Gridded Ion Thrusters Research Articles

Update on Diagnostics for DLR‘s Electric Propulsion Test Facility

DLR has operated since the end of 2011 a dedicated electric propulsion (EP) test facility, the High Vacuum Plume Test Facility Göttingen – Electric Thrusters (STG-ET). This EP test facility has now accumulated several years of thruster testing experience. Special features characterize this facility tailored to electric space propulsion testing. In the first place, a large vacuum chamber mounted on a low vibration foundation characterizes the facility. A performant pumping system ensures a good vacuum, and a target with low sputtering coating dumps the beam power. The vacuum chamber is 12m long and has a diameter of 5m. With respect to thruster tests, the design focus is on accurate thrust measurement, plume diagnostics, and plume interaction with spacecraft components. Electric propulsion thrusters have to run for thousands of hours, and with this the facility is prepared for long-term experiments. The present paper gives an overview of the facility, and describes the vacuum chamber, pumping system, diagnostics, and experiences with these components. After commissioning of STG-ET, the main focus was put on the testing of Hall-effect and gridded ion thrusters.

Advanced Electric Propulsion Diagnostic Tools at IOM

Recently, we have set up an Advanced Electric Propulsion Diagnostic (AEPD) platform [1], which allows for the in-situ measurement of a comprehensive set of thruster performance parameters. The platform utilizes a five-axis-movement system for precise positioning of the thruster with respect to the diagnostic heads. In the first setup (AEPD1) an energy-selective mass spectrometer (ESMS) and a miniaturized Faraday probe for ion beam characterization, a telemicroscope and a triangular laser head for measuring the erosion of mechanical parts, and a pyrometer for surface temperature measurements were integrated. The capabilities of the AEPD1 platform were demonstrated with two electric propulsion thrusters, a gridded ion thruster RIT 22 (Airbus Defence & Space, Germany, [13]) and a Hall effect thruster SPT 100D EM1 (EDB Fakel, Russia, [1,4]), in two different vacuum facilities.

A coupled performance and thermal model for radio-frequency gridded ion thrusters*

Recently proposed space missions such as Darwin, eLISA and NGGM have encouraged the development of electric propulsion thrusters capable of operating in the micro-Newton (μ N) thrust range. To meet these requirements, radio frequency (RF) gridded-ion thrusters need to be scaled down to a few centimeters in size. Due to the small size of these thrusters, it is important to accurately determine the thermal and performance parameters. To achieve this, a multi-physics performance model has been developed, composed of plasma discharge, 2D axisymmetric ion extraction, 3D electromagnetic and RF circuit models. The plasma discharge model itself is represented using 0D global, 2D axisymmetric and 3D molecular neutral gas, and Boltzmann electron transport sub-models. A 3D thermal model is introduced to determine the temperature distribution for various throttle points, using as inputs the plasma and electromagnetic field heating values obtained from the performance model. This also allows the validation of the performance model itself. Additionally, we analyze the effect the thruster’s temperatures play on the plasma properties/performance and vice versa. The model is based on the RIT 3.5 thruster developed for the NGGM mission geometry and predicts the RIT 3.5 experimental data within approximately 10%.

An advanced electric propulsion diagnostic (AEPD) platform for in-situ characterization of electric propulsion thrusters and ion beam sources

Experimental characterization is an essential task in development, qualification and optimization process of electric propulsion thrusters or ion beam sources for material processing, because it can verify that the thruster or ion beam source fulfills the requested mission or application requirements, and it can provide parameters for thruster and plasma modeling. Moreover, there is a need for standardizing electric propulsion thruster diagnostics in order to make characterization results of different thrusters and also from measurements performed in different vacuum facilities reliable and comparable. Therefore, we have developed an advanced electric propulsion diagnostic (AEPD) platform, which allows a comprehensive in-situ characterization of electric propulsion thrusters (or ion beam sources) and could serve as a standard on-ground tool in the future. The AEPD platform uses a five-axis positioning system and provides the option to use diagnostic tools for beam characterization (Faraday probe, retarding potential analyzer, ExB probe, active thermal probe), for optical inspection (telemicroscope, triangular laser head), and for thermal characterization (pyrometer, thermocamera). Here we describe the capabilities of the diagnostic platform and provide first experimental results of the characterization of a gridded ion thruster RIT-μX.

