Crandall, P., Wirz, R.E., “Air-breathing electric propulsion: mission characterization and design analysis.,” J Electr Propuls 1, 12 (2022); https://doi.org/10.1007/s44205-022-00009-8

Air breathing electric propulsion (atmosphere-breathing electric propulsion) (ABEP) has attracted significant interest as an enabling technology for long duration space missions in very low Earth orbit (VLEO) altitudes below about 300 km. The ABEP spacecraft and mission analysis model developed allows parametric characterization of key spacecraft geometry and thruster performance parameters such as spacecraft length-to-diameter, the ratio of solar array span to spacecraft diameter, thrust-to-power, effective exhaust velocity, and inlet efficiency. For the missions analyzed ABEP generally outperforms conventional electric propulsion (EP) below 250 km altitude. Using a 6U spacecraft architecture the model shows that below 220 km ABEP is the only viable propulsion option for desirable mission lifetimes. Parametric evaluations of key spacecraft and ABEP characteristics show that the most significant technological improvements to ABEP spacecraft performance and range of applicability for VLEO missions will come from advancements in inlet efficiency, low drag materials, solar array efficiency, and thrust-to-power.

Ottaviano A., Thuppul A., Hayes J., Dodson C., Li G., Chen Z., and Wirz, R.E., “In-situ Microscopy of Ion-Inducted Sputter Erosion of a Featured Surface,” Review of Scientific Instruments, 92, 073701 (2021); https://doi.org/10.1063/5.0043002

A novel method for the in situ visualization and profilometry of a plasma-facing surface is demonstrated using a long-distance microscope. The technique provides valuable in situ monitoring of the microscopic temporal and morphological evolution of a material surface subject to plasma–surface interactions, such as ion-induced sputter erosion. Focus variation of image stacks enables height surface profilometry, which allows a depth of field beyond the limits associated with high magnification. As a demonstration of this capability, the erosion of a volumetrically featured aluminum foam is quantified during ion-bombardment in a low-temperature argon plasma where the electron temperature is ∼7 eV and the plasma is biased relative to the target surface such that ions impinge at ∼300 eV. Three-dimensional height maps are reconstructed from the images captured with a long-distance microscope with an x–y resolution of 3 × 3 μm² and a focus-variation resolution based on the motor step-size of 20 μm. The time-resolved height maps show a total surface recession of 730 μm and significant ligament thinning over the course of 330 min of plasma exposure. This technique can be used for developing plasma-facing components for a wide range of plasma devices for applications such as propulsion, manufacturing, hypersonics, and fusion.

Samples S.A., Wirz R.E., “Parametric Analysis of High Delta-V CubeSat Missions with a Miniature Ion Thruster,” Journal of Spacecraft and Rockets, Vol. 58, No. 3, 2021; https://doi.org/10.2514/1.A34827

The increasing capabilities of the CubeSat platform have led to growing interest in performing more complex commercial and science missions with these miniature spacecraft. In particular, electric propulsion provides unprecedented mission Δ𝑉 and enables ambitious but low-cost Earth missions as well as lunar, asteroid, and interplanetary exploration. As a case study, a 6 U (where U represents a 10×10×10cm “unit”) CubeSat using the Miniature Xenon Ion Thruster was designed for a notional 3000m/s Δ𝑉 mission with a 2 kg, 1.6 U payload, resulting in a spacecraft wet mass of 11.9 kg and a burn duration of 15 months. This spacecraft is capable of up to 5.8km/s Δ𝑉 for a 0.5 U payload, as well as 2.1km/s for a 2 U payload. Parametric analyses with generalized electric thruster properties show that mission performance is sensitive to thruster 𝐼𝑠𝑝 and total efficiency 𝜂𝑇, with a 10% increase in efficiency resulting in a 16% decrease in burn time. This decrease in burn time is also possible by decreasing 𝐼𝑠𝑝, but it incurs mass and payload volume penalties. Parametric studies of neutralizer cathode properties show that neutralizer cathode selection is critical, and that such cathodes should be designed to require low power, low to zero flow rate, and long life for high-Δ𝑉 missions.

