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NASA NTRS1973

An 8-cm electron bombardment thruster for auxiliary propulsion

Thruster size, beam current level, and specific impulse trade-offs are considered for mercury electron bombardment ion thrusters to be used for north-south station keeping of geosynchronous spacecraft. An 8-cm diameter thruster operating at 2750 seconds specific impulse at thrust levels of 4.4 mN (1 m1b) to 8.9 mN (2 m6b) with a design life of 20,000 hours and 10,000 cycles is being developed. The thruster will have a dished two-grid system capable of thrust vectoring of + or - 10 degrees in two orthogonal directions. A preliminary thruster has been fabricated and tested; thruster performance characteristics have been determined at 4.45, 6.68, and 8.90 millinewtons.

Hudson, W. R.↗

NASA NTRS1976

Double ion production in mercury thrusters

The development of a model which predicts doubly charged ion density is discussed. The accuracy of the model is shown to be good for two different thruster sizes and a total of 11 different cases. The model indicates that in most cases more than 80% of the doubly charged ions are produced from singly charged ions. This result can be used to develop a much simpler model which, along with correlations of the average plasma properties, can be used to determine the doubly charged ion density in ion thrusters with acceptable accuracy. Two different techniques which can be used to reduce the doubly charged ion density while maintaining good thruster operation, are identified as a result of an examination of the simple model. First, the electron density can be reduced and the thruster size then increased to maintain the same propellant utilization. Second, at a fixed thruster size, the plasma density, temperature and energy can be reduced and then to maintain a constant propellant utilization the open area of the grids to neutral propellant loss can be reduced through the use of a small hole accelerator grid.

Peters, R. R.↗

NASA NTRS1995

Ion propulsion

An ion engine is a plasma thruster which produces thrust by extracting ions from the plasma and accelerating them to high velocity with an electrostatic field. The ions are then neutralized and leave the engine as high velocity neutral particles. The advantages of ion engines are high specific impulse and efficiency and their ability to operate over a wide range of input powers. In comparison with other electric thrusters, the ion engine has higher efficiency and specific impulse than thermal electric devices such as the arcjet, microwave, radiofrequency and laser heated thrusters and can operate at much lower current levels than the MPD thruster. However, the thrust level for an ion engine may be lower than a thermal electric thruster of the same operating power, consistent with its higher specific impulse, and therefore ion engines are best suited for missions which can tolerate longer duration propulsive phases. The critical issue for the ion engine is lifetime, since the prospective missions may require operation for several thousands of hours. The critical components of the ion engine, with respect to engine lifetime, are the screen and accelerating grid structures. Typically, these are large metal screens that must support a large voltage difference and maintain a small gap between them. Metallic whisker growth, distortion and vibration can lead to arcing, and over a long period of time ion sputtering will erode the grid structures and change their geometry. In order to study the effects of long time operation of the grid structure, we are developing computer codes based on the Particle-In-Cell (PIC) technique and Laser Induced Fluorescence (LIF) diagnostic techniques to study the physical processes which control the performance and lifetime of the grid structures.

Meserole, J. S.↗

NASA NTRS1989

Ion propulsion

An ion engine is a plasma thruster which produces thrust by extracting ions from the plasma and accelerating them to high velocity with an electrostatic field. The ions are then neutralized and leave the engine as high velocity neutral particles. The advantages of ion engines are high specific impulse and efficiency and their ability to operate over a wide range of input powers. In comparison with other electric thrusters, the ion engine has higher efficiency and specific impulse than thermal electric devices such as the arcjet, microwave, radio frequency, and laser heated thrusters and can operate at much lower current levels than the MPD thruster. However, the thrust level for an ion engine may be lower than a thermal electric thruster of the same operating power, consistent with its higher specific impulse, and therefore ion engines are best suited for missions which can tolerate longer duration propulsive phases. The critical issue for the ion engine is lifetime, since the prospective missions may require operation for several thousands of hours. The critical components of the ion engine, with respect to engine lifetime, are the screen and accelerating grid structures. Typically, these are large metal screens that must support a large voltage difference and maintain a small gap between them. Metallic whisker growth, distortion, and vibration can lead to arcing, and over a long period of time ion sputtering will erode the grid structures and change their geometry. In order to study the effects of long time operation of the grid structure, we are developing computer codes based on the particle-in-cell (PIC) technique and laser induced fluorescence (LIF) diagnostic techniques to study the physical processes which control the performance and lifetime of the grid structures.

