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.     

Development of life extension strategies for Australian military aircraft, using structural health monitoring of composite repairs and joints

The DSTO Centre of Expertise for Structural Mechanics (COE-SM) has recently developed methodologies for simulating structural health monitoring (SHM) systems for adhesively bonded composite repairs to Australian military aircraft. System design, interrogation strategy, and sensor placement are discussed, with particular emphasis on the development of techniques for embedding optical fibre sensors for optimal SHM system response.     

The effects of size and stacking sequence of composite laminated patch on bonded repair for cracked hole

The performance of a bonded repair for cracked holes has been studied using the three dimensional finite element method, linear elastic fracture mechanics and strain energy density theory. Increasing the composite patch size reduces the strain energy level at the crack tip; increasing the patch length normal to the crack is a better choice. The stacking sequences of the laminated patch have little influence on the strain energy distribution in the vicinity of the crack. To repair the cracked holes of aircraft components subjected to variable direction loading during flight, the orientations of the patch ply, 90° and ±45° with respect to the crack direction, are the optimum selection in bonded repairs.     

Damage Assessment of Composite Structures Using Digital Image Correlation

The steady increase of Carbon-Fiber Reinforced Polymer (CFRP) Structures in modern aircraft will reach a new dimension with the entry into service of the Boeing 787 and Airbus 350. Replacement of damaged parts will not be a preferable solution due to the high level of integration and the large size of the components involved. Consequently the need to develop repair techniques and processes for composite components is readily apparent. Bonded patch repair technologies provide an alternative to mechanically fastened repairs with significantly higher performance, especially for relatively thin skins. Carefully designed adhesively bonded patches can lead to cost effective and highly efficient repairs in comparison with conventional riveted patch repairs that cut fibers and introduce highly strained regions. In this work, the assessment of the damage process taking place in notched (open-hole) specimens under uniaxial tensile loading was studied. Two-dimensional (2D) and three-dimensional (3D) Digital Image Correlation (DIC) techniques were employed to obtain full-field surface strain measurements in carbon-fiber/epoxy T700/M21 composite plates with different stacking sequences in the presence of an open circular hole. Penetrant enhanced X-ray radiographs were taken to identify damage location and extent after loading around the hole. DIC strain fields were compared to finite element predictions. In addition, DIC techniques were used to characterise damage and performance of adhesively bonded patch repairs in composite panels under tensile loading. This part of work relates to strength/stiffness restoration of damaged composite aircraft that becomes more important as composites are used more extensively in the construction of modern jet airliners. The behaviour of bonded patches under loading was monitored using DIC full-field strain measurements. Location and extent of damage identified by X-ray radiography correlates well with DIC strain results giving confidence to the technique for structural health monitoring of bonded patches.     

Composite repairs to rib stiffened panels

There are several methods available for obtaining a first estimate for the design of composite repairs to cracks in thin metallic skins. One approach uses analytical formulae, which build on an analogy with the problem of a bonded symmetric lap joint (Rose LRF. An application of the inclusion analogy for bonded reinforcements. Int J Solid Struct 1981;17:827–838; Fredell RS. Damage tolerant repair techniques for pressurized aircraft fuselages. PhD Thesis, Delft University of Technology, June 1994; Baker AA, Jones R. Bonded repair of aircraft structures. Dordrecht: Martinus Nijhoff, 1988). The present paper extends this methodology to include the bonded repair to rib stiffened wing skins. To this end the present paper presents the results of a series of 3-D finite element analysis of various composite repair configurations. This study reveals that, after repair, the stress intensity factors exhibits an asymptotic behaviour as the crack length increases and that this stress intensity factor is a weak function of the stiffener spacing.     

Fatigue crack growth behavior of cracked aluminum plate repaired with composite patch

In this study, we investigated the fatigue crack growth behavior of cracked aluminum plate repaired with bonded composite patch especially in thick plate. Adhesively bonded composite patch repair technique has been successfully applied to military aircraft repair and expanded its application to commercial aircraft industry recently. Also this technique has been expanded its application to the repair of load bearing primary structure from secondary structure repair. Therefore, a through understanding of crack growth behavior of thick panel repaired with bonded composite patch is needed. We investigated the fatigue crack growth behavior of thick panel repaired with bonded composite patch using the stress intensity factor range (ΔK) and fatigue crack growth rate (da/dN). The stress intensity factor of patched crack was determined from experimental result by comparing the crack growth behavior of specimens with and without repair. Also, by considering the three-dimensional (3D) stress state of patch crack, 3D finite element analyses were performed to obtain the stress intensity factor of crack repaired by bonded composite patch. Two types of crack front modeling, i.e. uniform crack front model and skew crack front model, were used. The stress intensity factor calculated using FEM was compared with the experimentally determined values.     

