• Journal of Astronomy and Space Sciences
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• 제28권2호
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• pp.133-141
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• 2011

A satellite power system should generate and supply sufficient electric power to perform the satellite mission successfully during the satellite mission period, and it should be developed to be strong to the failure caused by the severe space environment. A satellite power system must have a high reliability with respect to failure. Since it cannot be repaired after launching, different from a ground system, the failures that may happen in space as well as the effect of the failures on the system should be considered in advance. However, it is difficult to use all the hardware to test the performance of the satellite power system to be developed in order to consider the failure mechanism of the electrical power system. Therefore, it is necessary to develop an accurate model for the main components of a power system and, based on that, to develop an accurate model for the entire power system. Through the power system modeling, the overall effect of failure on the main components of the power system can be considered and the protective design can be devised against the failure. In this study, to analyze the failure mode of the power system and the effects of the failure on the power system, we carried out modeling of the main power system components including the solar array regulator, and constituted the entire power system based on the modeling. Additionally, we investigated the effects of representative failures in the solar array regulator on the power system using the power system model.

• Journal of Power Electronics
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• 제16권2호
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• pp.666-674
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• 2016

In this paper, a high efficiency satellite electrical power supply system is proposed. The increased efficiency of the power supply system allows for downscaling of the solar array and battery weight, which are among the most important satellite design considerations. The satellite power supply system comprises two units, namely a generation unit and a storage unit. To increase the efficiency of the solar array, a maximum power point tracker (MPPT) is used in the power generation unit. In order to improve the MPPT performance, a novel algorithm is proposed on the basis of the hill climbing method. This method can track the main peak of the array power curve in satellites with long duration missions under unpredicted circumstances such as a part of the array being damaged or the presence of a shadow. A lithium-ion battery is utilized in the storage unit. An algorithm for calculating the optimal rate of battery charging is proposed where the battery is charged with the maximum possible efficiency considering the situation of the satellite. The proposed system is designed and manufactured. In addition, it is compared to the conventional power supply systems in similar satellites. Results show a 12% increase in the overall efficiency of the power supply system when compared to the conventional method.

• 대한전기학회:학술대회논문집
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• 대한전기학회 1995년도 하계학술대회 논문집 A
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• pp.198-201
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• 1995

The overall design criteria for the optimal design of a small LEO satellite power system are described in summary. The analysis result of the KITSAT-I whole orbit data suggests the efficient power opertion for KITSAT-I and also gives some crutial information for developing a new satellite power system.

• 항공우주시스템공학회지
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• 제8권1호
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• pp.42-47
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• 2014

The role of Electrical Power Subsystem (EPS) is to generate a power and distribute it to the electrical devices for the system operation. For on-orbit operation of cube satellite, it is also necessary to supply power to on-board mission devices as commercial satellite does. Recently, commercial EPS products dedicated for the cube satellite application has been developed and widely used for the power subsystem design. In this paper, a permanent magnet attitude stabilization method without external power has been introduced because it has advantage from power consumption point of view and the EPS design of cube satellite by applying the commercial EPS products has been introduced and investigated. This paper also deals with the specification of the commercial EPS products for the beginner of the cube satellite design.

• Journal of Power Electronics
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• 제6권1호
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• pp.18-28
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• 2006

The Electric Power Subsystem (EPS) is one of the most critical systems on any satellite because nearly every subsystem requires power. This makes the choice of power systems the most important task facing satellite designers. The main purpose of the Satellite EPS is to provide continuous, regulated and conditioned power to all the satellite subsystems. It has to withstand radiation, thermal cycling and vacuums in hostile space environments, as well as subsystem degradation over time. The EPS power characteristics are determined by both the parameters of the system itself and by the satellite orbit. After satellite separation from the launch vehicle (LV) to its orbit, in almost all situations, the satellite subsystems (attitude determination and control, communication and onboard computer and data handling (OBC&DH)), take their needed power from a storage battery (SB) and solar arrays (SA) besides the consumed power in the EPS management device. At this point (separation point, detumbling mode), the satellite's angular motion is high and the orientation of the solar arrays, with respect to the Sun, will change in a non-uniform way, so the amount of power generated by the solar arrays will be affected. The objective of this research is to select satellite EPS component types, to estimate solar array illumination parameters and to determine the efficiency of solar arrays during both detumbling and normal operation modes.

