• Aerospace Engineering and Technology
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• v.12 no.2
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• pp.57-63
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• 2013

The mission of a lot of satellites on geostationary orbit is the communication and/or the broadcasting. These satellites need a big power, so these have a large solar array. Recently, the new satellite for Earth environment monitoring is developing on geostationary orbit. The payload of Earth monitoring satellite requires better thermal condition on detector. Therefore this satellite uses a boom for the attitude stability instead of rejecting one-side solar array as a heat source. The other hand, it uses some momentum wheels being a more momentum capacity to control the large disturbance by solar pressure due to the asymmetric solar array configuration. In this paper, the analysis on the wheel allocation and the wheel off-loading for COMS is summarized and the results are verified by telemetry of COMS. COMS has no boom and a perfectly asymmetric solar array configuration, and it is operating well on geostationary orbit.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.322-325
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• 2007

This paper shows the development procedure of the low pressure LOX feed system which is used in the liquid rocket with a turbopump. Korean Air has cooperated with KARI in developing the LOX feed system to turbopump. The LOX feed system is characterized with cryogenic temperature and the thin-thickness tube for weight saving. The system in this project is composed with a main feed line and a recirculation line for the LOX temperature conditioning. Each piping system has many components, namely, bellows, filter, orifice, valves, flange and support. In this paper, system design & manufacturing, structural & thermal analyses, and component tests are explained. Finally, the system was assembled to the KARI's PTF test facility and functioned well to meet its required performance.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.11
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• pp.909-917
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• 2021

In this work, a one-dimensional combustor solver was constructed for the scramjet control m odel. The governing equations for fluid flow, Arrhenius based combustion kinetics, and the inje ction model were implemented into the solver. In order to validate the solver, the zero-dimensi onal ignition delay problem and one-dimensional scramjet combustion problem were considered and showed that the solver successfully reproduced the results from the literature. Subsequentl y, a ramjet analysis algorithm under subsonic speed conditions was constructed, and a study o n the inlet Mach number of the combustor was carried out through the thermal choking locatio ns at ram conditions. In such conditions, a model for precombustion shock train analysis was i mplemented, and the algorithm for transition section analysis was introduced. In addition, in or der to determine the appropriateness of the ram mode analysis in the code, the occurrence of a n unstart was studied through the length of the pseudo-shock in the isolator. A performance a nalysis study was carried out according to the geometry of the combustor.

• Journal of Astronomy and Space Sciences
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• v.24 no.2
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• pp.179-194
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• 2007

The design and analysis of satellite power subsystem is an important driver for the mass, size, and capability of the satellite. Every other satellite subsystem is affected by the power subsystem, and in particular, important issues such as launch vehicle selection, thermal design, and structural design are largely influenced by the capabilities and limitations of the power system. This paper introduces a new electrical power subsystem design program for the rapid development of LEO satellite and shows an example of design results using other LEO satellite design data. The results shows that the proposed design program can be used the optimum sizing and the analytical prediction of the on-orbit performance of satellite electrical power subsystem.

• Proceedings of the Materials Research Society of Korea Conference
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• 2009.11a
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• pp.24.2-24.2
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• 2009

