• Proceedings of the Korean Society For Composite Materials Conference
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• 2002.10a
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• pp.88-91
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• 2002

The two dimensional size effect of specimen gauge section (length x width) was investigated on the compressive behavior of a T300/924 [45/-45/0/90]3s, carbon fiber-epoxy laminate. A modified ICSTM compression test fixture was used together with an anti-buckling device to test 3mm thick specimens with a 30$\times$30, 50$\times$50, 70$\times$70, and 90mm$\times$90mm gauge length by width section. In all cases failure was sudden and occurred mainly within the gauge length. Post failure examination suggests that $0^{\circ}$ fiber microbuckling is the critical damage mechanism that causes final failure. This is the matrix dominated failure mode and its triggering depends very much on initial fiber waviness. It is suggested that manufacturing process and quality may play a significant role in determining the compressive strength. When the anti-buckling device was used on specimens, it was showed that the compressive strength with the device was slightly greater than that without the device due to surface friction between the specimen and the device by pretoque in bolts of the device. In the analysis result on influence of the anti-buckling device using the finite element method, it was found that the compressive strength with the anti-buckling device by loaded bolts was about 7% higher than actual compressive strength. Additionally, compressive tests on specimen with an open hole were performed. The local stress concentration arising from the hole dominates the strength of the laminate rather than the stresses in the bulk of the material. It is observed that the remote failure stress decreases with increasing hole size and specimen width but is generally well above the value one might predict from the elastic stress concentration factor. This suggests that the material is not ideally brittle and some stress relief occurs around the hole. X-ray radiography reveals that damage in the form of fiber microbuckling and delamination initiates at the edge of the hole at approximately 80% of the failure load and extends stably under increasing load before becoming unstable at a critical length of 2-3mm (depends on specimen geometry). This damage growth and failure are analysed by a linear cohesive zone model. Using the independently measured laminate parameters of unnotched compressive strength and in-plane fracture toughness the model predicts successfully the notched strength as a function of hole size and width.

• Proceedings of the Korean Society For Composite Materials Conference
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• 2004.10a
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• pp.61-64
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• 2004

The strength characteristics as well as deformation behaviors of honeycomb sandwich composite (HSC) structures were investigated under bending in consideration of various failure modes such as skin layer yielding, interface-delamination, core shear deformation and local buckling. Deformation behaviors of honeycomb sandwich plates were observed with various types of aluminum honeycomb core and skin layer. Their finite-element analysis simulation with a real model of honeycomb core was performed to analyze stresses and deformation behaviors of honeycomb sandwich plates. Its results were very comparable to the experimental ones. Consequently, the increase in skin layer thickness and in cell size of honeycomb core had dominant effects on the strength and deformation behaviors of honeycomb sandwich composites.

• Journal of Ocean Engineering and Technology
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• v.10 no.2
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• pp.35-41
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• 1996

This paper is aimed to solve local buckling problem that can frequently occur when the high speed ship's hull of sandwich structural type is crushed by rarbour and cargo. Experiment is performed on 36 specimens cut of 4-plates that made of sandwich type(Kevlar-Epoxy, Klegecell foam) and 16-Edgewise compressive test specimen, 16-Flatwise test specimen were tested by A.S.T.M. test method. The result of this study is analyzed and compared in test method and test jig to perorm Edgewise compressive test and Flatwise test.

• Composites Research
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• v.35 no.4
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• pp.269-276
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• 2022

In this paper, a composite stiffened panel was fabricated using a three-dimensional weaving method that can reduce the risk of delamination, and mechanical properties such as buckling load and natural frequency were investigated. The preform of the stringer and skin of the stiffened panel were fabricated in one piece using T800 grade carbon fiber and then, resin (EP2400) was injected into the preform. The compression test and natural frequency measurement were performed for the stiffened panel, and the results were compared with the finite element analyses. In order to compare the performance of 3D weaving structures, the stiffened panels with the same configuration were fabricated using UD and 2D plain weave (fabric) prepregs. Compared to the tested buckling load of the 3D woven panel, the buckling loads of the stiffened panels of UD prepreg and 2D plain weave exhibited +20% and -3% differences, respectively. From this study, it was confirmed that the buckling load of the stiffened panel manufactured by 3D weaving method was lower than that of the UD prepreg panel, but showed a slightly higher value than that of the 2D plain weave panel.

