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The purpose of this work is to describe the effects of the length of cracks and the patch size on the stress intensity factors in a bonded composite repair structure containing multiple site damage.

Finite element method was applied to simulate a bonded repair of a cracked aluminum plate with multiple site damage. A two‐dimensional three‐layer technique was utilized to model damage in a typical aluminum plate with collinear twin cracks.

This research has found that the stress intensity factors of collinear twin cracks can be reduced significantly through bonded composite repair, and their values strongly depend on the relative position of the cracks. Moreover, the composite patch should be 1.5 to two times longer than the crack length and the patch thickness should be 30‐40 percent of the plate thickness for the best repair performance.

Patch debonding can significantly reduce the repair efficiency and should be avoided if possible.

It is seen that, instead of the three‐dimensional finite element model, which is computationally intense, the two‐dimensional three‐layer finite element model has an adequate accuracy to obtain stress intensity factors in a bonded composite repair structure with multiple site damage.

Large structural objects, primarily concrete bridges, can be reinforced by gluing to their stretched surface tapes of fiber-reinforced polymer (FRP). The condition for this technology to work requires the quality of the bonding of FRP and the concrete to be perfect. Possible defects may arise in the phase of construction but also as a result of long-term fatigue loads. These defects having different forms of voids and discontinuities in the bonding layer are difficult to detect by optical inspection. This paper aims to describe the development of a rapid and nondestructive method for quantitative assessment of the debonding between materials.

The applied technique belongs to the wide class of active infrared (IR) thermography, the principle of which is to heat (or cool) the investigated object, and determine the properties of interest from the recorded, by an IR camera, temperature field. The methodology implemented in this work is to uniformly heat for a few seconds, using a set of halogen lamps, the FRP surface attached to the concrete. The parameter of interest is the thermal resistance of the layer separating the polymer tape and the concrete. The presence of voids and debonding will result in large values of this resistance. Its value is retrieved by solving an inverse transient heat conduction problem. This is accomplished by minimizing, in the sense of least squares, the difference between the recorded and simulated temperatures. The latter is defined as a solution of a 1D transient heat conduction problem with the already mentioned thermal resistance treated as the only decision variable.

A general method has been developed, which detects debonding of the FRP tapes from the concrete. The method is rapid and nondestructive. Owing to a special selection of the compared dimensionless measured and simulated temperatures, the method is not sensitive to the surface quality (roughness and emissivity). Measurements and calculation may be executed within seconds. The efficiency of the technique has been shown at a sample, where the defects have been artificially introduced in a controlled manner.

A quantitative assessment procedure which can be used to determine the extent of the debonding has been developed. The procedure uses inverse technique whose result is the unknown thermal resistance between the member and the FRP strip.

Various types of damage or cracking in the structural components of an airframe can occur during the service lifetimes of aging aircraft. These types of damage are commonly repaired with a patch that can be joined to the original structure by different techniques, e.g., riveting and bonding. The purpose of this paper is to describe the repair of a fatigue crack in the metallic wing structure of a jet trainer aircraft using an adhesively bonded boron composite patch.

The partial analytical design and numerical analysis of the repair is presented. Three different versions of the patch are quantitatively investigated. The efficiency of the designed adhesively bonded boron patch with the parent metallic structure is experimentally verified by panel tests, and two different patch geometries and two surface preparation techniques are investigated. The panels were designed, manufactured and tested as representative structures of the repaired structure.

Adhesively bonded composite repair increases the lifetime by at least one order compared with the non-repaired structure. Both surface preparations provide equivalent results. The repair lifetime is significantly influenced by the patch geometry, and the longer patch significantly increases the lifetime of the panel. The lifetime of the structure can be increased by ˜40-fold if the patch geometry is a rectangle with 1:1.5 proportions of the sides (length in the crack direction/length perpendicular to the crack propagation). The patch length in the crack direction should be twice that of the initial crack length. Additional patch length extension in the direction that is perpendicular to the crack propagation does not appear to be effective for significantly decreasing the stress intensity factor and patch efficiency. The repair also retards the crack propagation if the crack grows out of the patch. No significant disbonding was detected.

The work described in this paper provides information that is very useful for patch design and verification with relation to different patch geometries and technologies. The designed and verified repair has been successfully applied to an L-39 Czech aircraft structure.

This study provides an analysis of patch repair for cracked aircraft structures. Delamination is a type of damage that affects the patch's behavior. The purpose of this study is to assess the influence of delamination on repair performance.

An analytical and numerical study using the finite element method was conducted for a cracked plate repaired with a patch containing a pre-existing delamination defect. The method for defining the contact pair surfaces and modeling the delamination interaction within the patch interface is specified using the virtual crack closure technique (VCCT) approach.