Samples S.A., Wirz R.E., "Development of the MiXI Thruster with the ARCH Discharge", Plasma Res. Express, 2020, 2, 025008, https://doi.org/10.1088/2516-1067/ab906d

The Miniature Xenon Ion (MiXI) thruster with the Axial Ring-Cusp Hybrid 'MiXI(ARCH)' discharge was developed and operated with beam extraction at 1 kV. The thruster achieved 59% cathode-free total discharge efficiency at 23.7 mA xenon beam current with filament cathodes and low temperature operation, corresponding to a discharge loss of 226 W/A and propellant utilization of 72%. Thruster efficiency was observed to increase with increasing flow rate and decrease with increasing temperature up to thermal steady state. At thermal steady state, the thruster anode reached ∼320 °C due to the thermal isolation of the thruster head. Reducing the discharge chamber aspect ratio from 0.5 to 0.4 increased thermal steady state efficiency from 46% to 57% but required slow ramping of beam voltage and was limited to stable operation to above 0.5 sccm discharge propellant flow. In contrast to the 3-ring cusp configuration, MiXI(3-Ring), the performance is generally higher but is not able to achieve lower thrust levels and requires more complex start-up for stable operation. An analytical single-cell model was developed and applied to investigate internal processes of the MiXI(ARCH) discharge. The model emulated the effect of increasing flow on performance, indicating that the dominant loss mechanism is plasma electron current to the anode, in contrast to the 3-Ring geometry, which is dominated by primary electron losses. This model also matched trends reported in previous works of strongly increasing electron temperature and primary density with propellant utilization. Through this effort, the MiXI thruster's highest achievable total efficiency has been increased, and several mechanisms for further improved efficiency have been identified.

Dodson C.A., Jorns B.A., Wirz R.E., “Measurements of ion velocity and wave propagation in a hollow cathode plume,” Plasma Source Science and Technology, 28, 065009, 2019, https://iopscience.iop.org/article/10.1088/1361-6595/ab1c48

The mechanism responsible for the production of energetic ions in the plume of hollow cathodes for electric propulsion is still an open issue. These ions are of concern to cathode and thruster lifetime, particularly for cathodes operating at high (>20 A) discharge currents. Recent theoretical and experimental investigations suggest that there is a correlation between ion energy gain and ion acoustic turbulence. In this paper we present measurements of the evolution of the ion velocity distribution function in the near plume of a 100 A-class hollow cathode, operated in a regime in which the dominant mode is ion acoustic turbulence. Ion flow and thermal properties were related to measurements of the background plasma, fluctuation spectra, and dispersion relations obtained from an array of Langmuir probes. We found ions to flow outward from the cathode and accelerate downstream, to supersonic speeds, approximately aligned with the acoustic wave group velocity vectors. The directions of the ion flow and wave propagation were similar for a range of discharge currents and mass flow rates in the jet region of the plume. One operating condition showed a significant temperature increase, also in the direction of acoustic wave propagation, corresponding to the highest wave energy condition. These results are interpreted in the context of ion acoustic turbulence as a contributing mechanism for ion energy gain.

Dodson C.A., Perez-Grande D., Jorns B.A., Goebel D.M., Wirz R.E., "Ion Heating Measurements on the Centerline of a High-Current Hollow Cathode Plume", Journal of Propulsion and Power, 2018, https://doi.org/10.2514/1.B36788

An experimental investigation into the correlation between ion acoustic turbulence (IAT) and anomalous ion heating in the plume of a 100 A-class LaB6 hollow cathode is presented. Laser-induced fluorescence is employed to measure the ion velocity distribution function, and a translating ion saturation probe is used to quantify the spatial dependence of the IAT wave energy. It is found that over a range of flow rates and operating currents both the ion temperature and IAT energy increase downstream of the cathode in qualitatively similar ways. Both parameters also are shown to be impacted by operating conditions: the IAT energy and ion temperature decrease at higher flow rates and lower discharge currents. It is shown that the ratio between ion temperature and wave energy is related by a scaling parameter that depends on the background plasma parameters, and this relation is examined in the context of previous analytical work on IAT-induced ion heating.

