RF Ion Thruster Research Articles

Influence of anode temperature on ignition performance of the IRIT4-2D iodine-fueled radio frequency ion thruster

This paper reports the ignition performance of the iodine-fueled radio frequency (RF) ion thruster (IRIT) at different anode temperatures (). The experimental results show that the anode temperature plays important role on the ignition process of the IRIT. There were two characteristic temperatures related to the anode: the minimum ignition temperature () and the stable ignition temperature (), which were much lower than the pipeline temperature and the storage tank temperature. At < it failed to discharge. When ≤ < it was ignited with dramatical oscillations. At ≥ the discharge was stable in a large anode temperature range. The analysis showed that the different discharge phenomena at different anode temperatures were attributed to the change of iodine flow rate during the process of the iodine deposition-clogging and sublimation-clearing inside the thruster. The research helps improve the preheating design of the iodine-fueled electric thruster.

Evaluation of Ion Beam Behavior in 50 W Class RF Ion Thruster

Phenomenological behavior of ion beam acceleration through the grid system in 50 W class RF ion thruster has been investigated using PIC simulation and evaluated by experimental test using Faraday probe. Beam trajectory for various grid voltages reveals that the metal engine cover of the ion thruster which is needed to seal RF coil around the discharge chamber affects the beam divergence angle. Simulation result shows that the divergence angle increases by 10.52% mainly because of the larger radial electric field in the presence of the metal engine cover. The divergence angle increases as the accelerator grid voltage increases. The current density distribution measured by the Faraday probe shows a bigger divergence angle with the engine cover installed. For the test cases with mass flow rates from 3 sccm to 4 sccm at the RF power of about 50 W, the current density distribution exhibits the 2nd peak at the radial position about 4 cm from the centerline.

Non-invasive assessment of plasma parameters inside an ion thruster combining optical emission spectroscopy and principal component analysis

We present a non-invasive approach for determining plasma parameters such as electron temperature and density inside a radio-frequency ion thruster (RIT) using optical emission spectroscopy (OES) in conjunction with principal component analysis (PCA). Instead of relying on a theoretical microscopic model of the plasma emission to extract plasma parameters from the OES, an empirical correlation is established on the basis of conducting simultaneous OES and Langmuir diagnostics. The measured reference spectra are simplified and a PCA is performed. The PCA results are correlated with the plasma parameters of the Langmuir measurements yielding a one-to-one correspondence. This correlation allows us to derive the plasma parameters by analysis of a non-invasively determined emission spectrum without additional Langmuir measurements. We show how the plasma parameters can be calculated from OES measurements using this correlation. Under the assumption that the electronic system thermalizes on much shorter time scales than the period of the RF signal driving the thruster, we can also use time-resolved spectral data to determine the time evolution of plasma parameters. In future, this method may contribute to shorter test and qualification times of RITs and other ion thrusters.

Numerical simulation and experimental research of LRIT-30 radio frequency ion thruster

This study aimed at discussing the laws of the design parameters of a radio frequency (RF) ion thruster that influence the thruster performance, guide the performance optimization-oriented design of the thruster, and realize the high-accuracy continuous adjustment of the thrust performance. The key influencing parameters of the thrust performance were analyzed by the numerical simulation method. The influencing laws of RF parameters on key plasma parameters, as well as the thermal characteristics of the thruster under the rated parameters, were explored. Moreover, a LRIT-30 RF ion thruster was developed and subjected to the performance adjustment test. The research results demonstrated that the simulation model can describe the plasma parameter distribution of the discharge chamber and the thermal distribution of key components. The 3 cm RF ion thruster performs well under 2 MHz operational frequency. The RF power and gas flow rate are the key influencing factors of beam extraction and are appropriate for the accurate adjustment of parameters. The wide-range adjustment of thrust (0.5–2.3 mN) and specific impulse (869–2564 s) can be realized when the screen-grid voltage, decelerating screen voltage, RF power, and gas flow rate are 1500 V, −200 V, 40–65 W, and 0.4–0.8 SCCM, respectively. The performance indices of the proposed RF ion thruster are close to the international advanced level.

