Jin-Hwe Kweon

Failure Load Prediction of a Mechanically Fastened Composite Joint Subjected to a Clamping Force

As fiber reinforced composites have been widely used in aircraft and space structures, their joint design has become a very important research area because these joints are often the weakest parts of composite structures. In this article, the failure load of a mechanically fastened composite joint subjected to a clamping force was tested and predicted using the failure area index (FAI) method. The failure load of composite joints subjected to clamping forces on different geometric shapes and dimensions were predicted using the FAI method, and the results were compared with experimental results. From the tests and analyses, the failure load of a mechanically fastened composite joint subjected to a clamping force could be predicted within 23% via the FAI method.

Failure Load Prediction of Composite Joints using Linear Analysis

With the wide application of fiber-reinforced composite materials in aero-structures, robotic arms, and mechanical parts, the design of composite joints has become a very important research area. In this article, linear finite element analyses in which the pin of the composite joint is assumed to be a frictionless rigid body are performed and the failure load of the mechanically fastened composite joint is predicted by the failure area index (FAI) method. From these analyses, it has been found that the FAI method can predict the failure loads of composite joints to within 20.1%.

A Two-dimensional Progressive Failure Analysis of Pinned Joints in Unidirectional-Fabric Laminated Composites

A two-dimensional progressive failure analysis is conducted to predict the failure loads and modes of combined unidirectional-fabric laminated composite joints under pin-loading. The first objective of this study is to investigate the possibility of applying ACOS-J, a two-dimensional progressive damage analysis program, to the strength analysis of composite joints. The second objective is to study how the various failure criteria could be applied, separately or in combination, to pinned joints in combined unidirectional and fabric laminated composites. An eight-node laminated shell element is used for the finite element modeling. Post-failure stiffness is evaluated based on the complete unloading method combined with various failure criteria. A total of 52 specimens with nine different geometries were tested to obtain the experimental strength and failure mode. The results show that a finite element analysis based on the combined Yamada—Sun and Tsai—Wu criteria most accurately predicts the failure loads of the composite laminated joints.

Experimental and numerical study on the failure of sandwich T-joints under pull-off loading

In this study, the failure mechanism of sandwich-to-laminate T-joints under pull-off loading was investigated by experiment and the finite element method. A total of 26 T-joint specimens were manufactured and tested in order to investigate the effects of both adhesive thickness (0.4, 2.0, and 4.0 ㎜) and environmental conditions on the failure of the joints. The results showed that failure occurred mainly as intralaminar failure in the first layer of the sandwich face, which was contacted to the paste adhesive. The failure load did not significantly change with increasing adhesive thickness in both RTD (Room Temperature and Dry) and ETW (Elevated Temperature and Wet) conditions. In the case of ETW conditions, however, the failure load increased slightly with an increase in adhesive thickness. The joints tested in ETW conditions had higher failure loads than those tested in RTD conditions. In addition to the experiment, a finite element analysis was also conducted to investigate the failure of the joint. The stress inside the first ply of the sandwich face was of interest because during the experiment, failure always occurred there. The analysis results showed good agreement with the trend of experimental results, except for the case of the smallest adhesive thickness. The highest stress was predicted in the regions where initial failure was observed in the experiment. The maximum stress was almost constant when the adhesive thickness was beyond 2 ㎜.

Optimization of Composite Laminates Subjected to High Velocity Impact Using a Genetic Algorithm

In this study, a genetic algorithm was utilized to optimize the stacking sequence of a composite plate subjected to a high velocity impact. The aim is to minimize the maximum backplane displacement of the plate. In the finite element model, we idealized the impactor using solid elements and modeled the composite plate by shell elements to reduce the analysis time. Various tests were carried out to investigate the effect of parameters in the genetic algorithm such as the type of variables, population size, number of discrete variables, and mutation probability.

