Gridded Ion Thrusters Research Articles

In situ erosion measurement tools for electric propulsion thrusters: triangular laser head and telemicroscope

The lifetime of electric propulsion (EP) thrusters depends particularly on the erosion characteristics of operation relevant components, for instance, the grid hole erosion of gridded ion thrusters or the channel wall erosion of Hall effect thrusters. Here two tools for in situ erosion measurements are presented, a triangular laser head for surface profiling and a telemicroscope for high-resolution optical imaging. Both can give access to radial and axial erosion parameters. The measurements can be done in situ without the need for breaking the vacuum and dismounting the thruster, which reduces thruster testing time considerably. In situ measurements can also help to ensure reproducibility of thruster performance conditions and can improve statistics of thruster characterization. The present work describes the fundamentals of both techniques in detail, selected experimental setups are presented, their performance is characterized and critically evaluated. The capabilities and limitations related to erosion measurements of EP thrusters are, exemplary, demonstrated for a gridded ion thruster RIT-22 and a Hall effect thruster SPT-100D.

Ion Thruster Discharge Modeling with Adjustment of Coefficient of Anomalous Electron Conductivity

The results of plasma simulations in a discharge chamber of a Direct Current gridded ion thruster are dependent on the choice of anomalous electron conductivity model. Usually, additional electron diffusion is used to reproduce the experimental conductivity level. Up-to-date, the existing numerical models use different values of the additional diffusion coefficient, and this choice is often nonrigorous or insufficiently justified. In this research, the influence of the additional diffusion coefficient on the results of an axisymmetrical Full-Particle in Cell simulation of plasma in the discharge chamber of a Direct Current gridded ion thruster was investigated. The modeled thruster had a rated power of 3 kW, a thrust of 85 mN, and a specific impulse of 4500 s. The magnitude of the additional diffusion was varied from zero to the value of Bohm diffusion . gave the best agreement of the simulation results with the experimental data for the chosen thruster. A qualitative analysis of the plasma parameter distributions inside the discharge chamber at various anomalous electron conductivity coefficients is also provided. The profile of the ion current density through the ion-optical system obtained in the simulation is in sufficient agreement with experimental data.

Evaluation of solid-core thermal antimatter propulsion concepts

Energy released during antimatter annihilation can be used to heat a working fluid to directly produce thrust (an “antimatter rocket”) or run a thermodynamic cycle to generate power for conventional electric propulsion systems (antimatter power generation). For a solid reactor core, we evaluate and compare such antimatter propulsion concepts and assess the resulting performance for different propellants (hydrogen, water, methane, and carbon dioxide) and electric propulsion technologies (arcjets, Hall thrusters, and gridded ion thrusters). In contrast to positron–electron annihilation (which only produces gamma rays), complete stopping of relativistic pions released during antiproton–proton annihilation increases the required reactor core size by approximately an order of magnitude, resulting in significantly larger propulsion systems. Because of reactor core material considerations, the maximum temperature and specific impulse achievable in antimatter rockets is limited. By instead operating the core at a lower temperature to drive a closed thermodynamic Brayton cycle, antimatter power generation can run electric propulsion systems with higher specific impulses. Although this reduces the propellant mass needed for a given mission, the combination of thermodynamic cycle efficiency (around 30%) and electric thruster efficiency (typically between 30%–70%) increases the antimatter mass required by one to two orders of magnitude. The lower thrust of electric propulsion systems also leads to longer burn and maneuver times. Considering emerging industry trends, the use of renewable propellants offers large potential benefits to future antimatter propulsion systems that can refuel in space.

Formation and neutralization of electric charge and current of an ion thruster plume

A 3D hybrid model is introduced and applied to the simulation of the xenon plasma plume extraction, formation, and neutralization in a gridded ion thruster. The acceleration voltage is 1100 V and the inflow Xe+ per hole ranges from 0.07 to 0.92 μg s−1. While ions and neutrals are treated with a particle-in-cell formulation, electrons are modeled as two independent isothermal populations: one inside the discharge chamber and one in the plume. The definition of a thermalized potential allows to solve the electron currents in the high-conductivity limit of the Ohm’s law. The space charge neutralization distance is observed to be short and thus essentially independent of the acceleration grid-neutralizer distance, which is varied from 10 to 25 mm axially. However, this position strongly affects the electric current neutralization paths in the near plume for each ion beamlet. Electron inertial forces are shown to be comparable to collisional forces in certain plasma regions. A semi-analytical 1D fluid model of the plume, matched to the hybrid model, allows to complete the far plume expansion down to infinity. Grids with an infinite and finite number of apertures are simulated and compared with each other and with the 1D model. The numerically obtained divergence angle of the ion plume is compared with experimental measurements, observing relative errors of around 7% in the position of the optimal perveance, and smaller than 4% in the divergence angle average value.