The impact of neutralizer-free ignition of a radio frequency ion thruster on the lifetime of the ion optics system

In the initial stage of a radio frequency ion thruster (RIT) ignition, an influx of electrons is required from an external source into the discharge chamber and ionization of the neutral gas propellant. A neutralizer-free method for Townsend breakdown discharge ignition based on Paschen’s law was developed in this study. The feasibility of the ignition method was confirmed by performing thousands of ignition experiments. Metallic Molybdenum (Mo), pyrolytic graphite (PG) and Zr[Formula: see text]Ti[Formula: see text]Cu[Formula: see text]Ni[Formula: see text]Be[Formula: see text]alloy acceleration grids were prepared, and ignition-induced damage on the grids was investigated. A field-emission scanning electron microscope was used to inspect surface damage on the grids after multiple ignitions and to analyze the influence of the ignition method on the lifetime of the ion optical system. Grid materials for space missions that require multiple RIT ignitions (10[Formula: see text] should be high-strength blocks that are resistant to sputtering corrosion and high temperature.

Ground performance tests and evaluation of RF ion microthrusters for Taiji-1 satellite

The “Taiji-1” satellite is a test satellite for the verification of those key technologies involved in the “Taiji Program in Space”, China’s space gravitational wave detection project; and the spacecraft drag-free control technology is one of its key technologies used to improve the microgravity level of spacecraft. Thus, a demand for a high-precision and continuously adjustable micronewton-level thrust has been proposed on the spacecraft micro-propulsion system. In allusion to such a task, a set of micronewton-level RF ion propulsion system was designed based on the principle of self-sustained discharge of RF plasma so as to conduct studies on the parameter optimization and engineering of RF ion thruster, and an engineering prototype of the micronewton-level continuously adjustable RF ion thruster was successfully developed to meet the design index requirements. The operating parameters have been solidified through further studies on the extreme operating conditions of the engineering prototype, and a series of ground simulation tests of space environment were successfully passed. The ground test and calibration experiment results of the micro-propulsion engineering prototype show that the engineering prototype has fully met the requirements of the Taiji-1 mission, thus having laid a solid foundation for the successful space verification of the key technologies for the “Taiji-1” satellite.

Sputtering of Mo and Ag with xenon ions from a radio-frequency ion thruster.

The goal of this work is to set up an electric propulsion (EP) sputtering test section as a feasibility study for ground-based sputter testing of spacecraft materials with a radio-frequency ion thruster. Such experiments deliver valuable data, which are scarce but highly desired to model EP-based space missions, for example, with the Spacecraft Plasma Interaction System in order to predict the performance and lifetime of spacecraft components. This study assessed if sufficient testing conditions can be met to produce reliable experimental material data in the future. Therefore, the thruster was operated at ion energies of 1.5 and 1.8 keV, and a quartz crystal microbalance (QCM) was installed to detect sputter deposition rates. Molybdenum (Mo) and silver (Ag) were chosen as sputter targets. Wafer substrates served as a passive sampling method to characterize the composition of sputtered material by Rutherford backscattering spectrometry. Additionally, sputtering simulations matching the experimental conditions were performed with the software SDTrimSP. We obtained comparable experimental and computational data, as measured sputter deposition rates lie within the simulated order of magnitude and to some extent show the predicted angular dependence. Analysis of the deposited sputter material revealed the formation of metal oxides, which requires a future adaption of the material specific QCM settings. Furthermore, the cooling system of the QCM sensor head was not sufficient, limiting the comparability of results.

Radiative Heat Flux from Excited Atoms in a Radiofrequency Ion Thruster

Radiative Heat Flux from Excited Atoms in a Radiofrequency Ion Thruster

Ion thrusters for electric propulsion: Scientific issues developing a niche technology into a game changer.

