Crack Patching Research Articles

Crack growth induced delamination on steel members reinforced by prestressed composite patch

ABSTRACT Prestressed composite patch bonded on cracked steel section is a promising technique to reinforce cracked details or to prevent fatigue cracking on steel structural elements. It introduces compressive stresses that produce a crack closure effect. Moreover, it modifies the crack geometry by bridging the crack faces and so reduces the stress intensity range at the crack tip. Fatigue tests were performed on notched steel plate reinforced by CFRP strips as a step toward the validation of crack patching for fatigue life extension of riveted steel bridges. A crack growth induced debonded region in the adhesive‐plate interface was observed using an optical technique. Moreover, the size of the debonded region significantly influences the efficiency of the crack repair. Debond crack total strain energy release rate is computed by the modified virtual crack closure technique (MVCCT). A parametric analysis is performed to investigate the influence of some design parameters such as the composite patch Young's modulus, the adhesive thickness and the pretension level on the adhesive‐plate interface debond.

Delamination effects on cracked steel members reinforced by prestressed composite patch

Prestressed composite patch bonded on cracked steel section is a promising technique to reinforce cracked details or to prevent fatigue cracking on steel structural elements. It introduces compressive stresses that produce crack closure effect. Moreover, it modifies the crack geometry by bridging the crack lips and reduces the stress range at crack tip. Fatigue tests were performed on notched steel plate reinforced by CFRP strips as a step toward the validation of crack patching for fatigue life extension of riveted steel bridges. A debond crack in the adhesive–plate interface was observed by optical technique. Debond crack total strain energy release rate is computed by the modified virtual crack closure technique. A parametric analysis is performed in order to investigate the influence of some design parameters such as the composite patch Young’s modulus, the adhesive thickness and the pretension level on the adhesive–plate interface debond.

Dual contouring of hermite data

This paper describes a new method for contouring a signed grid whose edges are tagged by Hermite data (i.e; exact intersection points and normals). This method avoids the need to explicitly identify and process "features" as required in previous Hermite contouring methods. Using a new, numerically stable representation for quadratic error functions, we develop an octree-based method for simplifying contours produced by this method. We next extend our contouring method to these simpli£ed octrees. This new method imposes no constraints on the octree (such as being a restricted octree) and requires no "crack patching". We conclude with a simple test for preserving the topology of the contour during simplification.

Characterization of fracture behavior in repaired skin/stiffener structure with an inclined central crack

Finite element analysis for the stress intensity factor (SIF) at the skin/stiffener structure with inclined central crack repaired by composite stiffened panels is developed. A numerical investigation was conducted to characterize the fracture behavior and crack growth behavior at the inclined crack. In order to investigate the crack growth direction, maximum tangential stress (MTS) criterion are used. Also, this paper is to study the performance of the effective bonded composite patch repair of a plate containing an inclined central through-crack. The main objective of this research is the validation of the inclined crack patching design. In this paper, the reduction of stress intensity factors at the crack-tip and prediction of crack growth direction are determined to evaluate the effects of various non-dimensional design parameter, including; composite patch thickness and stiffener distance. We report, the results of finite element analysis on the stiffener locations and crack slant angles and discuss them in this paper. The research on cracked structure subjected to mixed mode loading is accomplished and concludes that more work using a different approaches is necessary. The authors hope the present study will aid those who are responsible for the repair of damaged aircraft structures and also provide general repair guidelines.

Stress analysis of a bonded repair over a corrosion grind-out using an inclusion model with a second ordered eigenstrain theory

