• Bulletin of the Korean Space Science Society
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• 2007.10a
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• pp.139-143
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• 2007

Electric propulsion has become a cost effective and sound engineering solution for many space applications. The success of SMART-1 and MUSES-C developed by European Space Agency (ESA) and Japan Aerospace Exploration Agency (JAXA) each proved that even small spacecraft could accomplish planetary mission with electric propulsion systems. A small electric propulsion system which is Hall effect thruster like SMART-1 is under development by SaTReC and GDPL (Glow Discharge Plasma Lab.) in KAIST for the next microsatellite, STSAT-3. To achieve optimized propulsion system, it is very necessary to understand plasma motions of Hall effect thruster. In this paper, we try to approach comprehensive plasma model with the particle simulation complementary to Particle In Cell (PIC) simulation. We think these two different approaches will help experimenters to optimize Hall effect thruster performances.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.2
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• pp.195-200
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• 2011

A low power Hall thruster is under development for orbit maintenance of a small Earth observing satellite. Both cylindrical and annular type thrusters were manufactured and tested to characterize the performance of cylindrical Hall thrusters. Results were described through comparative analyses. Cylindrical thrusters were manufactured in two different channel diameter dimensions, 28 mm and 50 mm. Thrust, ion velocity and ion current were measured in various operating conditions. The results show that cylindrical thrusters are more efficient in mass utilization and voltage utilization, but less efficient in current utilization than annular one.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.602-605
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• 2008

The anode sheath structure in the hollow anode of an anode-layer type Hall thruster was numerically computed using a fully kinetic 2D3V Particle-in-Cell and Direct Simulation Monte Carlo(PIC-DSMC) code. By treating both ions and electrons as particles, anode surface region, which is electrically non-neutral, was analyzed. In order to analyze in detail, the calculation code was parallelized using Message Passing Interface (MPI). The code successfully simulated the discharge current oscillation. In the low magnetic induction case, ion sheath appears in the anode surface because ionization is enough to maintain the plasma occurs in the anode hollow. As the magnetic induction increases, main ionization region move to outside of the anode. At the same time, anode sheath voltage decreases. In the high magnetic induction case, electron sheath appears on the anode surface periodically because the ionization occurs mainly in the discharge channel. This anode sheath condition shift can be explained using the simple sheath model.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.10
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• pp.791-799
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• 2020

Launched in October 2019, Northrop Grumman's MEV-1 was the world's first unmanned mission demonstrating the practical feasibility of on-orbit servicing. Although the concept of on-orbit servicing was proposed several decades ago, it has been developed to various mission concepts providing services such as orbit change, station keeping, propellant and equipment supply, upgrade, repair, on-orbit assembly and production, and space debris removal. The historical success of MEV-1 is expected to expand the market of on-orbit servicing for government agencies and commercial sectors worldwide. The on-orbit servicing essentially requires the utilization of a highly propellant efficient electric propulsion system due to the nature of the mission. In this study, the space mission analysis for a simple on-orbit mission involving Hall thruster is conducted, which is life extension mission for geostationary orbit satellites. In order to analyze the mission, design space exploration for various Hall thruster design variable combinations is performed. The values of design variables and operational parameters of Hall thruster suitable for the mission are proposed through design space analysis and optimization, and mission performance is derived. In addition, the direction of further improvement for the current on-orbit mission analysis process and space mission analysis involving Hall thruster is reviewed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.10
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• pp.992-997
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• 2010

The Propulsion Control Module(PCM) of Hall-thruster Propulsion System(HPS) for Science and Technology Satellite-3 (STSAT-3) has the flow control accuracy of less than ${\pm}$3% and the pressure control accuracy of less than ${\pm}$5%. The pressure controller adjusts pressure around the set point by using a Proportional Flow Control Valve (PFCV) and a high pressure transducer, while the flow controller regulates the flow rate using PFCV and the anode current telemetry of the Hall Thruster. The controllers are chosen as the Proportional and Integral(PI) type, and the PI gains are tuned based on the Matlab simulations. The result of the PCM test had the flow control accuracy of less than ${\pm}$1.87% and the pressure control accuracy of less than ${\pm}$5%. This paper describes the design, realization, and performance test results of the PCM.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.1
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• pp.84-89
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• 2011

A Xenon Feed System (XFS) has been developed for hall-effect thruster to small satellite space-propulsion system applications. The XFS delivers low pressure gas to the Anode and Cathode of thruster head unit from a xenon storage tank. Accurate throttling of the propellant mass flow rate is independently required for each channel of the thruster head unit. The mass flow rate to each channel is controlled using the accumulator tank pressure regulation through a micron orifice and isolation valve. This paper discusses the Xenon Feed System design including the component selections, performance estimation and functional test.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.11
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• pp.974-980
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• 2014

A power supply of cathode has been developed for hall-effect thruster for small satellite. A microwave cathode has been applied to the electric propulsion system and cathode power is necessary in oder to work securely. Anode current is varied by a flow rate controlled by anode tank pressure. Then cathode current has to be controled in proportion by anode current. So cathode power supply has been designed to offer a current proportional to anode current. Also cathode power has been tuned to work securely for cathode to emit more electron than anode within 0.03A. The function test of cathode power was performed by constructing an equivalent load for anode and cathode. It has been tested in a vacuum chamber in order to ensure a stable operation of the thruster. And it was confirmed that thruster normally has been operated in the space environment after the launch.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.245-250
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• 2004

The discharge current oscillation has been measured for various hollow anode widths and its axial positions using a 1㎾-class anode layer hall thruster. As a result, there were thresholds of magnetic flux density for stable discharge. The plasma structure inside the hollow anode was numerically analyzed using the fully kinetic 2D3V Particle-in-Cell (PIC) and Direct Simulation Monte Carlo (DSMC) methods. The results reproduced both stable and unstable operation modes. In the stable operation case, which corresponds to the case with low magnetic flux, the plasma penetrated into the hollow anode deeper than the case with higher magnetic flux density case. This suggests that comparably large substantial anode area should contribute to stable operation.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.4
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• pp.419-424
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• 2009

A Xenon feed system has been developed for a 300 W Hall-effect thruster intended for orbit maintenance of small satellite. The system can store about 2 kg of xenon gas at 150 bar and is capable of controlling the mass flow rate of the gas at 0.5 SCCM resolution. The performance of the system is verified with a laboratory experiment. It is confirmed that the operation of the feed system is successful at a pressure level of $1.0{\times}10^{-6}$ torr in the vacuum chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.282-286
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• 2004

Discharge current oscillation in a 20KHz range is a serious problem for Hall thruster performance, In our previous work, a possibility of controlling the oscillation amplitude by increasing the speed of neutral particles incoming to the ionization zone was predicted in our previous work. In this paper, the effects of diameter of anode orifice on the oscillation phenomena and the thrust performance were evaluated experimentally. The experimental results show that the measured amplitude of oscillation becomes smaller as the diameter of anode orifice. However, the larger orifice makes thrust performance lower. The results of numerical analysis of neutral particles show that these tendencies have much to do with neutral properties.