• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.9
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• pp.802-808
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• 2014

MEMS based solid propellant thruster researched for the purpose of an academic research will be verified at space environment through CubeSat program. For this, the temperature of the MEMS thruster should be within allowable operating temperature range by proper thermal control to prevent the ignition failure caused by ignition time delay and to guarantee the structural safety of the MEMS thruster in the low temperature. In this study, we proposed an effective thermal control strategy, that is to use micro-igniter as a heater and temperature sensor for active thermal control instead of using additional heater. The effectiveness of the strategy has been verified through on-orbit thermal analysis of CubeSats with MEMS thruster.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.4
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• pp.377-386
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• 2015

DubaiSat-2 is the first satellite developed in Korea equipped with a hall-effect thruster. In this paper, the performance of the DubaiSat-2 hall-effect thruster is verified by analyzing the orbit information of DubaiSat-2. The preparation and performance of orbit operations during 8 months after launch (2013.11.21., UTC) is emphasized and the effects of solar activity on orbit evolution is analyzed. In particular, the hall-effect thruster's thrust is estimated by analyzing difference between observed orbit evolution and predicted orbit. As a result, the estimated thrust is similar to the ground experiment result of 11 mN. The summarized result in this paper would be important reference to improve the stability and effectiveness of satellite operation during the early operation and normal mission lifetime in case of low Earth orbit satellites.

• Aerospace Engineering and Technology
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• v.1 no.1
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• pp.97-105
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• 2002

The mission life of a satellite determines the amount of fuel required on-board, while the total mass requirement limits the fuel to be loaded. Hence, for the design of thruster control loop, not only the satellite pointing accuracy but the saving of fuel is to be considered. In this paper, a two-step fuzzy controller is proposed for the thruster control loop to save fuel consumption. This approach combines requirements for pointing control accuracy with minimum fuel consumption into a fuzzy controller design. To demonstrate this approach, we have designed a fuzzy controller for the Sun Pointing Mode of KOMPSAT-1. The performance of this fuzzy controller design is compared with that of PD controller used for KOMPSAT-1.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.67-70
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• 2006

In this study micro-thrust measurement method in high vacuum chamber is introduced. This is important for the development of micro-thruster for micro-satellite applications. At Chungnam national University, high-vacuum experimental facility has been constructed to simulate space environment. And strain gauge besed micro-thrust measurement in vacuum chamber has been studied and discussed.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.2
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• pp.40-46
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• 2022

Hall thrusters are one of the electric propulsion, where ions are accelerated to generate thrust and are widely utilized in space missions due to their high specific impulses. Recently, as the utilization of small and micro satellites with the mass of similar or less than 100 kg is highly increasing, the importance of research and development of the low-power electric propulsion is also raised. In this study, we developed two sub-200 W or less class, laboratory model Hall thrusters and measured the thrust and analyzed the discharge characteristics. Consequently, we obtained 2.5-9.0 mN of thrust, 600-1,150 s of specific impulse, and 15-28% of anode efficiency at 50-175 W of anode power.

• Aerospace Engineering and Technology
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• v.8 no.1
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• pp.144-153
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• 2009

To study the plume effects in the rarefied region, the Direct Simulation Monte Carlo(DSMC) method is usually adopted because the plume field usually contains the entire range of flow regime from the near-continuum in the vicinity of nozzle exit through transitional state to free molecular at far field region from the nozzle. The objective of this study is to investigate the behaviors of a small monopropellant thruster plume in the rarefied region numerically using DSMC method. To deduce accurate results efficiently, the preconditioned scheme is introduced to calculate continuum flow fields inside thruster to predict nozzle exit properties used for inlet conditions of DSMC method. By combining these two methods, the rarefied flow characteristics of plume such as strong nonequilibrium near nozzle exit, large back flow region, etc, can be investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.9
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• pp.906-915
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• 2009

To study the plume effects in the vacuum area, the Direct Simulation Monte Carlo(DSMC) method is usually adopted because the plume field usually contains the entire range of flow regime from the near-continuum in the vicinity of nozzle exit through transitional state to free molecular at far field region from the nozzle. The objective of this study is to investigate the behaviors of a small monopropellant thruster plume in the vacuum area numerically using DSMC method. To deduce accurate results efficiently, the preconditioned scheme is introduced to calculate continuum flow fields inside thruster to predict nozzle exit properties used for inlet conditions of DSMC method. By combining these two methods, the vacuum flow characteristics of plume such as strong nonequilibrium near nozzle exit, large back flow area, etc, can be investigated.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.171.2-171.2
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• 2012

위성 개발에서 추력기는 위성의 경사각 및 고도 등의 궤도 제어 용도 이외에 위성 동작 초기 혹은 비상 상황에서 안정적인 전력 공급을 위한 자세 제어용 구동기로 사용되어야 하므로 매우 높은 신뢰성을 필요로 한다. 국내의 실용위성을 위해 개발되어 사용되고 있는 출력기는 1 파운드의 작은 용량으로 위성 운영에 일부 제약을 주게 된다. 본 논문은 위성 운영에 있어 반드시 필요한 궤도 천이 절차와 관련하여 기존에 사용된 절차를 보완하기 위한 방법에 대해 기술한다. 기존에 개발된 위성에서는 궤도 조정을 위한 자세 변화에 추력기를 사용하였다. 그러나 위성의 무게가 커짐에 따라 자세 변환을 위한 시간이 오래 걸려 궤도 조정 효율이 떨어지는 요인이 되고 있다. 뿐만 아니라, 자세 변화 과정에서 벡터 방향의 추력으로 인해 원하지 않는 궤도 변화가 생기므로 정밀 궤도 결정에도 영향을 주게 된다. 최근에 개발된 위성의 경우, 위성의 기동 성능을 높이기 위해 고성능 반작용 휠을 사용하므로 이를 이용하여 궤도 천이 전에 자세 변화를 하도록 하고 있다. 이러한 방법을 적용한 결과, 정밀 궤도 결정에 도움이 될 뿐만 아니라 자세 변화로 인한 연료 소모를 줄이는 효과도 있어 위성의 수명 연장에 도움이 되는 것으로 확인되었다.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.30-37
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• 2013

A ground hot-firing test was conducted to take out the optimal design configurations for the catalyst bed of liquid-monopropellant hydrazine thruster which could be used for primary engine or attitude control thruster of space vehicles. Performance characteristics with the variation of thrust-chamber length are investigated in terms of thrust, specific impulse, chamber pressure, characteristic velocity, and hydrazine decomposition rate. Additionally, the correlations between propellant-supply pressure and performance parameters are given. As results, increase of catalyst-bed length leads to performance degradation in this test condition, and also decreases propellant consumption efficiency with the supply pressure variation.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.5
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• pp.65-72
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• 2021

700 W class laboratory model Hall thruster, which can be used for the orbit control or station keeping of small satellites, was developed. The size of the discharge channel was determined using a scaling law, and the magnetic field was designed to be symmetric with respect to the midline of the discharge channel and to be maximized outside the discharge channel. Base pressure of a vacuum chamber was maintained below 2.0×10^-5 Torr during experiments, and the thrust was measured by a thrust stand. The anode flow rate and coil current were varied with the fixed anode voltage at 300 V. Under the operation condition at 2.36 mg/s anode flow rate and 2.4 A coil current, performance was optimized as 38 mN thrust, 1,540 s total specific impulse, and 50 % anode efficiency at 620 W anode power.