• Composites Research
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• v.15 no.4
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• pp.23-31
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• 2002

The two dimensional size effect of specimen gauge section ($length{\;}{\times}{\;}width$) was investigated on the compressive behavior of a T300/924 $\textrm{[}45/-45/0/90\textrm{]}_{3s}$, carbon fiber-epoxy laminate. A modified ICSTM compression test fixture was used together with an anti-buckling device to test 3mm thick specimens with a $30mm{\;}{\times}{\;}30mm,{\;}50mm{\;}{\times}{\;}50mm,{\;}70mm{\;}{\times}{\;}70mm{\;}and{\;}90mm{\;}{\times}{\;}90mm$ gauge length by width section. In all cases failure was sudden and occurred mainly within the gauge length. Post failure examination suggests that $0^{\circ}$ fiber microbuckling is the critical damage mechanism that causes final failure. This is the matrix dominated failure mode and its triggering depends very much on initial fiber waviness. It is suggested that manufacturing process and quality may play a significant role in determining the compressive strength. When the anti-buckling device was used on specimens, it was showed that the compressive strength with the device was slightly greater than that without the device due to surface friction between the specimen and the device by pretoque in bolts of the device. In the analysis result on influence of the anti-buckling device using the finite element method, it was found that the compressive strength with the anti-buckling device by loaded bolts was about 7% higher than actual compressive strength. Additionally, compressive tests on specimen with an open hole were performed. The local stress concentration arising from the hole dominates the strength of the laminate rather than the stresses in the bulk of the material. It is observed that the remote failure stress decreases with increasing hole size and specimen width but is generally well above the value one might predict from the elastic stress concentration factor. This suggests that the material is not ideally brittle and some stress relief occurs around the hole. X-ray radiography reveals that damage in the form of fiber microbuckling and delamination initiates at the edge of the hole at approximately 80% of the failure load and extends stably under increasing load before becoming unstable at a critical length of 2-3mm (depends on specimen geometry). This damage growth and failure are analysed by a linear cohesive zone model. Using the independently measured laminate parameters of unnotched compressive strength and in-plane fracture toughness the model predicts successfully the notched strength as a function of hole size and width.

• Proceedings of the Korean Vacuum Society Conference
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• 1999.07a
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• pp.134-134
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• 1999

스퍼터링 소스는 전자기 박막 등 기능성 박막을 비롯하여 결질피막, 장식성 피막등의 제조에 이용되는 것으로 각종 증발원 중에서 가장 널리 사용되는 증발원이다. 70년대 이후 스퍼터링 소스는 마그네트론 스퍼터링으로 대표되는 방식이 사용되어 왔으며 지금까지도 가장 일반적인 방식이 되어 왔다. 마그네트론 스퍼터링 증발원은 증발율에서는 기술적인 향상이 이루어진 반면 이온화율의 향상은 그다지 이루어지지 않아 경질피막과 같은 화합물 피막의 특성 향상에는 한계를 드러내게 되었다. 그러다가 186년 Window 등 이 자장의 세기를 변형시킨 비평형 마그네트론 소스(Unbalanced Magnetron;UBM)를 처음 발표하여 이온화율의 향상이 가능하다는 것이 알려지면서 이에 대한 많은 연구가 진행되었다. UBM 소스는 마그네트론 스퍼터링 소스의 외부에 전자석을 설치하여 기판에 흐르는 이온의 양을 증가시킴으로써 소스와 기판사이의 거리를 증가시킬 수 있고 따라서, 복잡한 형상의 부품코팅이 가능하며 피막 특성을 향상시킬 수 있는 장점이 있다. 본 연구에서는 UBM 스퍼터링 소스를 설계, 제작하여 그 특성을 다양한 측면에서 조사하였다. 특히, 자작의 최적 설계를 통해 전자석의 조건을 도출하였음, Dual UBM 소스의 특성을 동시에 조사하였다. 자기장의 simulation에는 Quick field 프로그램을 이용하였고 기존의 방식과의 비교를 통해 최적의 조건을 도출하였다. 이를 바탕으로 inner pole의 크기를 30mm, outer pole의 크기를 26mm로 고정하여 설계하였고, 외부에 전자석이 설치된 UBM 소스를 제작하였다. 본 UBM 소스는 4" 타겟을 사용할 수 있으며 전자석의 조건을 10A까지 변화시켜 자기장의 세기를 변화시킬 수 있게 하였다. 제작된 소스의 동작조건 설정과 최적화를 위한 스퍼터링 장치를 함께 제작하여 UBM 소스의 최적 동작 조건을 도출하였다. 전자석의 전류가 4.5A일 때 Inner Pole과 Outer Pole의 자기장의 세기가 도일함을 알 수 있었다. 기판과 타겟의 거리가 200mm일 경우에 기판에 흐르는 전류밀도는 2mA/cm2이상이 됨을 확인하였다. 이 결과는 기존의 마그네트론 소스가 기판과 타겟사이의 거리가 100mm일 때 1mA/cm2 정도가 되는 것과 비교하면 이온화율이 획기적으로 향상된 것임을 알 수 있다.수 있다.