Recommended Practice for Use of Faraday Probes in Electric Propulsion Testing

Faraday probes are a common plasma diagnostic used to determine the local ion charge flux of electric propulsion plumes. Standard practices, guidelines, and recommendations are provided for experimental methods and analysis techniques that aim to standardize community practices, to mitigate test environment effects, and to reduce systematic measurement error in order to improve plume predictions in the space environment. The approaches are applicable to time-averaged plasma properties in the near-field and far-field of electric propulsion plumes, with emphasis on Hall effect thrusters and gridded ion thrusters. Considerations for other electric propulsion technologies are provided, including electrosprays, arcjets, and electromagnetic thruster concepts. These test strategies are expected to increase the quality of comparisons between different thrusters and vacuum environments, thereby broadening the applicability of ground-based measurements and enhancing the fidelity for on-orbit predictions and modeling validation.

Near-field plume properties of an ion beam formed by alternating extraction and acceleration of oppositely charged ions

This paper is devoted to study the expansion of a beam composed of packets of positively and negatively charged ions generated by alternating extraction and acceleration. This beam is extracted from an ion–ion plasma, i.e. the electron density is negligible compared to the negative ion density. The alternating acceleration of ions is ensured by two grids placed in the ion–ion plasma region. The screen grid in contact with the plasma is biased with a square voltage waveform while the acceleration grid is grounded. A two-dimensional particle-in-cell (2D-PIC) code and an analytical model are used to study the properties of the near-field plume downstream of the acceleration grid. It is shown that the possible operating bias frequency is delimited by an upper limit and a lower one that are in the low MHz range. The simulations show that alternating acceleration with bias frequencies close to the upper frequency limit for the system can achieve high ion exhaust velocities, similar to traditional gridded ion thrusters, and with lower beam divergence than in classical systems. Indeed, ion–ion beam envelope might be reduced to 15° with 70% of ion flux contained within an angle of 3°. Thus, this alternating acceleration method is promising for electric space propulsion.

Global model of an iodine gridded plasma thruster

Most state-of-the-art electric space propulsion systems such as gridded and Hall effect thrusters use xenon as the propellant gas. However, xenon is very rare, expensive to produce, and used in a number of competing industrial applications. Alternatives to xenon are currently being investigated, and iodine has emerged as a potential candidate. Its lower cost and larger availability, its solid state at standard temperature and pressure, its low vapour pressure and its low ionization potential make it an attractive option. In this work, we compare the performances of a gridded ion thruster operating separately with iodine and xenon, under otherwise identical conditions using a global model. The thruster discharge properties such as neutral, ion, and electron densities and electron temperature are calculated, as well as the thruster performance parameters such as thrust, specific impulse, and system efficiencies. For similar operating conditions, representative of realistic thrusters, the model predicts similar thrust levels and performances for both iodine and xenon. The thruster efficiency is however slightly higher for iodine compared with xenon, due to its lower ionization potential. This demonstrates that iodine could be a viable alternative propellant for gridded plasma thrusters.

Implementation and verification of a hybrid performance and impedance model of gridded radio-frequency ion thrusters

In this paper, we show the development steps for an iterative performance and impedance model of gridded radio-frequency (RF) ion thrusters. The input parameters are equivalent to those of the real propulsion system; i.e., coil current, propellant mass flow, and extraction grid voltages. Therefore, the model is easily validated and verified by experimental data and can furthermore be used to optimize the overall thruster performance. The model predicts volume-averaged plasma parameters such as electron temperature, conductivity, total pressure, and ionization fraction as well as thruster performance data like generated thrust, specific impulse, and mass and electrical efficiency. The above mentioned plasma parameters are obtained as functions of the discharge chamber’s geometry by using a charge balance equation that relates generated ions to ions lost at the chamber’s walls. The plasma related quantities influence the electromagnetic field penetration which is here evaluated by means of a diffusion equation for the vector potential. The vector potential is obtained by a 3D Finite-Difference-Method on a cubic and rectangular grid which, in principle, offers the opportunity to have arbitrary plasma chamber and coil geometries. An actual ion thruster geometry is evaluated in this study in favor of experimental verification of the numerically obtained data. The thruster’s coil generates highly asymmetric electromagnetic fields which motivates the use of a three-dimensional solver. A dissipation model based on Ohm’s law and Poynting’s theorem is used to determine the absorbed power within the discharge. To obtain a stable solution, the electromagnetically absorbed power is equated to the power lost due to elastic and inelastic collisions and electron wall flux. This whole process is iteratively repeated until the degree of ionization converges within a given threshold. To relate the stable discharge parameters to the thruster performance an extraction model based on a modified version of Child-Langmuir’s law is used. Within the model, the extracted ion beam current is calculated as a function of extraction aperture, applied extraction voltage, and plasma sheath size. A particle balance equation is then used to iteratively compute the total pressure combining neutral, electron, and ion pressures. After convergence, the plasma impedance, typically described by a transformer model, is transformed to its equivalent series representation as seen by the RF source at its terminals. Furthermore, thruster performance data are evaluated which gives rise to use this model for optimization purposes. It can thus be regarded as a toolbox for virtual prototyping.