The problem of an isotropic sheet with an internal elliptical grind-out repaired with an octagonal bonded patch is analyzed, following Rose's two-stage analysis procedure [Theoretical analysis of crack patching, in: A.A. Baker, R. Jones (Eds.), Bonded Repair of Aircraft Structures, Kluwer Academic Publishers, Dordrecht, 1988] for a crack patching problem. For simplicity, the bending effect is ignored in the present analysis. In the first stage, an infinite sheet reinforced by an octagonal patch under a prescribed far-field stress and a uniform thermal field is analyzed using an inclusion model with a second ordered eigenstrain theory, without considering a grind-out cavity. The stresses ( σ ij *) inside the patched area at a center of the prospective grind-out location are calculated and later used as far field boundary conditions for a second (stage II) problem. In the second stage, the patch is assumed to be infinite and an integral part of a sheet. Stage II analysis then involves solving a problem of an infinite patched sheet containing an embedded elliptical grind-out cavity under far field stresses and initial strain field. The latter problem is also solved using an inclusion model with a second ordered eigenstrain. This approach had been used previously by the present authors very successfully in solving for a similar repair problem but with an elliptical patch and for a grind-out with a uniform depth [Analytical approach to bonded repair of elliptical dent, corrosion grind-out and cut-out, Theoretical and Applied Fracture Mechanics, 2000, submitted] in a relatively thin sheet. However, our recent 3-D finite element study showed that the use of such inclusion analogy in a stage II stage analysis would not yield sufficiently accurate results for a repair problem of a very deep grind-out cavity in a thick sheet. The original approach given in (Duong et al., loc. cit.) therefore has been modified in this paper. The present paper also deals in particular with additional complexities such as octagonal shaped patch, a grind-out cavity with a spherical depth, and effects of thermal stresses due to curing and/or cruising.

Design guidelines for composite patches bonded to cracked aluminum substrates

Design guidelines are proposed for boron fiber-reinforced epoxy composite patches bonded to cracked aluminum substrates. Stress analysis was performed on the composite patch bonded to an uncracked aluminum substrate. The load transfer via the adhesive and the stress levels in the aluminum substrate and the patch were determined. The local stress and stress distribution in the repaired V-shape side cracked substrates were obtained. The effects of crack length and patch geometries including the length and number of plies on the failure behavior of the repair were identified. The results were validated by static tensile tests. It was found that the theoretical calculations were in agreement with the experimental results. This indicates that the proposed analysis can be used as design guidelines for composite repairs of cracked structures. These guidelines coupled with a simple surface preparation technique developed in this work might aid composite patch installation in service.

Fracture mechanics analysis on the bonded repair of a skin/stiffener with an inclined central crack

The enhancement of the service life of damaged or cracked structures is currently a major issue for researchers and engineers. An analysis for the stress intensity factor (SIF) at a skin/stiffener structure with an inclined central crack repaired by composite stiffened panels is developed. A numerical investigation was conducted to characterize the fracture behavior and crack growth behavior at an inclined crack. In order to investigate the crack growth direction, maximum tangential stress (MTS) criterion are used. Also, this paper studies the performance of an effective bonded composite patch repair of a plate containing an inclined central through-crack. The main objective of this research is the validation of an inclined crack patching design. In this paper, the reduction of SIFs at the crack-tip and prediction of crack growth direction are determined in order to evaluate the effects of various non-dimensional design parameters including: composite patch thickness and stiffener distance. The results of finite element analysis on the stiffener locations and crack slant angles are discussed herein. Research on cracked structure subjected to mixed-mode loading is reported and concludes that more work using a different approach is necessary. The authors hope the present study will aid those who are responsible for the repair of damaged aircraft structures and also provide general repair guidelines.

Analytical approach to bonded repair of elliptical dent, corrosion grind-out, and cut-out

The problem of an isotropic sheet with an internal elliptical dent, corrosion grind-out, or cut-out repaired with an elliptical bonded patch is analyzed, following the similar two-stage analysis originally proposed by Rose [International Journal of Solid and Structures 17 (1981) 827; in: Bakers, Jones (Eds.), Bonded Repair of Aircraft Structures, Martinus Nijhoff, Dordrecht, 1988, p. 77] for a crack patching problem. For simplicity, the bending effect is ignored in the analysis. In the first stage, an infinite sheet reinforced by an elliptical patch under a prescribed far-field stress is analyzed using the inclusion analogy, without considering the dent, grind-out or cut-out. The constant stresses inside the patched area ( σ 0 ij ) are then calculated and later used as the far-field boundary conditions for the second (stage II) problem. In the second stage, the patch is assumed to be infinite and an integral part of the sheet. Stage II analysis then involves solving a problem of an infinite patched sheet containing a circular dent, grind-out, or cut-out under far-field stresses σ 0 ij . The latter problem is also solved using the inclusion analogy. Because the patch in a typical design is much larger than the damage area, the solutions of the first and second problems are approximately the same as the solutions of the original problem inside and outside the patched area, respectively.

