DMT) was being developed in collaboration between GRC and the Aerojet Corporation. An ion optics assembly electrostatically identical to the one being wear tested was successfully vibration tested. 6 A thermo -mechanical model of the DMT was also developed. The results of the 2000 hr wear test would then support and augment a detailed design process and significantly accelerate the delivery of high -fidelity hardware. However, before the assembly of the HiPEP DMT, NASAs Project Prometheus redirected the HiPEP design work to support the design of the Herakles ion thruster, which was in part based on advances made under the HiPEP program. The HiPEP thruster was designed to accommodate a large range of operational requirements and to facilitate the future development of higher -power ion thrusters. To this end, it has rectangular discharge chamber and incorporates pyrolytic graphite (PG) ion optics. HiPEP versions have been successfully operated with both dc (i.e., hollow cathode-based) and microwave discharges. 7,8 Operation at 25 kW over a specific impulse range of 6000 to 9000 s using a dc discharge was demonstrated. 8 Following the performance demonstration of the HiPEP thruster with PG ion optics and a dc discharge, 9 the thruster entered a 2000 hr wear test. The objectives of the test were to

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 missio n 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 expl oration 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 predict ed hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.

Environmental Testing of NASA's Evolutionary Xenon Thruster Prototype Model 1 Reworked Ion Engine

NASA’s Evolutionary Xenon Thruster propulsion system is an advanced ion propulsion system that is oriented toward robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, a power processing unit, feed system components, and a thruster gimbal. The prototype model engine, PM1, was subjected to qualification-level environmental testing to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified, including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration, which led to abbreviated thermal testing, and the generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings and renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10:0 grms for 2min in each of the three axes. Thermalvacuum testing included three thermal cycles from 120 to 215 C with hot engine restarts. Thruster performance was nominal throughout the test program, with minor variations in a few engine operating parameters likely caused by facility effects. There were no significant changes in engine performance as characterized by the engine operating parameters, ion optics performance measurements, and beam current density measurements, indicating no significant changes to the hardware as a result of the environmental testing. The redesigned cathode heater terminations successfully survived the vibration environments. NASA’s EvolutionaryXenonThruster PM1Rengine and the breadboard gimbal were found to be well designed tomeet environmental requirements based on the results reported herein.

A Facility and Ion Thruster Back S putter Survey for Higher Power Ion Thrusters

One of the significant difficulties in ground testing ion thrusters is the back sputtering of material from the vacuum facility walls resulting from beam impingement. This effect becomes even more pronounced with the development of higher power thrusters. Two major effects result from the backsputtered material. The first is a mask ing of wear rates present on the optics. A second effect is the accumulation of carbon on surfaces that can lead to spalling of flakes that could cause an electrical short. This paper evaluates how the amount of backsputtering changes for different facil ities and different ion thrusters. Results from a model are presented along with some comparison to existing data.

Annular-Geometry Ion Engine: Concept, Development Status, and Preliminary Performance

This publication discusses the concept, projected capabilities, technology development plan, and preliminary performance data obtained for an Annular-Geometry Ion Engine (AGI-Engine). The AGI-Engine is the basis for a new class of Next-Generation Electric Propulsion Thrusters under investigation at NASA Glenn Research Center. The AGI-Engine holds the promise of achieving substantial increases in input power (>10X) and power density (2-3X) relative to conventional ion thrusters at specific impulse values of interest for near-term mission applications.