Patino M.I., Wirz R.E., "Characterization of Xenon Ion and Neutral Interactions in a Well-Characterized Experiment", Physics of Plasma, Vol. 25, 2018, 062108, https://doi.org/10.1063/1.5030464

Interactions between fast ions and slow neutral atoms are commonly dominated by charge-exchange and momentum-exchange collisions, which are important to understanding and simulating the performance and behavior of many plasma devices. To investigate these interactions, this work developed a simple, well-characterized experiment that accurately measures the behavior of high energy xenon ions incident on a background of xenon neutral atoms. By using well-defined operating conditions and a simple geometry, these results serve as canonical data for the development and validation of plasma models and models of neutral beam sources that need to ensure accurate treatment of angular scattering distributions of charge-exchange and momentum-exchange ions and neutrals. The energies used in this study are relevant for electric propulsion devices ∼1.5 keV and can be used to improve models of ion-neutral interactions in the plume. By comparing these results to both analytical and computational models of ion-neutral interactions, we discovered the importance of (1) accurately treating the differential cross-sections for momentum-exchange and charge-exchange collisions over a large range of neutral background pressures and (2) properly considering commonly overlooked interactions, such as ion-induced electron emission from nearby surfaces and neutral-neutral ionization collisions.

Dankongkakul B., Wirz R.E., "Axial Ring-Cusp Hybrid (ARCH) Plasma Discharge Design: An Approach to Highly Efficient Miniature-Scale Ion Sources", Plasma Sources Science and Technology, 2018, https://iopscience.iop.org/article/10.1088/1361-6595/aae63c

The miniaturization of conventional direct-current ion sources is predominantly restricted by efficiency limitations associated with the increased surface area-to-volume ratio of smaller-scale discharge chambers—reducing the effective confinement length of the high-energy 'primary' electrons that is necessary for efficient plasma generation. The Axial Ring-Cusp Hybrid (ARCH) plasma discharge addresses this scaling limitation by using a new approach that combines magnetic and electrostatic confinement to decouple the primary and plasma electrons loss mechanisms. Simulated ion thruster performance measurements show that the ARCH discharge may be capable of achieving a discharge loss and a propellant mass utilization of 175 eV/ion and 0.87, respectively. These estimates are supported by full internal maps of the plasma properties, including the electron energy distribution function, inside the discharge chamber. The measurements show highly effective confinement of the primary electrons, high average plasma electron temperatures of ∼5 eV, and low plasma sheath potential relative to the anode—attributes generally found only in efficient conventional-scale discharges with good overall plasma confinement. As such, the new ARCH discharge design approach may allow miniature ion thrusters to achieve the performance and efficiency levels similar to those of highly efficient conventional ion thrusters.

Dankongkakul B., Wirz R.E., "Miniature Ion Thruster Ring-Cusp Discharge Performance and Behavior", Journal of Applied Physics, Vol. 122, 023208, 2017, https://doi.org/10.1063/1.4995638

Miniature ion thrusters are an attractive option for a wide range of space missions due to their low power levels and high specific impulse. Thrusters using ring-cusp plasma discharges promise the highest performance, but are still limited by the challenges of efficiently maintaining a plasma discharge at such small scales (typically 1–3 cm diameter). This effort significantly advances the understanding of miniature-scale plasma discharges by comparing the performance and xenon plasma confinement behavior for 3-ring, 4-ring, and 5-ring cusp by using the 3 cm Miniature Xenon Ion thruster as a modifiable platform. By measuring and comparing the plasma and electron energy distribution maps throughout the discharge, we find that miniature ring-cusp plasma behavior is dominated by the high magnetic fields from the cusps; this can lead to high loss rates of high-energy primary electrons to the anode walls. However, the primary electron confinement was shown to considerably improve by imposing an axial magnetic field or by using cathode terminating cusps, which led to increases in the discharge efficiency of up to 50%. Even though these design modifications still present some challenges, they show promise to bypassing what were previously seen as inherent limitations to ring-cusp discharge efficiency at miniature scales.

Araki S.J., Wirz R.E., "Cell-centered Particle Weighting Algorithm for PIC Simulations in a Non-uniform 2D Axisymmetric Mesh", Journal of Computational Physics, Vol. 272, pp. 218–226, Sept. 2014, https://doi.org/10.1016/j.jcp.2014.04.037

Standard area weighting methods for particle-in-cell simulations result in systematic errors on particle densities for a non-uniform mesh in cylindrical coordinates. These errors can be significantly reduced by using weighted cell volumes for density calculations. A detailed description on the corrected volume calculations and cell-centered weighting algorithm in a non-uniform mesh is provided. The simple formulas for the corrected volume can be used for any type of quadrilateral and/or triangular mesh in cylindrical coordinates. Density errors arising from the cell-centered weighting algorithm are computed for radial density profiles of uniform, linearly decreasing, and Bessel function in an adaptive Cartesian mesh and an unstructured mesh. For all the density profiles, it is shown that the weighting algorithm provides a significant improvement for density calculations. However, relatively large density errors may persist at outermost cells for monotonically decreasing density profiles. A further analysis has been performed to investigate the effect of the density errors in potential calculations, and it is shown that the error at the outermost cell does not propagate into the potential solution for the density profiles investigated.