Indirect low-thrust trajectory optimization with gridded ion thruster model

The work deals with indirect optimization of minimum-time and minimum-fuel interplanetary trajectories when gridded ion thruster models are considered. Using an accurate model of solar electric propulsion is beneficial in preliminary mission design, and allows including operational constraints. The maximum thrust and the specific impulse are expressed as a function of thruster input power, which is achieved by means of point-fitting lines that match the performance capabilities of the thrusters. Minimum-time and minimum-fuel problems are formulated to be solved by indirect optimization. In order to increase the accuracy and robustness of the shooting procedure, analytic Jacobians are derived, and a hybrid switching detection technique is used to improve the integration accuracy for minimum-fuel problems. Two examples of Earth-to-Mars transfer and Near-Earth rendezvous mission using the realistic NASA’s Evolutionary Xenon Thruster (NEXT) are given to substantiate the feasibility and effectiveness of the proposed method.

Investigation of variable aperture on the performance and lifetime of ion thruster

Beam flatness is an important parameter that determines the performance and the lifetime of a gridded ion thruster. To improve the beam flatness of the 30 cm (LIPS-300) ion thruster, variable aperture ion optics that adapts to the decreasing ion density as the radius increases is proposed. It is the ion optics that the screen grid surface is divided into several zones, where the aperture diameter in each zone is determined by the ion density and the electron temperature upstream of the screen grid. The beam current density in the central area is artificially reduced. A particle in cell-Monte Carlo collision model is applied in this work to investigating the effect of variable aperture on the perveance and the maximum beam current per aperture by simulating the extraction, focusing and acceleration processes of ions. Taking into account the engineering implementability, the screen grid surface is divided into four zones. The hole diameter in each zone is decreased from 1.95 mm to 1.8 mm, 1.9 mm, 1.8 mm and 1.7 mm, respectively. The simulation results show that the maximum ion density in the center area of grid is decreased by 10.6% and 6.99%, while it is increased by 6.49% and 22.3% in the edge region, respectively. The beam flatness of the variable aperture ion optics is improved from 0.69 to 0.88. The erosion rate is decreased by 31.9%, but the total beam current is also decreased by 7.15%. The simulation results can provide a valuable reference of the development of the ion thruster.

Plasma hysteresis caused by high-voltage breakdown in gridded microwave discharge ion thruster μ10

A cluster of gridded microwave discharge ion thrusters μ10 contributed to the success of the deep space exploration missions of the Japanese asteroid explorers Hayabusa and Hayabusa2. To meet the demands of the next mission DESTINY+, the maximum ion beam current was increased to 200 mA from the 170 mA used in Hayabusa2 by redesigning the magnetic field and grid configuration. However, this thruster exhibited the plasma hysteresis caused by a high-voltage breakdown. Specifically, the ion beam current (i.e. thrust) has two modes at a given propellant flow rate, namely high- and low-current modes. In nominal operation, the thruster operates in high-current mode. However, once a high-voltage breakdown occurs, the ion beam current transitions to and remains in low-current mode, which significantly decreases thrust efficiency. In this paper, based on an investigation of the physical mechanism that causes the hysteresis, we proposed a recovery method to return the thruster operation to high-current mode. In the investigation, we focused on the plasma mode transition, where plasma is generated inside the waveguide. We experimentally investigated the hysteresis by varying the grid configuration and the waveguide geometries while maintaining the magnetic field. The results show that the neutral density and plasma conditions inside the waveguide are strongly related to the hysteresis, which indicates that the main cause of the hysteresis is plasma generation inside the waveguide.

A global model study of argon plasma chemistry used as propellant of a gridded ion thruster

The search for new propulsion devices for corrections and adjustments in the orbits of space systems composed of constellations of satellites in specific arrangements has progressively increased with the growth of requests for data transmission with greater speed and capacity. Electric propulsion is the best solution to reduce the propellant mass, in addition to allowing long mission times. Among the various types of thrusters, electrostatic thrusters have a better-defined physics, making possible the development of computational models for the study and optimization of their operational parameters. A global model was develed for a cylindrical ion thruster with the inductively coupled plasma (ICP) generator originated by the circulation of a radio-frequency current in an inductive coil that surrounds the thruster, having at the output a polarized grid extraction system accelerating the ionized argon gas (propellant). The global model is formed by charged and neutral particle balance equations and energy equations of electrons and gas. Neutral and excited species are accelerated by drift out of the thruster and ionized species accelerated for the reason of the electric field generated by the pair of polarized grids in direct current. The most in-depth study of argon plasma chemistry found that, among the various parameters that interfere with species density, the number of excited species has a direct influence on ionization, although it does not interfere much with ion density, what can be translated into a higher rate ion replacement, which is extremely important to gain thrust and specific impulse.

