• Korean Journal of Remote Sensing
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• v.18 no.1
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• pp.53-59
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• 2002

The MSC is a payload on the KOMPSAT-2 satellite to perform the earth remote sensing. The instrument images the earth using a push-broom motion with a swath width of 15 km and a GSD(Ground Sample Distance) of 1 m over the entire FOV(Field Of View) at altitude 685 km. The instrument is designed to haute an on-orbit operation duty cycle of 20% over the mission lifetime of 3 years with the functions of programmable gain/offset and on-board image data compression/storage. The MSC instrument has one channel for panchromatic imaging and four channel for multi-spectral imaging covering the spectral range from 450nm to 900nm using TDI(Time Belayed Integration) CCD(Charge Coupled Device) FPA(Focal Plane Assembly). The MSC hardware consists of three subsystem, EOS(Electro Optic camera Subsystem), PMU(Payload Management Unit) and PDTS(Payload Data Transmission Subsystem) and each subsystems are currently under development and will be integrated and verified through functional and space environment tests. Final verified MSC will be delivered to spacecraft bus for AIT(Assembly, Integration and Test) and then COMSAT-2 satellite will be launched after verification process through IST(Integrated Satellite Test). In this paper, the introduction of MSC, the configuration of MSC electronics including electrical interlace and design of CEU(Camera Electronic Unit) in EOS are described. MSC Operation parameters induced from the operation concept are discussed and analyzed to find the influence of system for on-orbit operation in future.

• Journal of Korea Society of Industrial Information Systems
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• v.18 no.5
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• pp.81-92
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• 2013

The satellite includes many release mechanisms such as solar array deployment, antenna deployment, cover to protect contamination in scientific equipment, pyro value of the propulsion subsytem, and bypass device in Li-Ion cell module. A drive the initiators is a critical to the successful mission because the initiators of release mechanism driving by the pyrotechnic circuit is operated in single short. The pyrotechnic circuit has to provide switching network for safety. A typical switching network has defect consisting of high current rating fire switch to handle switching transient current during fire the initiator. The pyrotechnic circuit is required some form of power conditioning to reduce the peak power demanded from the bus if the initiators are to be fired from the main bus. This paper design a pyrotechnic circuit synchronized to the fire-command to activate the fire switch to overcome use high current rating fire switch to handle switching transient current during fire the initiator. The pyrotechnic circuit provides a current limited widow pulse for fire current synchronized to the fire-command to insure that fire switch will only carry the current but never switch it. The current limited widow pulse for fire current can be possible to use low current rating and light mass switch in switching network. The current limit function in the pyrotechnic circuit reduces supply voltage to initiator and provides the effect of power conditioning function to reduce peak bus power. The pyrotechnic circuit to apply satellite development on geostationary orbit is verified the function by test in development model.

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

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.7
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• pp.586-593
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• 2014

The CNUSAIL-1 project aims to develop and operate a 3U-sized cube satellite with solar sail mechanism. The primary mission is to successfully deploy the solar sail in a low earth orbit, and the secondary mission is to collect the scientific data for the effect of the solar sail deployment and operation on orbit maneuver and attitude change of the cube satellite. For this, the bus system will collect and transmit the dynamic data of the satellite and the visual images of the solar sail operation. This paper describes solar sail mission and conceptual design of CNUSAIL-1. The actuation/operation of the solar sail and the bus system are preliminarily designed in terms of attitude control system, communication system, electrical power system, command and data handling system, structure and thermal control system is designed.