Failure Load Prediction by Damage Zone Method for Single-lap Bonded Joints of Carbon Composite and Aluminum

A damage zone method based on 3D finite element analysis was proposed to predict the failure loads of single-lap bonded joints with dissimilar composite-aluminum materials. To simulate delamination failure, interply resin layers between any two adjacent orthotropic laminas of composite adherend were assumed with a thickness of one-tenth of a composite lamina. Geometrically nonlinear effects due to the large rotation of the single-lap joint were included in the analysis. Analysis also considered the material nonlinearity of the aluminum adherend due to the stress exceeding yield level. Based on the experimental observation that the failure modes of the specimens were dominated by delamination and debonding, the Ye-criterion was applied to account for the out-of-plane failure of composite adherend and the Von Mises strain criterion was applied for the adhesive layer. The failure indices were multiplied to the predicted damage zone as a weight factor and the calculated damage zones were divided by an area or volume considering the joint geometry. Predicted failure loads show deviation within 18% from experimental results for nine different bonding lengths or adherend thicknesses.

Structural Analysis of a Composite Target-drone

A finite element analysis for the wing and landing gear of a composite target-drone air vehicle was performed. For the wing analysis, two load cases were considered: a 5g symmetric pull-up and a -1.5g symmetric push-over. For the landing gear analysis, a sinking velocity of 1.4 m/s at a 2g level landing condition was taken into account. MSC/NASTRAN and LS-DYNA were utilized for the static and dynamic analyses, respectively. Finite element results were verified by the static test of a prototype wing under a 6g symmetric pull-up condition. The test showed a 17% larger wing tip deflection than the finite element analysis. This difference is believed to come from the material and geometrical imperfections incurred during the manufacturing process.

Numerical Modeling of Combined Matrix Cracking and Delamination in Composite Laminates Using Cohesive Elements

Sub-laminate damage in the form of matrix cracking and delamination was simulated by using interface cohesive elements in the finite element (FE) software ABAQUS. Interface cohesive elements were inserted parallel to the fiber orientation in the transverse ply with equal spacing (matrix cracking) and between the interfaces (delamination). Matrix cracking initiation in the cohesive elements was based on stress traction separation laws and propagated under mixed-mode loading. We expanded the work of Shi et al. (Appl. Compos. Mater. 21, 57–70 2014) to include delamination and simulated additional [45/−45/0/90]s and [02/90n]s {n = 1,2,3} CFRP laminates and a [0/903]s GFRP laminate. Delamination damage was quantified numerically in terms of damage dissipative energy. We observed that transverse matrix cracks can propagate to the ply interface and initiate delamination. We also observed for [0/90n/0] laminates that as the number of 90° ply increases past n = 2, the crack density decreases. The predicted crack density evolution compared well with experimental results and the equivalent constraint model (ECM) theory. Empirical relationships were established between crack density and applied stress by linear curve fitting. The reduction of laminate elastic modulus due to cracking was also computed numerically and it is in accordance with reported experimental measurements.

A Numerical Study of the High-Velocity Impact Response of a Composite Laminate Using LS-DYNA

The failure of a Kevlar29/Phenolic composite plate under high-velocity impact from an fragment simulation projectile was investigated using the nonlinear explicit finite element code, LS-DYNA. The composite laminate and the impactor were idealized by solid elements, and the interface between the laminas was modeled as a tiebreak type in LS-DYNA. The interaction between the impactor and laminate was simulated using a surface-to-surface eroding contact algorithm. When the stress level meets the given failure criteria, the layer in the element is eroded. Numerical results were verified through existing test results and showed good agreement.

Crippling analysis of composite stringers based on complete unloading method

This paper addresses a nonlinear finite element method for the crippling analysis of composite laminated stringers. Composite stringer is idealized by the nine-node laminated shell element based on the first order shear deformation theory. The stiffness degradation by the local failure is simulated by the complete unloading method. A modified arc-length algorithm is incorporated into the nonlinear finite element method to trace the post-failure equilibrium path after local buckling. Finite element results are compared with those by experiments and show the excellent agreement. A parametric study is performed to assess the effect of the flange-width, web-height, and stacking sequence on the buckling, local buckling, and crippling stresses of stringers.