• 한국전산유체공학회:학술대회논문집
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• 2006.05a
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• pp.61-62
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• 2006

• 제어로봇시스템학회:학술대회논문집
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• 1998.10a
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• pp.53-57
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• 1998

Star tracker placement configuration is proposed and the properness of the placement configuration is verified for star tracker's sun avoidance angle requirement. Precision attitude determination system is successfully designed using a gyro-star tracker inertial reference system for a candidate LEO spacecraft. Elaborate kalman filter formulation for a spacecraft is proposed for covariance analysis. The covariance analysis is performed to verify the capability of the proposed attitude determination system. The analysis results show that the attitude determination error and drift rate error are good enough to satisfy the mission of a candidate spacecraft.

• Aerospace Engineering and Technology
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• v.11 no.1
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• pp.42-48
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• 2012

The telemetry data received from satellite in real-time are used to monitor LEO satellite during the AIT (Assembly, Integration & Test) phase and the mission operation phase after launch. However, it is impossible to check all the incoming telemetry data from satellite in real time in order to detect abnormality of satellite quickly. Especially, the contact time of LEO satellite is limited because of its orbital characteristics. So the anomaly state of the LEO satellite should be detected and resolved during the contact time. Therefore, all incoming spacecraft telemetry data must be selected and manipulated in MIMIC. It is used in order to display summarized information about spacecraft in a visualized way that is quickly and easily understood. That is, it provides essential function to monitor a satellite both in orbit and during testing. In this paper, the design and development of MIMIC currently used in KOMPSAT, a LEO Earth observation satellite is described in detail. In future work, we plan to enhance MIMIC in order to improve user-friendliness and efficiency.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.85-89
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• 2006

Propulsion System provides the required velocity change impulse for orbit transfer from parking orbit to mission orbit and three-axis vehicle attitude control impulse. New LEO Satellite propulsion system (PS) will be an all-welded, monopropellant hydrazine system. The PS consists of the subassemblies and components such as Thrusters, Propellant Tank, Pressure Transducer, Propellant Filter, Latching Isolation Valves, Fill/Drain Valves, interconnecting propellant line assembly, and thermal hardwares for operation-environment control of the PS. In this study, preliminary design process of LEO Satellite propulsion system will be summarized.

• Aerospace Engineering and Technology
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• v.11 no.2
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• pp.33-38
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• 2012

A LEO Satellite performs automatic initial operations by FSW after separation from a launch vehicle. After initial operation by FSW is finished, preparation for normal operation is performed by ground during bus initial activation and checkout phase. First of all, we check state of health of the satellite including solar array deployment status. After then, each unit of spacecraft bus is activated and checked. After activation and checkout of every units used for normal operation, we check maneuver performance for imaging mission and orbit maintenance performance. Because the Bus IAC is performed during limited ground contact time, every detailed procedure must be designed considering ground contact. Therefore, the Bus IAC procedure is separated into several parts based on ground contact duration. In addition, the procedures for every possible operation including expected situation as results of IAC procedures and unexpected contingency situation must be prepared. The contingency operation is also designed based on ground contact duration. The LEO satellite was successfully launched and the Bus IAC was successfully performed. In this paper, we explain design concepts and execution results of Bus IAC.

• Journal of Satellite, Information and Communications
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• v.6 no.1
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• pp.1-5
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• 2011

In recent space industry, It has become a major trend to launch lunar exploration satellites to extend activities in the deep space environment. In this paper, a link budget analysis is carried out for the lunar exploration satellite. One of the major difference between the lunar satellite and LEO spacecraft lies in the orbit parameters. The vast distance between spacecraft and the Earth station imposes a challenging task for the spacecraft designers in terms of achieving stable communication link budget. The satellite tool kit software has been adopted to simulate the lunar exploring satellite. The relative distance between the spacecraft and the ground stations are tracked and the communication link budget is calculated accordingly.

• Aerospace Engineering and Technology
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• v.10 no.1
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• pp.193-200
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• 2011

The external disturbance torque acting on a low earth orbit spacecraft was analyzed. For the Earth pointing attitude, the maximum torque to the spacecraft is about $8.3{\times}10^{-4}$ Nms and the momentum accumulated for an orbit is about 1.4 Nms and for the Sun pointing attitude, the maximum torque to the spacecraft is about $1.6{\times}10^{-3}$ Nms and the momentum is accumulated about 3.0 Nms in the spacecraft body reference frame. The analysis results confirm that the size of magnetic torquer selected previously for the satellite is sufficient to manage the accumulated momentum by considering the dumping capacity for an orbit.

• Advances in aircraft and spacecraft science
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• v.9 no.4
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• pp.279-300
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• 2022

This paper analyzes the possibilities of a new method to using TLE data for detecting satellite maneuvers. The method has a number of advantages over other methods that are designed to detect maneuvers. It allows not only to detect maneuvers, but also to get a more complete picture of the maneuver. In particular, the method makes it possible to estimate the moments of the beginning and end of the maneuver, calculate the changes in the orbital elements, evaluate the tangential and binormal components of the impulse, and finally, calculate the impulse of the satellite obtained as a result of the maneuver. To demonstrate in detail the capabilities of the algorithm, the proposed method was applied to one of LEO satellites - TIANHUI-1 satellite. After the efficiency of the method was proved, this method was applied to the China Space Station - TIANHE-1 (CSS), Starlink-1095 and Starlink-2305 satellites. The maneuvers of the CSS and Starlink-1095 satellite during their close encounter on 1 July, 2021, and the CSS and Starlink-2305 satellite during their close encounter on 21 October, 2021 are analyzed in detail. The minimum distances between the CSS and Starlink satellites at the moments of their maximum approaches are estimated. An estimate of the computation time of this algorithm is given, and the possibility of its use for monitoring maneuvers or other anomalous orbital changes of a large number of satellites in near real-time is shown. It is assumed that on the basis of this method, a service for monitoring satellite maneuvers can be created.

• International Journal of Aeronautical and Space Sciences
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• v.6 no.1
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• pp.61-70
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• 2005

The currently developed propulsion system(PS) is composed of propellant tank, valves, thrusters, interconnecting line assembly and thermal hardwares to prevent propellant freezing in the space environment. Comprehensive engineering analyses in the structure, thermal, flow and plume fields are performed to evaluate main design parameters and to verify their suitabilities concurrently at the design phase. The integrated PS has undergone a series of acceptance tests to verify workmanship, performance, and functionality prior to spacecraft level integration. After all the processes of assembly, integration and test are completed, the PS is integrated with the satellite bus system successfully. At present, the severe environmental tests have been carried out to evaluate functionality performances of satellite bus system. This paper summarizes an overall development process of monopropellant propulsion system for the attitude and orbit control of LEO(Low Earth Orbit) observation satellite from the design engineering up to the integration and test.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.2
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• pp.141-148
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• 2015

This paper presents a sun-synchronous orbit design which effectuvely includes the requirements derived from spacecraft to ground station contact and spacecraft to target image accessibility. For this purpose, operation parameters of multiple spacecraft are defined as Contact Overlap, Contact Overlap Gap, Access Overlap, Access Overlap Gap. These parameters are used to form a Figure of Merit that reflects the operational requirements. The Figure of Merit is optimized to increase the efficiency of operating multiple spacecraft in constellation and is used to determine the operational orbit of each spacecraft that constitutes the constellation.