The transition from old space to new space along with increasing commercialization has a major impact on space flight, in general, and on electric propulsion (EP) by ion thrusters, in particular. Ion thrusters are nowadays used as primary propulsion systems in space. This article describes how these changes related to new space affect various aspects that are important for the development of EP systems. Starting with a historical overview of the development of space flight and of the technology of EP systems, a number of important missions with EP and the underlying technologies are presented. The focus of our discussion is the technology of the radio frequency ion thruster as a prominent member of the gridded ion engine family. Based on this discussion, we give an overview of important research topics such as the search for alternative propellants, the development of reliable neutralizer concepts based on novel insert materials, as well as promising neutralizer-free propulsion concepts. In addition, aspects of thruster modeling and requirements for test facilities are discussed. Furthermore, we address aspects of space electronics with regard to the development of highly efficient electronic components as well as aspects of electromagnetic compatibility and radiation hardness. This article concludes with a presentation of the interaction of EP systems with the spacecraft.

Performance of a 4 cm iodine-fueled radio frequency ion thruster

The performance of an iodine radio ion thruster with a 4 cm diameter (IRIT4) was studied experimentally in this paper. Regulation of the mass flow rates of the iodine propellant is achieved by using a temperature control method of the iodine reservoir. Performance of the thruster using iodine as propellants is obtained at different total thruster powers of 40.6–128.8 W, different grid voltages of 1000–1800 V and the iodine flow rate of 100 μg s−1. Results show that thrust and specific impulse increase approximately linearly with the increasing total thruster power and the screen grid voltage. The thrust of 2.32 mN and the specific impulse of 2361 s are obtained at the nominal total thruster power of 95.8 W and the screen grid voltage of 1800 V. It is also indicated that performance of the iodine propellant is comparable to that of the xenon propellant; and a difference between them is that the iodine thrust is slightly higher than xenon when the total thruster power is more than 62 W. At the nominal 95.8 W total thruster power, the thrust values of them are 2.32 mN and 2.15 mN respectively, and the thrust-to-power ratios of them are 24.2 mN kW−1 and 23.5 mN kW−1, respectively.

Mathematical model of radio-frequency ion thruster with an additional magnetostatic field

The paper presents an engineering simplified mathematical model for estimating integral characteristics of radio-frequency ion thrusters with an additional magnetostatic field. In addition to the integral characteristics calculation, the proposed model allows to obtain the distributions of some plasma local parameters in a gas-discharge chamber of thruster – atoms and charged particles concentrations, and temperature of electrons. In a calculation, axisymmetric geometry of main elements of a thruster structure is considered. Two approaches to taking into account the influence of an additional magnetic field on the distribution of charged particles concentration are implementing. Each of them allows to estimate integral characteristics of radio-frequency ion thrusters with acceptable accuracy. The comparison of the experimental data with simulation results showed that error of thruster’s parameters estimation does not exceed 20-25%, which is sufficient for preliminary designing of new radio-frequency ion thrusters models with an additional source of magnetostatic field.

Experimental Study on the Effects of Discharge Chamber Length on 5 cm Radio-Frequency Ion Thruster

For keeping microgravity level of microgravity and space science satellites, a kind of electric thruster, the radio-frequency ion thruster (RIT) has been designed and applied. The RIT produces thrust through ionizing neutral particles in the discharge chamber and extracting ions by the ion optics system. The length of discharge chamber affects the ionization of neutral particles as well as the performance of RIT. In this paper, an experimental study has been carried out for analyzing the influence of discharge chamber length. In the experiment, the RIT-5 with various discharge chamber lengths, 27.6 mm, 30.0 mm, 31.9 mm, 35.5 mm, 40.0 mm, has different performances under the same condition. With the same radio frequency power and propellant flow rate, the RIT-5 has the optimal performance in beam current and efficiency when the discharge chamber length is 30.0 mm or 31.9 mm.