Mao H-S., Wirz R.E., Goebel D.M., "Plasma Structure of Miniature Ring-Cusp Ion Thruster Discharges", Journal of Propulsion and Power, Vol. 3(3), pp. 628-636, June 2014, https://doi.org/10.2514/1.B34759

Previous miniature ion thruster studies have demonstrated impressive performance using ring-cusp discharges. These studies suggest that the magnetic field must be sufficiently strong to increase primary electron confinement times for ionization, but weak enough to allow plasma electrons to escape and maintain the plasma potential necessary for ionization. To investigate these phenomena, an experiment was developed to allow detailed measurements of the internal structure and characteristics of a miniature ring-cusp discharge. These measurements provide spatially resolved values for plasma density, electron temperature, and plasma potential along a meridian plane. The magnetic field configuration is arranged as a quasi-periodic domain in order to generalize the findings to all multipole discharges. The results show that the magnetic field strength drives the plasma structure, and the dependence on discharge power can be removed with proper scaling of the plasma parameters. The stronger magnetic field results in a higher peak plasma density, but relatively low discharge utilization efficiency. In addition, the potential measurements indicate the likely onset of discharge instability. In contrast, the weaker magnetic field, or baseline configuration, better uses the volume of the chamber. This leads to a higher and more uniform density near the downstream end of the discharge where ion extraction would occur, implying superior discharge utilization.

Conversano R.W., Wirz R.E., "Mission Capability Assessment of CubeSats Using a Miniature Ion Thruster", Journal of Spacecraft and Rockets, Vol. 50(5), pp. 1035-1046, April 2013, https://doi.org/10.2514/1.A32435

The successful miniaturization of many spacecraft subsystems make CubeSats attractive candidates for evermore-demanding scientific missions. A three-cell CubeSat employing the miniature xenon ion thruster, which features high efficiency and impulse capability, yields a unique spacecraft that can be optimized for a variety of missions ranging from significant inclination changes in a low Earth orbit to lunar transfers. A nominal configuration of a high-Δ𝑉 CubeSat has a dry mass of approximately 6.3 kg, including a 0.75 kg payload, margins, and contingencies. Depending on the thruster and propellant tank configuration, this CubeSat is capable of delivering mission Δ𝑉 values from 1000 to over 7000m/s, enabling low-Earth-orbit inclination change missions and lunar missions. A parametric analysis on a three-cell high-Δ𝑉 CubeSat bus revealed that a range of payload volumes (up to nearly 1.4 units) and masses (up to nearly 6 kg) can be accommodated depending on the Δ𝑉 requirements and mission type. Additionally, this analysis showed that a high-Δ𝑉 three-cell CubeSat in a 600 km low Earth orbit can be designed to provide an inclination change of over 80 deg.

Araki S.J., Wirz R.E., "Ion–Neutral Collision Modeling Using Classical Scattering with Spin-Orbit Free Interaction Potential", IEEE Transactions on Plasma Science, Vol. 41(3), pp. 470-480, Feb. 2013, https://doi.org/10.1109/TPS.2013.2241457

A particle-in-cell Monte Carlo collision model is developed to explore dominant collisional effects on high-velocity xenon ions incident to a quiescent xenon gas at low neutral pressures. The range of neutral pressure and collisionality examined are applicable for electric propulsion as well as plasma processing devices; therefore, the computational technique described herein can be applied to more complex simulations of those devices. Momentum and resonant charge-exchange collisions between ions and background neutrals are implemented using two different models, classical scattering with spin-orbit free potential and variable-hard-sphere model, depending on the incident particle energy. The primary and charge-exchange ions are tracked separately, and their trajectories within a well-defined “Test Cell” domain are determined. Predicted electrode currents as a function of the Test Cell pressure are compared with electrode currents measured in an ion gun experiment. The simulation results agree well with the experiment up to a Test Cell pressure corresponding to a mean free path of the Test Cell length and then start to deviate with increasing collisionality at higher pressures. This discrepancy at higher pressures is likely due to the increasing influence of secondary electrons emitted from electrodes due to the high-velocity impacts of heavy species (i.e., beam ions and fast neutrals created by charge-exchange interaction) at the electrode surfaces.