Numerical analysis of the patch shape effects on the performances of bonded composite repair in aircraft structures

In this study, the three-dimensional finite element method is used to analyze the effects of the patch shape on the efficiency end the durability of bonded composite repairs of aircraft structures. The stress intensity factor at the crack tip is used as fracture criteria. The determination of this factor allows us to estimate the repair efficiency. The analysis of the stresses distribution in the adhesive layer allows us to estimate the durability of the adhesion between the damaged plate and the composite patch. The obtained results show that the repair performances are closely related to the patch shape. It was demonstrated that the rectangular shape of the patch could be improved using an “H” shape of the patch. This last shape could also be improved using an arrow shape.     

Mechanical properties of steel fibre reinforced and rubberised cement-based mortars

Previous studies demonstrated that crack cutting bonded cement-based repairs is highly detrimental to the durability of such applications. Laboratory tests and field experience showed that fibre reinforcement allowing the control of the crack opening and assuring the structural continuity is a solution to enhance the durability of bonded cement-based repairs. In other respect, recent work pointed out that the use of rubber aggregates obtained from grinding end-of-life tyres is a suitable solution to improve the strain capacity of cement-based materials. The present contribution focuses on the synergetic effect of rubber aggregate incorporation and of fibre reinforcement from the point of view of the use of the composite in the repair work application.     

Towards a practical structural health monitoring technology for patched cracks in aircraft structure

Although bonded composite patches often offer a far more effective repair than conventional mechanically fastened patches, full credit can not be given for their effectiveness in reducing crack growth when used to repair flight safety structure. This is because of the lack of non-destructive techniques to detect bonds with poor long-term integrity and ability to predict the likelihood and rate of patch disbonding. Structural health monitoring (SHM) may overcome this limitation, since patch health can be monitored on a continuous basis.After discussing some of the requirements and options for structural health monitoring of bonded repair patches, a case history is presented on the application of a simple strain-based SHM approach for monitoring the boron/epoxy patch repair of a critical fatigue crack in an F-111C wing. The effectiveness of the strain-based SHM approach is demonstrated, and improvements which would reduce its limitations and raise its practicality to the level (TRL8) where it could be used for in- flight application are discussed.Conventional strain gauges were used in the SHM system which, although found effective, have several limitations. A study on the option of using Bragg grating optical fibres as strain sensors is briefly described. Finally brief details are provided on an alternative longer-term approach and study to use acousto-ultrasonics as the basis of the SHM system.     

Comparison of the effectiveness of boron/epoxy and graphite/epoxy patches for repaired cracks emanating from a semicircular notch edge

The adhesively bonded composite patch repair technique has been used to restore or extend the service life of the cracked aluminium structural components because of its efficiency. In this study, the finite element method is used to analyse the performance of the different bonded composite patches at a semicircular lateral notch and the repair of cracks emanating from this kind of notch. The knowledge of the stress distribution in the neighbourhood of the cracks is important for the analysis of their repair according to the geometry of the patch. The effects of the mechanical and geometrical properties on the variation of the stress intensity factor in the crack tip were highlighted. The effects of the adhesive properties and of the patch size on the stress intensity factor variation at the crack tip in mode I were also highlighted. The comparison between the double and single patch repairs is also given in this study. The results obtained show that the stress intensity factor of the crack tip repaired by two composite patches, is reduced to a half compared to the one that is repaired only by one patch. The orientation of fibres possessing a higher rigidity perpendicularly to the crack propagation considerably influences the reduction of the stress intensity factor. The adhesive properties must be optimised in order to increase the performance of the patch repair or the reinforcement.     

Compressive behaviour of CFRP laminates repaired with adhesively bonded external patches

Composite patches can be used to reinforce and repair both cracked composite and metallic aircraft structures. The repair of a composite structure with a composite patch may use mechanical fastening, which often introduces undesirable stress concentrations or adhesive bonding, external or flush patches. To ensure a reliable and durable bond, various parameters such as the quality of surface preparation and the design of the composite patch (size, shape, stiffness) are very important. This paper describes the testing of bonded external patch repaired CFRP laminates loaded in compression. It is found that the critical failure mechanism is fibre microbuckling in the 0° plies accompanied by matrix cracking and delamination, triggered by failures at the adhesive/adherend interface. A three-dimensional finite element analysis is performed to estimate the stress field in the repaired region. The calculated stresses are then used with the maximum stress and average stress failure criteria to predict damage initiation, mode and location. Carefully designed external patch repairs can recover more than 80% of the undamaged compressive strength.     