A global model study of argon plasma chemistry used as propellant of a gridded ion thruster

The search for new propulsion devices for corrections and adjustments in the orbits of space systems composed of constellations of satellites in specific arrangements has progressively increased with the growth of requests for data transmission with greater speed and capacity. Electric propulsion is the best solution to reduce the propellant mass, in addition to allowing long mission times. Among the various types of thrusters, electrostatic thrusters have a better-defined physics, making possible the development of computational models for the study and optimization of their operational parameters. A global model was developed for a cylindrical ion thruster with the inductively coupled plasma (ICP) generator originated by the circulation of a radio-frequency current in an inductive coil that surrounds the thruster, having at the output a polarized grid extraction system accelerating the ionized argon gas (propellant). The global model is formed by charged and neutral particle balance equations and energy equations of electrons and gas. Neutral and excited species are accelerated by drift out of the thruster and ionized species accelerated for the reason of the electric field generated by the pair of polarized grids in direct current. The most in-depth study of argon plasma chemistry found that, among the various parameters that interfere with species density, the number of excited species has a direct influence on ionization, although it does not interfere much with ion density, what can be translated into a higher rate ion replacement, which is extremely important to gain thrust and specific impulse.

Water and xenon ECR ion thruster—comparison in global model and experiment

Gridded ion thrusters are one of the most commonly used types of electric propulsion, and alternative propellants have been studied for miniature ion thrusters to meet the demand of propulsion systems for micro-/nano-satellites. Water is a candidate as an alternative non-pressurized propellant for a CubeSat thruster. It is consistent with the CubeSat concept of short-term and low-cost development. In this paper, the characteristics of a miniature water ion thruster were compared with those of a xenon one using a global model and experiments. The dependence of the performance on the mass flow rate and the input microwave power was examined, and the effects of dissociation and doubly charged ions were directly measured by a quadrupole mass spectrometer. The estimates on the model were compared against experimental results for both propellants, and the performance of the thruster operating on xenon propellant was compared to the performance operating on water propellant. In the comparison between the estimates and the experimental results, the two differences were discussed: the one between water and xenon and the other from the experimental result in both cases. A performance decrease in the propellant utilization efficiency and the specific impulse cannot be avoided when using water as a propellant in an ion thruster. However, the ion production cost did not increase, and it showed the capability of water ion thruster for CubeSat application taking advantage of safety, low cost, non-pressurized system, and human-friendliness of water when used as a propellant.

Survey and Performance Evaluation of Small-Satellite Propulsion Technologies

The growing interest in small satellites (smallsats) is primarily a function of their affordability and versatility across a wide range of space mission applications. For these reasons, smallsats have found valuable applications in government, industry, and academic settings. The continued advancement of smallsats depends on the ability of the aerospace industry to supply affordable, reliable, and efficient miniaturized spacecraft thrusters. Choosing a suitable propulsion system for a smallsat mission involves tradeoffs between performance, cost, and reliability. This study compares the advertised performance of existing chemical, cold gas, and electric propulsion systems across two representative smallsat missions with the goal of providing mission-enabling information to the smallsat research community. Results show that electric propulsion systems are the top performers for both missions. The required wet mass for electrospray thrusters and pulsed plasma devices demonstrated low sensitivity to increasing orbit lifetime, increasing less than 0.5 kg over year increases in orbit lifetime, making them the top-performing systems in a low-impulse long-duration mission. Because of their characteristically high specific impulse, gridded ion thrusters emerged as the top-performing systems in a high-impulse interplanetary mission with delivered mass capability decreasing less than 25 kg for a delta-V increase of 2000 m/s.

Ion beam diagnostic for the assessment of miniaturized electric propulsion systems.

