• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.7-14
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• 2011

Various sizes of hydrazine monopropellant thruster have been used on satellite and space launcher vehicle. The test and handling procedure of hydrazine monopropellant thruster are usually difficult because of the toxicity of hydrazine and its decomposition product gases. Therefore, the numerical analysis can help understand the effects of various design parameters and can reduce the time as well as expenses. In this study, the numerical analysis is performed by modelling the catalyst bed as one dimensional porous medium. Thereby, resulting physical phenomena are examined by considering the variation of catalyst bed characteristics incurred by catalyst granule failure.

• Journal of Aerospace System Engineering
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• v.11 no.5
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• pp.65-75
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• 2017

This study proposes development of 3U CubeSat standard platform whose function and performance are verified via KAUSAT-5 development. 3U CubeSat platform specification was selected for the design of 3U Cubesat standard platform by examining existing CubeSat and state-of-art technology, and consequently a universally usable 3U CubeSat platform was designed. Standard platform was manufactured in 1.5U size and developed with a modular concept to be able to add and expand payloads and ADCS actuators for meeting the user's needs. In addition, in case of the power system, the solar panel, the battery, and the deployment mechanism are designed to be configured by the user. In the mechanical system design of a standard platform, subsystem and micro equipment functions/performance could be integrated and miniaturized on micro-sized PCBs and maximized electrical capability to accommodate multiple payloads. In the development of the 3U CubeSat, the satellite platform adopts the developed standard platform, which can reduce the cost and schedule for the whole satellite development by reducing the additional function verification.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.6
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• pp.455-466
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• 2020

In this study, we built a thermal model for SNIPE 6U nano-satellite which has scientific mission for measuring science data in near Earth space environment and described thermal design based on the thermal model. And the validity of the thermal design was verified through the on-orbit thermal analysis. The thermal design was carried out mainly on the passive thermal control techniques such as surface finishes, insulators, and thermal conductors in consideration of the characteristics of the nano-satellite. However, the components with narrow operating temperature range and directly exposed to the orbital thermal environments, such as a battery and thrusters, are accomodated with heaters to satisfy the temperature requirements. On-orbit thermal analysis conditions are based on the basic orbital conditions of the satellite, and thermal analysis was performed for Normal mode, Launch & Early Orbit Phase (LEOP), Safehold mode, and Maneuver mode which are classified by the power consumption and the attitude of the satellite according to the mission scenario. The analysis results for each mode confirmed that every component satisfies the temperature requirement. In addition, the heater capacity and duty cycle of the battery and thruster were calculated through the analysis results of the Safehold mode.

• Proceedings of the KIEE Conference
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• 2008.10b
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• pp.357-358
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• 2008

본 논문은 GPS를 이용한 무인 비행체의 자율비행에 관한 연구를 다루었다. 비행체의 종류는 크게 고정익기와 회전익기로 나뉘는데 본 연구에서는 회전익기의 형태를 가진 Quarter Vehicle을 사용하였다. Quarter Vehicle은 4개의 프로펠러에 의한 양력과 회전 반발력으로 비행을 한다. 이때의 양력은 수평면에 대해 수직으로 추력을 발생시키므로 다른 비행체보다 불안정하며 이를 안정하게 제어하기 위해 관성 센서를 적용하여 균형을 유지한다. 본 연구에서는 UAV의 자세와 균형을 안정적으로 유지하기 위해 관성센서를 적용하였으며 GPS를 활용하여 독립적인 자율비행이 가능하도록 하였다. 정확한 위치정보를 제공하는 GPS는 3개 이상의 위성으로부터 시간 및 위치 정보를 받아 좌표를 계산하고 비행체의 위치, 속도 및 방향을 결정하여 자율 비행이 가능하도록 한다. 또한 초형 지자기센서를 비행체에 적용하여 GPS의 방향 정보를 보완하도록 하였다. 본 논문에서는 무인 비행체의 자율비행의 기초가 되는 위치좌표 계산을 위한 GPS의 적용 방법과 비행경로계획 알고리즘을 제시 한다.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.36 no.11
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• pp.1399-1404
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• 2012

Because satellites operate in space, it is impossible to repair them when they malfunction. Therefore, to ensure the normal function of the payload used in the satellites, accurate assembly and installation of parts are crucial. To prevent abnormal functioning in the extreme environments during launch and in space, it is essential to test changes at the parts and system levels by performing alignment measurement before and after the launch environment test and the space environment test. Recently, noncontact three-dimensional precision machinery for medium- and large-sized parts has been developed. One of these is the theodolite measurement system, which is widely used in the aerospace industry. This study measures the angle of the dual thruster module that is used to control the attitude of KOMPSAT by using a theodolite, and alignment measurement and a reliability analysis are performed.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.198.2-198.2
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• 2012

