• 천문학논총
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• 제20권1호
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• pp.11-20
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• 2005

We present measurements of interstellar $H_2$ absorption lines in the continuum spectra of 54 early-type stars in the Galactic disk and halo and 3 stars in the Magellanic Clouds. The data were obtained with the Berkeley Extreme and Far-Ultraviolet Spectrometer (BEFS), part of the ORFEUS telescope, which flew on the ORFEUS-SPAS I and II space-shuttle missions in 1993 and 1996, respectively. The spectra extend from the interstellar cutoff at $912{\AA}$ to about $1200{\AA}$ with a spectral resolution of ${\sim}3000$ and statistical signal-to-noise ratios between 10 and 65. Assuming a velocity profile derived from optical observations (when available), we model the column densities N(J) of the rotational levels J = 0 through 5 for each line of sight. Our data reproduce the relationships among molecular and total hydrogen column density, fractional molecular abundance, and reddening first seen in Copernicusobservations of nearby stars (Savage et al. 1977). The results show that most of these molecular clouds have $H_2$ total column densities between $10^{15}cm^{-2}$ and $10^{21}cm^{-2}$, and kinetic temperatures from 21 K to 232 K, with average of 89 K, consistent with the result of Copernicus (Savage et al. 1977).

• Journal of Astronomy and Space Sciences
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• 제32권3호
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• pp.257-268
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• 2015

Characteristics of delta-V requirements for deploying an impactor from a mother-ship at different orbital altitudes are analyzed in order to prepare for a future lunar CubeSat impactor mission. A mother-ship is assumed to be orbiting the moon with a circular orbit at a 90 deg inclination and having 50, 100, 150, 200 km altitudes. Critical design parameters that are directly related to the success of the impactor mission are also analyzed including deploy directions, CubeSat flight time, impact velocity, and associated impact angles. Based on derived delta-V requirements, required thruster burn time and fuel mass are analyzed by adapting four different miniaturized commercial onboard thrusters currently developed for CubeSat applications. As a result, CubeSat impact trajectories as well as thruster burn characteristics deployed at different orbital altitudes are found to satisfy the mission objectives. It is concluded that thrust burn time should considered as the more critical design parameter than the required fuel mass when deducing the onboard propulsion system requirements. Results provided through this work will be helpful in further detailed system definition and design activities for future lunar missions with a CubeSat-based payload.

• 대한원격탐사학회:학술대회논문집
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• 대한원격탐사학회 2007년도 Proceedings of ISRS 2007
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• pp.65-68
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• 2007

FORMOSAT-3 was launched in April of 2006. It consists of six low earth orbit (LEO) satellites that will be eventually deployed to an orbit at 800 km height. Its scientific goal is to utilize the radio occultation (RO) signals to measure the bending angles when the GPS signals transect the atmosphere. The bending angle is then used to infer atmospheric parameters, including refractivity, temperature, pressure, and relative humidity fields of global distributions through inversion schemes and auxiliary information. The expected number of RO events is around 2500 per day, of which 200 events or so fall into the polar region. Consequently, the FORMOSAT-3 observations are expected to play a key role to improve our knowledge in the weather forecasting and space physics research in the polar region. In this paper, we use temperature profiles retrieved from FORMOSAT-3 RO observations to study the climatology of gravity wave activity in the polar region. FORMOSAT-3 can provide about 200 RO observations a day in the polar region, much more than previous GPS RO missions, and, hence, more detailed climatology of gravity wave activity can be obtained.

• 한국소음진동공학회논문집
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• 제17권12호
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• pp.1208-1216
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• 2007

Turbomachinery such as turbines, pumps and compressors, which are installed in transportation systems, including aircrafts, ships, and space vehicles, etc., often perform crucial missions and are exposed to potential dangerous impact environments such as base-transferred shock forces. To protect turbomachinery from excessive shock forces, it may be needed to accurately analyze transient responses of their rotors, considering the dynamics of mount designs to be applied. In this study a generalized FE transient response analysis model, introducing relative displacements, is proposed to accurately predict transient responses of a flexible rotor-bearing system with mount systems to base-transferred shock forces. In the transient analyses the state-space Newmark method of a direct time integration scheme is utilized, which is based on the average velocity concept. Results show that for the identical mount systems considered, the proposed FE-based detailed flexible rotor model yields more reduced transient vibration responses to the same shocks than a conventional simple model, obtained by treating a rotor as concentrated lumped mass, equivalent spring and a damper or Jeffcott rotor model. Hence, in order to design a rotor-bearing system with a more compact light-weighted mount system, preparing against any potential excessive shock, the proposed FE transient response analysis model herein is recommended.

• Journal of Power Electronics
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• 제9권4호
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• pp.631-642
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• 2009

The sun is a useful reference direction because of its brightness relative to other astronomical objects and its relatively small apparent radius as viewed by spacecrafts near the Earth. Most satellites use solar power as a source of energy, and so need to make sure that solar panels are oriented correctly with respect to the sun. Also, some satellites have sensitive instruments that must not be exposed to direct sunlight. For all these reasons, sun sensors are important components in spacecraft attitude determination and control systems. To minimize components and structural mass, some components have multiple purposes. The solar cells will provide power and also be used as coarse sun sensors. A coarse Sun sensor is a low-cost attitude determination sensor suitable for a wide range of space missions. The sensor measures the sun angle in two orthogonal axes. The Sun sensor measures the sun angle in both azimuth and elevation. This paper presents the development of a model to determine the attitude of a small cube-shaped satellite in space relative to the sun's direction. This sensor helps small cube-shaped Pico satellites to perform accurate attitude determination without requiring additional hardware.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2010년도 제35회 추계학술대회논문집
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• pp.682-685
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• 2010

