Search results

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.

The purpose of this paper is twofold: first, to observe an influence of different Composite Fibre-Reinforced Polymer (CFRP) patches, whose application to metals is very easy, in suppling and significantly elongating the service time; and second, the numerical calculation of the reduced stress intensity factor (SIF) range for strengthened cracked steel specimens.

One of the successful strengthening methods is the CFRP patching along the fatigue crack paths. The presented approach has been studied and discussed in this paper on the background of the numerical and experimental data. As it was expected, the proposed strengthening method is efficient and promising in case of the “immediate” repairs of critical members with cracks. The manufacturing process of specimens and test methodology as well as numerical approach to calculate SIFs for various reinforcements of steel specimens are presented. For this purpose, the Extended Finite Element Method was involved and described.

The main mechanism of fatigue crack growth retardation is associated with local ΔK reduction due to CFRP patches; any type of reinforcement results in an increase in a_f and a significant decrease in SIF values. The beach-marking method is described as a good, reliable and comprehensive method to capture the crack propagation in structures consisting of various materials and could be applied successfully for mixed mode testing.

A detailed experimental-numerical approach for fatigue crack growth in long-term operated structures made of steel is presented. The strengthening methodology is presented with consideration of the various CFRP patches configurations.

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.

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.

Bonded composite patches are ideal for aircraft structural repair as they offer enhanced specific properties, case‐tailored performance and excellent corrosion resistance. Bonding minimizes induced stress concentrations unlike mechanical fastening, whilst it seals the interface between the substrate and the patch and reduces the risk of fretting fatigue that could occur in the contact zone. The purpose of this paper is to assess the electrochemical corrosion performance and the environmentally induced mechanical degradation of aerospace epoxy adhesives when carbon nanotubes (CNTs) are used as an additive to the neat epoxy adhesive.

The galvanic effect between aluminium substrates and either plain or CNT enhanced carbon fibre composites, was measured using a standard galvanic cell. Also, rest potential measurements and cyclic polarizations were carried out for each of the studied systems. The effect of the CNT introduction to a carbon fiber reinforced plastic (CFRP) on the adhesion efficiency, before and after salt‐spraying for 10, 20 and 30 days, was studied. The adhesion efficiency was evaluated by the single lap joint test.

The corrosion behaviour of the system is polymer matrix type dependent. CNT introduction to a CFRP may induce small scale localized degradation.

This paper fulfills an identified need to study how the shear strength and the response to galvanic corrosion are affected by epoxy resins modified by carbon nanotubes.

The purpose of this study is to focus on the developments of carbon fiber reinforced polymer (CFRP) panels with stepwise graded properties on adhesive layer. The various arranges of the graded properties of the adhesive layer have been checked according to experimental results of the literatures and based on applicability.

The finite element (FE) models and experimental modal tests of the manufactured CFRP sandwich panel specimens have been investigated. The core thickness, core density and orientation of the fiber direction of the sandwich panel face – sheets have been parametrically checked based on modal behavior. Two fully free and fully clamped boundary conditions (BC) have been checked in stepwise graded adhesive zone (SGAZ) cases and first five non-zero natural frequencies (NF) have been compared. Dynamic response of the SGAZ includes modal analysis and transient dynamic loading have been performed numerically with ABAQUS 6.12 well-known FE code.

The first non-zero NF of SGAZ Case 4 was 11.69 per cent higher than homogenous Case 2 and 7.06 per cent lower than Case 1 in fully free boundary conditions. A total of 26.38 per cent is the greatest discrepancy between fist five non-zero NFs of all cases with two BCs (Case 1 vs Case 2 in fully clamped BC). Maximum structural damping behavior and minimum stress picks have been studied during transient dynamic loading analysis of CFRP panel with SGAZ. SGAZ Case 3 (middle adhesive with lower modulus) has increased the maximum structural damping while reducing the minimum out of plain tip displacements during transient dynamic loading by 111.26 per cent in comparison with homogenous Case 2. Also, Case 3 has reduced the Mises stress picks on the adhesive region by 605.68 per cent.