Meserole, J. S.↗

NASA NTRS1973

Performance optimized, small structurally integrated ion thruster system

A 5-cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. As optimized with a conventional ion extraction system, the system demonstrates a thrust T = 0.47 mlb at a beam voltage of 1600 V, total mass efficiency of 76%, and electrical efficiency of 56%. Under the subject contract effort, no significant performance change was noted for operation with two dimensional electrostatic thrust-vectoring grids. Structural integrity with the vectoring grids was demonstrated for shock (+ or - 30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.2 cm long by 13.4 cm flange bolt circle, with a mass of 9.0 Kg, including 6.8 Kg mercury propellant.

Hyman, J., Jr.↗

NASA NTRS1973

An 8-cm electron bombardment thruster for auxiliary propulsion

Thrustor size, beam current level, and specific impulse tradeoffs are considered for mercury electron bombardment ion thrustors to be used for north-south station keeping of geosynchronous spacecraft. An 8-cm-diam thrustor operating at 2750 sec specific impulse at thrust levels of 4.4 mN (1 mlb) to 8.9 mN (2 mlb) with a design life of 20,000 hours and 10,000 cycles is being developed. The thrustor will have a dished two-grid system capable of thrust vectoring of plus or minus 10 deg in two orthogonal directions. A preliminary thrustor has been fabricated and tested; thrustor performance characteristics have been determined at 4.45, 6.68, and 8.90 millinewtons.

Hudson, W. R.↗

NASA NTRS1973

A 9700-hour durability test of a five centimeter diameter ion thruster

A modified Hughes SIT-5 thrustor has been life-tested at the Lewis Research Center. The final 2700 hours of the test are described with a charted history of thrustor operating parameters and off-normal events. Performance and operating characteristics were nearly constant throughout the test except for neutralizer heater power requirements and accelerator drain current. A post-shutdown inspection revealed sputter erosion of ion chamber components and component flaking of sputtered metal. Several flakes caused beamlet divergence and anomalous grid erosion, causing the test to be terminated. All sputter erosion sources have been identified and promising sputter resistant components are currently being evaluated.

Nakanishi, S.↗

NASA NTRS2004

Experimentally Determined Plasma Parameters in a 30 cm Ion Engine

Single planar Langmuir probes and fiber optic probes are used to concurrently measure the plasma properties and neutral density variation in a 30cm diameter ion engine discharge chamber, from the immediate vicinity of the keeper to the near grid plasma region. The fiber optic probe consists of a collimated optical fiber recessed into a double bore ceramic tube fitted with a stainless steel light-limiting window. The optical fiber probe is used to measure the emission intensity of excited neutral xenon for a small volume of plasma, at various radial and axial locations. The single Langmuir probes, are used to generate current-voltage characteristics at a total of 140 spatial locations inside the discharge chamber. Assuming a maxwellian distribution for the electron population, the Langmuir probe traces provide spatially resolved measurements of plasma potential, electron temperature, and plasma density. Data reduction for the NSTAR TH8 and TH15 throttle points indicates an electron temperature range of 1 to 7.9 eV and an electron density range of 4e10 to le13 cm(sup -3), throughout the discharge chamber, consistent with the results in the literature. Plasma potential estimates, computed from the first derivative of the probe characteristic, indicate potential from 0.5V to 11V above the discharge voltage along the thruster centerline. These values are believed to be excessively high due to the sampling of the primary electron population along the thruster centerline. Relative neutral density profiles are also obtained with a fiber optic probe sampling photon flux from the 823.1 nm excited to ground state transition. Plasma parameter measurements and neutral density profiles will be presented as a function of probe location and engine discharge conditions. A discussion of the measured electron energy distribution function will also be presented, with regards to variation from pure maxwellian. It has been found that there is a distinct primary population found along the thruster centerline, which causes estimates of electron temperature, electron density, and plasma potential, to err on the high side, due this energetic population. Computation of the energy distribution fimction of the plasma clearly indicates the presence of primaries, whose presence become less obvious with radial distance from the main discharge plume.

