• 제어로봇시스템학회:학술대회논문집
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• 2004.08a
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• pp.1966-1971
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• 2004

As the remote sensing satellite technology grows, the acquisition of accurate attitude and position information of the satellite has become more and more important. Due to the data processing limitation of the on-board orbit propagator and attitude determination algorithm, it is required to develop much more accurate orbit and attitude determination, which are so called POD (precision orbit determination) and PAD (precision attitude determination) techniques. The sensor and attitude dynamics simulation takes a great part in developing a PAD algorithm for two reasons: 1. when a PAD algorithm is developed before the launch, realistic sensor data are not available, and 2. reference attitude data are necessary for the performance verification of a PAD algorithm. A realistic attitude dynamics and sensor (IRU and star tracker) outputs simulation considering their physical characteristics are presented in this paper, which is planned to be used for a PAD algorithm development, test and performance verification.

• Advances in aircraft and spacecraft science
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• v.4 no.6
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• pp.701-709
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• 2017

One of the components that used in the satellite thermal control subsystem is the Mechanically Pumped Fluid Loop (MPFL) system; this system mostly used in geosynchronous orbit (GEO) satellites, and can transfer heat from a hot point to a cold point using the fluid which circulated in a closed loop. Heat radiates to the deep space at the cold plate to cool down the fluid temperature. In this research, the radiative heatexchanger (RHX) for a MPFL system is optimized. The genetic algorithm has been used for minimizing the total mass and pressure drop by considering a constant transferred heat rate at the heat exchanger. The optimization has been done in two cases. In case I, two parameters are considered as a goal function, so optimization is performed using NSGA-II method. Results of optimization are shown in the pareto diagram. In case II, the diameter of pipe is considered constant, so the optimized value for distances of the parallel pipes is obtained by using the genetic algorithm, in which the system has the least total mass. Results show that in the RHX, by increasing the pipe diameter, pressure drop decreases and total mass increases. Also by considering a constant value for pipe diameter, an optimum distance between pipes and pipe length are obtained in which the system has a minimum mass.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.8
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• pp.85-90
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• 2003

To control spacecraft including satellite, we should understand precisely the performance of propulsion system and the program logic with appropriate format for satellite operations. In this study, the thruster performance functions was generated by using the best curve fitting for performance data from bi-propellant thrusters. Detailed thruster performance data are, in general, company proprietary information, therefore real firing tests were performed to understand the basic characteristics of the performance curve. Experimental rocket motor utilize liquid oxygen and kerosine as propellant and designed average thrust was 100 pound.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.10
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• pp.92-98
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• 2006

An 1 lbf of NASA standard monopropellant thruster, MRE-1, is used for KOMPSAT (Korea Multi-Purpose Satellite) which is launched in 2006 and provides reliable and cost-effective means for attitude and maneuvering control system. The monopropellant thruster obtains required thrust by thermal decomposition process of propellant through catalyst bed. During firing, the decomposition plays a role of a heat source that may occur an excessive radiation heat transfer to peripheral structures and electronics in relatively low temperature condition.Therefore, the radiation heat shield is needed to prevent the critical radiative heat exchange between thruster and satellite during firing. This paper summarizes an overall development process of radiation heat shield from the design engineering up to the manufacturing.

• Journal of Astronomy and Space Sciences
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• v.15 no.2
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• pp.459-473
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• 1998

As the satellite system becomes more complex, the probability of unpredictable failures may be increased due to design inadequacy, experience deficiency, lack of problem recognition. Poor quality control, improper testing, and workmanship fault. Consequently, these problems can lead to the reduction or end of the satellite mission lifetime. This article addresses general satellite failure modes and factors influencing satellite mission life. The mission life factors of LEO sun-synchronous KOMPSAT spacecraft are investigated, in which its mission life is predicted based on these factors. Since the end of mission due to random failures is not predictable, the predictable mission life factors such as power budget, propellant budget, battery charging/discharging cycle, radiation effects payload reliability, single point failure, and redundancy are primarily investigated.