An automated plume diagnostic system is presented for two-dimensional mapping of the ion beam produced by using miniaturized electric propulsion devices. The instrument is designed to operate with noble and reactive gases and consists of a semi-circular probe array with a unique angular distribution following a Gaussian shape for improved beam profile resolution. To demonstrate the operation of the diagnostic system, a series of experiments are conducted with a low-power gridded ion thruster using xenon and iodine propellants. The two-dimensional current density maps generated are used to calculate important beam properties for performance estimates, including the total ion current and the effective divergence angle, and we show how the measurements can be used to assess and optimize the design of ion optics for gridded ion thrusters. All current density measurements with the new instrument are corrected for secondary electron emission (SEE) due to ion bombardment, and using a separate, specially designed, electrostatic probe, we present the first measurements of the SEE yield for iodine ions impacting molybdenum and SAE 304 stainless steel surfaces.

Future Directions for Electric Propulsion Research

The research challenges for electric propulsion technologies are examined in the context of s-curve development cycles. It is shown that the need for research is driven both by the application as well as relative maturity of the technology. For flight qualified systems such as moderately-powered Hall thrusters and gridded ion thrusters, there are open questions related to testing fidelity and predictive modeling. For less developed technologies like large-scale electrospray arrays and pulsed inductive thrusters, the challenges include scalability and realizing theoretical performance. Strategies are discussed to address the challenges of both mature and developed technologies. With the aid of targeted numerical and experimental facility effects studies, the application of data-driven analyses, and the development of advanced power systems, many of these hurdles can be overcome in the near future.

SUPPORTING ADDITIVE MANUFACTURING TECHNOLOGY DEVELOPMENT THROUGH CONSTRAINT MODELLING IN EARLY CONCEPTUAL DESIGN: A SATELLITE PROPULSION CASE STUDY

AbstractFunction and constraints modelling are implemented to design two gridded ion thrusters for additive manufacturing (AM). One concept takes advantage of AM design freedom, disregarding AM limitations and is not feasible. The other concept considers AM limitations and is manufacturable and feasible. Constraints modelling highlights AM capabilities that can be improved, showing where future investment is needed. Constraints representation can also support the creation of technology development roadmaps able to identify areas of AM technologies that must be improved.

Particle-in-Cell simulation of an Ion Gun

In this paper, we intend to study the plasma profiles at the exit of an ion gun. Ion guns are close to the electrostatic ion thrusters. The ions are allowed to escape through an electrostatic potential difference of 100 V. The plasma density is considered to be a variable. A code is employed, based on Particle-in-cell method to study the plasma behavior at the exit of the ion gun. The code is written in 2D (r, z) coordinates. Initially, the ion velocities are sampled out of a Maxwellian distribution. However, the velocity evolves under the electrostatic potential difference with time. Computationally a sparse gridding is assumed. Various plasma profiles such as the density of ions and maximum z-directional velocity are calculated. The maximum z-velocity has been found to be on the order of 20 km/sec.

Experimental study of ion optics with square apertures for high-power ion thrusters.

Large-diameter ion-optical systems are widely used in gridded ion thrusters for missions in near-earth space and for deep-space research. Ion sputtering of the accelerator grid is the main factor that limits the thruster's life. The use of carbon-based materials with a lower sputtering rate compared to metal grids allows the thruster to increase its life more than four times. The most resistant to mechanical loads are grids made of carbon-carbon composite materials. Traditionally, ion optics grids use circular-shaped apertures with the centers located in the nodes of the hexagonal pattern. However, with this arrangement of holes, all carbon fibers in the carbon-carbon composite are cut, and this weakens the mechanical strength of the grid. Previously, the Keldysh Research Centre carried out computational studies of a new form of apertures, in which it is possible to preserve uncut fibers and thereby increase the mechanical strength of grids made of carbon-carbon composites. This paper presents the results of a comparative experimental study of the operational characteristics of ion optics with the traditional round shape of apertures and with the new shape of apertures, which are squares with rounded corners. Measurements of the accelerator grid currents depending on the accelerating voltage have shown that both traditional and new ion-optical systems have close perveance limits in which there is no direct ion impingement on the accelerator grid.

Azimuthal ion drift of a gridded ion thruster

We report the experimental and simulated azimuthal ion velocities of a gridded ion thruster, which generates a roll torque around the thrust axis. Laser-induced fluorescence spectroscopy was applied to two microwave ion thrusters with opposite magnetic polarities. A comparison of the measured results revealed a net misalignment of the grid optics and showed that the ions are continuously accelerated from inside the discharge chamber towards a direction downstream of the grid optics. To investigate the effect of the electromagnetic field, the authors conducted a two-dimensional particle-in-cell Monte Carlo collision (2D-PIC-MCC) numerical simulation. The numerical simulation agrees with the measurements and reveals that the ions are azimuthally accelerated by a gradient B drift, curvature drift, E × B drift and the Lorentz force. The reproduced roll torque is 3.1 ± 2.3 μNm and arises due to the mechanical tolerance of the grid optics. The roll torque shows good agreement with the result observed in the space operation. Therefore, the roll torque can be predicted by using our experiment and simulation.