Indium Antimonide(InSb)는 $3{\sim}5\;{\mu}m$대 적외선 감지영역에서 기존 HgCdTe(MCT)를 대체할 물질로 각광받고 있다. 1970년대부터군사적 용도로 미국, 이스라엘 등 일부 선진국에서 연구되기 시작했으며,이온주입, MOCVD, MBE 등 다양한 공정을 통해 제작되어 왔다. InSb 적외선 감지소자는 $3{\sim}5{\mu}m$대에서 HgCdTe와 성능은 대등한데 반해, 기판의 대면적화와 저렴한 가격, 우주공간 및 야전에서 소자 동작의안정성 등으로 InSb적외선 감지기는 냉각형 고성능 적외선 감지영역에서 HgCdTe를 대체해 가고 있다. 하지만 InSb는 77 K에서 0.225eV의 작은 밴드갭을 갖고 있기 때문에 누설전류로 인한 성능저하가 고질적인문제로 대두되었고, 이를 해결하기 위한 고품질 절연막 연구가 InSb적외선 수광 소자 연구의 주요이슈 중 하나가 되어왔다. 그 동안 PECVD, photo-CVD, anodic oxidation 등의 공정을 이용하여 $SiO_2$, $Si_3N_4$, 양극산화막(anodic oxide) 등 다양한 절연막에 대한 연구가 진행되었고[1,2], 절연막과 반도체 사이 계면에서의 열확산을 억제하여 계면트랩밀도를 최소화하기 위한 공정개발이 이루어졌다[3]. 하지만 InSb 적외선 감지기술은 국방 및 우주개발의 핵심기술중 하나로 그 기술의 이전이 엄격히 통제되고 있으며, 현재도 미국과 이스라엘, 일본, 영국 등 일부 선진국 만이 기술을 확보하고 있고, 국내의 경우 연구가 매우 취약한 실정이다. 따라서 본 연구에서는 InSb 적외선 감지기의 암전류를 제어하기 위한 낮은 계면트랩밀도를 갖는 절연막 증착 공정을 찾고자 하였다. 본 연구에서는 n형 (100) InSb 기판 ($n=0.2{\sim}0.85{\times}10^{15}cm^{-3}$ @ 77K)에 PECVD를 이용하여 $SiO_2$, $Si_3N_4$ 등을 증착하고 절연막으로서 이들의 특성을 비교 분석하였다. $SiO_2$는 160, 200, $240^{\circ}C$에서 $Si_3N_4$는 200, $300^{\circ}C$에서 증착하였다. Atomic Force Microscopy(AFM) 사진으로 확인한 결과, 모든 샘플에서표면거칠기가 ~2 nm의 평탄한 박막을 얻을 수 있었다. Capacitance-Voltage 측정(77K)을 통해 절연막 특성을 평가하였다. $SiO_2$와 $Si_3N_4$ 모두에서 온도가 증가할수록 벌크트랩밀도가 감소하는 경향을 볼 수 있었는데, 이는 고온에서 증착할 수록 박막 내의 결함이 감소했음을 의미한다. 반면계면트랩밀도는 온도가 증가함에 따라, 1011 eV-1cm-2 대에서 $10^{12}eV^{-1}cm^{-2}$ 대로 증가하였는데, 이는 고온에서 증착할 수 록 InSb 표면에서의 결함은 증가하였음을의미한다. 암전류에 큰 영향을 주는 것은 계면트랩밀도 이므로, $SiO_2$와 $Si_3N_4$ 모두 $200^{\circ}C$이하의 저온에서 증착시켜야 함을 확인할 수 있었다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.7
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• pp.586-593
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• 2014

The CNUSAIL-1 project aims to develop and operate a 3U-sized cube satellite with solar sail mechanism. The primary mission is to successfully deploy the solar sail in a low earth orbit, and the secondary mission is to collect the scientific data for the effect of the solar sail deployment and operation on orbit maneuver and attitude change of the cube satellite. For this, the bus system will collect and transmit the dynamic data of the satellite and the visual images of the solar sail operation. This paper describes solar sail mission and conceptual design of CNUSAIL-1. The actuation/operation of the solar sail and the bus system are preliminarily designed in terms of attitude control system, communication system, electrical power system, command and data handling system, structure and thermal control system is designed.

• Journal of Astronomy and Space Sciences
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• v.22 no.3
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• pp.329-352
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• 2005

Long term observations of full-disk Lyman-o irradiance have been made by the instruments on various satellites. In addition, several sounding rockets dating back to the 1950s and up through the present have measured the $Lyman-{\alpha}$ irradiance. Previous full disk $Lyman-{\alpha}$ images of the sun have been very interesting and useful scientifically, but have been only five-minute 'snapshots' obtained on sounding rocket flights. All of these observations to date have been snapshots, with no time resolution to observe changes in the chromospheric structure as a result of the evolving magnetic field, and its effect on the Lyman-o intensity. The $Lyman-{\alpha}$ Imaging Solar Telescope(LIST) can provide a unique opportunity for the study of the sun in the $Lyman-{\alpha}$ region with the high time and spatial resolution for the first time. Up to the 2nd year development, the preliminary design of the optics, mechanical structure and electronics system has been completed. Also the mechanical structure analysis, thermal analysis were performed and the material for the structure was chosen as a result of these analyses. And the test plan and the verification matrix were decided. The operation systems, technical and scientific operation, were studied and finally decided. Those are the technical operation, mechanical working modes for the observation and safety, the scientific operation and the process of the acquired data. The basic techniques acquired through the development of satellite based solar telescope are essential for the construction of space environment forecast system in the future. The techniques which we developed through this study, like mechanical, optical and data processing techniques, could be applied extensively not only to the process of the future production of flight models of this kind, but also to the related industries. Also, we can utilize the scientific achievements which are obtained throughout the project And these can be utilized to build a high resolution photometric detectors for military and commercial purposes. It is also believed that we will be able to apply several acquired techniques for the development of the Korean satellite projects in the future.