태양 전지판의 전개 여부는 저궤도 위성의 발사 성공 여부를 판단하는 가장 중요한 항목 중 하나이다. 태양 전지판이 성공적으로 전개되어야만 태양 지향 자세제어에 의해 위성 운용에 필요한 전력 생성이 가능하기 때문이다. 그러므로 발사 후 지상국 교신을 통해 최우선적으로 태양 전지판의 전개 여부를 판단한다. 태양 전지판의 전개 여부는 다양한 실패 상황에 가정해 총 5가지 조건을 통해 판단한다. 첫째, SAR1, SAR2의 입력 전류가 모두 0.8A보다 커야 한다. 만약 하나라도 0.8A 미만이라면 한 개 이상의 태양 전지판이 전개되지 않고 1번 태양 전지판이 태양 지향을 하지 못하는 상황이다. 둘째, SAR1 입력 전류와 SAR2 입력 전류의 값이 유사해야 한다. 만약 입력 전류 값이 크게 차이가 난다면 2번과 3번 태양 전지판 중 하나만 태양 지향을 하는 경우이다. 셋째, CSSA#5 출력 전류가 3.2mA보다 커야 한다. 만약 3.2mA보다 작다면 2번과 3번 태양 전지판의 전개가 실패하고 1번 태양 전지판이 태양 지향을 하는 경우 또는 1번 태양 전지판이 전개 실패하고 태양 지향을 하는 경우이다. 넷째, S/C Roll, Pitch, Yaw rate이 모두 0.2 deg/sec 보다 작아야 한다. 만약 body rate이 크다면 1번 태양 전지판의 전개 실패 상황을 예상할 수 있다. 다섯째, 각 태양 전지판의 온도 차이가 $35^{\circ}C$ 보다 작아야 한다. 만약 온도 차이가 크다면 1번 태양 전지판 전개 실패 상황에서 2번과 3번 태양 전지판이 태양 지향을 하는 경우이다. 총 다섯 가지의 조건을 모두 만족해야만 태양 전지판이 성공적으로 전개되었다고 판단한다. 태양 전지판의 전개 판단은 위성이 발사체에서 분리되고 약 4500초 이후 시점에 스발바드 지상국과의 교신을 통해 확인되었다. 이 시점의 SAR1 입력 전류는 약 2.00A, SAR2 입력 전류는 약 1.93A였기 때문에 모두 0.8A보다 크고 서로 유사한 값임을 확인했다. CSSA#5의 출력 전류는 약 3.5mA의 값을 나타냈다. S/C Roll rate은 -0.0084 deg/sec, Pitch rate은 -0.0072 deg/sec, Yaw rate은 -0.0303 deg/sec의 값을 나타냈다. 각 태양 전지판의 최대 온도 차이는 $7.7^{\circ}C$의 값을 나타냈다. 5가지 조건을 모두 만족함으로써 태양 전지판 전개는 성공적으로 수행된 것으로 판단했다.

• Journal of Astronomy and Space Sciences
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• v.14 no.2
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• pp.312-319
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• 1997

The magnetometer is one of the most important payloads for scientific satellite to monitor the near-earth space environment. The electromagnetic variations of the space environment can be observed with the electric and magnetic field measurements. In practice, it is well known that the measurement of magnetic fields needs less technical complexities than that of electric fields in space. Therefore the magnetometer has long been recognized as one of the basic payloads for the scientific satellites. In this paper, we discuss the scientific fluxgate magnetometer which will be on board the KITSAT-3. The main circuit design of the present magnetometer is based on that of KITSAT-1 and -2 but its facilities have been re-designed to improve the resolution to about 5nT for scientific purpose. The calibration and noise level test of this circuit have been performed at the laboratory of the Tierra Tecnica company in Japan.

• Economic and Environmental Geology
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• v.37 no.4
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• pp.401-411
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• 2004

The Earth's total magnetic field was calculated from on board TAM(Three-Axis Magnetometer) observations of KOMPSAT-1 satellite between June 19th and 21st, 2000. The TAM's telemetry data were transformed from ECI(Earth-Centered Inertial Frame) to ECEF(Earth-Centered Earth-Fixed Frame) and then to spherical coordination. Self-induced field from the satellite bus were removed by the symmetric nature of the magnetic field. The 2-D wavenumber correlation filtering and quadrant-swapping method were applied to eliminate the dynamic components and track-line noise. To test the validity of the TAM's geomagnetic field, ${\phi}$rsted satellite's magnetic model and IGRF2000 model were used for statistical comparison. The correlation coefficients between KOMPSAT-1/${\phi}$rsted and KOMPSAT-1/IGRF2000 models are 0.97 and 0.96, respectively. The global spherical harmonic coeffi-cient was then calculated from the KOMPSAT-1 data degree and order of up to 19 and compared with those from IGRF2000, $\phi$rsted, and CHAMP models. The KOMPSAT-1 model was found to be stable to degree & order of up to 5 and it can give new information for the low frequency components of the global geomagtic field.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.292-295
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• 2004

Firing performance test of hydrazine decomposition catalyst which is used in mono-propellant thruster of satellite and launcher was peformed. Equipment for catalyst test was developed and with this equipment reaction delay time, catalyst activity, granule stability of the catalyst firing performance was measured and analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.281-285
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• 2007

The flow characteristics in the thruster should be analyzed considering its geometry and the pressure ratio to estimate its performance and etc. This paper suggests the computational result of an axisymmetric real nozzle for the altitude control of a satellite to find out the application limit that the assumption of continuum mechanics holds. The steady non-reacted compressible flow field in the unstructured grid system is computed and analyzed with varying the environmental pressure (or the degree of vacuum) under the fixed pressure ratio in a real thruster of which the area ratio of exit to throat is 56. The assumption of the continuum mechanics is not approved when the environmental pressure is reduced less than $10^{-3}$ atm.