• Journal of the Institute of Electronics Engineers of Korea SC
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• v.47 no.3
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• pp.19-27
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• 2010

KSLV(Korea Space Launch Vehicle)-I is designed as a launch vehicle to enter a 100 kg-class satellite to the LEO(Low-Earth Orbit). Attitude angles of the upper-stage, including roll, pitch and yaw are controlled by cold gas thruster system using nitrogen gas. The thruster for the KSLV-I is the main pneumatic valve in the RCS(Reaction Control System). In this paper, the design, function tests, and environment tests of the thruster for KSLV-I are described. The developed thrusters are experimentally evaluated and successfully passed the required qualification and acceptance tests.

• Aerospace Engineering and Technology
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• v.8 no.2
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• pp.122-126
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• 2009

KSLV(Korea Space Launch Vehicle)-I is designed as a launch vehicle to enter a 100 kg-class satellite to the LEO(Low-Earth Orbit). Attitude angles of the upper-stage, including roll, pitch and yaw are controlled by cold gas thruster system using nitrogen gas. To verify the flow rate of the gas charging system and to prepare a nitrogen gas charging scenario, the development of a gas charging simulator for RCS(Reaction Control System) is required. This paper describes the orifice design, development, and test of the gas charging simulator for RCS of KSLV-I.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.1
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• pp.95-102
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• 2006

The PPT being currently developed for the flight model represents a significant leap in techniques and technology compared to the previous flight ones. The electrical energy to be charged in the pulsed plasma thruster (PPT) is a very important aspect to provide an uniform impulse bit ,, and a specific impulse ,, for satellite attitude control. In this paper, we propose a nonlinear control technique and a stability analysis based on the Lyapunov function for the pulsed plasma thruster. Specifically, the proposed control law guarantees to charge and discharge the electrical energy generated from the power processing unit (PPU) within the specified time.

• Advances in aircraft and spacecraft science
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• v.3 no.4
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• pp.427-445
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• 2016

Star sensors are the attitude estimation sensors of the satellite orbiting in its path. It gives information to the control station on the earth about where the satellite is heading towards. It captures the images of a predetermined reference star. By comparing this image with that of the one captured from the earth, exact position of the satellite is determined. In the process of imaging, stray lights are eliminated from reaching the optic lens by the mechanical enclosures of the star sensors called Baffles. Research in space domain in the last few years is mainly focused on increased payload capacity and reduction in launch cost. In this paper, a star sensor baffle made of Aluminium is considered for the study. In order to minimize the component weight, material wastage and to improve the structural performance, an alternate material to Aluminium is investigated. Carbon Fiber Reinforced Polymer is found to be a better substitute in this regard. Design optimisation studies are carried out by adopting suitable design modifications like implementing an additional L-shaped flange, Upward flange projections, downward flange projections etc. A better configuration of the baffle, satisfying the design requirements and achieving manufacturing feasibility is attained. Geometrical modeling of the baffle is done by using UNIGRAPHICS-Nx7.5(R). Structural behavior of the baffle is analysed by FE analysis such as normal mode analysis, linear static analysis, and linear buckling analysis using MSC/PATRAN(R), MSC-NASTRAN(R) as the solver to validate the stiffness, strength and stability requirements respectively. Effect of the layup sequence and the fiber orientation angle of the composite layup on the stiffness are also studied.

• Journal of computational fluids engineering
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• v.21 no.2
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• pp.112-119
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• 2016

NEXTSat-1 is the next-generation small-size artificial satellite system planed by the Satellite Technology Research Center(SatTReC) in Korea Advanced Institute of Science and Technology(KAIST). For the control of attitude and transition of the orbit, the system has adopted a RHM(Resisto-jet Head Module), which has a very simple geometry with a reasonable efficiency. An axisymmetric model is devised with two coil-resistance heaters using xenon(Xe) gas, and the minimum required specific impulse is 60 seconds under the thrust more than 30 milli-Newton. To design the module, seven basic parameters should be decided: the nozzle shape, the power distribution of heater, the pressure drop of filter, the diameter of nozzle throat, the slant length and the angle of nozzle, and the size of reservoir, etc. After quasi one-dimensional analysis, a theoretical value of specific impulse is calculated, and the optima of parameters are found out from the baseline with a series of multi-physical numerical simulations based on the compressible Navier-Stokes equations for gas and the heat conduction energy equation for solid. A commercial code, COMSOL Multiphysics is used for the computation with a FEM (finite element method) based numerical scheme. The final values of design parameters indicate 5.8% better performance than those of baseline design after the verification with all the tuned parameters. The present method should be effective to reduce the time cost of trial and error in the development of RHM, the thruster of NEXTSat-1.