Numerical mathematical model for calculating ion density in the gas-discharge chamber of a radio-frequency ion thruster

The mathematical model for calculation of the ion density in the gas-discharge chamber of a radio-frequency ion thruster is developed. The model is based on the assumption of higher potential of the central plasma volume comparing to the chamber walls due to higher mobility of electrons. Firstly, the initial approximation for the plasma potential is set and parameters of the ion trajectories in quasistationary electric field are computed. Then, from the assumption of keeping the total ion flow for each trajectory, the ion density is computed as well as the next approximation for plasma potential. The computation is performed iteratively until the change in the ion density is less than the predefined small value. The model was verified by comparing the results of numerical simulation with the analytical solution for spherical volume. The high accuracy of the numerical model is shown. Also, the ion density in the discharge chamber of a radio-frequency ion thruster was calculated with the uniform distribution of neutral atom density and coefficient of ionization rate. It was shown that at such conditions the ion density has a maximum in central part of the discharge chamber and it decreases towards the walls.

Better Prediction of the Performance of a Radio-frequency Ion Thruster

A zero-dimensional analytical model for an RF discharge ion thruster has been improved to predict the thruster performance better. Improvements were made by incorporating most physical phenomena expected in an RF ion thruster, such as secondary electron emission, double ionization, a variable Clausing factor, grid optical transparency, and an ion confinement factor affected by the electromagnetic field. Clausing factors were calculated for each flow rate by using a Monte-Carlo technique. The grid optical transparency was calculated using an ion optics simulation. The ion confinement factor was also calculated for each flow rate by using the calculated magnetic field strength obtained from magnetic field simulations. The ion confinement factor turned out to have the greatest effect on the results while minor corrections were achieved by other processes. Comparison with previously reported analytical solutions and experimental data showed improved prediction of the thruster performance for various size and power ranges.

Global model of a radio-frequency ion thruster based on a holistic treatment of electron and ion density profiles

We present a global model of a radio-frequency ion thruster. The model takes into account radial and axial density distributions for electrons and ions of the plasma inside the ionization vessel. These spatial distributions are based on analytical equations and heuristic assumptions and are used self-consistently in all conservation equations. They are considered in the 3D computation of electromagnetic fields and used to calculate the induced power generated by the coil current. We also consider the spatial ionization and excitation inside the plasma volume in the context of energy and charge conservation. Furthermore, the model includes effects of local charge and power losses on the walls. The extraction grid system is modeled in detail describing each extraction channel separately. The spatial dependence of the electron and ion density profile also leads to a radially varying ion beam current and ion focus across the grid system. Therefore, the parameters of each beamlet differ and need to be described individually by the 3D ion extraction code. An extension of the extraction code also simulates the neutral gas transmission coefficient of the aperture system. This approach enables us to determine the neutral gas density inside the ionization vessel as well as the neutral gas losses. The peripheral electric losses in the coil, the RF cables and the radio-frequency generator are derived by a circuit model.

Combined Chemical–Electric Propulsion for a Stand-Alone Mars CubeSat

Stand-alone interplanetary CubeSats require primary propulsion systems for orbit maneuvering and precise trajectory control. The current work focuses on the design and performance characterization of the combined chemical–electric propulsion systems that shall enable a stand-alone 16U CubeSat mission on a hybrid high-thrust–low-thrust trajectory from a supersynchronous geostationary transfer orbit to a circular orbit about Mars. The high-thrust chemical propulsion is used to escape Earth and to initiate stabilization at Mars. The low-thrust electric propulsion is used in heliocentric transfer, ballistic capture, and circularization. For chemical propulsion, design and performance characteristics of a monopropellant thruster and feed system using ADN-based FLP-106 propellant are presented. For electric propulsion, a performance model of an iodine-propelled inductively coupled miniature radiofrequency ion thruster is implemented to calculate the variation of thrust, specific impulse, and efficiency with input power. A power-constrained low-thrust trajectory optimization using the thruster performance model is pursued to calculate the transfer time, , and the required propellant mass for fuel-optimal and time-optimal transfers. Overall, the combined chemical–electric systems yield a feasible propulsion solution for stand-alone CubeSat missions to Mars that balances propellant mass and transfer time.

