• Journal of Aerospace System Engineering
• /
• v.8 no.1
• /
• pp.42-47
• /
• 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.

• International Journal of Aerospace System Engineering
• /
• v.3 no.2
• /
• pp.31-37
• /
• 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).

• Proceedings of the KIPE Conference
• /
• 2002.07a
• /
• pp.681-684
• /
• 2002

In this paper, electrical power subsystem(EPS) simulator for KOMPSAT 2 has been analyzed designed and simulated by object oriented design(OOP) method. To design EPS simulator, the EPS modules, which modeled solar array, solar array regulator, deployment device controller, battery, power control unit, and EPS control unit, are modeled. To verify the EPS simulator, the modules has been simulated. By OOP designs, the EPS simulator is very powerful because this method is applicable to design other EPS simulator.

• Aerospace Engineering and Technology
• /
• v.1 no.1
• /
• pp.72-83
• /
• 2002

Required power and solar array sizing of KOMPSAT-2 have been analyzed by ASTRIUM and KARI in November, 2000. There are Electrical Power Subsystem(EPS) design discrepancies between ASTRIUM and Korea Aerospace Research Institute(KARI) according to heritage program, EPS operation concepts, power source and the characteristic of the electrical boxes. To design the power system of KOMPSAT-2, ASTRIUM has used the EPS design of the CHAMP and GlobalStar program. But SSTI, TOMS-EP and KOMPSAT-1's design concepts has been used for KOMPSAT-2 EPS design by the KARI. To get the design conclusion, there are many trade-off meetings for the EPS sizing using each sides' heritage program and EPS operation concept. And the EPS design factors and approaching methods have been reviewed and discussed. In addition the EPS design results from ASTRIUM and KARI are summarized in this paper.

• Proceedings of the KSRS Conference
• /
• v.1
• /
• pp.220-223
• /
• 2006

The COMS(Communication, Ocean and Meteorological Satellite) EPS(Electrical Power Subsystem) is derived from an enhanced Eurostar 3000 EPS which is fully autonomous operation in normal conditions or in the event of a failure and provides a high level of reconfiguration capability and flexibility. This paper introduces the COMS EPS preliminary design result. The COMS EPS consists of a battery, a solar array wing, a PSR(Power Supply Regulator), a PRU(Pyrotechnic Unit), a SADM(Solar Array Drive Mechanism) and relay and fuse brackets. This can offer a bus power capability of 3 kW. The solar array is made of a deployable wing with two panels. One type of solar cells is selected as GaAs/Ge triple junction cells. Li-ion battery is base lined with ten series cell module of five cells in parallel. PSR associated with battery and solar array generates a power bus fully regulated 50 V. Power bus is centralised protection and distribution by relay and fuse brackets. PRU provides power for firing actuators devices. The solar array wing is routed by the SADM under control of the AOCS(Attitude Orbit Control Subsystem). The control and monitoring of the EPS especially of the battery, is performed by the PSR in combination with on-board software.

• Journal of Astronomy and Space Sciences
• /
• v.24 no.2
• /
• pp.179-194
• /
• 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 Korea Society of Industrial Information Systems
• /
• v.19 no.1
• /
• pp.31-41
• /
• 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.

• Proceedings of the KIEE Conference
• /
• 1996.11a
• /
• pp.355-358
• /
• 1996

Electrical Power System (EPS) of Electric Vehicle which consists of batteries, motor and driving subsystem, has been modeled. A battery model is modeled with an electrical circuit representing a characteristics of real battery. Driving subsystem is modeled as three different level namely exact, average and functional models. Load profile includes road information, speed profile and EV mechanical parameters, which are incorporated into a reference torque in the driving subsystem model. A system model is integrated to simulate the performance of electric vehicle such as energy balance, battery status, and electrical stress of each subsystem.

• Journal of Satellite, Information and Communications
• /
• v.1 no.2
• /
• pp.95-100
• /
• 2006

The COMS(Communication, Ocean and Meteorological Satellite) EPS(Electrical Power Subsystem) is derived from an enhanced Eurostar 3000 version. Eurostar 3000 EpS is fully autonomous operation in nominal conditions or in the event of a failure and provides a high level of reconfigure capability. This paper introduces the COMS EPS preliminary design result. COMS EPS consists of a battery, a solar arrat wing, a PSR(Power Supply Regulator), a PRU(Pyrotechnic Unit), a SDAM(Solar Array Drive Mechanism) and relay and fuse brackets. COMS EPS can offer a bus power capability of 3 kW. The solar array is made of a deployable wing with two panels. One type fo solar cells is selected ad GaAs/Ge triple junction cells. Li-ion battery is base lined with ten series cell module of five cells in parallel. PSR associated to battery and solar array wing generates a power bus fully regulated at 50 V. Power bus os centralized protection and distribution by relay and fuse brackets. PRU provides power for firing actuarors devices. The solar array wing is rotated by the SADM under control of the attitude orbit control subsystem. The control and monitoring of the EPS, especially of the battery, is performed by the PSR in combination with the on-board software.

• Proceedings of the KIPE Conference
• /
• 2005.07a
• /
• pp.360-362
• /
• 2005

In this paper, standard model for electrical power subsystem of a satellite simulator is presented and analyzed. The main purpose of standard model simulator is to promote platform independency, interoperability and reusability of simulation models. And, EPS simulator prototype model is proposed using the SMP2 standard.