• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.3
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• pp.204-211
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• 2007

Experiments were conducted to investigate the failure and strengths of carbon composite-to-aluminum single-lap bonded joints with 5 different bonding lengths. Joint specimens were fabricated to have secondary bonding of laminate and aluminum with a film type adhesive, FM73m. Tested joints have the bonding strengths between the values of aluminum-to-aluminum joints and composite-to-composite joints. In the joints with bonding length-to-width ratio smaller than 1, the strength decreases as the bonding length increases. In the joints with the ratio larger than 1, however, the strength converges to a constant value. Final failure mode of all the specimens was delamination. To use the maximum strength of the adhesive, it is important to design the joint to have strong resistance to delamination.

• Composites Research
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• v.36 no.5
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• pp.303-309
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• 2023

In this paper, a Progressive Damage and Failure Analysis (PDFA) modeling method was developed using ABAQUS/EXPLICIT to predict in-plane damage and delamination for Open-Hole Compression (OHC) testing. The proposed PDFA model was constructed based on Hashin criteria and cohesive behavior. The strength and stiffness of OHC specimens with three types of stacking sequences [(45/-45/0₂)₃]_s , [(45/0/-45/90)₃]_s and [45/-45/0/45/-45/90/(45/-45)₂]_s were compared to comprehensively evaluate the validity of the Finite Element(FE) model of PDFA. The strength and stiffness of the OHC specimens were predicted relatively well, with less than a percentage error 10.0 %. For the numerical simulation case for each layup, the damage initiation/evolution of OHC specimens were evaluated for delamination and tension/compression matrix damage before and after failure.

• Composites Research
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• v.13 no.1
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• pp.40-49
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• 2000

The objective of this research is to develop the impact monitoring techniques providing impact identification and damage diagnostics of smart composite laminates susceptible to impacts. This can be implemented simultaneously by using the acoustic waves by the impact loads and the acoustic emission waves from damage. In the previous research, we have discussed the impact location detection process in which impact generated acoustic waves are detected by PZT using the improved neural network paradigm. This paper describes the implementation of time-frequency analysis such as the Short-Time Fourier Transform (STFT) and the Wavelet Transform (WT) on the determination of the occurrence and the estimation of damage.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.7
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• pp.51-59
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• 2004

An experimental procedure to identify failure modes of impact damage using sensor signals and to analyze their general features is examined. A series of low-velocity impact tests from low energy to damage-induced high energy were performed on the instrumented drop weight impact tester to monitor the stress wave signals due to failure modes such as matrix cracking, delamination, and fiber breakage. The wavelet transform(WT) and Short Time Fourier Transform(STFT) are used to decompose the piezoelectric sensor signals in this study. The extent of the damage in each case was examined by means of a conventional ultrasonic C-scan. The PVDF sensor signals are shown to carry important information regarding the nature of the impact process that can be extracted from the careful signal processing and analysis.

• Composites Research
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• v.29 no.4
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• pp.209-215
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• 2016

The failure of composite sandwich structures with thickness and material variation was studied. The main body of the structure is sandwich plate made of the carbon composite face and Aluminum honeycomb core. It is connected with composite laminated flange without core through transition region of tapered sandwich panel with foam core. Tension and compression tests were conducted for the total of 6 panels, 3 for each. Test results showed that the panels under compression are vulnerable to the face failure along the material discontinuity line between two different cores. However the failure load of which panel does not show such failure can carry 16% more load and fails in honeycomb core and face debonding. For the tensile load, the extensive delamination failure was observed at the corner radius which connects the panel and the flange. The average failure load for compression is about 7 times the tensile failure load. Accordingly, these sandwich structures should be applied to the components that endure the compressive loadings.

• Composites Research
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• v.15 no.1
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• pp.41-52
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• 2002

This paper presents the thermo-elastic analysis for searching the behavior of carbon/carbon brake system during the braking period and the 3-D stress analysis to find the shape of the brake disk which is safe to the failure. The mechanical properties of the carbon/carbon brake disk were measured for both in-plane and out of plane directions. The mechanical properties were used as the input of the thermo-elastic analysis and 3-dimensional stress analysis for the brake disk. The gap between rotor clip and clip retainer is an important parameter in the loading transfer mechanism of the rotor disk. The change of gap was considered both the mechanical deformation and thermal deformation. Because the rotor clip and clip retainers were not contacted, they were excluded from the analysis model. Rotor disk was modeled by using the cyclic symmetry condition. The contact problems between rotor clip and key drum as well as between rotor disk and rotor were considered. From the results of the 3-D stress analysis, the stress concentration at the key hole of the brake disk was confirmed. The stress distributions were studied thor the variation of the rotation angle of the contact surface and the radius of curvature at the key hole part.

• Composites Research
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• v.18 no.2
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• pp.20-29
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• 2005

To evaluate the flexural deformation and strength of composite motor case above the glass transition temperature$(T_g),\;170^{\circ}C$, of resin material, a finite element analysis(FEA) model in which material non-linearity and progressive failure mode were considered was proposed. The laminated flexural specimens which have the same lay-up and thickness as the composite motor case were tested by 4-point bending test to verify the validity of FEA model. Also. mechanical properties in high temperature were evaluated to obtain the input values for FEA. Because the material properties related to resin material were highly deteriorated in the temperature range beyond $T_g$, the flexural stiffness and strength of laminated flexural specimen in $200^{\circ}C$ were degraded by also $70\%\;and\;80\%$ in comparison with normal temperature results. Above $T_g$, the failure mode was changed from progressive failure mode initiated by matrix cracking at $90^{\circ}$ ply in bottom side and terminated by delamination at the center line of specimen to fiber compressive breakage mode at top side. From stress analysis, the progressive failure mechanism was well verified and the predicted bending stiffness and strength showed a good agreement with the test results.

• 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.14 no.5
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• pp.46-53
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• 2001

In this research, damage behavior of singly oriented ply (SOP) fiber metal laminate (FML) subjected to concentrated load was studied. The static indentation tests were conducted to study fiber orientation effect on damage behavior of FML. During the static indentation tests, acoustic emission technique (AE) was adopted to study damage characteristics of FML. AE signals were obtained by using AE sensor with 150kHz resonance frequency and the signals were compared with indentation curves of FML. The damage process of SOP FML was divided by three parts, i.e., crack initiation, crack propagation, and penetration. The AE characteristics during crack initiation show that the micro crack is initiated at lower ply of the plate, then propagate along the thickness of the plate with creating tiber debonding. The crack grow along the fiber direction with occurring 60∼80dB AE signal. During the penetration, the fiber breakage was observed. As fiber orientation increases, talc fiber breakage occurs more frequently. The AE signal behaviors support these results. Cumulative AE counts could well predict crack initiation and crack propagation and AE amplitude were useful for the prediction of damage failure mode.

• 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.