Two- and three-dimensional geometrical nonlinear finite elements for analysis of adhesive joints

Special two- and three-dimensional adhesive elements have been developed for stress and displacement analyses in adhesively bonded joints. Both the 2-D and 3-D elements are used to model the whole adhesive system: adherends and adhesive layer. In the 2-D elements, adherends are represented by Bernoulli beam elements with axial deformation and the adhesive layer by plane stress or plane strain elements. The nodes of the plane stress–strain elements that lie in the adherend–adhesive interface are rigidly linked with the nodes of the beam elements, resulting in the offset nodes which coincide with the midplanes of the adherends. The 3-D elements consist of shell elements that represent the adherends and solid brick elements to model the adhesive. This technique results in smaller models with faster convergence than conventional 3-D finite element models. The resulting mesh can represent arbitrary beam- or plate-like geometries, which are a large part of adhesive joint designs. This model can include debonds as well as cracks within the adhesive, therefore it can be used for durability analysis of adhesive joints. Since large displacements are often observed in adhesively bonded joints, geometric nonlinearity is modeled. 2-D and 3-D stress analyses of single lap joints are presented. Important 3-D effects can be appreciated. A stress analysis of a crack patch geometry is presented.

Stress concentrations around bonded repairs

Ageing aircraft fuselages require safe and damage‐tolerant repair techniques. Bonded fibre–metal laminates (FML) combine a small mismatch in coefficients of thermal expansion (CTE) with the cracked skin, excellent fatigue properties and high strength, and are therefore promising materials for bonded repairs. This paper focuses on local stress concentrations in the fuselage skin at the edge of a bonded FML repair, and presents analytical, FEM and experimental results of single and multiple repairs in close proximity. Repairs on both flat uniaxially loaded skins and curved biaxially loaded skins were considered. The main objective of this research is the validation of the crack patching design program CalcuRep. The model on which CalcuRep is based, the ‘Rose model’, was extended and used for the analytical prediction of skin stresses in the specimens. It appeared that, in order to prevent fatigue critical stresses in the skin next to the patch, one should carefully choose the patch material and geometry. Furthermore, in a specific repair configuration with multiple patches, a minimum separation is required to prevent high stress concentrations in the skin between the patches. It was confirmed that, in most cases, a biaxial stress field has a favourable effect on the stress in the skin along the patch edge, compared with the uniaxial situation. With both the FEM and the extended Rose model, the stresses in the skin next to the patch could be predicted with sufficient accuracy. It is the intention to implement the extended Rose model in a new version of CalcuRep.

An Integrated Bonded Repair System: A Reliable Means of Giving New Life to Aging Airframes

The application of bonded composite repairs to restore cracked or corroded metallic airframes has seen rapid expansion recently, following twenty years of technology development. However, the lack of standardized bonded repair methods and certification concepts impedes wider application of the technique. Existing bonded repair technologies are maturing rapidly, but several specialities (design and analysis, surface preparation, installation, inspection, training and certification) must be integrated into a seamless package before civil and military airworthiness authorities will embrace bonded 'crack patching.' This paper outlines efforts by the United States Air Force to develop an integrated system to allow aircraft maintenance engineers to cover the complete life cycle of a bonded repair. The successful construction of this integrated system will speed the widespread acceptance of the technology and help to assure the continued airworthiness of the aging aircraft fleet.

Experimental Verification of Rose's Constant K Solution in Bonded Crack Patching

Bonded composite patches have been used for two decades to extend the lives of fatigue-damaged F-16, F-111, B-1B, C-141B, and many other aircraft. One of the key features of the technology is extremely slow crack growth under the bonded repair. Researchers have performed hundreds of experiments on repaired cracked panels, and have reported near-constant crack growth rates for a variety of relatively thin sheet (t < 3 mm or 0.125 inch) configurations and constant amplitude load cases. Constant crack growth rates rely on the existence of a constant crack tip cyclic stress intensity factor, Δ K, underneath the patch.

