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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 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 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 purpose of this study is experimental research of the mechanical behavior of slab reinforced by cork composite patch submitted to an eccentric progressive compressive load applied to on impact rectangle of dimensions 28 × 23 cm². An analytical model and numerical modeling by finite elements are performed. This study is motivated by the evaluation of the effectiveness of this type of partial reinforcement to improve strength and ductility. The results are given by load-displacement curves, tensile damages cartography and ultimate strength histogram.

In experimental protocol, the following two parameters have been considered: the dimensions of the patch and the eccentricity of the load. The sections of the patches are calculated so that the ratio (XP/YP) patch is proportional to the ratio (LD/lD), with a step of 6 cm longitudinally and 4 cm transversely. Several dimensions patches are considered: (6 × 4) cm², (12 × 8) cm² and (18 × 12) cm². The eccentric punching loading test was performed with an eccentricity of the load (1/3) L’ and (2/3) L’ compared to the center of gravity of the slab. Taking into account the eccentricity of the load in estimating the rupture strength, the equations are developed. Thus, numerical simulations are carried, to extract tensile damages cartography.

The results show that the rupture begins with the appearance of cracks in the unreinforced area. For an eccentricity of 1/3L’, the best strength/section ratio is obtained for patch (12 × 8) cm², whereas for an eccentricity de 2/3L’, the patch (6 × 4) cm² gives a better resistance. The results highlight the influence of the composite on the ultimate load. The force-displacement relations are little modified in the elastic phase. The experimental results have been compared with the theoretical models showing a good correlation.

The strength and ductility are depended on the dimensions of the patch and the eccentricity of the load. The use of a patch to cover the most stressed area, in the event of an eccentric axial load is a very economical solution compared to the total reinforcement. The damage field shows that the evolution of cracks depends on dimensions and the position of the patch. Indeed, the eccentricity of the vertical load induces an additional bending moment that will influence the fracture surface. The rupture load and ultimate displacement increase with the surface of the patch.

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.

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.

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.

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.

The purpose of this study concerns numerical studies and experimental validation of the mechanical behavior of hybrid specimens. These kinds of composite specimens are made up of thin carbon and glass substrates on which some Macro Fiber Composite^® (MFC) piezoelectric patches are glued. A proper design and manufacturing of the hybrid specimens as well as testing activities have been performed. The research activity has been carried out under the FutureWings project, funded by the European Commission within the 7th Framework.

The paper describes the basic assumptions made to define specimen geometries and to carry out experimental tests. Finite element (FE) results and experimental data (laser technique measurements) have been compared: it shows very good agreement for the displacements’ distribution along the specimens.

Within the objectives of the project, the study of passive and active deformation characteristics of the hybrid composite material has provided reference technical data and has allowed for the correct adaptation of the FE models. More in particular, using the hybrid specimens, both the bending deformations and the torsion deformations have been studied.

The deformation capability of the hybrid specimens will be used in the development of prototypical three-dimensional structures, that, through the electrical control of the MFC patches, will be able to change the curvature of their cross section or will be able to change the angle of torsion along their longitudinal axis.

The design of nonstandard specimens and the tests executed represent a novelty in the field of structures using piezoelectric actuators. The numerical and experimental data of the present research constitute a small step forward in the field of smart materials technology.

Among various efforts pursued to produce fuel efficient vehicles, light weight engineering (i.e. the use of low‐density structurally‐efficient materials, the application of advanced manufacturing and joining technologies and the design of highly‐integrated, multi‐functional components/sub‐assemblies) plays a prominent role. In the present work, a multi‐disciplinary design optimization methodology has been presented and subsequently applied to the development of a light composite vehicle door (more specifically, to an inner door panel). The door design has been optimized with respect to its weight while meeting the requirements /constraints pertaining to the structural and NVH performances, crashworthiness, durability and manufacturability. In the optimization procedure, the number and orientation of the composite plies, the local laminate thickness and the shape of different door panel segments (each characterized by a given composite‐lay‐up architecture and uniform ply thicknesses) are used as design variables. The methodology developed in the present work is subsequently used to carry out weight optimization of the front door on Ford Taurus, model year 2001. The emphasis in the present work is placed on highlighting the scientific and engineering issues accompanying multidisciplinary design optimization and less on the outcome of the optimization analysis and the computational resources/architecture needed to support such activity.