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Composite stiffened panels always have defects at the bonding surface in preparation,which accelerates the debonding of the interface between skin and stringer,greatly affects the bearing capacity of the structure. The model of stiffened panel with prefabricated debonding at the interface of stiffened strip and skin was established with both cohesive zone model（ CZM）and virtual crack closure technique（ VCCT） in ABAQUS,and the rationality of model was verified by experiments. The distribution of energy release rate in prefabricated debonding leading edge was given and the effect of debonding position and area on stiffness and carrying capacity was analyzed. The results indicate that the contribution of type II energy release rate to damage initiation is the greatest; the bearing capacity of the prefabricated debonding structure is stronger than that is in the middle when it is located in outside; the carrying capacity of the structure decrease with the increase of prefabricated debonding area.

as multi-spar composite box has been used for many airplanes,in order to assure structure security,it is necessary to study the mechanical properties of multi-spar composite box. Aiming at the buckling and postbuckling problems,buckling behavior and damage process of multi-spar composite box under bending-shearing load is researched in this paper. A self-defined nonlinear degradation rule was presented to simulate the propagation of the interface elements which were discretized with solid elements. In this model,the onset and propagation of damage can be predicted by quadratic failure criteria and mixed-mode strain energy release rate criteria respectively. cohesive elements are used to simulate the adhesive layer between skin and flange.composite structures were discretized with continuum shell elements. In this paper,a failure algorithm based upon Puck’s failure theory has been developed. Based on ABAQUS,a material degradation rule of composite was presented. The process of damage initiation,propagation and failure of the stiffened composite panels was simulated in detail. At last,a good agreement between tests and numerical prediction was observed.

The progressive damage analysis technique was studied by using finite element method and test validation of the repaired aircraft wing composite panel with discrete-source damage. Firstly,an analysis model adopting the two dimensional LaRC05 and B-K failure criteria was established to indentify the initial damage of composite ply and interface. Then,a user defined subroutine UMAT of material failure criteria and degradation rule based on commercial software ABAQUS was programmed and the load carrying capacity of the repaired aircraft wing composite panel was analyzed. The numerical results indicated that: the FEM analysis method presented in this paper was viable for the large-scale bolted repair structure. The comparison between analysis and test indicates that the load carrying capacity and failure mode of numerical results agree well with test results. The errors of predicted load carrying capacity are within 5%.

Since the wing is the principle bearing structure of aircraft,wing skin is repaired by using mechanically fastened joints. In this paper,scheme design,experimental verification,finite element analysis and design improvement investigation of aircraft composite wing panel with discrete-source damage were carried out. Firstly,the specimen with discrete-source damage and the repaired which used fastened joint repair method were designed and manufactured. Experimental verification of scheme design was performed. Load-carrying capacity of the repaired specimen increases about 124%,compared to the unrepaired,which fully meet the design requirements. Then,for further analysis,two different specimens’ FEM models were established.The results showed that the initial repaired scheme was conservative,and the cost of structure weight was increased. Finally,three improved design configurations were proposed,and the load-carrying capacity of configuration II increased by 4. 7%,its structure weight decreased by 3. 9%. This repaired scheme provides reference for the repair design of aircraft composite wing panel with discrete-source damage.

Fastened joint repair is a main repair method for composite structure with advantages of easy disassembling,high load transferring capacity. It is important to evaluate the residual strength in repair design process. In this paper,the tensile load bearing capacity of metal riveted repair laminates were evaluated by numerical simulation using progressive failure analysis method. Laminate was modeled using shell element whereas the fasteners were modeled by beam elements. MPC technique was adopt to simulate the constraint between fastener and hole. The USDFLD subroutine was used to degrade material stiffness by multi-lever manner while damage occurred. The failure analysis results including failure mode,failure load are in good agreement with test results. The errors of predicted failure loads are with 6%,which shows the proposed method is helpful to composite fastened joint repair design.

For the optimal design of the embedded methods and size parameters of buckypaper in the buckypaper/SMP composites, FLUENT is used to analyze the effect of shape and thickness of buckypaper on the temperature distribution of the buckypaper/SMP composites in the heating process. The analysis show that, compared with composites reinforced by flat buckypaper, the maximum and average temperature of the composites reinforced by pulse bending buckypaper are lower, the heating rate is slowly and the time to reach the steady state is shorter in the heating process under the same working conditions, but the temperature distribution is more even. It is concluded that the shape and thickness of buckypaper can change the unit volume heat of buckypaper, thereby affecting the temperature field distribution of buckypaper/SMP composites. When the unit volume heat of buckypaper is bigger, the heat production is bigger, the maximum and average temperature are higher, and the temperature distribution is more uneven. Because the heat dissipation rate is lower than the increase rate of heat production, it takes longer time for buckypaper/SMP to reach the steady state. [ABSTRACT FROM AUTHOR]

The influences of porosity, impact energies and moisture contents on the residual tensile strength and bending strength after impact of CFRP laminates are evaluated. CFRP laminates of three porosity levels were immersed in water for 7, 14 days and moisture saturation. Impact tests on the CFRP laminates of three porosity levels with different immersion time were conducted at five impact energy levels of 3 - 15J at room temperature, followed by the residual tensile test and bending test performed on the specimens. The experimental results show that the residual tensile strength and bending strength after impact decreased significantly with the increasing impact energy. For the same impact energy, no obvious effects of void contents on the residual tensile strength and bending strength after impact were observed. The delamination induced by the impact would reduce the influence of void on the residual strength of the impacted specimens. [ABSTRACT FROM AUTHOR]