Tarzi Z., Speyer J., Wirz R.E., "Fuel Optimum Low-Thrust Elliptic Transfer using Numerical Averaging", Acta Astronautica, Vol. 86, pp. 95-118, June 2013, https://doi.org/10.1016/j.actaastro.2013.01.003

Low-thrust electric propulsion is increasingly being used for spacecraft missions primarily due to its high propellant efficiency. As a result, a simple and fast method for low-thrust trajectory optimization is of great value for preliminary mission planning. However, few low-thrust trajectory tools are appropriate for preliminary mission design studies. The method presented in this paper provides quick and accurate solutions for a wide range of transfers by using numerical orbital averaging to improve solution convergence and include orbital perturbations. Thus, preliminary trajectories can be obtained for transfers which involve many revolutions about the primary body. This method considers minimum fuel transfers using first-order averaging to obtain the fuel optimum rates of change of the equinoctial orbital elements in terms of each other and the Lagrange multipliers. Constraints on thrust and power, as well as minimum periapsis, are implemented and the equations are averaged numerically using a Gausian quadrature. The use of numerical averaging allows for more complex orbital perturbations to be added in the future without great difficulty. The effects of zonal gravity harmonics, solar radiation pressure, and thrust limitations due to shadowing are included in this study. The solution to a transfer which minimizes the square of the thrust magnitude is used as a preliminary guess for the minimum fuel problem, thus allowing for faster convergence to a wider range of problems. Results from this model are shown to provide a reduction in propellant mass required over previous minimum fuel solutions.

Mao H-S., Wirz R.E., "Quasi-equilibrium Electron Density Along a Magnetic Field Line", Applied Physics Letters, Vol. 101(22), 224106, Nov. 2012, https://doi.org/10.1063/1.4768301

A methodology is developed to determine the density of high-energy electrons along a magnetic field line for a low- plasma. This method avoids the expense and statistical noise of traditional particle tracking techniques commonly used for high-energy electrons in bombardment plasma generators. By preserving the magnetic mirror and assuming a mixing timescale, typically the elastic collision frequency with neutrals, a quasi-equilibrium electron distribution can be calculated. Following the transient decay, the analysis shows that both the normalized density and the reduction fraction due to collision converge to a single quasi-equilibrium solution.

Wirz R.E., Anderson J., Katz I., "Time-Dependent Erosion of Ion Optics", Journal of Propulsion and Power, Vol. 27(1), pp. 211-217, Feb. 2011, https://arc.aiaa.org/doi/pdf/10.2514/1.46845

The accurate prediction of ion thruster life requires time-dependent erosion estimates for the ion optics assembly. Such information is critical to end-of-life mechanisms such as electron backstreaming. A two-dimensional ion optics code, CEX2D, was recently modified to handle time-dependent erosion, double ions, and multiple throttle conditions in a single run. The modified code is called CEX2D-T. Comparisons of CEX2D-T results with the NASA solar electric propulsion technology application readiness (NSTAR) thruster life demonstration test and extended life test results show good agreement for both screen and accelerator grid erosion, including important erosion features such as chamfering of the downstream end of the accelerator grid and reduced rate of accelerator grid aperture enlargement with time. The influence of double ions on grid erosion proved to be important for simulating the erosion observed during the NSTAR life demonstration test and extended life test.

