• 전기전자학회논문지
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• 제24권3호
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• pp.917-920
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• 2020

본 논문에서는 Ku대역 통신시스템을 사용하는 큐브 위성에 장착이 가능한 평면배열안테나 있어, isoflux 방사특성을 유도하기위한 급전신호 가중치 분석에 관하여 연구하였다. 평면배열안테나는 큐브위성 장착성을 고려하여 8*8개의 방사소자를 평면 배열하였으며, 각 축 기준 방사소자 간 간격은 0.5 lambda로 설정하였다. isoflux 방사특성 유도를 위한 급전신호 가중치는 저대역 통과필터 설계에 활용되는 신호처리 기법을 이용하여 산출하는 방법을 제시하였다. 본 논문에서 제안된 방법의 분석 결과 원활한 isoflux 패턴을 유도할 수 있음을 확인하였다.

• 천문학회보
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• 제46권1호
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• pp.58.4-59
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• 2021

This presentation introduces Korea's SNIPE (Small scale magNespheric and Ionospheric Plasma Experiment) mission, formation flying CubeSat constellation. Observing particles and waves on a single satellite suffers from inherent space-time ambiguity. To observe spatial and temporal variations of the micro-scale plasma structures on the topside ionosphere, four 6U CubeSats (~ 10 kg) will be launched into a polar orbit of the altitude of ~500 km in 2021. The distances of each satellite will be controlled from 10 km to more than 100 km by formation flying algorithm. The SNIPE mission is equipped with identical scientific instruments, solid-state telescope, magnetometer, and Langmuir probe. All the payloads have a high temporal resolution (sampling rates of about 10 Hz). Iridium modules provide an opportunity to upload changes in operational modes when geomagnetic storms occur. SNIPE's observations of the dimensions, occurrence rates, amplitudes, and spatiotemporal evolution of polar cap patches, field-aligned currents (FAC), radiation belt microbursts, and equatorial and mid-latitude plasma blobs and bubbles will determine their significance to the solar wind-magnetosphere-ionosphere interaction and quantify their impact on space weather.

• Advances in aircraft and spacecraft science
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• 제7권3호
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• pp.271-290
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• 2020

The paper describes the analysis of deployment strategies and trajectories design suitable for executing the inspection of an operative spacecraft in orbit through re-usable CubeSats. Similar missions have been though indeed, and one mission recently flew from the International Space Station. However, it is important to underline that the inspection of an operative spacecraft in orbit features some peculiar characteristics which have not been demonstrated by any mission flown to date. The most critical aspects of the CubeSat inspection mission stem from safety issues and technology availability in the following areas: trajectory design and motion control of the inspector relative to the target, communications architecture, deployment and retrieval of the inspector, and observation needs. The objectives of the present study are 1) the identification of requirements applicable to the deployment of a nanosatellite from the mother-craft, which is also the subject of the inspection, and 2) the identification of solutions for the trajectories to be flown along the mission phases. The mission for the in-situ observation of Space Rider is proposed as reference case, but the conclusions are applicable to other targets such as the ISS, and they might also be useful for missions targeted at debris inspection.

• 한국항공우주학회지
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• 제45권2호
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• pp.114-123
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• 2017

Little Intelligent Nanosatellite of KAIST(LINK)는 카이스트 항공우주 시스템 및 제어연구실(ASCL)에서 국제협력 프로젝트인 QB50의 일환으로 개발한 2U 크기의 큐브위성이다. QB50 프로젝트의 과학적 임무 목적은 열권 및 이온층 대기를 관측하는 것이며, 2017년 1분기부터 국제우주정거장에서 순차적으로 발사될 예정이다. 상기의 임무 구현을 위해 최종 비행모델을 개발하였으며, 수락 수준(Acceptance level)의 요구조건에 따라 환경시험 및 기능시험을 수행하여 위성의 성능 및 설계의 검증을 완료하였다. 본 논문에서는 비행모델 개발과 진동시험 및 열진공시험 결과에 대해 기술하였다.

• 항공우주시스템공학회지
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• 제8권1호
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• pp.36-41
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• 2014

KAFASAT is a CubeSat which has a high level mission of testing the possibility of establishing the LEO satellite constellation providing the role of communication nodes and quasi-realtime image recognition of battlefield in accordance with the aspect of future-war-environment. The high level mission is developed using the Pugh selection method, which is one of system engineering tools. In order to accomplish the high level mission objectives and deduce engineering level requirements, system engineering tools such as Analytic Hierarchy Process and Quality Function Deployment are used. The subsystem synthesis in the context of system engineering process is done using a developed integrated design environment. The paper also includes the conceptual design results of the KAFASAT, which can be used as a baseline for upcoming preliminary design.

