Ikkoh Funaki

Influences of electrical conductivity of wall on magnetohydrodynamic control of aerodynamic heating

Influences of the electrical conductivity of the wall of a space vehicle on the control of the aerodynamic heating in Earth-reentry flight by applying the magnetic field are numerically examined using an axisymmetric two-dimensional (r-z) thermochemical nonequilibrium magnetohydrodynamic computational fluid dynamics code. Numerical results show that when the wall of an axisymmetric blunt body is assumed to be an insulating wall, applying a dipole-type magnetic field with r and z components pushes the bow shock wave away from the blunt body and reduces the aerodynamic heating. On the other hand, when the wall is assumed to be a conducting wall, the aerodynamic heating cannot be reduced by applying the magnetic field. This is because the strong Hall electric field on the r-z plane cannot be obtained in the case of the conducting wall, so that the large electric current density in the azimuthal direction cannot be obtained and the shock wave cannot be pushed away from the blunt body.

Two-Dimensional Magnetohydrodynamic Simulation of a Magnetic Sail

A magnetic sail (Magsail) is a unique deep-space propulsion system that captures the momentum of the solar wind by a large artificial magnetic field produced around a spacecraft. To clarify the momentum transfer process from the solar wind to the spacecraft, we simulated the interaction between the solar wind and the artificial magnetic field of the Magsail using the magnetohydrodynamic model. The result showed the same plasma flow and magnetic field as those of the magnetic field of the Earth; when the solar wind passes a bow shock, the solar wind is decelerated and deflected because the solar wind cannot penetrate into the magnetic field, which is called the magnetosphere around the spacecraft. The change of the solar-wind momentum resulted in a pressure distribution along the magnetopause, which is the boundary between the solar-wind plasma and the magnetosphere. The pressure on the magnetopause is then transferred to the spacecraft via the Lorentz force between the induced current along the magnetopause and the current along the coil of the spacecraft. The simulation successfully demonstrated that the change of the momentum of the solar wind is transferred to the spacecraft via the Lorentz force, and the drag coefficient of the Magsail was estimated to be 0.9 ± 0.1 when the magnetic dipole is parallel to the solar wind.

Plasma Characterization of a 10-cm Diameter Microwave Discharge Ion Thruster

Plasma characterization was conducted for an electron-cyclotron-resonance (ECR) type ion thruster. For a 10-cm diameter microwave discharge ion source consisting of two samarium cobalt magnet rings surrounding a centered waveguide for launching microwaves, plasma profiles were found to have severely non-uniform distributions, with localized plasma found near the magnet rings. This localized plasma is mainly produced in the magnetic flux tubes between the two ring magnets, where electrons gain microwave energy as they pass the ECR line during the bouncing movement between magnetic mirrors. To obtain a low-cost microwave ion source, this type of ionization mechanism can be exploited. When introducing microwaves through a low magnetic field boundary, however, it is impossible to eliminate the accessibility difficulty related to the cutoff density, which results in a plasma below the cutoff density. Because of the accessibility difficulty, in this work, only a relatively small ion beam current density of 1.8 mA/cm2 was achieved.

Development of Electrodeless Plasma Thrusters With High-Density Helicon Plasma Sources

Helicon plasma sources are very useful in many aspects and are applicable to many fields across science and technology, as they can supply high-density (~1013 cm-3) plasmas with a broad range of external operating parameters. In this paper, developed, featured sources with various sizes are characterized along with discussions on their particle production efficiency. This paper aims to develop systems that can realize schemes with completely electrodeless plasma production and acceleration. This is expected to mitigate the existing problems of the finite lifetimes inherent in electric plasma propulsion tools. Experimental and theoretical approaches that implement such schemes are presented.

Electrode Configuration Effect on the Performance of a Two-Dimensional Magnetoplasmadynamic Arcjet

Thrust performance and internal plasma flowfield of a 1-MW class self-field magnetoplasmadynamic (MPD) arcjet were measured to evaluate their dependence on the cross-sectional geometry of the electrodes. A multichannel two-dimensional MPD arcjet in quasisteady operation was used to visualize the two-dimensional flowfield and reveal the correlation between the internal flowfield and the thrust performance. The experimental results for six different electrode configurations show that the thrust performance strongly depends on the thruster chamber cross-sectional geometries for the 7sp range of interest, 1000-3000 s. The cathode length determined the engine performance, regardless of the anode geometry. In particular, the convergent-divergent anode with a short cathode showed the best performance. The superior acceleration mechanism of the short cathode was explained on the basis of two-dimension al plasma distributions such as discharge current contours and plasma density obtained by Mach-Zehnder interferometry. A dense plasma region near the tip of the short cathode was observed and subsequent expansion guided by the diverging nozzle can enhance aerodynamic acceleration, which contributes to large thrust generation.