• Journal of Astronomy and Space Sciences
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• v.15 no.1
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• pp.251-260
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• 1998

We consider the motion of the subsystems as separate bodies as well as the entire satellite for the attitude and orbit control of a communication satellite by multi-body modeling technique. Thus, the system, can be applied to a general communication satellite as well as a specific communication satellite, i. e. Koreasat I,II. The simulation results can be viewed by two-dimensional graphics and three-dimensional animation. The graphical user interface(GUI) makes its usage much simpler. We have simulated a couple of scenarios for Koreasat I,II which are being operated as geostationary communication satellites to verify the system performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.77-84
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• 2006

With the development of micro/nano spacecraft, concepts of micro propulsion are introduced for orbit transfer and drag compensation as well as attitude control. Micro solid propellant thruster has been attention as one of possible solution for micro thruster. In this paper, micro solid propellant thruster is introduced and research on basic components of a micro solid propellant thruster is reported. Micro Pt igniter was fabricated through negative patterning and quantitative effect of geometry was estimated. The characteristic of HTPB/AP solid propellant was investigated to measure the homing velocity. A combustion chamber was fabricated by means of anisotropic etching of photosensitive glass. Finally, micro solid propellant thrusters having various geometries were fabricated and tested.

• Advances in aircraft and spacecraft science
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• v.1 no.1
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• pp.69-91
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• 2014

Numerical investigations are presented, which show that a back-flow flap can improve the dynamic stall characteristics of oscillating airfoils. The flap was able to weaken the stall vortex and therefore to reduce the peak in the pitching moment. This paper gives a brief insight into the method of function of a back-flow flap. Initial wind tunnel experiments were performed to define the structural requirements for a detailed experimental wind tunnel characterization. A structural integration concept and two different actuation mechanisms of a back-flow flap for a helicopter rotor blade are presented. First a piezoelectric actuation system was investigated, but the analytical model to estimate the performance showed that the displacement generated is too low to enable reliable operation. The seond actuation mechanism is based on magnetic forces to generate an impulse that initiates the opening of the flap. A concept based on two permanent magnets is further detailed and characterized, and this mechanism is shown to generate sufficient impulse for reliable operation in the wind tunnel.

• Proceedings of the KSRS Conference
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• v.2
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• pp.791-793
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• 2006

In the satellite camera, the incoming light source is converted to electronic analog signals by the electronic component for example CCD (Charge Coupled Device) detectors. The analog signals are amplified, biased and converted into digital signals (pixel data stream) in the video processor (A/Ds). The outputs of the A/Ds are digitally multiplexed and driven out using differential line drivers (two pairs of wires) for cross strap requirement. The MSC (Multi-Spectral Camera) in the KOMPSAT-2 which is a LEO spacecraft will be used to generate observation imagery data in two main channels. The MSC is to obtain data for high-resolution images by converting incoming light from the earth into digital stream of pixel data. The video data outputs are then MUXd, converted to 8 bit bytes, serialized and transmitted to the NUC (Non-Uniformity Correction) module by the Hotlink data transmitter. In this paper, the video data streams, the video data format, and the image data processing routine for satellite camera are described in terms of satellite camera control hardware. The advanced satellite with very high resolution requires faster and more complex image data chain than this algorithm. So, the effective change of the used image data chain and the fast video data transmission method are discussed in this paper

• Current Industrial and Technological Trends in Aerospace
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• v.8 no.1
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• pp.62-76
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• 2010

In this paper, recent researches about the unmanned lunar lander development are studied. Objectives of this study are to derive research trends and to identify key design activities especially in early design phase of unmanned lunar lander. Case study covers SELENE-2 of Japan, LEDA and MoonNEXT of ESA, and small and modular spacecraft approach of NASA. Lunar lander concepts proposed for the International Lunar Network Anchor Nodes are also studied. For each lunar lander program, mission requirements are summarized and mission design results are reviewed. Approaches of safe lunar landing including design of navigation, guidance and control, combination of sensors, derived sensor and propulsion performance requirements are also analyzed.