Haruki Takegahara

Development of a Novel Power Processing Unit for Hall Thrusters

A newly developed power processing unit (PPU) offers the advantages of smaller size and lighter weight than conventional PPUs. The thrust performance of a magnetic layer type Hall thruster developed at Kyushu University with this new PPU was investigated; it showed a good performance as compared with conventional power supplies. The thrust to power ratio was improved to 58 mN/kW at discharge voltage of 150 V and anode xenon mass flow rate of 1.0 mg/s.

Feasible Study on Low Frequency Plasma Jet as Ignition System for HAN based Propellant

A low frequency (LF) plasma jet is proposed as an ignition system candidate for hydroxyl ammonium nitrate (HAN) based propellant, especially focused on SHP163, as substitution of conventional hydrazine. Plasma generation capability and power consumption characteristics were investigated as functions of frequency, applied voltage, and distance between high-voltage and ground electrodes. LF plasma jet in itself was generated at 5 Hz of frequency, 5 KV of applied voltage, and 5 mm of electrodes distance; and its power consumption was 16 W. At lower frequency and higher voltage, plasma generation capability was increased. Power consumption was decreased at lower frequency, lower applied voltage, and shorter electrode distance. At same applied voltage, lower power consumption was obtained at lower frequency and shorter electrodes distance. Additionally, LF plasma jet was applied to initiate SHP163. 1.1 × 10 -2 g/s of mass reduction rate was obtained at 5 Hz of frequency and 5 KV of applied voltage, and its power consumption was 30 W. This result indicates that LF plasma jet has excellent possibility to be a good reaction initiation/enhancement system.

Experimental Investigation of Inductively Coupled Plasma Cathode for the Application to Ion Thrusters

The inductively coupled plasma cathode (ICP/C) has been developed as electron source for ion thrusters to liberate the thrusters from the limitations of hollow cathodes. The power consumption efficiency of this cathode was not sufficient for practical use in space. Therefore, in order to improve the cathode performance, its ignition and electron emission characteristics were investigated experimentally as functions of orifice dimensions, ion collector shape and total number of coils in this study. It becomes clear that inner pressure and electric field in the discharge vessel are dominant factors of plasma ignition in ICP/C. Thus, the pressure rise by changing orifice dimensions and the RF voltage rise by increasing total number of coils cause the enhancement of ICP/C ignition capability. Electron emission current in ICP/C is limited by collected ion current in the ion collector, which depends on the ion saturation current at the ion sheath formed on the collector surface. Additionally, the ion saturation current strongly depends on electron number density, vessel inner pressure and ion collection area. Increase of total number of coils causes the enhancement of ignition capability and the improvement of electron emission performance. Moreover it is desirable for ICP/C lifetime to increase the number of coils. As the improved cathode performance, 0.52 A of anode current was obtained at 24 W of RF power, 0.15 mg/s of xenon mass flow rate and 40 V of anode voltage.

Numerical Study of Beam Extraction Phenomena in an Ion Thruster

Ion beam extraction phenomenon is one of the most important processes in an ion propulsion system because ion beam current and beam accelerating voltage dominate thrust. In order to evaluate the beam extraction capability, which is determined by the interrelation between discharge plasma properties and beam accelerating conditions, there is a need to understand the problems of ion sheath formation. This study has been performed to develop a numerical simulation code which is capable of investigating the ion beam extraction phenomena in the ion thruster. In the calculation, both ions and electrons were treated as particles in a particle-in-cell routine to determine the ion sheath self-consistently. The calculation results of electric potential contours and extracted currents were compared with the experimental ones and they were found to be in good agreement. In addition, it was confirmed that the ion beam extraction capability can be simulated in a wide range of beam accelerating voltage.

Guest Editorial Introduction to the Special Issue on Plasma Propulsion

This Special Issue is dedicated to the physics, technology, and application of plasma propulsion for spacecraft. The field of plasma propulsion includes a broad variety of thrusters that can achieve high propellant exhaust velocity, thereby offering a large mass savings for space vehicles compared to chemical (combustion) rockets. These thrusters are broadly categorized by their propellant acceleration mechanism into three groups: 1) electrothermal; 2) electrostatic; and 3) electromagnetic [1] . Research into plasma propulsion dates back several decades, with a first application in space in 1964 on the Soviet Zond-2, which used an ablative pulsed plasma thruster to control the spacecraft orientation. Today, plasma propulsion is a very rapidly growing area of plasma science and technology, with over 250 operational spacecraft employing plasma propulsion in a variety of applications. Many new plasma thrusters have been recently developed, including numerous successful attempts to scale previously known systems to lower and higher power levels, ranging from a few watts to over 100 kW [2] , [3] . As space exploration shifts toward small and efficient satellites, or micro and nanosatellites [4] , [5] , there are many near-future space missions involving science, military, and commercial payloads utilizing micro and nanosatellite platforms. These platforms require very small levels of thrust for very fine attitude control, high resolution. Earth imaging and astronomy, as well as very precise positioning requirements and other very precise positioning requirements for spacecraft performing formation flying and interferometry missions. Miniaturized propulsion systems are required to satisfy these emerging needs for both the low-thrust missions and the propulsion on small-sized spacecraft.

Basic Characteristics of Discharge Plasma Ignition System for 1N-class RCS Thruster with Green Monopropellant

A new ignition system utilizing discharge plasma for reaction control system (RCS) thrusters with green monopropellant is designed and evaluated experimentally in this study. The discharge plasma ignition system laboratory model (DPI-LM) is designed for one of hydroxyl ammonium nitrate (HAN) based monopropellant, SHP163; moreover, the DPI-LM is in substitution for conventional solid catalyst. Objectives of this study are (1) to design and build of DPI-LM and (2) evaluate basic propellant ignition characteristics in terms of successful and stable propellant ignition conditions, power consumption, and fundamental lifetime estimation. In addition, in order to generate discharge plasma prior to propellant ignition, a noble gas is used. Effect of noble gas type, argon and helium, on propellant ignition characteristics are also evaluated. Argon gas shows better propellant ignition with wide ranges of argon and SHP163 mass flow rates. It is considered that the propellant ignition strongly connects to discharge plasma diffusion condition prior to ignition. The power consumption at an argon mass flow rate of 0.075 g/s and a SHP163 mass flow rate of 0.3 g/s is approximately 270 W. The electrode degradation as a function of accumulated experiment time is evaluated as simplified lifetime estimation. The results of the degradation is only 0.1 % in electrode mass at 2000s , and the stable propellant ignition keeps at an accumulated time of 2000 s.

Basic Characteristics of Reaction Initiation System Using Dishcharge Plasma for 1N-class Thruster with Green Propellants

A new reaction initiation (ignition) system using discharge plasma is proposed for 1Nclass reaction control system (RCS) thruster with green monopropellant, especially, one of hydroxyl ammonium nitrate (HAN) based liquid monopropellant, SHP163. This reaction initiation system is designed in substitution for conventional catalytic decomposition system. In this study, fundamental reaction initiation characteristic of SHP163 by discharge plasma, based on total amount of energy applied, were investigated. Better propellant reaction initiation was achieved at larger number of discharge attributed from frequency even though levels of applied energy were same. It became clear that reaction initiation characteristics depend on not only total amount of energy but also number of single discharge. Additionally, a laboratory model of new reaction initiation system was designed and built. Fundamental characteristics based on power consumption were investigated. The new reaction initiation system was demonstrated with 100 to 5000 Hz of frequency and 3.2 to 27.8 mg/s of helium mass flow rate. At any cases in this experiment, stable discharge plasma generation was observed, and highest power consumption was 35 W.