Hybrid Particle-in-Cell Simulations of Magnetic Sail in Laboratory Experiment

Magnetic sail is a propellantless propulsion system proposed for an interplanetary space flight. The propulsive force is produced by the interaction between the magnetic field artificially generated by a hoop coil equipped with the magnetic sail and the solar wind. Three-dimensional hybrid particle-in-cell simulations are performed to elucidate the plasma flow structure around the magnetic sail and to measure the propulsive force of the magnetic sail. We report the characteristics of the magnetosphere, such as the profile of the magnetic field, the thickness of the magnetopause current layer, and the predicted thrust value obtained by simulations, which agree well with laboratory experiment when simulations are carried out by considering the ion-neutral collision effect. The hybrid particle-in-cell simulation carried out without considering the collisional effect gave a thrust value of 3.5 N, which can be applied to the thrust evaluation of the magnetic sail in a magnetosphere with size of 300 km in a collisionless interplanetary space.

Overdense Plasma Production in a Low-power Microwave Discharge Electron Source

Plasma characterization of a low-power microwave discharge electron source was conducted. The electron source, which was developed for the neutralization of the 150 mA-class ion beam exhausted from an ion thruster, consists of a small discharge chamber of 18 mm diameter, into which an L-shape antenna is directly inserted into the magnetic circuit comprised of permanent magnets and iron yokes. An overdense plasma production for the 4.2 GHz microwave was observed for an input power range from 3 to 26 W and for the mass flow rate of 0.5–2.0 sccm. In such a wide range, the plasma density inside the discharge chamber can be proportionally increased as the microwave input power. This is because the direct insertion of the microwave antenna into the ECR magnetic field removes the accessibility difficulty of the microwave, and enables energy transmission from the antenna to the plasma even in the overdense mode. In addition, high-energy electrons above the ionization energy were observed for the large microwave input power above 10 W, and these electrons from the antenna also contribute to plasma production.

Sub-milli-newton class miniature microwave ion thruster

A miniaturized microwave ion source with a 1.6-cm beam diameter grid system was designed and then evaluated experimentally. Based on the HAYABUSA μ10 neutralizer, we fabricated a small 18-nun-diam discharge chamber, into which 4.2 GHz microwaves were launched through an L-shaped antenna that was located in a magnetic field created by permanent magnets and iron yokes. Ion beams were emitted from the small discharge chamber when operated with a grid system whose respective hole diameters of the screen grid and acceleration grid were 0.72 and 0.43 mm, and the total number of grid holes was 211. For a beam voltage of 1500 V and a microwave input power of 10 W, the typical thruster performance was thrust of 0.34 mN, a thrust/power ratio of 16 mN/kW, propellant utilization efficiency of 68%, and a specific impulse of 3200 s. If we were able to further reduce the ion production cost (circa 3000 W/A in the current experiment), this thruster would be a candidate for main propulsion of a small satellite or precise attitude control of various sizes of satellites.

Numerical Analysis of a Two-Dimensional Magnetoplasmadynamic Arcjet

The effect of electrode cone guration on thrust characteristics of a two-dimensional magnetoplasmadynamic (MPD) arcjet was numerically investigated. A simple magnetohydrodynamics (MHD) model was developed and the numerical results were compared with the experimental data for several electrode geometries. To understand the features of the e owe eld, we introduced a magnetosonic Mach number, which is dee ned as local velocity divided by a propagation speed of the MHD disturbance. Based on the magnetosonic Mach number distribution of the e owe eld, the model can explain the thrust characteristics of the MPD arcjet, especially the superiority of a short cathode under various anode cone gurations. Because the electromagnetic thrust is unaltered for the same anode cone guration, the electrothermal component of thrust makes a difference between the long and the short cathodes. With a short cathode cone guration, the large heat deposition near the cathode tip, which is inevitable to MPD arcjets, can be cone ned in the submagnetosonic region where the local e ow is accelerated to magnetosonic velocity. Then the thermal deposition into the submagnetosonic region can be efe ciently recovered through transmagnetosonic acceleration, resulting in a large thrust generation.

Comparison of Simulated Plasma Flow Field in a Two-Dimensional Magnetoplasmadynamic Thruster With Experimental Data

Comprehensive comparisons of the numerically simulated results of plasma flow fields in a 100-kW-class 2-D magnetoplasmadynamic thruster with the available experimental data are conducted. The propellant is argon of 1.25 g/s, and the discharge current is varied from 8 to 12 kA. The physical model includes a nonequilibrium single level of ionization and a collisional radiative model for argon ion to assess the reaction processes in detail. The data we mainly compared are the current path, electron number density, and electron temperature. There is qualitative agreement between the calculated and experimental results except for the electron temperature. In order to explain the disagreement of the electron temperature, we estimate the excitation temperature from the distributions of the excited ions in 4s and 4p states, the radiation of which was employed to determine the electron temperature in the experiment. As a result, it is found that the calculated excitation temperature becomes close to the measured result and that the plasma deviates from the partial local thermodynamic equilibrium near the anode surface. Regarding the thrust and thrust efficiency, their features against variation of the discharge current are well captured by the simulation, although they are slightly overestimated compared with the measured values.