Optimization of discharge parameters in an inductive RF ion thruster prototype

This paper presents the results of an experimental study of inductive RF ion thruster prototype parameters, which have demonstrated a strong dependence of the prototype characteristics on the length of the gas-discharge chamber and the operational frequency of the RF generator. The best result was obtained when the prototype was operated at a frequency of 4 MHz with a gas-discharge chamber 5 cm long. The experiments demonstrated that there is a significant resource for improving thruster parameters associated with an increase in the fraction of the RF power absorbed by the plasma. The latter can be achieved by superimposing an external magnetic field on the discharge. In parallel with the results of the experiments, this paper presents numerical calculations based on the model taking into account the presence of a parasitic capacitive channel of the RF power input into the discharge, as well as power losses in the external circuit. Calculations have shown that the experimentally observed effects are associated with the effect of the capacitive RF power input channel, as well as the frequency dependence of the plasma's ability to absorb the RF power.

Direct-current current transformer for the measurement of an electric propulsion ion beam.

An important task in the development of electric propulsion devices is ground-based testing. In order to characterize electric thrusters, their generated ion beams are regarded as figures of merit and have to be well-known and understood. We present an approach to measure DC ion beam currents generated by electric thrusters by means of an inductive direct-current current transformer. In contrast to commonly used devices like Faraday cups, this device can measure the beam current non-intrusively by means of magnetic field coupling rather than charge collection. This paper shows the development of a prototype sensor which can measure currents in the typical range for electric thrusters with high resolution and satisfactory accuracy. A detailed SPICE model to assist during hardware development is also introduced and verified by test case measurements. The prototype has been tested and validated with a radio-frequency ion thruster. Its readout shows very good agreement with the output of an analytical model which computes a Gaussian-shaped ion beam in the far-field based on experimental input data.

The Plasma Ionization Mechanism for a Radio Frequency Ion Thruster with a Flat Coil Inductor

A model of discharge for a radio frequency (RF) ion thruster with a flat induction coil is proposed. The configuration of magnetic fields and currents in 3D geometry is described. The skin layer boundary is determined by the condition at which the azimuthal current in plasma becomes zero under a transformer coupling with the inductor. Variation of radial magnetic field and azimuthal current at the generator driving frequency induces currents at the driving and second harmonic frequencies. Plasma becomes ionized due to magnetized collisional plasma electrons in mobile current layers. With operation at the optimal driving frequency, magnetic field of a definite sign fills the entire gas-discharge space in the course of convective diffusion within the current half-cycle. In the next half-cycle, it is filled with magnetic field of the opposite sign. Ohmic losses in the current layer are compensated in the course of field annihilation. The concentration of particles captured by the mobile layer was determined from the equality of “frosted” electron flows entering and leaving the layer in the course of single coulomb collisions with scattering at the pinch-angle. The azimuthal current of the fundamental harmonic component is determined by mobile layer currents outside the skin layer. The power losses are compensated by magnetic energy. The simulated results are compared with experimental data. The model of physical processes in radio frequency discharge with a flat induction coil has substantial peculiarities as compared with the discharge model using a cylindrical induction coil.

Radio-Frequency Ion Thruster with Additional Magnetic Field: Experimental Investigation

Performances of a radio-frequency ion thruster with an additional external magnetic field in the area of the high-frequency discharge are experimentally investigated and the results of the investigation are presented. The calculated distribution of a magnetic field generated by an additional magnetic system for the current in the winding 1 A is given. How an additional external magnetic field influences on the ion current is examined. Performances of a laboratory model of a radio-frequency ion thruster with and without a constant external magnetic field are compared. Modes of thruster operation and current values in a magnetic system’s windings under which the increment of ion current is minimal are determined.