• Composites Research
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• v.15 no.3
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• pp.18-29
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• 2002

The bending vibration and thermal flutter instability of spacecraft booms modeled as circular thin-walled beams of closed cross-section and subjected to thermal radiation loading is investigated in this paper. The thin-walled beam model incorporates a number of nonclassical effects of transverse shear, primary and secondary warping, rotary inertia and anisotropy of constituent materials. Thermally induced vibration response characteristics of a composite thin walled beam exhibiting the circumferantially uniform system(CUS) configuration are exploited in connection with the structural flapwise bending-lagwise bending coupling resulting from directional properties of fiber reinforced composite materials and from ply stacking sequence. The numerical simulations display deflection time-history as a function of the ply-angle of fibers of the composite materials, damping factor, incident angle of solar heat flux, as well as the boundary of the thermal flutter instability domain. The adaptive control are provided by a system of piezoelectric devices whose sensing and actuating functions are combined and that are bonded or embedded into the host structure.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.1
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• pp.56-64
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• 2002

Thermally induced vibration response of composite thin-walled beams is investigated in this paper. The flexible spacecraft appendages modeled as thin-walled beam incorporates a number of nonclassical effects of transverse shear, primary and secondary warping, rotary inertia and anisotropy of constitute materials. Thermally induced vibration responds characteristics of a composite thin walled beam exhibiting the circumferantially uniform system(CUS) configuration are exploited in connection with the structural flapwise bending lagwise bending coupling resulting from directioal properties of fiber reinforced composite materials and ply stacking sequence. A coupled thermal structure gradient is investigated.

• Aerospace Engineering and Technology
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• v.4 no.1
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• pp.129-142
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• 2005

A semi-monocoque truncated cone structure, which is a main structure for the payload adapter of KSLV-I, was designed. Static test was performed to confirm the reliability of the cone structure under the design loads. Strains and displacements are measured during four load cases; the compressive axial, pure bending, pure shear, and combined loading conditions. The results showed that the cone structure satisfies the design requirements. An equivalent axial load was applied to the cone structure so that the global buckling of the cone structure occurred. The measured buckling load was compared with the predicted one by finite element method. The results show a good agreement.

• The Journal of the Acoustical Society of Korea
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• v.39 no.4
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• pp.342-350
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• 2020

Acoustic and vibration research activities for the structural development of Korean space launch vehicles are introduced in this paper. Various dynamic loads exerted on a launch vehicle during its operation are summarized. The acoustical design method of payload fairings which protect satellites from harsh launch environment was reviewed. Several acoustic research activities were performed to enhance the analytical prediction ability during the development period of the Naro and the Nuri launcher. Specifically, the following research activities are reviewed: a test and vibro-acoustic analysis of composite cylinders whose layup properties are varied, a research on low-frequency acoustic load reduction by an acoustic resonator array and an acoustic test on the cylinder part of the Naro payload fairing. A vibro-acoustic analysis result for the Nuri launcher was introduced and predicted acoustic and vibration levels and measured ones are shown to be in a good agreement.

• Journal of Astronomy and Space Sciences
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• v.15 no.2
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• pp.509-523
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• 1998

Before an actual military communications satellite is designed and constructed, a feasibility study should take place. The basic functions of such system can be observed and demonstrated in an X-Band payload simulator. For this purpose a Payload Concept Model (PCM) for X-Band payload subsystem has been developed to simulate the workings of an actual military communications payload. This paper explains and illustrates the mechanical design, manufacture, and integration of the PCM. Basic RF tests also have been performed in order to verify the design requirement of the system. The results demonstrate successful development of the PCM and operation without RF losses.

• 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.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.133.1-133.1
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• 2012

전자파설계는 위성의 전력시스템, 통신시스템 뿐만 아니라 구조체 등 위성시스템 전반에 걸쳐서 종합적으로 고려가 되어야 하며, 이를 위해서는 개발 초기단계에서부터 시스템 설계에 반영되어야 한다. 위성시스템의 상세 설계가 끝난 후에는 시스템에 구현된 전자파 설계의 적합성을 검증하여야 하며 이는 해석 및 시험을 통해 이루어진다. 본 논문에서는 저궤도 중형급 위성이 우주환경에서 전자파적합성을 이루기 위한 설계 기법 및 전자파환경에 대한 적합성 검증과정에 대해서 다루고 있다. 저궤도 중형급 위성시스템에 대하여 구조물의 전자기적 특성을 정의하는 것부터 우주환경에서 위성의 RF호환성에 이르기까지 부품단위에서부터 시스템 수준까지의 전자파 설계 기준과 각 단계별로 전자파적합성을 검증하기 위한 방법 및 절차에 대해서 기술한다.

• Journal of Aerospace System Engineering
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• v.11 no.1
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• pp.1-7
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• 2017

While a satellite is carried into orbit by a launch vehicle, it is exposed to the severe launch environment with random vibrations and shock. Accordingly, these vibration sources affect electronic equipment, particularly the printed circuit board (PCB) in the satellite. When the launch load impacts the PCB, it causes negative behavior. This causes perpendicular bending around the boundary of fixation points that finally leads to the failure of solder joints, lead wires, and PCB cracks. To overcome these issues, the electronic equipment design must meet reliability requirements. In this paper, Steinberg's method is used to derive allowable and maximum deflection to verify design from a life perspective concerning the control board of the Antenna Pointing Driver (APD) mounted on KOMPSAT-3.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.2
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• pp.157-164
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• 2012

The design strategy of the multi-functional structure is that the electrical components and the circuits are directly put on their supporting structural panel in which the radiation shields and the thermal control functions are integrated. Applying the multi-functional structure reduces the total mass and size of the space system and makes it possible to lower launch cost. In present study the performance of thermal radiation for six types of multi-functional structure are investigated by the numerical method. The effect of the rib configuration on heat transfer for the multi-functional-structure is not important alone but is meaningful considering with the structural stiffness, difficulty of manufacturing and mass increase. In heat spreading point of view, the thickness of the outer conductive layer is important rather than the rib configuration and the trade-off study with the mass and thickness is required for optimum design.

• Composites Research
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• v.23 no.6
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• pp.47-54
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• 2010

Thermal vibration characteristics of a thin walled composite beam attached on spacecraft are investigated in this paper. The composite beam is assumed to have a thin CUS(circumferentially uniform stiffness) wall and modeled with several composite materials which are already space qualified such as T300/Epoxy, YS90AEpoxy. Steady state angle and peak to peak error in spacecraft attitude angle and tip displacement of composite beams are evaluated as a performance index for thermal vibration characteristics. Evaluation results shows that composite beam made out of YS90A has nearly 2 times better than T300 in terms of peak to peak attitude error angle.