Yuichiro Aoki

Damage propagation in CFRP laminates subjected to low velocity impact and static indentation

This paper describes a damage accumulation mechanism in cross-ply CFRP laminates [02/902]2S subjected to out-of-plane loading. Drop-weight impact and static indentation tests were carried out, and induced damage was observed by ultrasonic C-scan and an optical microscope. Both tests gave essentially the same results for damage modes, sizes, and load-deformation history. First, a crack occurred in the bottom 0° layer accompanying some delamination along the crack caused by bending stress. Then, transverse cracks occurred in the middle 90° layer with decreasing contact force between the specimen and the indenter. Measured local strains near the impact point showed that the stress state changed from a bending dominant state to an in-plane tensile dominant state. A cohesive interface element was used to simulate the propagation of multiple delaminations and transverse cracks under static indentation. Two types of analytical models are considered, one with multiple delaminations and the other with both multiple delaminations and transverse cracks. The damage obtained for the model with only multiple delaminations was quite different from that obtained from the experiment. However, the results obtained from the model with both delaminations and transverse cracks well explain the characteristics of the damage obtained in the experiment. The existence of the transverse cracks is essential to form the characteristic impact damage.

Novel Test Method for Mixed Mode II and III Interlaminar Fracture Toughness

A novel test method is proposed to evaluate mixed mode II and III interlaminar fracture toughness of composite laminates. The method is a well-improved test method of the split cantilever test method and easy to conduct, while a computational analysis is necessary to specify the critical toughness values. The aim of the test is to obtain mixed mode fracture toughness. The twisting force that exists in the case of the split cantilever test method is avoided by making the specimen symmetric by introducing two symmetrically located delaminations. By the present method, the two cracks grow stably keeping their configuration with the increase of applied displacement, that is, stable self-similar crack growth is realized. The energy release rate can be numerically evaluated from the experimentally obtained data, such as, the applied load (or applied displacement) as well as the crack configuration. The energy release rate varies along the curved crack front and the dependence of the mixed mode critical energy release rate on the ratio of the mode components can be obtained by one test trial. The method can be conveniently used to roughly estimate the failure criterion for mode II and III mixed mode Interlaminar fracture.

Fracture Resistance of Carbon/Epoxy Composite Laminates under Mixed-Mode II and III Failure and Its Dependence on Fracture Morphology

The fracture resistance of composite laminates under mixed-mode II and III stable damage propagation conditions and its dependence on the fracture morphology were experimentally studied. Cross-ply carbon/epoxy composite specimens were tested by the double notched split cantilever beam (DNSCB) test method. Complex damage consisting of delaminations and shear-induced matrix cracks grew in the middle of the specimen. The delaminations propagated outside the complex damage region. The energy release rate was decomposed to transverse and parallel components to the fiber, G ST and G SL. The component G ST was almost zero at the simple delamination edges and very high in the region of complex damage. The fracture resistances in the regions of complex damage and simple delaminations were numerically evaluated. The fracture resistance was 1.5–2 times higher in the complex damage region than in the simple delamination region.

Damage analysis in composite laminates by using an interface element

A numerical analysis is performed to study an impact damage accumulation problem in composite laminates. A special three-dimensional interface element is developed based on the cohesive model. The element enables the mesh independent configurations of a crack to be calculated and can treat contact problems of the delaminated area. The energy stored in the element per unit area is defined as a function of continuous relative displacements of the delaminated surface. The energy stored before the perfect separation of the interface is equal to the interlaminar fracture toughness. The softening cohesive relations between the tractions and the relative displacements are given by differentiating the energy function with respect to the relative displacements. The maximum value of traction may coincide with interfacial bonding strength. The element is incorporated in a commercially available finite element code. Crack propagation problems under pure Mode I and Mode II loading conditions for three-dimensional models are calculated to show the validity of the present element. The convergence of solution with mesh refinement is examined. The analytical solutions converge smoothly and agree well with the theoretical ones. The present method may be a good tool to simulate the damage accumulation problem of CFRP laminates.

DURABILITY AND DAMAGE TOLERANCE EVALUATION OF VARTM COMPOSITE WING STRUCTURE

Durability and damage tolerance of subcomponent and full-scale wing box structure fabricated by VaRTM are evaluated. Fatigue spectrum with load enhancement factor was applied to the test articles for 1 DSO of 40,000 flights. The Mini-TWIST fatigue spectrum is used for both tests. Then, impact damages are given to the skin stiffened by co-cured stringer and typical skin part by drop-weight to create the delamination. After that, impact damage growth is evaluated during 1 DSO fatigue spectrum and optimal inspection interval is examined. Finally, residual strength of structures is verified by ultimate load test with 150% design limit load. Applied strain level for Subcomponents are intentionally higher than original one in order to evaluate the structural performance in more critical condition. Non-destructive inspection is carried out by 3D ultrasonic scan system with multiple-array sensors to evaluate delamination growth. In Subcomponent test, stringer run-out shows local out-of-plane deformation and that causes disbonding of stringer termination. The disbonding area gradually increases during 1 DSO fatigue test. However, the structure did not show any degradation of structural performance. The damage tolerance tests verify that impact-induced delaminations have not grown throughout the 1 DSO. In the final ultimate load test, the load bearing-capabilities of present VaRTM wing structure have been verified and the structure could survive for 4 seconds without any detrimental deformation and damage growth.

A Study of Quality Assurance of VaRTM Composite Wing Structure

[Abstract] Vacuum assisted resin transfer molding (VaRTM) method is one of the candidates to achieve low cost aircraft structure. A development of 6 m span VaRTM wing box project is conducted to demonstrate VaRTM technology and to make clear the certification procedure of VaRTM by Japan Aerospace Exploration Agency (JAXA). The project result will eventually support the policy and guidance procedures of Japan Civil Aviation Bureau (JCAB), and also a certification procedure of aircraft companies in Japan. This 6 m VaRTM made wing box demonstrator is assumed a 30 passenger aircraft and a static test of this wing box is planed in fiscal year of 2007. A 2 m wingspan specimen was made as an experimental production which contains every technical component to be solved before manufacture of 6 m wing box specimen. Technical expertise and the remaining issues were obtained from this 2 m specimen.

Damage accumulation in composite laminates during quasi-static transverse loading

A mechanism of damage accumulation in composite laminates subjected to a quasi-static concentrated load is numerically studied by using the finite element method to disclose the impact damage problem. The energy release rate distributions are calculated along the delamination edges in square composite laminates with various stacking sequences. Not only circular delaminations but also more realistic impact damage models are investigated to explain the reason why the typical impact damage must be created. The effects of transverse cracks and stacking sequence on the energy release rate are discussed.