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.

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

NASA is investigating high power, high specific impulse propulsion technologies that could enable ambitious flights such as multi-body rendezvous missions, outer planet orbiters and interstellar precursor missions. The requirements for these missions are much more demanding than those for state-of-the-art solar-powered ion propulsion applications. The purpose of the NEXIS program is to develop advanced ion thruster technologies that satisfy the requirements for high power, high specific impulse operation, high efficiency and long thruster life. The nominal design point for the NEXIS thruster is 20 kWe at a specific impulse of 7500 s with an efficiency over 78% and a xenon throughput capability of greater than 2000 kg. These performance and throughput goals will be achieved by applying a combination of advanced technologies including a large discharge chamber, erosion resistant carbon-carbon grids, an advanced reservoir hollow cathode and techniques for increasing propellant efficiency such as grid masking and accelerator grid aperture diameter tailoring. This paper provides an overview of the challenges associated with these requirements and how they are being addressed in the NEXIS program.

An overview of the Nuclear Electric Xenon Ion System (NEXIS) program

In the MUCES-C mission conducted by JAXA (Japan Aero Exploration Agency), a microwave neutralizer is mounted with a microwave ion engine on the HAYABUSA space probe. The neutralizer consists of an L-shaped antenna to inject microwaves and samarium cobalt magnets to provide ECR (electron cyclotron resonance). Plasma production of a higher density than the cutoff density is expected in the discharge chamber, but the neutralizer is so small that high-precision measurements using a probe are difficult. To clarify the plasma production mechanism in the microwave neutralizer, numerical analysis was conducted using a code coupling PIC (particle-in-cell) method, and a FDTD (finite-difference-time-domain) method. This paper describes effects caused by varying magnetic field configuration and antenna position in the neutralizer. The calculation results show that bringing the antenna closer to the ECR region is effective for plasma production.

Analysis of electron and microwave behavior in microwave discharge neutralizer

Mu Space Engineering Spacecraft-C (MUSES-C) was launched into the transfer orbit in May 9, 2003. Microwave discharge ion engines are used as its primary propulsion system. The Data Relay Test Satellite (DRTS) with 1.8kW/250mN-class hydrazine DC arcjet thruster systems was launched on September 10, 2002, and now on service on the geostationary orbit. Ion engine systems onboard Engineering Test Satellite VIII (ETS-VIII) will be also launched in 2004 by the H-IIA rocket. Basic re searches of many kinds of electric thrusters are also being carried out in lots of academic places.

/pdf/an-overview-of-electric-propulsion-activities-in-japan-5aicz2t8zu.pdf

An overview of electric propulsion activities in japan

The electron cyclotron resonance microwave discharge ion thruster μ10, installed in the Japanese asteroid explorer Hayabusa, experienced trouble during a round-trip voyage between Earth and an asteroid due to degradation of the neutralizer. The neutralizer is a critical component limiting the thruster lifetime. The mechanism of the performance degradation is investigated here to find out methods to extend its lifetime for next-generation μ10 thrusters. Fault tree analysis during on-the-spot inspections of the neutralizer, previously used for 20,000 h endurance testing on the ground and in experimental simulations, leads to the following proposed degradation mechanism. The internal-surface materials are sputtered by doubly ionized xenon and accumulate on other surfaces, forming thin films containing iron. Next, thermal cycles result in the peeloff of the thin films due to the difference between the thermal expansions of the films and the base plates. Finally, thin films containing iron (that is, ferromagne...

Performance Degradation of a Spacecraft Electron Cyclotron Resonance Neutralizer and Its Mitigation

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.

Electron Emission Mechanism of Microwave Discharge Neutralizer

Space tether system has wide variety of applications for technological and scientific missions such as momentum exchange orbit transfer by releasing payload connected with the tether, artificial gravity generation by rotating two bodies connected with the tether and scientific inspections of the upper atmosphere by multi probes on the long tether. The electrodynamic tether is a kind of space tether application that possesses the conducting tether and plasma contactors. It produces thrust on the earth orbit by the interaction between the current through the tether and the geomagnetic field. So, it can be used for orbit raising thrust force or orbit lowering drag force depending on the direction of the current. It also generates electrical power (electromotive force) by the conducting tether across the geomagnetic field. The electrodynamic tether also has the potential to provide a near-propellantless propulsion system, using the energy from sunlight converted to electrical energy. Therefore, the electrodynamic tether system can be applicable for various missions such as station-keeping, power generation, orbit raising, and the deorbit of largesized debris when some practical problems such as the tether cutting would be solved.

Evaluation of Plasma Contactor Ground Experiments for Electrodynamic Tether

It is expected that an orbit transfer vehicle using an ion thruster (IOTV) can obtain a large payload ratio, because the ion thruster has a high specific impulse (several thousands seconds). However, orbit transfer missions by means of the IOTV require large trip times, therefore it is necessary for the ion thruster to have a sufficient endurance performance. In this study, the trajectory of the solar powered IOTV was analyzed in consideration of the influence of earth's eclipse, then the payload capability of the IOTV and the steady/endurance performance of the ion thruster were evaluated. For results, it was found that the IOTV could achieve approximately double the payload mass of the OTV using a chemical rocket, and a required durability to the ion thruster was comparable to one that has used for ETS-VI.

Application Study of Ion Thruster to Orbit Transfer Vehicle.

A high current compact ECR ion source

A preliminary life test of a low-power DC arcjet anode was conducted using a 500-W-class laboratory-model arcjet. After a continuous operation of 90-h, we compared the degradation of two types of anodes: one is a conventional pure tungsten anode, which is made of multiple fine grains of several tens of micrometer in size; another is a tungsten anode made of only four coarse grains. Both materials showed nearly the same amount of degradation from the viewpoint of so-called a constrictor closure phenomenon, in which the radius of the constrictor decreases during the continuous operation. However, micrograph analyses showed different degradation features for the two anodes: for the conventional tungsten anode, grains were severely embrittled in the recrystalization processes after the thruster operation accompanied by many cracks or even drops out of grains; in contrast, the anode made of four coarse grains showed a few cracks only along the grain boundaries, and its thrust performance was unchanged before and after the 90-h test. Hence there is a possibility to suppress the grain embrittlement by completely removing these grain boundaries from the tungsten anode by making the anode from one huge grain.

/pdf/a-study-of-degradation-mechanism-of-a-low-power-dc-arcjet-21v2ct7qe3.pdf

Haruki Takegahara

Papers

Electron Emission Mechanism of Microwave Discharge Neutralizer

Evaluation of Plasma Contactor Ground Experiments for Electrodynamic Tether

Application Study of Ion Thruster to Orbit Transfer Vehicle.

A high current compact ECR ion source

A Study of Degradation Mechanism of a Low-Power DC Arcjet Anode