The their results revealed that strain energy release

The composite technology is widely used in aerospace applications to
harness its specific characteristics. Thin walled composite panels are among the most utilized
structural elements in aerospace structures. They are made of advanced
composite parts/components, joined together to form the structure of the required
geometry and are subjected to any combination of in-plane, out-of-plane and
shear loads during its application. Adhesive
joints are employed in advanced aerospace composite structure assemblies to
realize light weight structures. However, composite structures are susceptible to performance
reduction in the presence of defects, due to manufacturing issues or damages
during service loads.  Debonds occurring in the adhesively bonded composite
joints due to manufacturing issues have great impact
on the load carrying capability and they are a major concern in aerospace
structures. Hence it necessitates the assessment of damage tolerance of bonded
composite joints with inherent flaws (like debonds) in the design of modern
aerospace vehicles where adhesive joints are widely used.

Early
studies on composite skin-stiffener debonded configurations are reported by
Wang et al. 1. Bolotin 2 reviewed
and presented the most of the aspects concerning the delaminations and other
interlaminar crack-like defects in composite structures. Yap et al. 3 have proposed a comprehensive
finite element method to study the effect of skin-to-stiffener debonding. Pradhan
et al. 4 carried out parametric
studies on debonding in adhesively bonded composite joints and their results
revealed that strain energy release rate is sensitive to the orientation of
fibers in the composite adherends.

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Studies
on buckling of composite plates having debonds were reported in literature. Kim
and Kwon 5,6 have studied the effect of open disbonds in the composite
sandwich panel flange joint under compressive loading environment. Kwon and Kim
7 have reported that debond growth initiation was strongly affected by
adhesive fracture toughness and flange stiffness. Debond length and flange
width strongly affect buckling but were found to mildly influence debond growth
initiation. The significance of adhesive thickness and curing conditions on the
critical energy release rate also envisaged. Mikulik et al.
8 employed fracture mechanics based crack tip element methodology to predict
the skin-to-stiffener separation. Similar studies on debonding
in adhesively bonded composite stiffened panels were reported by various
investigators 9-11. Few review articles on
damage tolerant analysis were reported in literature 12-16. da Silva et al. 17, 18 presented an extensive
survey on the analytical models for adhesively bonded joints both single and
double lap joints and their review reveals that almost all models for lap
joints are two dimensional and linear elastic for both adherends and adhesive.

Damage tolerant design is very challenging and requires
expertise in damage mechanics, fracture mechanics, structural mechanics,
material science, and physics to guide the experimental and analytical work. To
acquire the knowledge on damage tolerant design, it is essential to know the
effect of active debonds on the bonded joints during loading. Though the
studies on the adhesively bonded joints with debonds are reported in
literature, it is evident that literature on onset of growth of closed debonds on
the adhesively bonded joints are limited. This
motivated the study on the structural response of adhesively bonded composite
joints containing closed (embedded) active debonds. The finite element (FE)
tool has been proved useful in predicting the behavior of composite structures
in the presence of active defects. One of the most popular methods implemented
in FE tool to analyze debond/crack propagation is Virtual Crack Closure
Technique, based on fracture mechanics concepts, which is detailed by Krueger 19,20.
The method allows obtaining the strain energy release rates and is based on the
assumption that, when a crack grows the energy released in this process is
equal to the work necessary to close the crack to its initial length before
propagation. The objective of the present work is to perform a comprehensive study
on the onset of initiation of debond growth in adhesively bonded composite
joints under compressive load. The influence of parameters such as laminate
sequence, debond location, size and its shapes (square and circular) are
investigated using VCCT with mixed-mode failure criteria.