AN ANALYTICAL MODEL FOR STRESS ANALYSIS AND FAILURE PREDICTION OF BONDED JOINTS WITH TAPERED EDGES

An analytical model has been developed for the analysis of bonded composite-to-metal joints subject to axial tensile or compressive load on the substrate. Two joint configuration are considered which simulate bonded composite patch repair such as fatique life enhancement and crack patching. While the metallic susstrate is of uniform thickness, the composite patch has a thickness tapered from the central portion to its edges. The model calculates axial forces/stresses in the substrate and patch, shear stress in the adhesive, and peel stresses on the substrate-adhesive and patch-adhesive interfaces. The residual stresses after curing due to thermal mismatch between the substrate and patch are considered. Based on the stress results obtained and the strength allowables of the joint member materials, the joint failure load is predicted and the failure mode is examined. Eight typical failure modes are considered and the associated safety margins are determined.

Delamination effects in fuselage crack patching : Gut, C. Society for the Advancement of Material and Process Engineering, 1161 Parkview Drive, P.O. Box 2459, Covina, CA 91722, U.S.A. 1996. 903–914 Conference: Materials and Process Challenges: Aging Systems, Affordability, Alternative Applications. Vol. 41-I, Anaheim, CA, U.S.A. 24–28 Mar. 1996

The present paper addresses the fatigue characteristics of two cellular foam core materials as used in load carrying sandwich structures. The fatigue loading studied is a constant shear stress which corresponds to the main type of loading that the core in a sandwich structure exhibits. Based on results from testing stress-life curves are presented for a number of stress ratios (R) and are fitted to a simple two-parameter Weibull function. The influence of the R-value is emphasized and standard Haigh diagrams are constructed. It is seen that the fatigue behaviour of the core materials investigated herein well can be described in manners similar to that of classical metal fatigue. A thorough investigation is performed on the behaviour of fatigue crack formation and growth.

Multilayer hybrid-stress finite element analysis of a patched crack

An assumed hybrid-stress finite element model together with two types of composite multilayer elements (MLTUP and MLTPH) is developed to study the crack patching problem. Three-dimensional interlaminar stresses including transverse normal and shear stresses are accounted for. The assumed stress field satisfies: (1) the equilibrium conditions within each layer, (2) the interlaminar traction reciprocity condition at interlaminar boundaries and (3) traction-free boundary conditions on the top and bottom surfaces of the laminate. Since the stress parameters assumed and the number of nodes taken are independent of the number of layers, the multilayer elements devised are applicable to the laminate with a large number of layers. The singly and doubly patched cracked panels under uniaxial tension are studied. At the same time, with a bending field induced, the strength can be predicted accurately without using the bending correction factors. In addition, the maximum stress intensity factor is found at the unpatched free surface, and an unbalanced composite patch is devised to extend the safe life of the cracked panel. Excellent agreemenets between the present results and referenced solutions show the high accuracy and validity of the present technique.

A finite element model for single-sided crack patchings

A finite element model is established for analyzing the behavior of cracked plates which are repaired with a single-sided patch. The formulation is based on the Reissner-Mindlin plate theory with an assumed variation of the transverse shear and normal stresses through the thickness of the cracked -plate and patch. The generalized stress-strain relations relating the transverse shear stress resultants and the adhesive stresses to the displacements of the plate and patch are established by using a variational principle. By means of the finite element model presented herein, single-sided crack patching problems can be solved with a reasonable estimate of the adhesive stresses and the stress intensity factor. Numerical examples are provided to illustrate the effects of the patch size on the stress intensity factor in the cracked plate and the stress distribution in the adhesive layer, and compared with results from the previous analysis.

