In the process of fabrication, installation, maintenance, and use, the laminate structures of composite will inevitably experience impact from foreign objects, resulting in delamination damage. In the aerospace vehicle service process, the fatigue damage evolution and failure mechanism become intricate due to the coupling and influence of various impact-induced damage modes under complex fatigue loads. This paper introduces initial impact damage into two typical laminate structures using the drop hammer impact method. Based on A-scan and C-scan technologies, the damage modes were obtained, and compressive buckling fatigue tests were carried out. It was found that after one million cycles, no damage propagation was observed, and there were no significant changes in the stiffness, buckling load, and load-bearing capacity of the specimens. This verifies the design concept of static coverage fatigue and non-propagation of damage under compression loads for skin laminates, providing basis and support for the composite design of aircraft.
Composite structures used in aircraft need to deal with a wide variety of climatic conditions, such as high temperature, cold temperature, and hydrothermal environment, which probably affects their bearing capability and life expectancy. Additionally, tapered composite laminates are commonly adopted to accommodate the need for varying thicknesses in aircraft structures. Consequently, environmental influence on the tapered composite structures has become a major concern in the utilization of composites. In this work, an experimental program is executed on the tapered composite structures in room temperature atmosphere (RTA), cold temperature dry (CTD), and elevated temperature wet (ETW) conditions separately. The results indicate that the large stress concentration induced by the presence of discontinuous plies in the tapered section motivates the occurrence of delamination at a relatively low load. Environmental exposure does not have a distinct effect on the delamination initiation load. In the hydrothermal environment, the degradation of tensile strength is negligible, and the delamination process as well as the failure pattern of ETW specimens are similar to RTA ones. However, cryogenic conditions result in a severely deteriorating effect on the tensile strength. The failure pattern of CTD specimens exhibits a brittle fracture characteristic and no obvious yield process is observed.
The load-bearing capacity of the stiffener runout structure is significantly reduced by interface debonding due to load eccentricity and stiffness discontinuity. This study investigates that failure behavior of the stiffener runout with crack stop bolts under tension load considering the influence of debonding at the skin/stiffener interface. Finite element models of pre-disbonding between stiffener and skin interface of stiffener runout with bolts were established using the ABAQUS platform. The bolt reinforcement area was modeled using the Global Bolted Joint Model (GBJM). The effect of the prefabricated debonding area on the damage initiation and structural load carrying capacity was analyzed, and the bolt transfer loads of the structure with different prefabricated debonding areas were compared. The results demonstrate that the damage initiation load increases with the expansion of the interface debonding area, and the damage load of the structure is almost independent of the interface debonding area. Additionally, the larger the intact area of the stiffener-skin bonding interface, the smaller the total load transferred through the crack stop bolt.
According to the two failure mechanisms of fatigue fracture of fasteners, the fatigue damage evolution equation of microcrack initiation and propagation of fasteners is proposed, and the crack initiation life and crack propagation life models are deduced. After superposition, the fatigue model of fastener fracture failure is obtained. Based on the fatigue test results, two widely used fatigue life models (exponential and power functions) are compared. The results show that the exponential type and power function type only fit the test data without considering the different failure modes of the structure. Based on the failure mechanism, the fitting results of the model are in good agreement with the experimental results, and have great reference value in the design of fatigue structure.
Liquid Composite Molding (LCM) processes are more cost-effective compared to autoclave-cured prepreg, but an independent preforming step is typically required to convert 2D fabric blanks into complex 3D shapes prior to molding. Numerical models are therefore important to predict the formation of defects during the design phase, in order to ensure the quality of final composite components. A macroscopic finite element model was employed to predict the forming behavior of multi-layered biaxial Non-Crimp Fabrics (NCF) during the press tool forming using a hemispherical punch. The forming behavior of the NCF was predicted by simulations considering the bending stiffness of the NCF, enabling fabric wrinkling to be simulated. Simulation results indicate a correlation between fabric wrinkling and the in-plane shear deformation of fabrics. The severity of wrinkles was also influenced by the layup sequence. Compared with the single orientation layups, more wrinkles were predicted for the layup comprising plies stacked at different orientations due to the dissimilar shear deformation between these plies.
Based on digital image correlation method (DIC), the shear stability of two configurations of composite stiffened plates was experimentally studied. The test results show that the buckling load and failure load of I-stiffened panels are 158.2 kN and 282.1 kN respectively, and the main failure models are the debonding of the truss skin interface and skin folding; the buckling load and failure load of M-stiffened panels are 181.9 kN and 292.3 kN respectively, and the main failure modes are consistent with that of T-stiffened panels. Then, a progressive damage failure model of stiffened panels was established to simulate the buckling and post-buckling behaviour. The results show that the buckling mode, buckling load, failure load and failure mode are consistent with the experimental results.
Access to the requested content is limited to institutions that have purchased or subscribe to SPIE eBooks.
You are receiving this notice because your organization may not have SPIE eBooks access.*
*Shibboleth/Open Athens users─please
sign in
to access your institution's subscriptions.
To obtain this item, you may purchase the complete book in print or electronic format on
SPIE.org.
INSTITUTIONAL Select your institution to access the SPIE Digital Library.
PERSONAL Sign in with your SPIE account to access your personal subscriptions or to use specific features such as save to my library, sign up for alerts, save searches, etc.