Wirz R.E., Katz I., Goebel D., Anderson J., "Electron Backstreaming Determination for Ion Thrusters", Journal of Propulsion and Power, Vol. 27(1), pp. 206-210, Feb. 2011, https://arc.aiaa.org/doi/pdf/10.2514/1.46844

Electron backstreaming in ion thrusters is caused by the random flux of beam electrons past a potential barrier established by the accelerator grid. A technique that integrates this flux over the radial extent of the barrier reveals important aspects of electron backstreaming phenomena for individual beamlets, across the thruster beam, and throughout thruster life. For individual beamlets it was found that over 99% of the electron backstreaming occurs in a small area at the center of the beamlet that is less than 20% the area of the beamlet at the potential barrier established by the accelerator grid. For the thruster beam it was found that over 99% of the backstreaming current occurs inside of r = 6 cm for the over 28 cm diameter NSTAR grid. Initial validation against extended life test data for the NSTAR thruster shows that the technique provides the correct behavior and magnitude of electron backstreaming limit, 𝑉 𝑏 𝑠 V bs ​ . From the sensitivity analyses it is apparent that accelerator grid chamfering due to sputter erosion contributes significantly to the sharp rise in electron backstreaming limit observed in the extended life test, but does not explain the rise in grid ion transparency. Reduction of the grid gap over the life of the thruster also contributes to increases in electron backstreaming limit and increases in ion transparency. Screen grid erosion contributes generally to rises in 𝑉 𝑏 𝑠 V bs ​ and grid ion transparency, but for the assumptions used herein, it appears to not have as much of an effect as chamfering or grid gap change. Overall, it is apparent that accelerator grid chamfering, grid gap change, and screen grid erosion are important to the increase in electron backstreaming observed during the NSTAR extended life test.

Wirz R.E., Anderson J.R., Goebel D.M., Katz I., "Decel Grid Effects on Ion Thruster Grid Erosion", IEEE Transactions on Plasma Science, Vol. 36(5), pp. 2122-2129, Oct. 2008, https://doi.org/10.1109/TPS.2008.2001041

Jet Propulsion Laboratory (JPL) is currently assessing the applicability of the 25-cm Xenon Ion Propulsion System (XIPS) as part of an effort to infuse low-cost technically mature commercial ion thruster systems into NASA deep space missions. Since these mission require extremely long thruster lifetimes to attain the required mission DeltaV, this paper is focused on understanding the dominate wear mechanisms that effect the life of the XIPS three-grid system. Analysis of the XIPS three-grid configuration with JPL's CEX3D grid erosion model shows that the third ldquodecelrdquo grid effectively protects the accel grid from pits and grooves erosion that is commonly seen with two-grid ion thruster grid systems. For a three-grid system, many of the charge-exchange ions created downstream of the grid plane will impact the decel grid at relatively low energies ( ~25 V), instead of impacting the accel grid at high energies ( ~200 V), thus reducing overall erosion. JPL's CEX3D accurately predicts the erosion patterns for the accel grid, although it appears to overpredict the pits and grooves erosion rates due, mainly, to uncertainties in incident energies and angles for sputtering ions and since it does not account for local redeposition of sputtered material. Since the model accurately simulates the erosion pattern but tends to overpredict the erosion rates for both the two- and three-grid sets, this comparative analysis clearly shows how the decel grid significantly suppresses the erosion of the downstream surface of the accel grid as observed in experimental tests. The results also show that the decel grid has a relatively small effect on barrel erosion (erosion of the aperture wall) and erosion of the upstream surface of the accel grid. Decreasing the accel grid voltage of the XIPS can reduce barrel (and total) erosion of the accel grid and should be considered for high-DeltaV missions.

Wirz R.E., Goebel D.M., "Effects of Magnetic Field Topography on Ion Thruster Discharge Performance", Plasma Sources Science and Technology, Vol. 17, 035010, June 2008, https://doi.org/10.1088/0963-0252/17/3/035010

Traditional magnetic field design techniques for dc ion thrusters typically focus on closing a sufficiently high maximum closed magnetic contour, Bcc, inside the discharge chamber. In this study, detailed computational analysis of several modified NSTAR thruster 3-ring and 4-ring magnetic field geometries reveals that the magnetic field line shape as well as Bcc determines important aspects of dc ion thruster performance (i.e. propellant efficiency, beam flatness and double ion content). The DC-ION ion thruster model results show that the baseline NSTAR configuration traps the primary electrons on-axis, which leads to the high on-axis plasma density peak and high double ion content observed in experimental measurements. These problems are further exacerbated by simply increasing Bcc and not changing the field line shape. Changing the field line shape to prevent on-axis confinement (while maintaining the NSTAR baseline Bcc) improves thruster performance, improves plasma uniformity and lowers double ion content. For these favorable field line geometries, we observe further improvements to performance with increased Bcc, while maintaining plasma uniformity and low double ion content. These improvements derive from the fact that the field lines guide the high-energy primaries to regions where they are most efficiently used to create ions while a higher Bcc prevents the loss of ions to the anode walls. Therefore, it is recommended that the ion thruster designer first establish a divergent field line shape that ensures favorable beam flatness, low double ion content and reasonable performance; then the designer may adjust the Bcc to attain desirable performance and stability for the target discharge plasma conditions.