Effect of Changing Source Capillary Radius on Bulk Flow Parameter Scaling Laws for Hypersonically Expanding Arc-Ablated Polycarbonate Plasma for Fusion and Space Applications

Capillary pulsed electrothermal plasma source coupled to a linear converging–diverging nozzle was previously found to be adequate as a simulator for the supersonic bulk flow patterns of the aerosol particulates generated following a hard disruption event that damages the plasma facing components in a fusion reactor. A similar system can assume an important role in the computational studies related to space electric propulsion and electrostatic ion thrusters. Scaling laws for the prediction of bulk temperature, density, pressure, plasma bulk velocity and Mach number as a function of the axial length traversed measured from the source exit had been proposed previously for an arc-ablated polycarbonate plasma jet evolving into a vacuum chamber via the transition nozzle. In the present work, the source capillary radius has been varied and with the sectional lengths remaining fixed, the converging and diverging angles have also been varied with it, as the diverging exit Mach number was kept constant at a high-hypersonic value ~18. The effect of changing source capillary radius as well as the transition region geometry on the aforesaid bulk flow parameters has been studied computationally and new scaling laws have been proposed.

A collisionless plasma thruster plume expansion model

A two-fluid model of the unmagnetized, collisionless far region expansion of the plasma plume for gridded ion thrusters and Hall effect thrusters is presented. The model is integrated into two semi-analytical solutions valid in the hypersonic case. These solutions are discussed and compared against the results from the (exact) method of characteristics; the relative errors in density and velocity increase slowly axially and radially and are of the order of 10−2–10−3 in the cases studied. The plasma density, ion flux and ambipolar electric field are investigated. A sensitivity analysis of the problem parameters and initial conditions is carried out in order to characterize the far plume divergence angle in the range of interest for space electric propulsion. A qualitative discussion of the physics of the secondary plasma plume is also provided.

The PEGASES Gridded Ion-Ion Thruster Performance and Predictions

The plasma propulsion with electronegative gases (PEGASES) thruster is a gridded ion thruster that accelerates alternately positively and negatively charged ions to provide thrust. Over the last few years, various prototypes have been tested, adequate diagnostics have been developed, and analytical models and simulations are made to better understand and control the physics involved. Here, we present the state-of-the-art in the PEGASES development as of 2013, and discuss its possible future in space.

Coincident ion acceleration and electron extraction for space propulsion using the self-bias formed on a set of RF biased grids bounding a plasma source

We propose an alternative method to accelerate ions in classical gridded ion thrusters and ion sources such that co-extracted electrons from the source may provide beam space charge neutralization. In this way there is no need for an additional electron neutralizer. The method consists of applying RF voltage to a two-grid acceleration system via a blocking capacitor. Due to the unequal effective area of the two grids in contact with the plasma, a dc self-bias is formed, rectifying the applied RF voltage. As a result, ions are continuously accelerated within the grid system while electrons are emitted in brief instants within the RF period when the RF space charge sheath collapses. This paper presents the first experimental results and a proof-of-principle. Experiments are carried out using the Neptune thruster prototype which is a gridded Inductively Coupled Plasma (ICP) source operated at 4 MHz, attached to a larger beam propagation chamber. The RF power supply is used both for the ICP discharge (plasma generation) and powering the acceleration grids via a capacitor for ion acceleration and electron extraction without any dc power supplies. The ion and electron energies, particle flux and densities are measured using retarding field energy analyzers (RFEA), Langmuir probes and a large beam target. The system operates in Argon and N2. The dc self-bias is found to be generated within the gridded extraction system in all the range of operating conditions. Broad quasi-neutral ion-electron beams are measured in the downstream chamber with energies up to 400 eV. The beams from the RF acceleration method are compared with classical dc acceleration with an additional external electron neutralizer. It is found that the two acceleration techniques provide similar performance, but the ion energy distribution function from RF acceleration is broader, while the floating potential of the beam is lower than for the dc accelerated beam.

