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A fully kinetic 2D axisymmetric Particle-in-Cell (PIC) model is used to examine the effects of background facility pressure on the plasma transport and propulsive efficiency of magnetic nozzles. Simulations are performed for a low-power (150 W class) cathode-less radio-frequency (RF) plasma thruster, operating with xenon, between background pressures up to 1 0 − 2 Pa and average electron discharge temperatures of 4–16 eV. When the electron temperature within the near-plume region reaches 8 eV, a decisive reduction in performance occurs: at 1 0 − 2 Pa, in-plume power losses surpass 25% of the discharge energy flux. Given that the ionisation energy for Xe is 12 eV, the 8 eV threshold indicates that a consistent percentage of electrons has energy enough to trigger ionisation. On the other hand, when the temperature is below such threshold, the primary collisions are charge-exchange and inelastic ion scattering, and the power loss remains less than 10%. It is established that losses in the considered thruster are significant if the facility pressure is greater than 1 0 − 3 Pa, at absorbed powers larger than 130 W. At the nominal 150 W, this results in a 15% thrust reduction. When facility pressure is taken into consideration over ideal vacuum simulations, numerical error is reduced to < 30% when compared to experimental thrust measurements at 1 0 − 3 Pa. • Investigation of facility effects on cathode-less RF thruster performance. • Particle-in-Cell (PIC) simulation of magnetic nozzle expansion into background gas. • Up to 15% thrust reduction at 150 W beyond 1 0 − 3 Pa. • Electron inelastic collisions can cause up to 70% of total in-plume power losses. • Experimental results/setup should therefore take into account facility pressure. [ABSTRACT FROM AUTHOR]

Ignition of combustible air/propane mixture in a high-speed flow with a microwave streamer discharge is considered. The experimental results of structure of streamer discharge and quantities at various pressures are presented. Pressure and temperature measurements in various sections of the streamer discharge in high-speed flow are analysed. The results obtained demonstrate the possibility of ignition of air/propane combustible mixture. The main physical mechanisms responsible for heating the discharge plasma to the ignition temperature of the combustible mixture and their characteristic time scales are discussed. The minimum time and minimum level of microwave radiation required to ignite a combustible mixture under various conditions are determined. • Streamer discharge structure at various pressures in high-speed flow is presented. • Streamer discharge provides reliable ignition in a high-speed flow. • Minimum time and pulse energy required to ignite a mixture are determined. [ABSTRACT FROM AUTHOR]

The study of ignition of flammable mixtures by a microwave discharge is of interest for the design of propulsion systems with increased reliability and possibility of the use of lean fuel/air mixtures. A mathematical model for estimating plasma heating and conversion caused by a microwave subcritical streamer discharge is developed. Pressure of the medium is 13.3 kPa, temperature is 300 K. Plasma is created by microwave discharge with flat mirror and initiator (thin antenna). The initiator is hollow, fuel mixture (propane/air) is pumped through it to the exterior medium. Power of microwave radiation is 1.5 kW. Air flows around the initiator with velocity varying from 11 to 500 m/s. The plasma region and its conductivity are set based on experimental statistics, this is a key feature that reduces the consumption of computing resources. The mathematical model consists of three main stages. At the first stage, the Boltzmann equation for the electron gas in unperturbed medium is solved. At the second stage, the Helmholtz equations and plasma domain are considered for microwave system under study. At the third stage, the Navier–Stokes and transport equations with external heat power and additional plasma reactions for a compressible medium are solved. The results of numerical calculations are compared with the data of a physical experiment. The proposed model gives approximate estimations of discharge parameters and flow quantities, while the requirements for computational resources and time are significantly reduced in comparison to traditional models. • Numerical model of microwave discharge ignition is developed. • The model is validated against experimental data. • Streamer discharge provides reliable ignition in supersonic flow. [ABSTRACT FROM AUTHOR]

