• Journal of Astronomy and Space Sciences
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• v.24 no.2
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• pp.179-194
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• 2007

The design and analysis of satellite power subsystem is an important driver for the mass, size, and capability of the satellite. Every other satellite subsystem is affected by the power subsystem, and in particular, important issues such as launch vehicle selection, thermal design, and structural design are largely influenced by the capabilities and limitations of the power system. This paper introduces a new electrical power subsystem design program for the rapid development of LEO satellite and shows an example of design results using other LEO satellite design data. The results shows that the proposed design program can be used the optimum sizing and the analytical prediction of the on-orbit performance of satellite electrical power subsystem.

• Journal of Aerospace System Engineering
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• v.8 no.1
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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 Korea Society of Industrial Information Systems
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• v.19 no.1
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• pp.31-41
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• 2014

The satellite on geostationary orbit accommodates multiple payloads into a single spacecraft platform and launched in June 26, 2010. The Electrical Power Subsystem provides a fully regulated power bus at $50V_{DC}$ in sunlight and eclipse conditions. The electrical power required to the satellite is generated by a solar array wing and the energy is stored by a Li-Ion battery with a capacity of 192.5Ah. This paper selects the main design parameters, compares and analyzes with the results at ground test and in orbit operation to apply this performance evaluation of the Electrical Power Subsystem to next satellite design on geostationary orbit. The Electrical Power Subsystem is demonstrated nominal behavior without significant degradation through the performance evaluation from design to in orbit operation.

• Journal of Power Electronics
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• v.6 no.1
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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.

• The Transactions of The Korean Institute of Electrical Engineers
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• v.57 no.4
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• pp.543-550
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• 2008

This paper proposes an voltage and reactive power control methodology, which is motivated towards implementation in the korea power system. The main voltage control devices are capacitor banks, reactor banks and LTC transformers. Effects of control devices are evaluated by local subsystem's cost computations. This local subsystem is decided by 'Tier' and 'Electrical distance' in the whole system. The control objective at present is to keep the voltage profile within constraints with minimum switching cost. A robust control strategy is proposed to make the control feasible and optimal for a set of power-flow cases that may occur important event from system. This studies conducted for IEEE 39-bus low and high voltage contingency cases indicate that the proposed control methodology is much more effective than PSS/E simulation tool in deciding switching of capacitor and reactor banks.

• Journal of Satellite, Information and Communications
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• v.2 no.2
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• pp.36-40
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• 2007

EGSE is used to check out satellite payload during the development prior to launch. The EGSE represented in this paper is a test system for Ka band communication transponder of COMS. The EGSE consist of two subsystems as CTS subsystem and PCTS subsystem. Communication Test subsystem (CTS) performs satellite transponder RF performance testing, data analysis and trending. Most of transponder RF performances are automatically tested by the CTS subsystem. Power, Command & Telemetry subsystem (PCTS) monitor telemetry messages from the transponder and send tele-commands to satellite transponder for the configuration change. PCTS also provide simulated S/C power to the transponder during the ground validation testing. The EGSE test functions are verified by the transponder simulator testing and will be used for the flight model transponders testing.

• Proceedings of the KSR Conference
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• 2002.10a
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• pp.113-118
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• 2002

Vibrational characteristics of coupled beam ＆ plate system are considered on simple system, which consists of plates(2-subsystem) and beams(4-subsystem), using experimental statistical energy analysis(ESEA). First, damping and coupling loss factors of the system are determined by power injection method (PIM). Then, energy distribution of all the subsystem is estimated from the power balance equation. Finally, these quantities are compared with measured energy. The correlation of measured and estimated results for the sample problem is reasonably good.

• International Journal of Aerospace System Engineering
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• v.3 no.2
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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).

• Aerospace Engineering and Technology
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• v.12 no.2
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• pp.14-23
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• 2013

Fault management design of the satellite describes preparations for failures which can occur during operational phase. Fault management design contains detection and isolation function of anomaly, and also it contains function to maintain the satellite in safe condition until the ground station finds out a cause of failure and takes a countermeasure. Unlike normal operation, safing operation is automatically performed by Power Control and Distribution Unit and Integrated Bus Management Unit which loads Flight Software without intervention of ground station. Since fault management operation is automatical, fault management logic and functionality of relevant hardware should be thoroughly checked during ground test phase, and error which is similar to actual should be carefully applied without damage. Verification test for fault management design is conducted for various subsystems of satellite. In this paper, we show the design process of fault management design verification test for Electrical Power Subsystem and Attitude and Orbit Control Subsystem of Low Earth Orbit satellite flight model and the test results.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2005.11a
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• pp.79-83
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• 2005

SEA is a useful tool to predict noise and vibration response in high frequency region but has a weak point not to be able to express modal behavior in low frequency region. For a structure with middle subsystem having relatively higher modal density than excited subsystem and receiving subsystem, we studied the possibility that the modal behavior of receiving subsystem can express by considering finite mobility of excited subsystem. For a simply three-coupled beams which is chosen for feasibility study, the response of receiving beam was investigated with varying the length & area moment of inertia of middle beam. In case that the middle beam has relatively higher modal density than exciting beam, the application to finite mobility of excited beam led to express modal behavior of receiving beam relatively well.