西山 和孝

The air-breathing ion engine (ABIE) is a new type of electric propulsion system to be used to compensate the aerodynamic drag of the satellite orbiting at extremely low altitudes. 140 to 240km. To save the propellant mass for a long operation lifetime, it inhales the high-speed low-density atmosphere surrounding the satellite and use it as the propellant of ion engines. Since feasibility and performance of the ABIE depend strongly on the compression ratio and the air-intake efficiency, numerical analysis has been performed by means of the direct-simulation Monte-Carlo method to clarify the characteristics of the air-intake performance in highly rarefied flows. Influences of the aspect-ratio of the air-intake duct, the angle of attack, and the flight altitude are investigated.

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. © 2002 American Institute of Aeronautics and Astronautics, Inc.

Radiated electric field emissions from the prototype model of the Ion Engine System (IES) of the MUSES-C mission were measured in accordance to MIL-STD-461 E. The average noise level exceeded the narrowband specification at frequencies less than 5 MHz. The microwave discharge neutralizer generates a broadband noise and narrowband oscillations which have a fundamental frequency of about 160 kHz and are accompanied by its harmonics up to the 5th. The leakage of the 4.25 GHz microwave for plasma production and its second harmonic were 65 dB and 35 dB above specification, respectively. The X-band receiver onboard the MUSES-C measured the noise from the IES at the up-link frequency of 7.2 GHz through a horn antenna. This susceptibility test proved that the microwave discharge ion thruster will never interfere the deep space microwave communication.

Noise and oscillatory behavior of a plasma column produced in front of the microwave discharge neutralizer developed for MUSES-C mission were experimentally investigated. Radiated electric field emissions were measured following to MIL-STD-461 E. The average noise level exceeded the narrowband specification by 30 dBμV/m at frequencies less than 5 MHz. Noise in electron emission current was also measured by using a current probe and a spectrum analyzer, and was compared with the noise of a hollow cathode. The microwave discharge neutralizer generates a broadband noise and oscillations which have a fundamental frequency of about 160 kHz and are accompanied by its harmonics up to the 5th. Considering the dependence on the diameter of the plasma column, they are probably the radial oscillation modes of ion acoustic waves. Although the hollow cathode shows nearly the same noise level at frequencies less than 1 MHz, intense oscillation exists in the 1-10 MHz range, which is generated by the keeper plasma.

Institute of Space and Astronautical Science has developed a cathodeless microwave discharge neutralizer for the microwave ion engine system. In order to clarify the electron emission mechanism of the microwave neutralizer, electron current characteristics as well as plasma parameters inside the neuralizer were measured. The electron current was greatly influenced by the material of the neutralizer orifice. Among several materials with different work functions, an orifice made of tungsten impregnated with barium-oxygen showed the best performance. It is concluded that by adopting a low-work-function material, the secondary-electron emission by the singly-charged-ion impact was emphasized, and thereby the electron current was increased.

As one of malfunctions on the ion thruster the lack of neutralization is investigated by means of the ground experiment using a 2cm microwave discharge ion thruster and a miniature model similar to the MUSES-C spacecraft. Electron leak form the plasma beam to the high voltage solar array caused a slight charging. No electron emission resulted in the full charging and pulled back the extracted ions to the thruster again, when the so-called virtual anode was formed in front of the decel grid. Based on the experimental simulation a design philosophy on the ion thruster and the spacecraft is described. © 2001 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

MUSES-C space mission requires the ion engine system (IBS) 16,000 hours in the operational time. The ion thruster is driven by 4.2GHz microwave in the result that the hollow cathodes are completely eliminated. The life critical parts are screened to the electro-static grid system and the microwave feeder of the neutralizer. The engineering model of the microwave discharge ion thruster was evaluated for the lifetime by several methods: the numerical simulation, accelerated wear tests and the real time endurance test. Combining these techniques MUSESC/ IES is concluded to have a good capability to endure the required life. © 2000 The American Institute of Aeronautics and Astronautics Inc. All rights reserved.

An electron cyclotron resonance (ECR) ion source was designed and its performance was studied in comparison with a DC ion source. Ion production cost of 340 V/ion for ECR ion source and 189 V/ion for DC ion source were obtained at the pressure of 0.5 m Torr and the discharge power of 50 W. A luminosity distribution of the Ar-II line was imaged through an interferometer filter and a luminous arch hill was observed near the ECR region, where the Plasma was resonantly generated. Ion wall loss measurement showed that the ion production cost became worse due to the excessive ion wall loss near the cusped magnets, where the plasma confinement was inferior to that of the DC ion source.