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The purpose of this paper is to use the fiber metal laminates (FML) composites as a patch for repairing a single notched specimen made of AL1035 aluminum alloy. The FML composite patch was bonded on one side of the cracked specimens by adhesive Araldite 2015. Then the fatigue crack growth tests were conducted on the specimens and the effects of both FML patch lay-up sequence and pre-crack angle on the fatigue life were investigated. Finally, the effect of repairing on the fracture parameters (SIF and crack propagation direction) at the crack front has also been calculated using three-dimensional finite element analysis.

The fatigue crack growth tests were conducted on the specimens and the effects of both FML patch lay-up sequence and pre-crack angle on the fatigue life were investigated.

The results show that the fatigue life of the patched specimens with inclined crack increased approximately 2-6.02 times compared to the un-patched specimens. In addition, the fatigue crack growth rate decreased significantly when the patch was used. Generally, the FML patch with Plate-Fiber-Fiber-AL lay-up has more efficiency than other lay-up sequences.

Recently, composite patches are used in the structure repair processes to increase the service life of cracked components. The bonded patch method is one of the efficient methods among repairing methods. Today, the FMLs are used in the aircraft structures as a replacement of high-strength aluminum alloys due to their lightweight and high-strength properties. Many researches have been performed on single and double side repaired panels using composite patches. In this study, the FML composites have been used as a patch for repairing a single notched specimen made of AL1035 aluminum alloy.

The maintenance of aircraft structure with lower cost is one of the prime concerns to regulatory authorities. The carbon fiber-reinforced polymer (CFRP) patches are widely used to repair the cracked structure. The demands and application of CFRP compel its price to increase in the near future. A distinct perspective of repairing the cracked aluminum panel with the hybrid composite patch is presented in this paper. The purpose of this paper is to propose an alternative patch material in the form of a hybrid composite patch which can provide economical repair solution.

The patch hybridization is performed by preparing the hybrid composite from tows of carbon fiber and glass fiber. Rule of hybrid mixture and modified Halpin–Tsai’s equation are used to evaluate the elastic constant. The stress intensity factor and interfacial stresses are determined using finite element analysis. The debonding initiation load is evaluated after testing under mode-I loading condition.

The hybrid composite patch has rendered the adequate performance for reduction of stress intensity in the cracked panel and control of interfacial stresses in the adhesive layer. The repair efficiency and repair durability of the composite patch repair was ensured by incorporation of the hybrid composite patch.

The studies involving patch hybridization for the application of composite patch repair are presently lacking. The influence of the patch stiffness, methodology to prepare the hybrid composite patch and effects of hybridization on the performance of composite patch repair is presented in this paper.

The purpose of this paper is to study the deterministic elastic buckling behavior of cylindrical fiber–metal laminate panels subjected to uniaxial compressive loading and the investigation of GLAss fiber-REinforced aluminum laminate (GLARE) panels using probabilistic finite element method (FEM) analysis.

The FEM in combination with the eigenvalue buckling analysis is used for the construction of buckling coefficient–curvature parameter diagrams of seven fiber–metal laminate grades, three glass-fiber composites and monolithic 2024-T3 aluminum. The influences of uncertainties concerning material properties and laminate dimensions on the buckling load are studied with sensitivity analyses.

It is found that aluminum has a stronger impact on the buckling behavior of the fiber–metal laminate panels than their constituent uni-directional or woven composites. For the classical simply supported boundary conditions, it is found that there is an approximately linear relation between the buckling coefficient and the curvature parameter when the diagrams are plotted in double logarithmic scale. The probabilistic calculations demonstrate that there is a considerable probability to overestimate the buckling load of GLARE panels with deterministic calculations.

In this study, the deterministic and probabilistic buckling response of fiber metal laminate panels is investigated. It is shown that realistic structural uncertainties could substantially affect the buckling strength of aerospace components.

Fibre metal laminates were developed at Delft University during the last two decades as a family of new hybrid materials consisting of bonded thin metal sheets and fibre/adhesive layers. This laminated structure provides the material with excellent fatigue, impact and damage tolerance characteristics and a low density. While the 20 per cent weight reduction was the prime driver behind the development of this new family of materials, it turns out that additional benefits like cost reduction and an improved safety level have become more and more important. The combination of these aspects in one material makes fibre metal laminates a strong candidate material for fuselage skin structures of the new generation of high capacity aircraft. The focus on this application currently leads to industrialization and qualification that makes this material available to the aircraft designer.

Since failure of laminated composites by delaminations is common, the purpose of this paper is to present a numerical procedure to check the stability of delaminations in fiber metal laminate (Glare), with different possible damage configurations, under uni-axial tension. Deformation behavior of the laminate is also examined. Influence of the type and the extent of damage, represented by varying sizes and number of delaminations, on delamination driving force and laminate deformation is found.

Delaminated Glare is modeled by finite element method. Interface cohesive elements are used to model the delaminations. Finite element results provide the deflection/deformation characteristics of the laminate. Driving forces of delaminations are estimated by J integrals that are numerically obtained over cyclic paths near delamination tips. Laminates with different types of delaminations are also fabricated and externally delaminated for measurement of their interlaminar fracture toughness. The delamination is considered to be stable if its driving force is less than corresponding interlaminar fracture toughness of the laminate.

Delaminations are found to be stable in laminates with lower number of delaminations and unstable in laminates with higher number of delaminations. Increase in size of delaminations increases the deformations but reduces the delamination driving force whereas increase in number of delaminations increases both deformations and driving forces. The trends change in case of laminates with symmetrical damage. Shape of delamination is also found to influence the deformations and driving forces. The finite element model is validated.

There is scope for validating the numerical results reported in the paper by theoretical models.

Checking the stability of delaminations and their effect on deformation behavior of the laminate helps is assessment of safety and remaining life of the laminate. If failure is predicted, preemptive action is taken by using repair patch ups at identified critical locations in order to avoid failures in service conditions.

The paper offers the following benefits: use of cohesive zone method that is readily possible in finite element procedures and is relatively simple, fast and reasonably accurate is demonstrated; suitability of using J integrals over paths crossing non-homogeneous and property mismatched material layers is tested; and influence of the type and the extent of damage in the laminate on its deformation behavior and delamination driving forces is found. This type of work has not been reported so far.