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The effects of antenna shape and magnetic field configuration on thrust performance in a 30W class miniature microwave discharge ion thruster were investigated with the objective of improving its thrust performance. The ion beam currents were measured for three different antenna configurations: disc, star, and cross. The results showed that the optimum antenna shape depends on the desired mass flow level because of the tradeoff between microwave-plasma coupling efficiency and surface area recombination. The ion beam current with various magnitude of magnetic field was measured, since magnetic confinement and microwave-plasma coupling could be affected by the magnetic field configuration. The ion beam current was increased with increasing the number of magnets, though under too much magnitude of magnetic field, the thruster cannot be ignited. So, there is an optimum magnitude of magnetic field. The thrust performance of the miniature microwave discharge ion thruster, i.e., propellant utilization, and ion beam production cost were 0.97 and 600W/A, respectively, at mass flow rate of 0.018mg/s, and incident microwave power of 8W.

We detail the design, fabrication, and experimental and calculation results for three cavity solar thermal propulsion (STP) thrusters — medium, small, and very small — made of single-crystal molybdenum (SC-Mo), developed in the National Aerospace Laboratory of Japan (NAL). We obtained very high temperatures — 2,300 K for the propellant gas and 2,200 K on the outer surface of the thruster — at an appropriate propellant flow with the small thruster by solar-heating with a suitable concentrator of 1.05 m in diameter, which corresponds to an 800-second-class specific impulse thruster. Temperatures obtained in experiments were much lower than expected, however. To find points for improvement and evaluate thruster performance in space, we conducted a model calculation that confirmed the thruster could achieve an 800 to 900-second-class specific impulse with a 0.1–2 N class thrust magnitude in space. STP is thus a candidate for near-future propulsion in applications such as upper stages of orbit transfer vehicles.

MUSES-C is an interplanetary spacecraft which will be launched in 2002 with the mission for the sample return from an asteroid. MUSES-C has high voltage solar arrays operating at 120 V and Xe ion thruster system as the main propulsion system. Plasma interaction between the backflow plasma from the ion thruster plume and the solar array is investigated by computer simulation and laboratory experiments. The plasma density near the solar arrays is expected to be 1011 m-3∼1013 m-3. Due to the operation of ion thruster neutralizer, the solar array has a positive potential with respect to the plasma. The increase of electron current to the solar array, however, may drop the array potential. Once the array potential becomes negative, arcing is observed to occur even at the voltage of −120 V. Several methods to suppress the plasma interaction are proposed.

The behavior of exhausted plasma from ion thruster in the near earth orbit was analyzed to find its influence on the earth's environment. It was found by the analysis of the motion of beam plasma using PIC method that the quasineutral state of the exhausted plasma was broken and the charge separation occurred soon after their exhaust because their density much decreases by the beam divergence and the electric field in the plasma can not keep the electrons to stay in the plasma. So, the ions will be trapped by the geomagnetic field for a long period. It was also found by the analysis using the plasma density distribution model and the ion thruster mission analysis model that the plasma components density distribution near the earth will change by the energy input of these trapped ions if the large scale operation of ion thruster is performed.

An ion thruster emits a high density plasma plume and the electromagnetic noise radiated from the plasma current may affect the spacecraft instrument. An electromagnetic particle simulation code has been developed to investigate the effects of the electromagnetic interference induced from the ion thruster plume. The electromagnetic waves are excited due to plasma instability caused by strong polarization field inside imperfectly neutralized plume. The strength and frequencies of electromagnetic noise radiated by ion thruster plume depend on the distribution of charged particles inside the plume, static magnetic field strength and its orientation.

This paper reports results of fundamental experiments using a continuous-wave CO2 laser thruster. The Laser-Sustained Plasma (LSP) was maintained under a pressure range from 2 to 7 atm with a laser power level of more than 300W. The location of LSP can be controlled by adjusting power level and focus position of the 1aser beam. Thruster performance was evaluated by measuring thrust using a load-cell sensor. Energy balance in the thruster was clarified by measuring heat loss and transparent laser loss.

Numerical analyses and experiments of a Hall thruster have been conducted not only to improve thruster performance but also to understand the plasma acceleration processes. 1kW class Hall thrusters were fabricated to measure thrust, ion beam current, and distributions of plasma density, electron temperature, and space potential in the acceleration channel. Two-dimensional ion flow model was made to analytically examine the plasma acceleration processes in the channel, and the analytical results are compared with the experimental ones.

資料番号: SA1160299000

We have investigated the plasma production in a micro RF ion thruster, where the plasma source is 5.0 mm in radius and 10 mm in length. To find the optimum condition to generate the RF plasma in the wide range of frequency f = 1-1000 MHz and pressure p = 0.1-10.0 Pa, we employ the equivalent circuit model, where the global model is also incorporated to obtain the plasma parameters self-consistently. The numerical results have indicated that capacitive coupling dominated over inductive coupling and the sheath resistance has a significant influence on the power coupling efficiency at lower frequency. The power coupling efficiency could be more than 80% at higher frequencies (> 30 MHz).

The research and development process of the ablative pulsed plasma thruster SIMP-LEX is summarized. By experiment, the magnetic field in between the electrodes was measured with an induction probe, and a high-speed camera was applied to estimate the propagation of the plasma plume. Further, the optimization process towards higher thrust efficiency by means of geometric and electrical modifications is described. By analysis, the calculation of magnetic field and change in inductance is summarized and the numerical models for the performance simulation of SIMP-LEX introduced. A peak value of 0.7 T was found for the magnetic field created during an 80 J discharge whereas a propagation velocity of 28 km/s was estimated for the first plasma plume. The optimization process towards a new engineering model yielded an increase of 40% in mean exhaust velocity to 26 km/s, and a maximum thrust efficiency of 32% was evaluated respectively. Calculation of magnetic field and change in inductance based on the Biot-Savart law yield input for the slug model, which can then predict the general performance parameters with a good accordance.

Ion thrusters are promising propulsion devices not only for standard-sized spacecraft but also for small spacecraft. However, difficulty of the development of a miniature neutralizer has been a critical problem for miniature ion thrusters. We have proposed a novel ion thruster system to solve this problem. That is a switching operation, where a plasma source can select ion beam extraction or electron emission by electrical switching. The major challenge of this study is to find the method to effectively performs both functions from one plasma source. To match the system to small spacecraft, the both modes must be conducted using low microwave power and low mass flow as low as 1.0 W and 0.15 sccm. Firstly, fundamental characteristics of electron emission were investigated using an orifice plate. Based on this result, secondly, a special grid system for the switching operation was designed and examined. This grid system has two-different-size apertures for ions and electrons respectively. This grid showed high ion beam extraction performance, that is, 4.1 mA ion beam current at 1.0 W microwave power input. It also showed much higher electron current than the old grid system, although the electron current of 1.0 mA has not reached the targeted current yet. As a result, the experiment showed that the special grid system proposed here is an effective method to realize the switching operation.

Several experimental and numerical simulation results for the development of miniature microwave discharge ion thrusters have been reported. The important points to improve the performance of the ion thruster are the shape of the antenna for microwave emission and the magnetic configuration in its thruster head. The extracted ion beam current of 12mA is obtained in conditions of input 2.45GHz microwave power of 8W and propellant (xenon) flow rate of 0.2 sccm with a star type antenna. A numerical simulation code coupling the particle-in-cell method and the finite difference time domain method was used to know plasma behavior in the discharge chamber of an ion thruster.