• Journal of Astronomy and Space Sciences
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• v.20 no.4
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• pp.339-350
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• 2003

Engineering qualification model payload for a communications and broadcasting satellite(CBS) was developed by ETRI from May, 2000 to April, 2003. For. the purpose of functional test and verification of the payload, a real-time hardware-in-the-loop(HITL) CBS simulator(CBSSIM) was also developed. We assumed that the spacecraft platform for the CBSSIM is a geostationary communication satellite using momentum bias three-axis stabilization control technique based on Koreasat. The payload hardware is combined with CBSSIM via Power, Command and Telemetry System(PCTS) of Electrical Ground Support Equipment(EGSE). CBSSIM is connected with PCTS by TCP/IP and the payload is combined with PCTS by MIL-STD-1553B protocol and DC harness. This simulator runs under the PC-based simulation environment with Windows 2000 operating system. The satellite commands from the operators are transferred to the payload or bus subsystem models through the real-time process block in the simulator. Design requirements of the CBSSIM are to operate in real-time and generate telemetry. CBSSIM provides various graphic monitoring interfaces and control functions and supports both pre-launch and after-launch of a communication satellite system. In this paper, the HITL simulator system including CBSSIM, communications payload and PCTS as the medium of interface between CBSSIM and communications payload will be described in aspects of the system architecture, spacecraft models, and simulator operation environment.

• Journal of Satellite, Information and Communications
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• v.10 no.4
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• pp.95-100
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• 2015

These days, satellite core technologies are being developed as a way to provide various information by considering simultaneously sending, wide area covering, highly precide, and anti-disaster technologies. Not only global positioning, and image but also space launcher, satellite bus, satellite payload, earth station are being convergently developed in a different technological field. Especially, it is required a lot of initial investing expenditure to provide the Earth observational information service based on the space technologies. Such a trend and change of satellite technologies Korea has realized the necessity for the domestic independent development of next generation earth observation satellites, and are preparing the profound items such as a detailed implementation plan for the efficient development project. Like the satellite advanced countries, it should be transparently carried out that an efficient implementation of the developing target related to the geostationary earth observation satellite development, establishment of technological auditing function and quality assurance system, implementation plan, progressing courses and results of the satellite development program by way of planning, evaluation and management. For these things cited above, it is necessary to operate systematically and continuously the professional structural system by the governmental department in order to control the geostationary earth observation satellite development project. Therefore, this study proposes a development project management improvement method of the Korea next generation geostationary earth observation satellite based on the development project management system of the domestic geostationary satellite system.

• 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.

• Aerospace Engineering and Technology
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• v.10 no.2
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• pp.20-28
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• 2011

The electro-optical payload of a low-earth orbit satellite is thermally decoupled with the bus, which supports a payload for a mission operation. The payload has a cooling unit of FPA(Focal Plane Assembly) which has a thermal behavior increasing its temperature instantly during an operation in order to dissipate a waste heat into the space. The FPA cooling unit should include a radiator and heatpipes with a sufficient performance in worst hot condition, and a heater design to maintain its temperature above a minimum allowable temperature in the worst cold condition. In this paper, we analyzed the thermal requirements and the heater design constraints from the thermal analysis results for the current thermal design of the FPA cooling unit and the design elements of the better heater design were found.

• Journal of computational fluids engineering
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• v.10 no.3 s.30
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• pp.63-69
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• 2005

The COMS, the first meteorological geostationary satellite in Korea, is under development by KARI. The radiator size and the heater power for the thermal control of COMS are calculated using an analytic method. The total radiator area of $4.85\;m^2$ and the total heater power of 794.77 W are determined at a conceptual design of COMS. The commercial software, SINDA and TRASYS, are utilized in order to compare and verify the analytic results. The results of on-orbit numerical simulation of cold and hot cases show that the radiator size and heater power obtained from the analytic method are appropriate to maintain COMS equipments within required temperature ranges.

• The Transactions of the Korean Institute of Power Electronics
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• v.10 no.4
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• pp.403-410
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• 2005

This Paper Presents development of solar may shunt switching unit with a fully regulated Power regulation for Geostationary Earth Orbit(GEO) satellite. This shunt switching unit comprises the solar may shunt modules that regulate the solar array power. These solar array shunt modules connect/disconnect the solar array segments to/from the bus through switching actions. And that is also possible simply extension to an existing design by FPGA control logic changing. In order to verify the proposed design, the control logic and worst case analysis are analyzed and the simulation and experimental results we shown.