Initial Performance Evaluation of a Gridded Radio Frequency Ion Thruster

Helicon plasma sources are devices that are capable of efficiently producing high-density plasmas. There is growing interest in using a helicon plasma source in space propulsion as a replacement to the direct current plasma discharge in ion engines. A radio frequency ion engine is developed that combines a helicon plasma source with electrostatic grids and a magnetically shielded anode. Thruster performance evaluation includes estimation of the discharge plasma ion density and electron temperature and measurement of the plume current density, grid currents, and thrust. The radio frequency ion engine is tested across the following operating parameter ranges: 343–600 W radio frequency power, 50–250 G magnetic field strength, argon flow rate, and 100–600 V discharge voltage range at an operating pressure of , discharge ion density and electron temperature range and 5.4–9.9 eV, respectively. Maximum beam current extracted is 120 mA at a 600 V discharge and is primarily limited by ion impingement on the grids due to insufficient grid potentials. The maximum measured thrust is 2.77 mN with an average uncertainty of . Further optimization of the thruster operating conditions and grid assembly can improve thruster performance.

Experimental validation of the dual positive and negative ion beam acceleration in the plasma propulsion with electronegative gases thruster

The PEGASES (Plasma Propulsion with Electronegative Gases) thruster is a gridded ion thruster, where both positive and negative ions are accelerated to generate thrust. In this way, additional downstream neutralization by electrons is redundant. To achieve this, the thruster accelerates alternately positive and negative ions from an ion-ion plasma where the electron density is three orders of magnitude lower than the ion densities. This paper presents a first experimental study of the alternate acceleration in PEGASES, where SF6 is used as the working gas. Various electrostatic probes are used to investigate the source plasma potential and the energy, composition, and current of the extracted beams. We show here that the plasma potential control in such system is key parameter defining success of ion extraction and is sensitive to both parasitic electron current paths in the source region and deposition of sulphur containing dielectric films on the grids. In addition, large oscillations in the ion-ion plasma potential are found in the negative ion extraction phase. The oscillation occurs when the primary plasma approaches the grounded parts of the main core via sub-millimetres technological inputs. By controlling and suppressing the various undesired effects, we achieve perfect ion-ion plasma potential control with stable oscillation-free operation in the range of the available acceleration voltages (±350 V). The measured positive and negative ion currents in the beam are about 10 mA for each component at RF power of 100 W and non-optimized extraction system. Two different energy analyzers with and without magnetic electron suppression system are used to measure and compare the negative and positive ion and electron fluxes formed by the thruster. It is found that at alternate ion-ion extraction the positive and negative ion energy peaks are similar in areas and symmetrical in position with +/− ion energy corresponding to the amplitude of the applied acceleration voltage.

Electrostatic Ion Thrusters ‐ Towards Predictive Modeling

AbstractThe development of electrostatic ion thrusters so far has mainly been based on empirical and qualitative knowhow, and on evolutionary iteration steps. This resulted in considerable effort regarding prototype design, construction and testing and therefore in significant development and qualification costs and high time demands. For future developments it is anticipated to implement simulation tools which allow for quantitative prediction of ion thruster performance, long‐term behavior and space craft interaction prior to hardware design and construction. Based on integrated numerical models combining self‐consistent kinetic plasma models with plasmawall interaction modules a new quality in the description of electrostatic thrusters can be reached. These open the perspective for predictive modeling in this field. This paper reviews the application of a set of predictive numerical modeling tools on an ion thruster model of the HEMP‐T (High Efficiency Multi‐stage Plasma Thruster) type patented by Thales Electron Devices GmbH. (© 2014 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

Ion production cost of a gridded helicon ion thruster

Helicon plasma sources are capable of efficiently ionizing propellants and have been considered for application in electric propulsion. However, studies that estimate the ion production cost of the helicon plasma source are limited and rely on estimates of the extracted ion current. The ion production cost of a helicon plasma source is determined using a gridded ion thruster configuration that allows accurate measurement of the ion beam current. These measurements are used in conjunction with previous characterization of the helicon plasma to create a model of the discharge plasma within the gridded thruster. The device is tested across a range of operating conditions: 343–600 W radio frequency power at 13.56 MHz, 50–250 G and 1.5 mg s−1 of argon at a pressure of 1.6 × 10−5 Torr-Ar. The ion production cost is 132–212 ± 28–46 eV/ion, driven primarily by ion loss to the walls and anode, as well as energy loss in the anode and grid sheaths.

