This paper proposes an electrothermal microplasma thruster using azimuthally symmetric microwave-excited plasmas, which consists of a microplasma source and a micronozzle. The microplasma source is made of a 10 mm long dielectric chamber of 2 mm in inner diameter covered with an electrically grounded metal, which produces high temperature plasmas at around atmospheric pressure. The micronozzle has a throat of 0.2 mm in diameter, which converts high thermal energy of plasmas into directional kinetic energy to produce the axial thrust. First, we have developed a numerical model for Ar microplasmas and micronozzle flows to estimate the thruster performance. The model consists of three modules: a volume-averaged global model, an electromagnetic model for microplasma sources, and a fluid model for micronozzle flows. Numerical results indicate that the microwave power absorbed in plasmas increases with microwave frequency f and relative permittivityεd of the dielectric chamber, to achieve the plasma density in the range 1014-1016 cm-3. A certain combination of the frequency and permittivity significantly increases the power absorption. The micronozzle flow was found to be very lossy because of high viscosity in thick boundary layers, implying that shortening the nozzle length with increasing half-cone angles suppresses the effect of viscous loss and thus enhances the thrust performance. A thrust of 2.5-3.5 mN and a specific impulse of 130-180 s were obtained for a given microwave power range (Pt <10 W), which is applicable to a station-keeping maneuver for microspacecraft less than 10 kg. Moreover, we have developed a microwave-excited microplasma source, which has a dielectric chamber of 10 mm length and 1.5 mm in inner diameter, where off-the-shelf mullite (εd ≈ 6) and zirconia (εd ≈ 12-25) tubes are employed. Experiments were performed at f = 2 and 4 GHz, Pt < 10 W, an Ar flow rate of 50 sccm, and a microplasma pressure of 10 kPa, where optical emission spectroscopy and Langmuir probe measurement were employed for the diagnostics of microplasmas. The measurements indicate that the Ar I emission intensity and plasma density ne increase with f and εd, and that the ne is in the range 1011-1013 cm-3. The rotational temperature Trot of N2 molecule in the added gas was in the range 1100-1500 K, and the specific impulse estimated from the temperature T = Trot was determined to be approximately 70 s from the model analysis.
We designed a pulsed plasma thruster (PPT) with a propellant feeding mechanism using two poly-tetra-fluoroethylene (PTFE) bars as propellants. Both the PPT and capacitors were mounted on a thrust stand with a 1-m-long arm developed for precise measurement of impulse bit. The initial thrust performance showed impulse bit per joule of 43-48 μNs/J, specific impulse of 470-500 s and thrust efficiency of 10-12% with stored energy of 4.5-14.6J. 10000-shots operation was conducted with a stored energy of 8.8J. As a result, each PTFE bar was supplied in a length of approximately 2 mm, and a total impulse of approximately 3.6 Ns was obtained, though both the impulse bit and the thrust efficiency gradually decreased with shot number because of uneven receding of the PTFE surface. In order to investigate physical phenomena in a whole system, an unsteady numerical calculation was performed, which simultaneously simulates unsteady phenomena of discharge, heat transfer to/inside the PTFE, ablation from the PTFE surface and plasma flow. The calculated discharge current waveform and the PTFE mass loss showed good agreements with experimental results. The uneven shape of the PTFE surface after 10000-shots operation was also qualitatively explained with the calculation.
The Hall-effect plasma accelerators are one of the most advantageous electric propulsion device in space. The Hall-effect plasma accelerators are generally categorized into magnetic-layer type and anode-layer type. The anode-layer type Hall-effect plasma accelerator is expected to be widely accepted as space thrusters with longer-lifetime and higher thrust performance. In this study, the fundamental performance characteristics of an anode-layer type Hall-effect plasma accelerator have been investigated. A 1kW-class anode-layer type Hall-effect plasma accelerator, TALT(Toyonaka Anode-Layer Thruster)-1, was operated under various magnetic fields, discharge voltages, propellant mass flow rates and anode positions and structures. The thrust performance was enhanced with positioning the anode front in the upstream and widening the anode propellant path width. The thrust and the specific impulse ranged from 12mN to 45mN and from 800s to 1739s, respectively, at the discharge voltages of 200-400V under the xenon mass flow rates of 1-3mg/s. The thrust efficiency reached 45%.
Recently, miniaturization and light weight conversion of the satellite itself have become necessary. And high improvement of efficiency of the thruster has become necessary with the substantial restriction of the electric energy which it can use. In this study, conversion utilization of general solar light energy to electric power was proposed as the method of obtaining energy at outer space. We observed the arc-jet thruster because the thrust density was high even in the thrusters and the system and structure was relatively simple. Digital control power supply to electric power was developed in order to convert low, and to miniaturize and lighten the arc-jet thruster. As a result, discharge starting reproducibility with single power supply was good and we succeeded in stable arc discharge in the low electric current territory.
In order to develop reusable space vehicle, it is important to ensure sufficient reliability of thermal protection systems under re-entry environments. For such a purpose, a sensor system to detect a recession rate of anti-oxidation SiC coating on carbon-carbon composite was attempted to be developed. This sensor consisted of multi-layered SiC/carbon coating on a SiC substrate, and high temperature oxidation damage of SiC in the multi-layered coating can be detected by the change of electric resistant of the coating caused by oxidation of carbon layers. In the present paper, conceptual design of the sensor was presented and several required technologies to develop the sensor were discussed. The discussion included how to form the multilayered coating and measuring technique of the electric resistance at high temperatures.
A quantitative analysis in erosion properties of polyimide and fluorinated ethylene propylene (FEP) under simulated low earth orbit space environment of atomic oxygen and ultraviolet was performed. An atomic oxygen environment was simulated by a laser detonation atomic oxygen beam source, whereas that of ultraviolet was made by a 172 nm-eximer light. Increase in erosion rate up to 400 % was observed on polyimide when relative ultraviolet intensity is high. In contrast, no synergistic effect was observed on FEP in tested conditions. X-ray photoelectron spectroscopy revealed that atomic oxygen did not accommodate on FEP surfaces even after atomic oxygen exposures, in contrast, increase in surface oxygen concentration was detected at the polyimide surfaces. These experimental results suggest that ultraviolet-induced desorption of reactive products, which was an origin of synergistic effect on polyimide, does not occur at the FEP surface.