• 한국항공우주학회지
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• 제46권10호
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• pp.867-875
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• 2018

종래의 큐브위성용 나일론선 절단방식 태양전지판 구속분리장치는 단순히 패널 평면상에 나일론선을 체결함에 따른 취약한 구속력으로 인해 태양전지판 면적이 증가함에 따라 발사하중에 대한 구조 건전성 확보에 한계가 존재한다. 본 연구에서는 전술한 종래 분리장치의 한계점 극복을 위해 Ball & Socket 접속부가 반영된 별도의 타원형 브라켓을 적용하여 높은 구속력, 전개 및 평면 방향 동시구속 및 체결작업의 용이성 등의 장점을 갖는 6U 큐브위성용 태양전지판 구속분리장치를 제안하였다. 상기 구속분리장치의 설계 방향성 파악을 위해 큐브위성용 태양전지판 조립체에 대한 발사하중을 고려한 구조해석을 실시하였다. 또한, 상이한 온도조건에서의 나일론선 두께 및 체결횟수에 따른 기능시험을 수행하여 제안된 구속분리장치의 유효성을 검증하였다.

• 항공우주시스템공학회지
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• 제15권1호
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• pp.86-94
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• 2021

PCB 기반 전개형 태양전지판은 경량화 및 전기적 연결의 용이성으로 주로 큐브위성에 사용되나, 태양전지판의 면적이 증가할수록 발사환경에서 유발되는 굽힘거동이 증가하기 때문에 태양전지셀의 구조건전성 보장에 한계가 있다. 종래에는 태양전지판의 강성증가를 통해 굽힘거동을 최소화하고자 알루미늄 및 복합재 기반의 보강재를 또는 태양전지판을 적용하였지만, 태양전지판의 부피 및 무게 증가로 제한적인 설계요구조건을 가진 태양전지판의 단점으로 작용한다. 본 연구에서는 점탄성 테이프로 다층 박판을 적층하여 고댐핑 특성 구현이 가능한 6U 규격의 고댐핑 적층형 태양전지판을 제안하였다. 제안된 태양전지판의 기본특성파악을 위해 자유감쇠시험을 수행하였으며, 인증수준의 발사진동시험을 통해 설계유효성을 입증하였다. 또한, 시험결과를 토대로 일반 PCB 태양전지판과 고댐핑 적층형 태양전지판의 진동특성을 예측하고 비교분석을 수행하였다.

• International Journal of Aerospace System Engineering
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• 제1권1호
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• pp.34-43
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• 2014

The STEP Cube Lab (Cube Laboratory for Space Technology Experimental Projects) is a 1U cube satellite developed by the Space Technology Synthesis Laboratory of Chosun University to be launched in 2015. Its mission objective is twofold: to determine which of the fundamental space technologies researched at domestic universities, will be potential candidates for use in future space missions and to verify the effectiveness of the technologies by investigating mission data obtained from on-orbit operation of the cube satellite. In this paper, a structural design concept based on the 1U standard to achieve the mission objective is introduced. The validity of the design has been demonstrated by quasi-static analysis and modal analysis. In addition, a non-explosive separation device triggered by burn wire heating, which is one of the main mission payloads is introduced.

• 천문학회보
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• 제39권2호
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• pp.110.1-110.1
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• 2014

The optics of Space telescope is one of the major parts of space mission used for imaging observation of astronomical targets and the Earth. These kinds of space mission have a bulky and complex opto-mechanics with a long optical tube, but there are attempts have been made to observe a target with a small satellite in many ways. In this paper, we describe an optical design of a reflecting telescope for use in a CubeSat mission. For this design, we adopt the off-axis segmented method of astronomical observation techniques based on the Ritchey-Chr$\acute{e}$tien type telescope. The primary mirror shape is a rectangle with dimensions of $8cm{\times}8cm$, and a secondary mirror has dimensions of $2.4cm{\times}4.1cm$. The focal ratio is 3 which can obtain a $0.3{\times}0.2$ degree diagonal angle in a $1280{\times}800$ CMOS color image sensor with a pixel size of $3{\mu}m{\times}3{\mu}m$. This optical design can capture a ${\sim}4km{\times}{\sim}2.3km$ area of the earth's surface at 700 km altitude operation. Based on this conceptual design, we will keep trying to study more for astronomical observation with Attitude control system.

• Journal of the Optical Society of Korea
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• 제17권6호
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• pp.533-537
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• 2013

Space telescope optics is one of the major parts of any space mission used to observe astronomical targets or the Earth. This kind of space mission typically involves bulky and complex opto-mechanics with a long optical tube, but attempts have been made to observe a target with a small satellite. In this paper, we describe the optical design of a reflecting telescope for use in a CubeSat mission. For this design we adopt the off-axis segmented method for astronomical observation techniques based on a Ritchey-Chr$\acute{e}$tien type telescope. The primary mirror shape is a rectangle with dimensions of $8cm{\times}8cm$, and the secondary mirror has dimensions of $2.4cm{\times}4.1cm$. The focal ratio is 3 which can yield a 0.383 degree diagonal angle in a $1280{\times}800$ CMOS color image sensor with a pixel size of $3{\mu}m{\times}3{\mu}m$. This optical design can capture a ${\sim}4km{\times}{\sim}2.3km$ area of the earth's surface at 700 km altitude operation.