Design of Emergency Aircraft Repairs Using Composite Patches

The increasing demand for fatigue life extension of both military and civilian aircraft has led to advances in repair technology for cracked metallic structures. Conventional structural repairs may significantly degrade the aircraft fatigue life and lower its aerodynamic performance. Adhesively bonded composite reinforcement is a new technology of great importance due to the remarkable advantages obtained, such as mechanical efficiency and repair time and cost reduction. In this article, bonded composite patch repairs were designed for quick application to aircraft under emergency conditions, such as aircraft battle damage repair (ABDR). A formulated method was developed, to be applied when damage has to be restored quickly, without restrictions to safety of flight. Different damage cases were investigated using finite-element analysis (FEA), taking into account specific parameters of the structure under repair. Based on the FEA results, a quick design procedure using composite patch repairs for the most frequent damage cases is proposed.     

Repair of floating offshore units using bonded fibre composite materials

On ships, tankers and similar vessels structural defects such as cracks and corrosion damage are typically repaired by welding. However, welding is unwanted hotwork on floating offshore units (FOUs) such as floating, production, storage and offloading (FPSO) and floating, storage and offloading (FSO) vessels because it requires shutdown of parts of the vessel thus resulting in expensive production delays. Bonded fibre composite material patch repairs can be used as an alternative to overcome the hazards of hotwork associated with welding. The patches are bonded over the defect and the integrity of the original structure is hence restored. The patch repair technology can also be utilised to provide upgrades, such as life extensions and higher design requirements.A recommended practice (RP) has been developed that describes requirements for patch repairs used in FOUs. To provide flexibility and to fit different repair needs, the RP defines a range of Repair Classes that can be used depending on the urgency of the repair and the need for optimisation. The qualification effort increases with the degree of optimisation. The requirements are based on an extensive test program carried out in the laboratory (reported elsewhere) as well as experience from full-scale repair demonstrators.Two full-scale repair demonstrators were carried out on actual FOUs to demonstrate the feasibility of the RP for bonded composite patch repair. In addition the demonstrators also showed the viability of using bonded composite repairs under harsh conditions encountered in oil and gas exploration and production environments. The first repair was carried out to arrest a fatigue crack that had developed from the corner of a door, while the second repair was carried out to restore material loss on a heavily pitted deck floor. Both repair cases are used as examples to demonstrate the proposed qualification procedures whilst at the same time discussing the challenges and potential applications of this patch repair technology for FOUs in the oil and gas industry.     

Phased‐array transducers for damage detection in aircraft structures*

The work outlined in this paper is a part of a wider effort aimed at the development of a built‐in system for real‐time damage detection in aircraft structures. The objective of this study was to show the feasibility of applying low‐frequency phased‐array transducers for damage detection in metallic structures with bonded repairs and, also, in structures made of composite materials. This paper is divided into two main portions: a component dedicated to the monitoring of crack growth under a simulated composite repair and a part describing an introductory study done for composite panels monitoring. The data acquisition system is also described. The results of the study described in this paper indicate that a low‐frequency, built‐in phased‐array system can successfully estimate the length of a crack growing under a bonded repair. Satisfactory results of monitoring simulated defects in composite structures are presented, as well.     

Bonded aircraft repairs under variable amplitude fatigue loading and at low temperatures

Bonded repairs can replace mechanically fastened repairs for aircraft structures. Compared to mechanical fastening, adhesive bonding provides a more uniform and efficient load transfer into the patch, and can reduce the risk of high stress concentrations caused by additional fastener holes necessary for riveted repairs. Previous fatigue tests on bonded Glare (glass‐reinforced aluminium laminate) repairs were performed at room temperature and under constant amplitude fatigue loading. However, the realistic operating temperature of ?40 °C may degrade the material and will cause unfavourable thermal stresses. Bonded repair specimens were tested at ?40 °C and other specimens were tested at room temperature after subjecting them to temperature cycles. Also, tests were performed with a realistic C‐5A Galaxy fuselage fatigue spectrum at room temperature. The behaviour of Glare repair patches was compared with boron/epoxy ones with equal extensional stiffness. The thermal cycles before fatigue cycling did not degrade the repair. A constant temperature of ?40 °C during the mechanical fatigue load had a favourable effect on the fatigue crack growth rate. Glare repair patches showed lower crack growth rates than boron/epoxy repairs. Finite element analyses revealed that the higher crack growth rates for boron/epoxy repairs are caused by the higher thermal stresses induced by the curing of the adhesive. The fatigue crack growth rate under spectrum loading could be accurately predicted with stress intensity factors calculated by finite element modelling and cycle‐by‐cycle integration that neglected interaction effects of the different stress amplitudes, which is possible because stress intensities at the crack tip under the repair patch remain small. For an accurate prediction it was necessary to use an effective stress intensity factor that is a function of the stress ratio at the crack tip R_{crack tip} including the thermal stress under the bonded patch.     