near grid plasma↗

NASA NTRS2021

Post-test Inspection of NASA’s Evolutionary Xenon Thruster Long Duration Test Hardware: Ion Optics

A Long Duration Test (LDT) was initiated in June 2005 as a part of NASA’s Evolutionary Xenon Thruster (NEXT) service life validation approach. Testing was voluntarily terminated in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse. The post-test inspection objectives for the ion optics were derived from the original NEXT LDT test objectives, such as service life model validation, and expanded to encompass other goals that included verification of in situ measurements, test issue root causes, and past design changes. The ion optics cold grid gap had decreased only by an average of 7% of pretest center grid gap, so efforts to stabilize NEXT grid gap were largely successful. The upstream screen grid surface exhibited a chamfered erosion pattern. Screen grid thicknesses were ≥ 86% of the estimated pretest thickness, indicating that the screen grid has substantial service life remaining. Deposition was found on the screen aperture walls and downstream surfaces that was primarily composed of grid material and back-sputtered carbon, and this deposition likely caused the minor decreases in screen grid ion transparency during the test. Groove depths had eroded through up to 35% of the accelerator grid thickness. Minimum accelerator aperture diameters increased only by about 5-7% of the pretest values and downstream surface diameters increased by about 24-33% of the pretest diameters. These results suggest that increasing the accelerator aperture diameters, improving manufacturing tolerances, and masking down the perforated diameter to 36 cm were successful in reducing the degree of accelerator aperture erosion at larger radii.

George C. Soulas↗

NASA NTRS1976

A 15,000-hour cyclic endurance test of an 8-centimeter-diameter electron bombardment mercury ion thruster

A laboratory model 8-cm thruster with improvements to minimize ion chamber erosion and peeling of sputtered metal was subjected to a cyclic endurance test for 15,040 hours and 460 restarts. A charted history of several thruster operating variables and off-normal events are shown in 600-hour segments at three points in the test. The transient behavior of these variables during a typical start-stop cycle is presented. Performance and operating characteristics were nearly constant throughout the test except for a change in the accelerator back-streaming limit. Findings of the post-test inspection confirmed most of the expected results. Charge-exchange ions caused normal accelerator grid erosion. The workability of the various design features have been substantiated, and attainable improvements in propellant utilization efficiency should significantly reduce accelerator erosion.

Nakanishi, S.↗

NASA NTRS2008

Status of the NEXT Ion Thruster Long-Duration Test After 10,100 hr and 207 kg Demonstrated

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 June 21, 2007, the thruster has accumulated 10,100 hr of operation 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 207 kg of xenon and demonstrated a total impulse of 8.5 106 N-s; the highest total impulse ever demonstrated by an ion thruster in the history of space propulsion. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Overall ion 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, and the decrease in cold grid-gap observed during the NSTAR Extended Life Test have been mitigated. This paper presents the status of the NEXT LDT to date.

Herman, Daniel A.↗

NASA NTRS2015

Electronegative Gas Thruster

The project is an international collaboration and academic partnership to mature an innovative electric propulsion thruster concept to Technology Research Level-3 (TRL-3) through direct thrust measurement. The project includes application assessment of the technology ranging from small spacecraft to high power. The Plasma propulsion with Electronegative GASES(PEGASES) basic proof of concept has been matured to TRL-2 by Ane Aanesland of Laboratoire de Physique des Plasma at Ecole Polytechnique. The concept has advantages through eliminating the neutralizer requirement and should yield longer life and lower cost over conventional gridded ion engines. The objective of this research is to validate the proof of concept through the first direct thrust measurements and mature the concept to TRL-3.