Making a stepwise graded adhesive region (without any added mass) has been shown that it is a novel and useful way to achieve a wide range of stiffness on CFRP panels.

Development of the sandwich panels with various stiffness and damping properties.

The effects of carbon fiber reinforced polymer (CFRP) reinforcement form, adhesive type and pre-crack width on failure mode, shear capacity, deflection response, CFRP strain response and crack patterns of strengthened specimens were investigated.

This paper presents a geopolymer adhesive that matches the performance requirements of CFRP adhesive, which is applied to pre-cracked beams reinforced with CFRP strips.

For specimens with varying structural properties, two failure modes, the CFRP-concrete interface substrate failure and the fracture failure of CFRP, are observed. Moreover, the shear capacity, ultimate deflection and bending stiffness of the U-shaped CFRP-strengthened beams are enhanced in comparison to the complete-wrapping CFRP-strengthened beams. With an increase in pre-crack width, the increase in shear capacity of RC beams shear-strengthened with CFRP strips is less than that of non-cracked beams, resulting in a limited influence on the stiffness of CFRP-strengthened beams. The comparison of experimental results showed that the proposed finite element model (FEM) effectively evaluated the mechanical characteristics of CFRP-strengthened RC beams.

Taking into consideration the reinforcement effect and the concept of environmental protection, the geopolymer adhesive reinforcement scheme is preferable to applying epoxy resin to the CFRP-strengthened RC beams.

Piezoelectric actuators and sensors are an invaluable part of lightweight designs for several reasons. They can either be used in noise cancellation devices as thin‐walled structures are prone to acoustic emissions, or in shape control approaches to suppress unwanted vibrations. Also in Lamb wave based health monitoring systems piezoelectric patches are applied to excite and to receive ultrasonic waves. The purpose of this paper is to develop a higher order finite element with piezoelectric capabilities in order to simulate smart structures efficiently.

In the paper the development of a new fully three‐dimensional piezoelectric hexahedral finite element based on the p‐version of the finite element method (FEM) is presented. Hierarchic Legendre polynomials in combination with an anisotropic ansatz space are utilized to derive an electro‐mechanically coupled element. This results in a reduced numerical effort. The suitability of the proposed element is demonstrated using various static and dynamic test examples.

In the current contribution it is shown that higher order coupled‐field finite elements hold several advantages for smart structure applications. All numerical examples have been found to agree well with previously published results. Furthermore, it is demonstrated that accurate results can be obtained with far fewer degrees of freedom compared to conventional low order finite element approaches. Thus, the proposed finite element can lead to a significant reduction in the overall numerical costs.

To the best of the author's knowledge, no piezoelectric finite element based on the hierarchical‐finite‐element‐method has yet been published in the literature. Thus, the proposed finite element is a step towards a holistic numerical treatment of structural health monitoring (SHM) related problems using p‐version finite elements.

The paper aims to present a design and numerical verification procedure of a composite casing of a microstrip antenna for an aerospace satellite.

The casing for the microstrip antenna was designed in a form of a laminate shell with variable number of layers of reinforcing fabric. The material properties, both static and dynamic, were determined experimentally and then exported to an environment of numerical analyses. The numerical modal analysis allows optimizing the geometry and lay-up of the casing in such a way that a number of modal shapes occurring in the operational frequency band was significantly reduced, several modal shapes with high displacement in flanges of the casing were eliminated and the values of natural frequencies were increased. A final model of the composite casing was subjected to two types of analyses which simulate typical operation conditions during spacecraft mission. These analyses contained thermomechanical quasi-static analyses with 12 loadcases and thermomechanical shock analyses with 9 loadcases, which simulate various mechanical and temperature conditions.

Results of the performed analyses were compared with safety margins determined by following requirements to spacecraft vehicles. The obtained results confirm the design feasibility, which allow considering the proposed design during manufacturing of a prototype in further studies.

Moreover, the presented results can be considered as a design methodology guideline, which can be helpful for engineers working in the aerospace industry.

The originality of the paper lies in the proposed design and verification procedure of composite elements subjected to operational loading during a spacecraft mission.