KSLV-I 나로호는 대한민국 최초의 위성발사체로 러시아와 공동개발 되었다. 2009년 8월과 2010년 6월에 있었던 1, 2차 발사는 위성을 궤도로 올리려는 목표 달성에는 실패하였지만 발사를 안전 관점에서 보았을 때는 인명이나 시설 손상과 같은 사고 없이 안전하게 이루어졌다. 안전한 발사를 이루기 위하여 우주발사체 사업단은 러시아의 카운터파트인 흐루니체프사와의 협력을 통해 발사체의 기술적 안전 프로그램을 개발한 바 있다. 본 논문에서는 KSLV-I에 적용된 기술적 안전 관리 프로그램을 소개하고자 한다.

• 대한임베디드공학회논문지
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• 제15권4호
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• pp.177-187
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• 2020

Virtualization with hypervisors is one of emerging topics in multicore processors for space. Hypervisors are software layers to make several independent virtualized environments on one processor. Since all hardware resources are virtualized and distributed only by hypervisors, overall performance of processors can be improved by fully utilizing the resources. However at the same time, there are overheads for virtualizing and distributing hardware resources. Satellites are one of hard real time systems, and performance degradation with overheads should be analyzed thoroughly. Previous research on the overheads focused on single core systems. Even the overheads were analyzed in multicore systems, SMP environment was not fully included. This paper builds SMP environment with XtratuM, one of hypervisors for space missions, and analyzes performance degradation with overheads. Two boards of GR712RC with 2 LEON3FT CPUs and GR740 with 4 LEON4 CPUs are used in experiments. On each board, SMP benchmark functions are executed on SMP environment with XtratuM and on that without XtratuM respectively. Results are analyzed to find timing characteristics including overheads. Finally, applicability of the XtratuM to flight software in SMP is also reviewed.

• 대한원격탐사학회지
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• 제15권4호
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• pp.367-375
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• 1999

This paper describes the development, operations, and applications of ROCSAT-l and its Ocean Color Imager (OCI) remote-sensing payload. It is the first satellite program of NSPO. The satellite was successfully launched by Lockheed Martin's Athena on January 26, 1999 from Cape Canaveral, Florida. ROCSAT-l is a Low Earth Orbit (LEO) experimental satellite. Its circular orbit has an altitude of 600km and an inclination angle of 35 degrees. The satellite is designed to carry out scientific research missions, including ocean color imaging, experiments on ionospheric plasma and electrodynamics, and experiments using Ka-band (20∼30GHz) communication payloads. The OCI payload is utilized to observe the ocean color in 7 bands (including one redundant band) of Visible and Near-Infrared (434nm∼889nm) range with the resolution of 800m at nadir and the swath of 702km. It employs high performance telecentric optics, push-broom scanning method using Charge Coupled Devices (CCD) and large-scale integrated circuit chips. The water leaving radiance is estimated from the total inputs to the OCI, including the atmospheric scattering. The post-process estimates the water leaving radiance and generates different end products. The OCI has taken images since February 1999 after completing the early orbit checkout. Analyses have been performed to evaluate the performances of the instrument in orbit and to compare them with the pre-launch test results. This paper also briefly describes the ROCSAT-l mission operations. The spacecraft operating modes and ROCSAT Ground Segment operations are delineated, and the overall initial operations of ROCSAT-l are summarized.

• Advances in aircraft and spacecraft science
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• 제6권5호
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• pp.427-442
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• 2019

Small satellites represent an emerging opportunity to realize a wide range of space missions at lower cost and faster delivery, compared to traditional spacecraft. However, small platforms, such as CubeSats, shall increase their actual capabilities. Miniaturized electric propulsion systems can provide the satellite with the key capability of moving in space. The level of readiness of miniaturized electric propulsion systems is low although many concepts have been developed. The present research intends to build a flexible test platform for the assessment of selected small propulsion systems in relevant environment at laboratory level. Main goal of the research is to analyze the mechanical, electrical, magnetic, and chemical interactions of propulsion systems with the modern CubeSat-technology and to assess the performance of the integrated platform. The test platform is a 6U CubeSat hosting electric propulsion systems, providing mechanical, electrical and data interfaces, able to handle a variety of electric propulsion systems, thanks to the ability to regulate and distribute electric power, to exchange data according to several protocols, and to provide different mechanical layouts. The test platform is ready to start the first verification campaign. The paper describes the detailed design of the platform and the main results of the AIV activities.

• 우주기술과 응용
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• 제3권4호
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• pp.322-332
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• 2023

Satellite attitude-control actuators are equipped with a reaction wheel for three-axis attitude control. The reaction wheel rotates a motor inside the actuator to generate torque in the vector direction. When using the reaction wheel, there are restrictions on the torque values generated as the motor rotates. The torque value of the reaction wheels mounted on small satellites is approximately 10 mNm, and high values are not used. Therefore, three-axis attitude control of a small satellite is possible using a reaction wheel, but this method is not suitable for missions that require rapid attitude control at a specific time. As a technology to overcome the small torque value of the reaction wheel, the control moment gyro (CMG) is currently in wide use as a rapid attitude-control actuator in space satellites. The CMG has an internal gimbal mounted at a right angle to the rotation motor and generates a large torque value. In general, when the gimbal operates, a torque value approximately 100 times greater is generated, making it suitable for rapid posture maneuvering. Currently, we are developing a technology for mounting a controlled moment gyro on a small satellite, and here we share the development status of an 800 mNm CMG.