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. • Enhanced thrusters have plasma hysteresis caused by a high-voltage breakdown. • The hysteresis is caused by plasma generation and extinction inside the waveguide. • Temporarily gas stop is effective for the recycling after high-voltage breaks down. • The hysteresis depends on the thruster geometry even with the same magnetic field. • Ion optics design is one of the most important parameters for the hysteresis. [ABSTRACT FROM AUTHOR]

A stabler ignition and combustion of the premixed air/fuel mixture remains a topic of great interest due to the high demand for effective and efficient propulsion systems in aerospace applications. Extensive experimental studies and numerical simulations have been carried out to understand and develop ignition systems using cold and non-thermal plasma in microwave streamer discharges. Use of sub-critical streamer discharge allows to ignite air/fuel mixture by an external electric field whose intensity is significantly lower than the intensity required to induce air breakdown. Ignition and combustion of air/propane mixture by microwave sub-critical streamer discharge in low- and high-speed flows is studied experimentally. Qualitative (flame structure) and quantitative (size of combustion region, temperature distributions) characteristics of the flame zone are provided for different flow speeds and equivalence fuel/air ratios, and conditions of discharge initiation with metallic microwave vibrator are studied. The streamer discharge makes it possible to create multiple ignition points providing almost instantaneous ignition of entire volume of the mixture. The results show that the sub-critical streamer discharge can potentially be used in various applications providing more rapid and efficient combustion. • Stable ignition of lean/rich mixtures in air flows is demonstrated • Streamer discharge provides instantaneous volumetric ignition of the mixture. • Effect of fuel/air ratio on combustion efficiency is investigated. • Streamer discharge yields improvement of combustion speed and efficiency. [ABSTRACT FROM AUTHOR]

The authors redesigned the discharge chamber for the μ10 microwave discharge ion thruster to improve its thrust performance and succeeded in enhancing the maximum beam current and thrust efficiency. However, it was found that the ion current ratio extracted from the discharge chamber with the redesigned configuration was lower than that obtained with the original configuration. To investigate the relationship between ion extraction and the magnetic field geometry, the ion loss current distribution in these two types of discharge chamber were measured by electrostatic probes. Using planar probes with a guard ring, the ion current that flowed into the wall was measured without disturbing the ion beam current. The results show that ionization occurs mainly near the upstream magnet. In addition, the ion flux on the sidewall in the redesigned discharge chamber is about 1.5-2 times larger than that in the original discharge chamber. This suggests that the distance between the edge of the plasmaproduction region and the chamber wall with consideration of the Larmor radius of ions is an important parameter in discharge chamber design. In addition, although the ion beam current showed a tendency to saturate at high microwave power, the ion loss to each part in the discharge chamber increased in proportion to input microwave power. The decrease in the extracted ion ratio in the redesigned discharge chamber is considered to be caused by a decrease in the electrostatic ion transparency of the screen grid. Therefore, in a well-tuned microwave discharge ion thruster, it is difficult to improve the thrust efficiency by increasing the discharge power. A design that suppresses the wall loss of ions is thus important. • Ion loss in two types of ECR ion thrusters was measured in each parameter. • Ion beam is saturated as increasing input microwave power. • Ion loss distribution depends on magnetic lines geometry near the wall. • Ions are weakly magnetized and lost on the order of their Larmor radius. • Reducing the loss region by strengthening the magnetic field near the wall is important. [ABSTRACT FROM AUTHOR]

In space operation of the microwave discharge ion thruster μ10 on the asteroid explorers Hayabusa and Hayabusa2, the propellant utilization efficiency deteriorated much more than in the ground endurance test. In this study, a fault tree analysis and experimental simulations of space operation were performed, focusing on the grid-derived internal carbon contamination. It was found that the performance deterioration due to the waveguide contamination matched that in Hayabusa2 qualitatively and quantitatively. Based on the experimental verification, the future performance is experimentally predicted. • Performance deterioration investigation for microwave ion thruster μ10 in space. • Fault tree analysis showed possible cause: grid-derived carbon contamination. • Experiments revealed carbon on waveguide decrease propellant utilization efficiency. • Successful thruster performance reproduction in "Hayabusa2" operation by experiment. • Prediction of future thruter performance based on experimental verification. [ABSTRACT FROM AUTHOR]