• Journal of Astronomy and Space Sciences
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• v.22 no.1
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• pp.13-20
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• 2005

Bound abrasive grinding is used for the initial fabrication phase of the precision aspheric mirrors for both space and ground based astronomical telescopes. We developed a new grinding optimization process that determines the input grinding variables for the target surface roughness, checks the grinding error magnitude in resulting surface roughnesses, and minimizes the required machining time. Using the machining data collected from the previous grinding runs and subsequently fed into the multivariable regression engine, the process has the evolving controllability that suggests the optimum set of grinding variables for each target surface roughness. The process model was then used for ten grinding experiments that resulted in the grinding accuracy of $=-0.906{\pm}3.38(\sigma)\;nm(Ra)$ for the target surface roughnesses of Zerodur substrate ranging from 96.1 nm (Ra) to 65.0 nm (Ra) The results imply that the quantitative process optimization technique developed in this study minimizes the machining time and offers the nanometric surface roughness controllability superior to the traditional, qualitative, craftsman based grinding process for the astronomical optical surfaces.

• The Journal of Korean Institute of Electromagnetic Engineering and Science
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• v.26 no.3
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• pp.227-247
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• 2015

An X-band $8{\times}16$ dual-polarized active phased array antenna system has been implemented based on the substrate integrated waveguide(SIW) technology having low propagation loss, complete EM shielding, and high power handling characteristics. Compared with the microstrip case, 1 dB less is the measured insertion loss(0.65 dB) of the 16-way SIW power distribution network and doubled(3 dB improved) is the measured radiation efficiency(73 %) of the SIW sub-array($1{\times}16$) antenna element. These significant improvements of the power division loss and the radiation efficiency using the SIW, save more than 30 % of the total power consumption, in the active phased array antenna systems, through substantial reduction of the maximum output power(P1 dB) of the high power amplifiers. Using the X-band $8{\times}16$ dual-polarized active phased array antenna system fabricated by the SIW technology, the main radiation beam has been steered by 0, 5, 9, and 18 degrees in the accuracy of 2 degree maximum deviation by simply generating the theoretical control vectors. Performing thermal cycle and vacuum tests, we have found that the SIW array antenna system be eligible for the space environment qualification. We expect that the high efficiency SIW array antenna system be very effective for high performance radar systems, massive MIMO for 5G mobile systems, and various millimeter-wave systems(60 GHz WPAN, 77 GHz automotive radars, high speed digital transmission systems).

• The Bulletin of The Korean Astronomical Society
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• v.42 no.2
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• pp.87.3-88
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• 2017

한국천문연구원은 차세대소형위성 1호의 근적외선 영상분광기 NISS (Near-infrared Imaging Spectrometer for Star formation history) 탑재체를 개발하여 2017년 6월 30일에 최종 비행모델을 납품하였고, 이 발표는 탑재체 NISS 구조체의 비행모델 개발 결과를 보고한다. NISS는 0.9 - 2.5um (R~20) 근적외선 파장에서 관측을 해야 하기 때문에, 구조체의 배경잡음을 없애기 위해서 200K까지 passive cooling으로 냉각되며, H2RG 검출기는 소형 냉동기에 의해 약 88K에서 운영된다. NISS 구조체의 passive cooling을 효율적으로 수행하기 위해서 방열판, Kevlar 지지대, MLI, 표면제어용 필름 등을 조립하였고, 실제 지상 시험을 통해서 그 성능을 확인하였다. NISS 구조체는 최종 시스템 조립 과정에서 전자부 하네스 조립을 함께 수행했으며, 온도 모니터링 센서를 부착하고 소형 냉동기 피드백 온도를 반복 시험을 통해서 결정하였다. NISS 구조체는 미러 및 렌즈를 지지하는 광기계부를 함께 포함하기 때문에 발사 및 우주환경에서 광학 성능을 유지하기 위한 설계를 거쳐서 제작 되었으며, 최종 시스템 검교정 시험, 진동 및 열진공 시험을 통해서 그 성능을 확인하였다. NISS를 탑재한 차세대소형위성 1호는 2018년 상반기에 미국의 Falcon 9 발사체에 실려서 발사될 예정이다.