The efficiency of the repair is measured in terms of the J-integral. The effects of delamination initiation, mechanical loading, crack length and patch stacking sequences are presented. It is noted that in mode I, delamination propagation is only significant at node A. The numerical results are in good agreement with those of the analytical solution found in the literature. It is observed that the patch's behavior is strongly dependent on loading, crack size and stacking sequences in terms of reducing the structure's lifespan, especially in the presence of delamination.

The numerical modeling presented by the VCCT approach is highly valuable for studying delamination evolution. The influence of loading, crack size and stacking sequences on repair performance is discussed in this work.

In this article, a novel hybrid composite patch consisting of unidirectional carbon fiber and glass fiber is considered for repair of the aircraft structure. The purpose of this paper is to assess the performance of hybrid composite patch repair of cracked structure and propose an optimized solution to a designer for selection of the appropriate level of a parameter to ensure effective repair solution.

Elastic properties of the hybrid composites are estimated by micromechanical modeling. Performance of hybrid composite patch repair is evaluated by numerical analysis of stress intensity factor (SIF), shear stress, and peel stress. Design of experiment is used to determine responses for a different combination of design parameters. The second-order mathematical model is suggested for SIF and peel stress. Adequacy of the model is checked by ANOVA and used as a fitness function. Multiobjective optimization is carried out with a genetic algorithm to arrive at the optimal solution.

The hybrid composite patch has maintained equilibrium between the SIF reduction and rise of the peel stress. The repair efficiency and repair durability can be ensured by selection of an optimum value of volume fraction of glass fiber, applied stress, and adhesive thickness.

The composite patch with varying stiffness is realized by hybridization with different volume fraction of fibers. Analysis and identification of optimum parameter to reduce the SIF and peel stress for hybrid composite patch repair are presented in this article.

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 study is to determine the optimal shape of a one-sided elliptical composite material patch of an adhesively bonded repair of cracked metal plates under biaxial stress.

The approach consists on determining the patch topology and adhesive thickness that minimize the stress intensity factor and the bending moment caused by the asymmetry of the repair by applying a differential evolution algorithm with a selection phase using the Deb’s rules.

The results demonstrate that an elliptical patch of major axis length equal to the plate width, and minor axis length equal to the crack length, with a thin adhesive thickness, provides the highest stress intensity factor and bending moment reduction, maximizing the fatigue life of the repair.

The results are limited to linear elastic behavior of the cracked plate and a fully rigid bond between the cracked plate and the patch. The effectiveness of the repair was verified by theoretical calculation of the fatigue life, thus experimental validation is still needed.

The results of this work can be applied to experimental validations of the effectiveness of the elliptical one-side composite bonded repairs, avoiding and extensive number of experiments, and also, encourage maintainers to explore on this technique that is more economical and easier to apply, in comparison to other repair techniques. By following the patch geometry recommendations proposed herein, it is analytically predicted that the fatigue life may increase by as much as 27 times that of the unpatched plate.

Currently, there are no detailed studies that assess one-side patch repair procedures, which require consideration of the bending moment and biaxial stress state, and therefore, the optimal patch geometry and adhesive thickness are unknown.

This paper comprehensively addresses the influence of chopped strand mat glass fiber-reinforced polymer (GFRP) patch configurations such as geometry, dimensions, position and the number of layers of patches, whether a single or double patch is used and how well debonding the area under the patch improves the strength of the cracked aluminum plates with different crack lengths.

Single-edge cracked aluminum specimens of 150 mm in length and 50 mm in width were tested using the tensile test. The cracked aluminum specimens were then repaired using GFRP patches with various configurations. A three-dimensional (3D) finite element method (FEM) was adopted to simulate the repaired cracked aluminum plates using composite patches to obtain the stress intensity factor (SIF). The numerical modeling and validation of ABAQUS software and the contour integral method for SIF calculations provide a valuable tool for further investigation and design optimization.

The width of the GFRP patches affected the efficiency of the rehabilitated cracked aluminum plate. Increasing patch width WP from 5 mm to 15 mm increases the peak load by 9.7 and 17.5%, respectively, if compared with the specimen without the patch. The efficiency of the GFRP patch in reducing the SIF increased as the number of layers increased, i.e. the maximum load was enhanced by 5%.

This study assessed repairing metallic structures using the chopped strand mat GFRP. Furthermore, it demonstrated the superiority of rectangular patches over semicircular ones, along with the benefit of using double patches for out-of-plane bending prevention and it emphasizes the detrimental effect of defects in the bonding area between the patch and the cracked component. This underlines the importance of proper surface preparation and bonding techniques for successful repair.

Bonded composite repair technology was pioneered at ARL and has been successfully used to repair a wide range of military and civilian aircraft. The technique has proved to be very cost effective and has been shown to increase the availability of aircraft by significantly reducing the repair time. ARL has undertaken collaborative work with a number of airline operators, the CAA and FAA and with Boeing to demonstrate the advantage of the technology when compared to conventional repair methods. The Defence Science and Technology Organisation (DSTO) has signed a licence agreement with an Australian company. Helitech P/L, to exploit this technology on a world‐wide basis.