Wirz R.E., Sullivan R., Przybylowski J., Silva M., "Hollow Cathode and Low-Thrust Extraction Grid Analysis for a Miniature Ion Thruster", International Journal of Plasma Science and Engineering, 2008, pp. 1-11, June 2008, http://dx.doi.org/10.1155/2008/693825

Miniature ion thrusters are well suited for future space missions that require high efficiency, precision thrust, and low contamination in the mN to sub-mN range. JPL's miniature xenon Ion (MiXI) thruster has demonstrated an efficient discharge and ion extraction grid assembly using filament cathodes and the internal conduction (IC) cathode. JPL is currently preparing to incorporate a miniature hollow cathode for the MiXI discharge. Computational analyses anticipate that an axially upstream hollow cathode location provides the most favorable performance and beam profile; however, the hot surfaces of the hollow cathode must be sufficiently downstream to avoid demagnetization of the cathode magnet at the back of the chamber, which can significantly reduce discharge performance. MiXI's ion extraction grids are designed to provide >3 mN of thrust; however, previous to this effort, the low-thrust characteristics had not been investigated. Experimental results obtained with the MiXI-II thruster (a near replica or the original MiXI thruster) show that sparse average discharge plasma densities of ∼5×1015–5×1016 m-3 allow the use of very low beamlet focusing extraction voltages of only ∼250–500 V, thus providing thrust levels as low as 0.03 mN for focused beamlet conditions. Consequently, the thrust range thus far demonstrated by MiXI in this and other tests is 0.03–1.54 mN.

Goebel D.M., Polk J.E., Sengupta A., Wirz R.E., "Increasing the Life of a Xenon-Ion Spacecraft Thruster", NASA Tech Brief, Vol. 31(11), NPO-43495, Nov. 2007, https://ntrs.nasa.gov/search.jsp?R=20100011234

A short document summarizes the redesign of a xenon-ion spacecraft thruster to increase its operational lifetime beyond a limit heretofore imposed by nonuniform ion-impact erosion of an accelerator electrode grid. A peak in the ion current density on the centerline of the thruster causes increased erosion in the center of the grid. The ion-current density in the NSTAR thruster that was the subject of this investigation was characterized by peak-to-average ratio of 2:1 and a peak-to-edge ratio of greater than 10:1. The redesign was directed toward distributing the same beam current more evenly over the entire grid andinvolved several modifications of the magnetic- field topography in the thruster to obtain more nearly uniform ionization. The net result of the redesign was to reduce the peak ion current density by nearly a factor of two, thereby halving the peak erosion rate and doubling the life of the thruster.

Goebel D., Wirz R.E., Katz I., "Analytical Ion Thruster Discharge Performance Model", Journal of Propulsion and Power, Vol. 23(5), pp. 1055-1067, Oct. 2007, https://doi.org/10.2514/1.26404

A particle and energy balance model of the plasma discharge in magnetic ring-cusp ion thrusters has been developed. The model follows prior work in the development of global zero-dimensional discharge models that use conservation of particles into and out of the thruster, conservation of energy into the discharge and out of the plasma in the form of charged particles to the walls and beam, and plasma radiation. The present model is significantly expanded over the prior art by closing the set of equations with self-consistent calculations of the internal neutral pressure, electron temperature, primary electron density, electrostatic ion confinement (due to the ring-cusp fields), plasma potential, discharge stability, and time-dependent behavior during recycling. The model only requires information on the thruster geometry, ion optics performance, and electrical inputs, such as discharge voltage and currents, to produce accurate performances curves of discharge loss vs mass utilization efficiency. The model has been benchmarked against several ion thrusters, and successfully predicts the thruster discharge loss as a function of mass utilization efficiency for a variety of thrusters. The discharge performance model will be described and results showing ion thruster performance and stability presented.