船木 一幸

The magneto-plasma sail (mini-magnetospheric plasma propulsion) produces the propulsive force due to the interaction between the artificial magnetic field around the spacecraft inflated by the plasma and the solar wind erupted from the Sun with a speed of 300-800 km/s. The principle of the magneto-plasma sail is based on the magnetic sail whose original concept requires a huge mechanical coil structure, which produces a large magnetic field to capture the energy of the solar wind. Meanwhile in the case of the magneto-plasma sail, the magnetic field will be expanded by the inertia of plasma flow to a few tens of km in diameter, resulting in a thrust of a few N. R.Winglee's group of the University of Washington originally proposed the idea of magnetic field inflation by the plasma. This paper investigates the characteristics of the magneto-plasma sail by comparing it with the other low-thrust propulsion systems (i.e., electric propulsion and solar sail), and the potential of its application to near future outer planet missions is studied. Furthermore, an engineering validation satellite concept is proposed in order to confirm the propulsion system specification and operation methodology. The main features are summarized as: The satellite mass is around 180kg assuming the H-IIA piggyback launch. 2) Since the magnetopause of the Earth magnetosphere is about 10Re at Sun side and the bow shock is located at about 13Re from the Earth, the satellite is injected into an orbit with 250km perigee altitude and 20 Re apogee distance where apogee is located at the Sun side. 3) The magneto-plasma sail is turned on only in the vicinity of apogee outside the Earth's magnetosphere. 4) The thrust is estimated by the orbit determination result, and the plasma wind monitor is installed on the satellite to establish the relationship between the solar wind and the thrust.

An ion beam optics for a 10-cm-diam 400-W-class microwave discharge ion thruster was fabricated and its applicability to a long-term space mission was demonstrated. The optics consists of three 1-mm-thick flat carbon-carbon composite panels with approximately 800 holes that were mechanically drilled and positioned with +/-0.02-mm accuracy. When mounted on an aluminum ring, spacing for the three grids was kept at 0.5 mm by three sets of spacers. The thruster produced an ion beam current of 140 mA with a microwave power of 32 W for plasma generation and a total acceleration voltage of 1.8 kV. Although the grid is sputtered by the impingement of slow ions produced in charge-exchange collisions between fast beam ions and neutral atoms leaking from the engine, the grid showed only slight damage even after an 18,000-h endurance test. Also, other qualification tests including a mechanical test under launch conditions as well as a thermal vacuum test simulating the spacecraft thermal environment were successfully completed. Hence, the grid system was qualified for spacecraft propulsion.

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.

2001，

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 I-sp 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 typo-dimensional 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 call enhance aerodynamic acceleration, which contributes to large thrust generation.

Operation of an ion engine in space requires an electron source to maintain the spacecraft's electrical neutrality. To neutralize exhausted ions, a low-power, compact microwave electron source was developed. An 18-mm-diameter noncavity-type discharge chamber consisting of an L-shaped antenna and a magnetic circuit was developed, and electron emission current of 100 mA was obtained at the power level of 10 W. maintaining the mass flow rate of xenon at 0.5 sccm.

The effect of electrode configuration 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 flowfield, we introduced a magnetosonic Mach number, which is defined as Local velocity divided by a propagation speed of the MHD disturbance, Based on the magnetosonic Mach number distribution of the flowfield, the model can explain the thrust characteristics of the MPD arcjet, especially the superiority of a short cathode under various anode configurations, Because the electromagnetic thrust is unaltered for the same anode configuration, the electrothermal component of thrust makes a difference between the long and the short cathodes, With a short cathode configuration, the large heat deposition near the cathode tip, which is inevitable to MPD arcjets, can be confined in the submagnetosonic region where the local now is accelerated to magnetosonic velocity, Then the thermal deposition into the submagnetosonic region can be efficiently recovered through transmagnetosonic acceleration, resulting in a large thrust generation.