• Proceedings of the Korean Vacuum Society Conference
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• 1999.07a
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• pp.135-135
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• 1999

It is well known that Raman spectroscopy is powerful tool in analysis of sp3/sp3 bonding fraction in diamond-like carbon(DLC) films. Raman spectra of DLC film is composed of D-peak centered at 1350cm-1 and G-peak centered at 1530cm-1. The sp3/sp3 fraction is qualitatively acquired by deconvolution method. However, in case of DLC film, it is generally observed that G-peak position shifts toward low wavenumber as th sp3 fraction increases. However, opposite results were frequently observed in ta-C films. ta-C film has much higher residual compressive stress due to its high sp3 fraction compared to the DLC films deposited by CVD method. Effect of residual stress on G-peak position is most recommendable parameter in Raman analysis of ta-C, due to its smallest fitting error among many parameters acquired by peak deconvolution of symmetric spectra. In current study, the effect of residual stress on Raman spectra was quantitatively evaluated by free-hang method. ta-C films of different residual stress were deposited on Si-wafer by modifying DC-bias voltage during deposition. The variation of the G-peak position along the etching depth were observed in the free-hangs of 20~30${\mu}{\textrm}{m}$ etching depth. Mathematical result based on Airy stress function, was compared with experimental results. The more reliable analysis excluding stress-induced shift was possible by elimination of the Raman shift due to residual compressiove stress.

• Proceedings of the KSME Conference
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• 2000.11a
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• pp.250-255
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• 2000

Damage induced by low velocity impact loading in aircraft composite laminates is the form of failure which is occurred frequently in aircraft. Low velocity impact can be caused either by maintenance accidents with tool drops or by in-flight impacts with debris. As the consequences of impact loading in composite laminates, matrix cracking, delamination and eventually fiber breakage for higher impact energies can be occurred. Even when no visible impact damage is observed, damage can exist inside of composite laminates and the carrying load of the composite laminates is considerably reduced. The reduction of strength and stiffness by impact loading occurs in compressive loading due to laminate buckling in the delaminated areas. The objective of this study is to determine inside damage of composite laminates by impact loading and to determine residual compressive strength and the damage growth mechanisms of impacted composite laminates. For this purpose a series of impact and compression after impact tests are carried out on composite laminates made of carbon fiber reinforced epoxy resin matrix with lay up pattern of $[({\pm}45)(0/90)_2]s$ and $[({\pm}45)(0)_3(90)(0)_3({\pm}45)]$. UT-C scan is used to determine impact damage characteristics and CAI(Compression After Impact) tests are carried out to evaluate quantitatively reduction of compressive strength by impact loading.