• Journal of Aerospace System Engineering
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• v.12 no.2
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• pp.1-6
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• 2018

The torque ripple that is generated by the irregularity of magnetic flux density on the BLDC motor in a satellite actuator degrades the satellite attitude control performance. In this paper, the performance analysis of permanent magnet configurations (shape, arrangement, and air gap) is simulated by the Finite Element Method (FEM) to find the appropriate combination of the configuration. The configuration is chosen by comparing between rectangular and arc-shaped permanent magnets and single-arrangement and dual-arrangement magnets. The performance is verified by a prototype.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.597-600
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• 2004

Fabrication technique and performance test of catalytic micro propulsion are treated based on MEMS technology. This propulsion is designed to use hydrogen peroxide as liquid mono-propellant for attitude control of pica-satellite. The propellant is fed into the micro reactor channel and decomposed into hot gas yielding controllable thrust by catalyst. In order to increase the efficiency of the reaction that depends on the contact area of propellant and catalyst, porous surface formation on the channel accompanied by platinum particle deposition has been performed using H$_2$PtCl$_{6}$ solution as a precursor. Several thrusters were fabricated in different concentration of H$_2$PtCl$_{6}$ solution to determine the best quantity of Pt particles. For the comparison of the performance of each thruster, the volume of oxygen generated by the decomposition of hydrogen peroxide and the thrust were measured.red.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.314-317
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• 2004

On 9/5/2003, the planet probe “HAYABUSA” as MUSES-C project was launched by The Institute of Space and Astronautical Science. “HAYABUSA” has microwave discharge ion engines and these engines are characterized by their high efficiency and specific impulse in comparison with chemical engine. A large ion engine can be used as a planet explorer, while a small ion engine can be used as attitude control of small satellite. We have been developing a high thrust density microwave discharge ion engine using “Multi-Monopole Antenna”. The performance of this engine are: ion cost of 344W/A, propellant utilization efficiency of 52% and thrust density of 0.055mN/$\textrm{cm}^2$ for Kr gas flow rate of 2.5sccm, microwave(2.45㎓) power of 32W and acceleration voltage of l.4㎸.

• Journal of Magnetics
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• v.21 no.4
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• pp.529-536
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• 2016

Magnetically suspended sensitive gyroscopes (MSSGs) provide an interesting alternative for achieving precious attitude angular measurement. To effectively reduce the measurement error caused by dynamic imbalance, this paper proposes a novel compensation method based on analysis and modeling of the error for a MSSG. Firstly, the angular velocity measurement principle of the MSSG is described. Then the analytical model of dynamic imbalance error has been established by solving the complex coefficient differential dynamic equations of the rotor. The generation mechanism and changing regularity of the dynamic imbalance error have been revealed. Next, a compensation method is designed to compensate the dynamic imbalance error and improve the measurement accuracy of the MSSG. The common issues caused by dynamic imbalance can be effectively resolved by the proposed method in gyroscopes with a levitating rotor. Comparative simulation results before and after compensation have verified the effectiveness and superiority of the proposed compensation method.

• Bulletin of the Korean Space Science Society
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• 2005.04a
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• pp.138-142
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• 2005

A reaction wheel, an actuator for satellite attitude control, produces disturbance torque and force as well as its axial control torque. The disturbances have an influence on the pointing stability of high precision satellites. The measurement of disturbances for such a satellite, therefore, is necessary. The wheel's rotation, however, causes the vibration of the table and its vibration induces measurement errors, especially large near the resonance frequency of the Measurement table. For the purpose of overcoming these defects, a calibration method using frequency compensation is suggested in this paper. Disturbance parameters are identified from data examined by frequency compensation. Measurement frequency range can be expanded far higher than the resonance frequency, since the degradation of data accuracy caused by its vibration is well alleviated even in the resonance area.