• 한국위성정보통신학회논문지
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• 제5권2호
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• pp.29-33
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• 2010

위성 산업은 상업적 군사적 유용성 등의 특성으로 인해 지속적인 발달이 이루어져 왔다. 이중에서도 위성의 전력 시스템은 위성의 수명에 직접적인 관련을 가지고 있으며, 실제 궤도상에서의 Test가 불가능한 특성을 보인다. 또한 저궤도 소형위성은 전력에 민감하므로, 효과적인 전력의 안정화 및 신뢰성 향상은 중요한 문제이다. 일반적으로 저궤도 위성의 전력은 각 부하의 특성에 따라 변환 및 제어가 이루어진다. 따라서 저궤도위성의 전력 시스템은 일반적으로 여러 단계의 전력 변환을 거치게 되므로, 신뢰성 향상을 위한 1차 및 2차 혹은 그 이상 단계의 Converter의 일반화 모델링 및 안정화를 위한 제어기 설계 및 외란에 의한 영향성의 분석이 요구된다. 본 논문에서는 저 궤도 위성 전력계 시스템의 전력 변환을 위한 Converter의 일반화 모델링을 통해 안정화 설계를 위한 파라미터를 추정하고 이를 통한 신뢰성 향상 및 최적 제어 방법에 대해 알아보기로 한다.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2003년도 The Fifth Asian Computational Fluid Dynamics Conference
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• pp.248-250
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• 2003

Thermal analysis is performed to protect the propulsion system of low-earth-orbit earth observation satellite from unwanted thermal disaster like propellant freezing. To implement thermal design adequately, heater powers for the propulsion system estimated through the thermal analysis are decided. Based on those values anticipated herein, the average power for propulsion system becomes 22.02 watts when the only one redundant catalyst bed heater is turned on. When for the preparation of thruster firing, 25.93 watts of the average power is required. All heaters selected for propulsion components operate to prevent propellant freezing meeting the thermal requirements for the propulsion system with the worst-case average voltage, i.e. 25 volts.

• Journal of Power Electronics
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• 제10권2호
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• pp.210-217
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• 2010

This paper introduces a lithium ion battery management system (BMS) for the STSAT-3 satellite. The specifications of a lithium ion battery unit are proposed to supply power to the satellite and the overall electrical and mechanical designs for a lithium ion battery management system are presented. The structural simulation results will be shown to confirm the behavior of both the BMS and the cells.

• 항공우주기술
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• 제11권1호
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• pp.1-6
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• 2012

인공위성의 전력 시스템을 보호하기 위해서 퓨즈나 LCL (Latching Current Limiter) 등을 부하와 전원 공급 장치 사이에 연결한다. 본 논문에서는 인공위성에서 장착되는 퓨즈의 용량 선정 기법에 대해 논하였다. 퓨즈 용량 선정 방법에 대해 기술하였고, 병렬 (Parallel) 퓨즈 연결 방식을 수학식으로도 단일 (Single) 퓨즈 연결 방식으로 나타낼 수 있음을 보였다. 또한, 퓨즈에 연결되는 부하의 정확한 전류 특성이 알려지지 않은 상태에서 적절한 퓨즈 선정 기법을 새롭게 제시하였다.

• International Journal of Aerospace System Engineering
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• 제3권2호
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• pp.31-37
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• 2016

STEP Cube Lab (Cube Laboratory for Space Technology Experimental Project) is a 1U standardized pico-class satellite. Its main mission objective is an on-orbit verification of five fundamental core space technologies. For assuring a successful missions of the STEP Cube Lab with five payloads, electrical power subsystem (EPS) shall sufficiently provide an electrical power to payloads and bus systems of the satellite during an entire mission life. In this study, a design process of EPS system was introduced including power budget analysis considering a mission orbit and various mission modes of the satellite. In conclusion, adequate EPS hardware in compliance with design requirements were selected. The effectiveness and mission capability of EPS architecture design were confirmed through an energy balance analysis (EBA).