Flutter analysis of anisotropic panels with patched cracks

A finite element approach is developed to analyze the panel flutter problems of thin plate-like composite panels with patched cracks. The panel studied is a compound structure that comprises three layers including a thin cracked panel, an adhesive, and a thin patch. A 48-degree-of-f reedom (DOF) triangular crack patching element is derived for numerically analyzing the compound structure. The aerodynamic pressure acting on the panel is estimated using linearized piston theory with aerodynamic damping effect neglected. By a proper tailoring of the materials, very high flutter and/or divergence boundary could usually be obtained for the composite panels. The existence of a crack usually reduces the aeroelastic stability boundary; however, some exceptions were found for the composite panel within certain range of specific filament orientation. The deterioration in flutter/divergence performance due to a crack can, in general, be cured by means of patching, and anisotropic patching is more effective as compared to the isotropic patching provided that the tailoring of the structural parameters is done correctly. I. Introduction A RELIABLE and accurate flutter analysis is of primary importance in the flight vehicle design. The onset of flutter would quickly develop into either catastrophic structural failure or undesirable limit cycle type oscillation. The local panel flutter of the vehicle could be essential for a global flutter to occur. For instance, the initiation of cracks in wing structure is not unusual as they may be produced by either molding, fatigue, or impact damage.1'2 These defects could induce extremely high stresses around the cracktip regions and, thus, precipitate structural failure. When cracks are present, this local structural problem may have significant influence on the local flutter/divergence of the panel structure. The onset of the local aeroelastic instability may cause the failure of the cracked panel and cumulatively affect the global structure as well as the airload distribution and finally result in

Isoparametric shear spring element applied to crack patching and instability: Chu, R. C. and Ko, T. C. Theor. Appl. Fract. Mech. May 1989 11, (2), 93–102

Polyaniline coatings of various controlled thicknesses are applied to 7075-T6 aluminum alloy substrates and evaluated for their corrosion protection ability in a corrosive saline environment using electrochemical techniques. The coating thicknesses are in the range of approximately 5 to 35 μm and they increase linearly with increasing quantities of PANI solution deposited on the substrate. The relationship between the coating thickness, roughness, and hydrophilicity is examined. Coatings ranging from 12 to 23 μm in thickness have the lowest roughness values and highest contact angles with water. The PANI-coated 7075 alloy electrodes with low roughness and high contact angles show better corrosion protection properties in saline solution (3.5% NaCl), as determined by electrochemical impedance spectroscopy (EIS) and voltammetric polarization measurements. Scanning electronic microscopy (SEM) images demonstrate the homogenous surface morphology of the PANI coatings on the micrometer scale. Even after prolonged exposure to corrosive media, no cracks were found.

Isoparametric shear spring element applied to crack patching and instability

In this work the isoparametric shear spring element is applied to the stress and energy analysis of a center-crack panel reinforced by a rectangular patch. In this model, only transverse shears are assumed to prevail in the adhesive layer. The stresses and crack-tip stress intensity factors are obtained for reinforcement on both sides and one side of the panel, and are found to be in agreement with those obtained by previous authors using the triangular shear spring element. Crack stability that tends to vary with patch thickness is determined from the local and global maximum of the minimum strain energy density function denoted, respectively, as [(d W/d V) min max] L at point L and [(d W/d V) min max] G at point G. The distance l between L and G gives the prospective path of subcritical crack growth and its magnitude provides a measure of the degree of crack stability. A patched panel with small l tends to be more stable than that with large l. By increasing the patch thickness beyond a certain value, l can be contained within the patch such that failure, if initiated, will be highly localized. Such a behavior is exhibited. Numerical results are provided to support the foregoing conclusion.

Fibre composite repair of cracked metallic aircraft components — practical and basic aspects

Crack patching, the use of advanced fibre composite patches (such as boron/epoxy or graphite/epoxy) bonded with structural film adhesives to repair cracks in metallic aircraft components, is a significant development in aircraft maintenance technology, offering many advantages over conventional repair procedures based on metallic patches and mechanical fasteners. This paper reviews selected theoretical and experimental aspects of Australian work on this topic and describes a preliminary design approach for estimating the minimum thickness patch that could be employed in a given repair situation. Finally, the paper provides a case study on our repair to the wing skin of Mirage III aircraft. Aspects discussed include evaluation of minimum cure and surface treatment conditions for adhesive bonding in repair situations, potential thermal and residual stress problems, resulting from patching, studies on overlap joints representing repairs and crack propagation behaviour in patched panels.