Modeling Neutral Densities Downstream of a Gridded Ion Thruster

Modeling Neutral Densities Downstream of a Gridded Ion Thruster

Modeling Neutral Densities Downstream of a Gridded Ion Thruster

The details of a model for determining the neutral density downstream of a gridded ion thruster are presented. An investigation of the possible sources of neutrals emanating from and surrounding a NEXT ion thruster determined that the most significant contributors to the downstream neutral density include discharge chamber neutrals escaping through the perforated grids, neutrals escaping from the neutralizer, and vacuum facility background neutrals. For the neutral flux through the grids, near- and far-field equations are presented for rigorously determining the neutral density downstream of a cylindrical aperture. These equations are integrated into a spherically-domed convex grid geometry with a hexagonal array of apertures for determining neutral densities downstream of the ion thruster grids. The neutrals escaping from an off-center neutralizer are also modeled assuming diffuse neutral emission from the neutralizer keeper orifice. Finally, the effect of the surrounding vacuum facility neutrals is included and assumed to be constant. The model is used to predict the neutral density downstream of a NEXT ion thruster with and without neutralizer flow and a vacuum facility background pressure. The impacts of past simplifying assumptions for predicting downstream neutral densities are also examined for a NEXT ion thruster.

Very high delta-V missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept

A new and innovative type of gridded ion thruster, the “Dual-Stage 4-Grid” or DS4G concept, has been proposed and its predicted high performance validated under an ESA research, development and test programme. The DS4G concept is able to operate at very high specific impulse and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. This makes it a possible candidate for ambitious missions requiring very high delta-V capability and high power. Such missions include 100 kW-level multi-ton probes based on nuclear and solar electric propulsion (SEP) to distant Kuiper Belt Object and inner Oort cloud objects, and to the Local Interstellar medium. In this paper, the DS4G concept is introduced and its application to this mission class is investigated. Benefits of using the DS4G over conventional thrusters include reduced transfer time and increased payload mass, if suitably advanced lightweight power system technologies are developed. A mission-level optimisation is performed (launch, spacecraft system design and low-thrust trajectory combined) in order to find design solutions with minimum transfer time, maximum scientific payload mass, and to explore the influence of power system specific mass. It is found that the DS4G enables an 8-ton spacecraft with a payload mass of 400 kg, equipped with a 65 kW nuclear reactor with specific mass 25 kg/kW (e.g. Topaz-type with Brayton cycle conversion) to reach 200 AU in 23 years after an Earth escape launch by Ariane 5. In this scenario, the optimum specific impulse for the mission is over 10,000 s, which is well within the capabilities of a single 65 kW DS4G thruster. It is also found that an interstellar probe mission to 200 AU could be accomplished in 25 years using a “medium-term” SEP system with a lightweight 155 kW solar array (2 kg/kW specific mass) and thruster PPU (3.7 kg/kW) and an Earth escape launch on Ariane 5. In this case, the optimum specific impulse is lower at 3500 s which is well within conventional gridded ion thruster capability.

Dormant Cathode Erosion in a Multiple-Cathode Gridded Ion Thruster

DOI: 10.2514/1.37031 A rectangular gridded ion thruster discharge chamber is investigated for operation with multiple discharge cathode assemblies. The multiple-cathode approach attempts to increase thruster throughput and lifetime by operating three discharge cathode assemblies sequentially, possibly providing a threefold increase in discharge chamber life. Previous multiple-cathode electric propulsion devices, such as the SPT-100, have shown dormantcathode erosion to be a life-limiting phenomenon. Similar results in a multiple-cathode discharge chamber may decrease the anticipated gain in discharge lifetime. To assess possible dormant-cathode sputtering erosion and to determine the operational configuration that minimizes this erosion, diagnostic cylinders are designed and used to measure plasma properties at the dormant-cathode locations. Each diagnostic cylinder appears similar to the active discharge cathode assembly, but is outfitted with Langmuir probes. Plasma properties are then used in a simple sputtering-erosion model to predict erosion of the dormant cathodes. Results indicate that the device should be operated at the 0 A electromagnet current configuration for minimum dormant-cathode erosion. For this optimum configuration,typicalnumberdensity,electrontemperature,andplasmapotentialvaluesare5:0 10 11 cm 3 ,5eV, and 27 V with respect to cathode common, respectively. The erosion model indicates that the dormant cathodes will suffer preoperation erosion, but the erosion rate is 26 times slower than the active discharge cathode assembly. Compared with a single-discharge-cathode-assembly thruster, the model predicts an increase in lifetime by a factor of 2.9 for a triple-discharge-cathode-assembly device.