Dankanich, John↗

NASA NTRS1984

Performance capabilities of the 12-centimeter Xenon ion thruster

The 8- and 12-cm mercury ion thruster systems were developed primarily to provide N-S station keeping of satellites with masses up to about 1800 to 3600 kg respectively. The on-orbit propulsion requirements of recently proposed Large Space Systems (LSS) are beyond the thrust capabilities of the baseline 8- and 12-cm thruster systems. This paper presents a characterization of the performance capabilities of the 12-cm Xenon ion thruster to enable an evaluation of its application to LSS auxiliary propulsion requirements. With minor thruster modifications and simplifications the thrust was increased to 64 mN, a factor of six over the baseline 12-cm mercury thruster performance. The thruster was operated over a range of specific impulse of about 2000 to 4000 seconds and at total efficiencies up to 68.0 percent. The operating levels reached in this study were found to be close to the operating limits of the thruster design in terms of perveance, grid breakdown voltage and thruster component temperatures such as those of the magnets and cathode baffle.

Mantenieks, M.↗

NASA NTRS1984

Performance capabilities of the 12-centimeter Xenon ion thruster

Mantenieks, M.↗

NASA NTRS1985

Ring cusp discharge chamber performance optimization

An experimental study of the effects of discharge chamber length and the locations of the anode, cathode and ring cusp within the chamber on the performance of an 8 cm dia. ring cusp thruster is described. As these lengths and positions are varied the changes induced in plasma ion energy costs, extracted ion fractions and ion beam profiles are measured. Results show that the anode may be positioned at any location along an 'optimum virtual anode' magnetic field line and minimum plasma ion energy costs will result. The actual location of this field line is related to a 'virtual cathode' magnetic field line that is defined by the cathode position. The magnetic field has to be such that the virtual anode field line intersects the grids at the outermost ring of grid holes to maximize the extracted ion fraction and flatten the ion beam profile. Discharge chamber lengths that were as small as possible in the test apparatus yielded the lowest extracted ion fractions.

Hiatt, J. M.↗

NASA NTRS2001

Performance Evaluation of Titanium Ion Optics for the NASA 30 cm Ion Thruster

The results of performance tests with titanium ion optics were presented and compared to those of molybdenum ion optics. Both titanium and molybdenum ion optics were initially operated until ion optics performance parameters achieved steady state values. Afterwards, performance characterizations were conducted. This permitted proper performance comparisons of titanium and molybdenum ion optics. Ion optics' performance A,as characterized over a broad thruster input power range of 0.5 to 3.0 kW. All performance parameters for titanium ion optics of achieved steady state values after processing 1200 gm of propellant. Molybdenum ion optics exhibited no burn-in. Impingement-limited total voltages for titanium ion optics where up to 55 V greater than those for molybdenum ion optics. Comparisons of electron backstreaming limits as a function of peak beam current density for molybdenum and titanium ion optics demonstrated that titanium ion optics operated with a higher electron backstreaming limit than molybdenum ion optics for a given peak beam current density. Screen grid ion transparencies for titanium ion optics were as much as 3.8 percent lower than those for molybdenum ion optics. Beam divergence half-angles that enclosed 95 percent of the total beam current for titanium ion optics were within 1 to 3 deg. of those for molybdenum ion optics. All beam divergence thrust correction factors for titanium ion optics were within 1 percent of those with molybdenum ion optics.

Soulas, George C.↗

NASA NTRS2009

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.↗

NASA NTRS1979

Sputtering in mercury ion thrusters

Ground-based tests of Hg ion thrusters have identified sputter erosion of thruster components as one of the main life limiting phenomena. Subsequent measurements have revealed that sputtering rates can be affected by background gases at pressures as low as 10 to the -10th torr. With the recent interest in thin film technology, sputtering in the presence of reactive gases has been studied in great detail. This paper presents the results of many of those studies and applies them to the sputtering of electric thrusters. A model, which assumes that chemisorption is the dominant mechanism, is applied to the sputtering rate measurements of the screen grid of a 30-cm thruster in the presence of nitrogen. The model utilizes inputs from a variety of experimental and analytical sources. The model of environmental effects on sputtering was applied to thruster conditions of low discharge voltage and a discussion of the comparison of theory and experiment is presented.

Mantenieks, M. A.↗