Masaki Yonemoto
(Nagoya University, Japan)

This paper reports operation characteristics of a steady-state, two-dimensional MPD thruster using a hollow cathode. Electrode degradation is one of the most serious problems in MPD (MagnetoPlasmaDynamic) thruster. In order to solve this problem, a tungsten cathode was replaced with a hollow cathode which has better electron emission performance and a longer lifetime. A rectangular configuration was employed for the acceleration section between electrodes. In order to evaluate thruster performance, the discharge voltage and thrust were measured with varying an applied magnetic field, discharge current, propellant mass flow rate and the inter-electrode distance.

Abdolrahim Rezaeiha
(Sharif University of Technology, Iran)

Pulsed Plasma Thruster (PPT) is one of the promising space propulsion devices for the micro- and nano-satellites as the following advantages: simplicity, lightweight, robustness, and low power consumption. Moreover, PPT can generate very small impulse bits with high specific impulse. Therefore PPT has great advantages as a precision attitude control system. Although PPTs and their relevant issues have been investigated to a certain degree in recent years, but as a result of the fact that PPT performance is a function of a lot of different parameters like geometrical dimensions, discharge energy, etc and also because of its sophisticated physics; it is felt that issues related to developing an optimized PPT needs to be investigated in greater detail for a PPT with smaller dimensions and better performance. Therefore a laboratory benchmark pulsed plasma thruster was designed, developed and successfully tested in the vacuum chamber at 10-6 mbar. Then a parametric study has been conducted on investigating the effect of discharge energy on PPT performance. The PPT impulse bit varied from about 0.4 mN-s to more than 1.3 mN-s while discharge energy changed from 10 J to 50 J. This paper reviews the results of the mentioned parametric study in brief.

Masamichi Naka
(Osaka Institute of Technology, Japan)

The Project of Osaka Institute of Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at Osaka Institute of Technology in 2007. In PROITERES, a nano-satellite with electrothermal pulsed plasma thrusters (PPTs) will be launched in 2011. The main mission is powered flight of small satellite by electric thruster. This study aims at improvement in performance by changing configuration of PPTs. The total impulse of about 5 Ns was achieved with a teflon cylindrical discharge room 9.0 mm in length and 1.0 mm in diameter in 53,000-shot operation with 2.43 J/shot. After lots of tests using the engineering model of PPT system, the development of the PPT system flight-model was completely finished.

Souichi Masui
(Kyushu Institute of Technology, Japan)

The pulsed plasma thruster (PPT), has attracted attention again as a micro-thruster because of its compactness, lightweight and comparatively low power consumption. Nevertheless, propellant utilization efficiency of conventional Teflon PPTs is relatively low among electric propulsion devices because the propellant originated from late-time ablation is not utilized well for thrust. The liquid propellant PPT (LP-PPT), where water or ethanol propellant is fed with an injector, is proposed to overcome the disadvantages and thrust measurement has shown that LP-PPT provided higher specific impulses than the conventional PPT. However, water requires temperature management for storage owing to relatively high freezing point. Moreover, even if ethanol, which has sufficiently low freezing point, is used as propellant, pressurant is necessary as well as water because the vapor pressures are deficient for self-pressurization. In this study, we have proposed to use dimethyl ether (DME) as propellant. DME, which has a freezing point of 131 K at 1 atm and a vapor pressure of 6 atm at 298 K, is stored in tanks as a liquid and requires no pressurant. We have designed a DME pulsed plasma thruseter and evaluated performance. Thrust measurement of a coaxial type thruster at a capacitor stored energy of 13 J.

Tony Schoenherr
(The University of Tokyo, Japan)

The pulsed magnetoplasmadynamic thruster SIMP-LEX is developed in collaboration at the universities in Stuttgart and Tokyo in order to serve as main propulsion system for the planned satellite missions Perseus and BW1. Within this project, the thruster's physics are investigated to gain information about the thrust-creating processes and to draw conclusions for the thruster. Investigation of the discharge behavior was followed by experiments using emission spectroscopy and interferometry. The change of heavy particle temperature and electron density as well as the evolution and propagation of the created species was accessed by these means. From these values, the electron temperature was calculated over time for different positions downstream of the plasma flow. The various spectral lines of the PTFE plasma were identified, and the composition of the plasma traced throughout the discharge. An example spectrum with identified species is shown in Fig. 1. Manyfold ionized species like C++ and C+++ indicate high energy densities, and the presence of Cu+ represents the erosion of the copper electrodes. Variations in discharge voltage of the thruster showed variations in the composition of the plasma as well as in the temperatures and electron densities. The results are compared and conclusions for the thruster drawn.

Keiji Hagiwara
(University of Tokai, Japan)

Experimental investigation of impulse-bit and propellant consumption rate, or mass shot, per single pulse discharge were conducted to characterize the thrust performance of the rectangular laser-electromagnetic hybrid acceleration thruster with various propellant materials. From the result, alumna propellant showed significantly superior performance. The largest values of the measured impulse-bit, specific impulse and thrust efficiency were 49 uNsec, 6,200 sec and 22%, respectively. To investigate this high Isp characteristic, speed of ions of plasma plumes exhausted from the thruster were measured with time-of-flight (TOF) measurement, where temporal variations of ion current with Faraday-cup were measured. Form the result, the maximum value of the ion speed of exhausted plasma was 42 km/sec at charge energy of 5.8 J. In addition, the observation of the temporal behavior of the exhausted plasma plumes was conducted with ICCD camera.