• Proceedings of the KIEE Conference
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• 1995.07b
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• pp.907-909
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• 1995

The SaTReC is to develop, deploy, and operate a low Earth orbiting small satellite system, KITSAT-3, carrying a remote sensing payload, a space science payload, and a data collection system. Through the development of KITSAT-3, the SaTReC is to demonstrate the small satellite system which provides highly accurate attitude control, high speed data transmission, and a unique spacecraft configuration and to provide educational opportunities to Korean space industries and research institute. The KITSAT-3 is expected to be launched in the beginning of 1997 by Chinese Long March IV as a secondary payload into about 800 km's sunsynchronous orbit.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.3
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• pp.45-50
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• 2004

The design of reaction wheel control logic is critical to achieve the spacecraft attitude stabilization and performance requirements for the successful mission. Due to various uncertainties on orbit there exist limitation to obtain the model parameters through the ground tests and to design the associated control logic. Thus, the model parameter correction using on-orbit data is essential to the control performance on orbit. This paper performs the system identification using KOMPSAT-1 telemetry data and extracts the model parameters of the reaction wheel. Moreover, the reaction wheel is remodeled and compared with the ground test results.

• Journal of Astronomy and Space Sciences
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• v.14 no.2
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• pp.386-394
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• 1997

In this study, a few of the modeling methods for flexible spacecraft were introduced and adopted to the modeling of a 3-axes stabilization satellite. The generated model was put into pre-built rigid body attitude control loop. A Lumped Parameter Model(Global Mode Model: GMM) was recommended for the absence of the Finite Element Method(FEM) model. Finally, GMM was compared with FEM in terms of designing a control filter. A 1st-order filter was designed to meet requirements of the controller since the new flexible model was applied, and that filter was added to motor controller and axis controller. MATLAB/Simulink was used as a tool for design and simulation of the control loop and filter.

• Aerospace Engineering and Technology
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• v.3 no.1
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• pp.97-102
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• 2004

Propulsion System provides the required velocity change impulse for orbit transfer from parking orbit to mission orbit and three-axis vehicle attitude control impulse. KOMPSAT-2 propulsion system(PS) is 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. This paper summarizes a development process of the liquid propulsion system from the design engineering up to the test and evaluation.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.10
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• pp.875-887
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• 2015

The optimal layout design is used in the development of various areas of industry. In the field of space systems, components must be placed properly in the limited space of spacecraft by considering mechanical, thermal and electrical interfaces. When applying optimal layout design, a proper, even ideal placement of components is possible in the limited space of a satellite platform. Through the optimal placement design, the minimized moment of inertia enhances efficient attitude control, rapid maneuver and mission performance of the satellite. This paper proposes 3D optimal layout design that minimizes the spacecraft's moment of inertia and effect of thermal dissipation between inner components as well as interference between inner components based on a cubic-structure satellite platform. This study proposes the new genetic algorithm for 3D optimal layout design of the satellite platform.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.593-596
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• 2004

We are developing a micro-solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft. The prototype has ø 0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 22 x 22 mm substrate. In previous studies, an impulse thrust of 4.6 x 10$^{-4}$ Ns was obtained in vacuum, but we found the problems of unacceptably low ignition success rate and incomplete combustion. This paper describes experiments to improve the ignition rate. In order to achieve this goal, we tried to solidify paste-like ignition aid (RK) on the ignition heaters with strong adhesion. To make the paste-like RK, isoamyl acetate was added to RK powder. We tested 9 rockets, but only 2 rockets were ignited with huge ignition energy. This is because the heat con-duction between the ignition heater and the RK was too low to ignite the RK, since dried RK had a lot of pores. Also, a large cavity was sometimes found just above the ignition heater.

• 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.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.4 no.5
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• pp.503-514
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• 2017

The attitude control of an aircraft is usually fulfilled by means of thrusters at high altitudes. Therefore, the possibility of using also aerodynamic surfaces would produce the advantage of reducing the amount of fuel for the thrusters to be loaded on board. For this purpose, Zuppardi already considered some aerodynamic problems linked to the use of a wing flap in a previous paper. A NACA 0010 airfoil with a trailing edge flap of 35% of the chord, in the range of angle of attack 0-40 deg and flap deflections up to 30 deg was investigated. Computer tests were carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km of Earth Atmosphere. The present work continues this subject, considering the same airfoil and free stream conditions but two flap extensions of 45% and 25% of the chord and two flap deflections of 15 and 30 deg. The main purpose is to compare the influence of the flap dimension with that of the flap deflection. The present analysis is carried out in terms of: 1) percentage variation of the global aerodynamic coefficients with respect to the no-flap configuration, 2) increment of pressure and heat flux on the airfoil lower surface due to the Shock Wave-Shock Wave Interaction (SWSWI) with respect to the same quantities with no SWSWI or in no-flap configuration, 3) flap hinge moment. Issues 2) and 3) are important for the design of the mechanical and thermal protection system and of the flap actuator, respectively. Under the above mentioned test and geometrical conditions, the flap deflection is aerodynamically more effective than the flap extension, because it involves higher variation of the aerodynamic coefficients. However, tests verify that a smaller deflection angle involves the advantage of a smaller increment of pressure and heat flux on the airfoil lower surface, due to SWSWI, as well as a smaller hinge moment.