Japan's second asteroid explorer Hayabusa2 was successfully launched on Dec 3, 2014, to return a sample from asteroid 162173 Ryugu by 2020. Four xenon ion thrusters based on electron cyclotron resonance discharge propelled the spacecraft for 547 h during its first year in space. Hayabusa2 completed an Earth gravity assist on Dec 3, 2015, followed by 798 and 2593 h of ion thruster operation, called the first and second transfer phases of delta-v, respectively. The third transfer phase of delta-v was conducted from Jan 10, 2018, to Jun 6, 2018, in which the final 2475-h ion thruster operation was executed before the rendezvous with Ryugu. The cumulative operating times for the four ion thrusters are 6,450, 11, 5,193, and 6418 h. This paper summarizes the 6515-h powered flight by the ion engine system, which produced 1015 m/s delta-v, in terms of thruster performance change, roll torques generated by various combinations of ion thrusters, and spacecraft surface erosion history measured by two quartz crystal microbalances located near the thrusters. In parallel with the space flight operation, an engineering model of the microwave discharge neutralizer has been under long-duration testing on the ground since 2012. It has accumulated 55,170 h of diode-mode operation as of Mar 15, 2019. • Hayabusa2, JAXA's asteroid explorer arrived at its destination, asteroid Ryugu. • A cluster of four microwave discharge ion engines worked very well. • The cumulative operation time was 6515 h, producing 1015 m/s delta-v. • An engineering model of the neutralizer has passed 55,170 h of ground operation. [ABSTRACT FROM AUTHOR]

The microwave cathode was developed as a neutralizer for the microwave ion thrusters of the Japanese asteroid explorers Hayabusa and Hayabusa2. Since it emits hundreds of mA of electron current, ion currents collect at the wall of the cathode, which causes fatal destruction due to sputtering. In an effort to reduce the sputtering voltage, this study investigates the effect of the strength of the magnetic field at the nozzle on the anode voltage. Firstly, a magnetic field is applied at the nozzle by a coil. Using the coil, decreasing the magnetic field intensity increases the electron density at the exit of the nozzle. It is presumed that the applied magnetic field facilitates the detachment of magnetic lines by the electrons inside the microwave cathode, resulting in a reduction of the anode voltage. By weakening the nozzle magnetic field, trapped electrons are reduced and the transportability to the outside is improved. Secondly, to realize the same magnetic field intensity achieved in the first experiment without any additional power consumption, the author proposes the use of a magnetic shield. The magnetic shield reduces the anode voltage from 37 V to 32 V at 180 mA, the nominal current of the flight model. Since the sputtering rate exponentially increases with the anode voltage, reducing the anode voltage through these techniques is effective in increasing the lifetime of the cathode. • The electron current of the microwave discharge cathode increased by 44%. • The improvement was achieved by weakening the nozzle magnetic field with a coil. • A similar effect was obtained by a redesigned magnetic circuit without power supply. • Langmuir probe measurements confirmed the increase of the electron number density. [ABSTRACT FROM AUTHOR]

There have been consistent efforts in developing more efficient combustion for propulsion systems. Ignition and combustion control using cold and non-thermal plasma in microwave discharges have become a major topic of interest. In this study, a microwave subcritical streamer discharge is used to initiate ignition and combustion of premixed air/fuel mixture in a long cylindrical tube. The streamer discharge is arising on the internal surface of the dielectric tube using a passive vibrator in a single pulse regime at atmospheric pressure and temperature. The propagation speed of the combustion front in the quartz cylindrical tube filled by the air/propane mixture is analyzed experimentally and numerically. The streamer discharge creating a multitude of ignition points provides practically instantaneous ignition of the mixture in the entire volume. The speed of streamer induced combustion front has been shown to be higher compared to that initiated by a spark. Increasing the length of streamer discharge leads to increasing the flame propagation speed. The combustion efficiency has also been shown to be higher when using the microwave streamer ignition. [ABSTRACT FROM AUTHOR]