Effect of initial damage level and patch configuration on the fatigue behaviour of reinforced steel plates

The application of carbon fibre reinforced polymer composites externally bonded on cracked steel plates is an effective system in extending the fatigue life of these structural elements. In particular, composite patches bonded on the crack tip region reduce the stress concentration and the crack opening displacement, leading to an extension of the fatigue life. In order to additionally show the effectiveness of this kind of reinforcing technique, experimental tests were performed at the laboratories of the Politecnico di Milano. Fatigue tests were executed on single edge notched tension specimens reinforced by pultruded strips bonded to a single side (non‐symmetric reinforcement). Different patch configurations (reinforcement stiffness and patch location) and initial damage levels were considered as parameters influencing the repair effectiveness in extending the fatigue life. The results showed that the use of carbon fibre reinforced polymer materials bonded around the tip region allows extending the fatigue life for different amount of initial damage level. Finally, this work provides some useful information for the more efficient repair configuration.     

Fracture toughness analysis of inclined crack in cylindrical shell repaired with bonded composite patch

Adhesively bonded composite patch repair has been widely used to restore or extend the service life of cracked structural components due to its efficiency and cost-effectiveness compared to mechanical repair technique. Current available knowledge on patch repair mainly focus on flat damaged structures and the corresponding analysis methods and empirical databases are computationally efficient. In contrast, only limited work has contributed to studying patch repair to curved damaged structures. Authors have developed an adhesive element in conjunction with a shell element to investigate the effect of curvature on the adhesive stresses and mode I fracture toughness of the cracked host shell in the curved repairs. In this paper, this technology is again employed to model an adhesively bonded composite patch repair to a cylindrical shell embedded with an inclined through-thickness crack. The total strain energy release rate (SERR), calculated by the modified virtual crack closure technique (VCCT), is used to evaluate the mix-mode fracture toughness of the damaged structure and further to estimate the efficiency of patch repair. An automatic mesh generation scheme is proposed to conduct a quick parametric analysis, which can also be used to structural optimization design of composite patch repair. The numerical results are presented to show the effect of curvature and inclined angle of the through-thickness crack on fracture toughness of the repaired structure subject to different loads.     

Comparison between rectangular and trapezoidal bonded composite repairs in aircraft structures: A numerical analysis

The optimization of the patch shape of bonded composite repair in aircraft structures is a good way to improve the repair performance. In this study, the three-dimensional finite element method is used to compare the repair performance of patches with rectangular and trapezoidal shapes in aircraft structures. The comparison is done by analysing the stress intensity factor (SIF) at the tip of repaired crack and the distribution of the adhesive stresses for the two patch shapes. The obtained results show that, when the crack length is ranged from 5 to 20 mm, the trapezoidal shape presents lower stress intensity factor at the crack tip, which is beneficial for the fatigue life and lower adhesives stresses, which is beneficial for the repair durability. These advantages disappear when the crack length reaches the value of 40 mm. It is also shown that the use of the trapezoidal shape reduce the mass of the patch, which can reduce the repair cost.     

Damage monitoring and analysis of composite laminates with an open hole and adhesively bonded repairs using digital image correlation

High performance composite materials, such as Carbon–Fibre Reinforced Plastic (CFRP) composites, are being increasingly used in aerospace industry, such as fuselage primary structures in Boeing 787 or Airbus 350, where high strength and stiffness are required at minimum weight [1]. The design of composite structures frequently includes discontinuities such as cut-outs for access and fastener holes for joining and they become critical regions under thermo-mechanical loading. Understanding of notched specimen behaviour is necessary for the design of complex structures where parts are mostly connected with bolts and rivets [2]. The effect of these discontinuities on the behaviour of composite materials is an important topic because it causes a relatively large reduction in strength compared to the unnotched laminate [3]. In the first part of the current work, the assessment of the damage process taking place in notched (open-hole) specimens under uniaxial tensile loading was studied. Two-dimensional (2D) and three-dimensional (3D) Digital Image Correlation (DIC) techniques were employed to obtain full-field surface strain measurements in carbon–fibre/epoxy M21/T700 composite plates with different stacking sequences in the presence of an open circular hole. Penetrant enhanced X-ray radiographs were taken to identify damage location and extent after loading around the hole. DIC strain fields were compared to numerical predictions. In the second part of the study, DIC techniques were used to characterise damage and performance of adhesively bonded patch repairs in composite panels under tensile loading. This part of work relates to strength/stiffness restoration of damaged composite aircraft that becomes more important as composites are used more extensively in the construction of modern jet airliners. In the current work, external bonded patches have been studied. Adhesively bonded repairs are the most common type of repair carried out with composite materials [1], [4]. The behaviour of bonded patches under loading was monitored using DIC full-field strain measurements. Location and extent of damage identified by X-ray radiography correlates well with DIC strain results giving confidence to the technique for structural health monitoring of bonded patches.