Azimuthal velocity measurement in the ion beam of a gridded ion thruster using laser-induced fluorescence spectroscopy

This paper presents the first laboratory-based study to measure the azimuthal velocities of ions in the beam of a gridded ion thruster. Through the operation of gridded ion thrusters in space, it has been confirmed that these thrusters cause an unexpected roll torque about the ion beam axis. To reveal the physical mechanism that produces this torque, laser-induced fluorescence spectroscopy has been applied to a microwave ion thruster that was installed in Japanese asteroid probes. This technique can be used to measure the azimuthal velocity by estimating the Doppler shift of the Xe II 5p4()6p 5/2 to Xe II 5p4()6s 3/2 transition at 834.659 nm. The measurement was conducted without a neutralizer cathode to avoid the possibility of the cathode affecting the trajectory of the ion beam. The measured velocity functions are the sum of the spectra of the high velocity beam ions and those of charge exchange ions. By deconvolving these spectra, the azimuthal velocities were successfully measured and were found to range from −700 to 620 m s−1 with an accuracy of ±25%. The measured azimuthal velocity profile was accurately reproduced by the simulated velocity profile obtained using a model, which includes the effects of the maximum possible misalignment of the accelerator grid with respect to the screen grid and the Lorentz force produced by the magnetic field leaked from the discharge chamber. A roll torque of 0.5 ± 0.1 μN m about the thrust axis was calculated from the velocity profile, which is lower than that reported in flight data, but additional mechanisms are suggested to explain this discrepancy.

High Fidelity Multi-Objective Design Optimization of a Downscaled Cusped Field Thruster

The Cusped Field Thruster (CFT) concept has demonstrated significantly improved performance over the Hall Effect Thruster and the Gridded Ion Thruster; however, little is understood about the complexities of the interactions and interdependencies of the geometrical, magnetic and ion beam properties of the thruster. This study applies an advanced design methodology combining a modified power distribution calculation and evolutionary algorithms assisted by surrogate modeling to a multi-objective design optimization for the performance optimization and characterization of the CFT. Optimization is performed for maximization of performance defined by five design parameters (i.e., anode voltage, anode current, mass flow rate, and magnet radii), simultaneously aiming to maximize three objectives; that is, thrust, efficiency and specific impulse. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the three design objectives and additional performance measures. The research indicates that the anode current and the Outer Magnet Radius have the greatest effect on the performance parameters. An optimal value for the anode current is determined, and a trend towards maximizing anode potential and mass flow rate is observed.

Electric propulsion reliability: Statistical analysis of on-orbit anomalies and comparative analysis of electric versus chemical propulsion failure rates

With a few hundred spacecraft launched to date with electric propulsion (EP), it is possible to conduct an epidemiological study of EP's on orbit reliability. The first objective of the present work was to undertake such a study and analyze EP's track record of on orbit anomalies and failures by different covariates. The second objective was to provide a comparative analysis of EP's failure rates with those of chemical propulsion. Satellite operators, manufacturers, and insurers will make reliability- and risk-informed decisions regarding the adoption and promotion of EP on board spacecraft. This work provides evidence-based support for such decisions. After a thorough data collection, 162 EP-equipped satellites launched between January 1997 and December 2015 were included in our dataset for analysis. Several statistical analyses were conducted, at the aggregate level and then with the data stratified by severity of the anomaly, by orbit type, and by EP technology. Mean Time To Anomaly (MTTA) and the distribution of the time to (minor/major) anomaly were investigated, as well as anomaly rates.The important findings in this work include the following: (1) Post-2005, EP's reliability has outperformed that of chemical propulsion; (2) Hall thrusters have robustly outperformed chemical propulsion, and they maintain a small but shrinking reliability advantage over gridded ion engines. Other results were also provided, for example the differentials in MTTA of minor and major anomalies for gridded ion engines and Hall thrusters. It was shown that: (3) Hall thrusters exhibit minor anomalies very early on orbit, which might be indicative of infant anomalies, and thus would benefit from better ground testing and acceptance procedures; (4) Strong evidence exists that EP anomalies (onset and likelihood) and orbit type are dependent, a dependence likely mediated by either the space environment or differences in thrusters duty cycles; (5) Gridded ion thrusters exhibit both infant and wear-out failures, and thus would benefit from a reliability growth program that addresses both these types of problems.