Abstract To improve the performance of the 10-cm-class microwave discharge ion thruster μ10 for use in future deep space exploration missions planned by the Japan Aerospace Exploration Agency (JAXA), a new discharge chamber was designed, and its performance was tested. The maximum beam current in the new discharge chamber geometry was 16% higher than that in the original geometry, which was used in the Hayabusa 2 space explorer, under the same discharge power. To investigate the reason for this performance improvement, the multi-charged ion ratio in the plume, the beam current density profiles, and the ion current in the discharge chamber were measured by probes. It was found that the multi-charged ion efficiency and the beam divergence efficiency in the redesigned configuration were not significantly different from those in the Hayabusa 2 configuration. This shows that the increase in the ion beam current enhances the thrust. In addition, it was confirmed that the total ion current inside the new discharge chamber is higher than that in the Hayabusa 2 configuration. The ion extraction efficiency, however, was lower than that in the Hayabusa 2 configuration. This suggests that the increase in the total ion current per unit of incident microwave power is the cause of the performance improvement. In the redesigned configuration, the thrust is 12.0 mN, the specific impulse is 3122 s, the discharge loss is 162 W/A, and the propulsion efficiency is 39.6% at the peak performance point. Highlights • An ion beam current increase of 16% was obtained by redesigning the discharge chamber of the μ10 thruster. • It was verified by E×B probe measurement that the beam current improvement was not due to multi-charged ion current. • It was confirmed by electrostatic probe measurement that total plasma current was increased. • A thrust of 12 mN and a total efficiency of 39.6% have been achieved at the peak performance point. [ABSTRACT FROM AUTHOR]

Abstract Fundamental characteristics of a microwave-discharge miniature ion thruster using water as a propellant, instead of the conventional xenon, are presented in this paper intended for future applications for CubeSats. Water has advantages in terms of safety, handling ability, and availability compared to any other propellant, and these features are especially important for CubeSats. The ion thruster consists of an ion source and a neutralizer, and both use an electron cyclotron resonance heating to generate the plasma. Water was fed to the ion thruster or the neutralizer as a vapor and the direct thrust measurement was conducted. As a result, the thrust correction factor of the ion source was 0.92 ± 0.074 with a beam acceleration voltage of over 0.80 kV, which was similar to the typical values for xenon ion thruster. The ion beam was extracted at the lower mass flow rate, the highest thrust was 164 ± 13.4 μN, and the highest specific impulse was 665 ± 59 s while neglecting the neutralizer. The electron emission current from the neutralizer was 9.5 ± 0.08 mA, which is sufficient to neutralize the ion beam. These experiments demonstrated the basic features of the miniature ion thruster using water and a future possibility of water propellant utilization for CubeSats. Highlights • Miniature ion thruster using water for CubeSats was characterized. • Direct thrust measurement and beam current measurement were conducted. • The highest thrust was 164 ± 13.4 μN with less than 2 W microwave input. • The performance dropped compared with the xenon case. • The lighter mass and lower ionization cross section would cause this drop. [ABSTRACT FROM AUTHOR]

The Japan Aerospace Exploration Agency launched an asteroid sample return spacecraft "Hayabusa2" on December 3, 2014 by the Japanese H2A launch vehicle. Hayabusa2 aims at the round trip mission to the asteroid 162173 Ryugu. Hayabusa2 successfully conducted the Earth gravity assist on December 3, 2015, and now the spacecraft is flying toward Ryugu with the microwave discharge ion engine as the means of propulsion. As of September 2016, 1346 h of the ion engine operation has been achieved as planned. Three touch downs/sample collections, one kinetic impact/crater generation, four surface rovers deployment and many other in-situ observations are planned in the asteroid proximity phase. The operation team will perform extensive operation practice/rehearsal using a hardware-in-the-loop simulator in the year 2017 to be ready for the asteroid arrival in the summer 2018. [ABSTRACT FROM AUTHOR]