Fatigue crack growth in hybrid boron/glass/aluminium fibre metal laminates

The crack growth behaviour of hybrid boron/glass/aluminium fibre metal laminates (FMLs) under constant‐amplitude fatigue loading was investigated. The hybrid FMLs consist of Al 2024‐T3 alloy as the metal layers and a mixture of boron fibres and glass fibres as the fibre layers. Two types of boron/glass/aluminium laminates were fabricated and tested. In the first type, the glass fibre/prepreg and the boron fibre/prepreg were used separately in the fibre layers, and in the second type, the boron fibres and the glass fibres were uniformly mingled together to form a hybrid boron fibre/glass fibre prepreg. An analytical model was also proposed to predict the fatigue crack growth behaviour of hybrid boron/glass/aluminium FMLs. The effective stress intensity factor at a crack tip was formulated as a function of the remote stress intensity factor, crack opening stress intensity factor, and the bridging stress intensity factor. The bridging stress acting on the delamination boundary along the crack length was also calculated based on the crack opening relations. Then, the empirical Paris‐type fatigue crack growth law was used for predicting the crack growth rates. A good correlation between the predicted and experimental crack growth rates has been obtained.     

Experimental characterization of the crack-tip-opening angle in fibre metal laminates

Although the crack-tip-opening angle (CTOA) has been shown to be well suited for modelling stable crack growth in monolithic sheet aluminium alloys, its applicability for fibre metal laminates has not been fully analyzed yet. Fracture test were performed on M(T) panels made of Glare 2-3/2-0.4, Glare 3-3/2-0.4 and laminated 2024-T3 3/2-0.4. Different fatigue pre-crack lengths were created to study the effect of bridging fibres on the CTOA measured on the external layer. The effect of bridging fibres resulted in small deviations of the CTOA vs. crack extension curve with respect to the reference panel made of metal laminate. The CTOA criterion could be successfully used for predicting the residual strength in fibre metal laminates.     

Fatigue behaviour of carbon fibre-reinforced aluminium laminates

A new hybrid composite (CARALL), consisting of thin layers of carbon fibre/ epoxy prepreg sandwiched between aluminium sheets, has been developed. It is shown that this class of materials offers higher modulus, higher tensile strength and lower density than 2024-T3 alloy in the longitudinal direction. Under tension-tension fatigue loading, the hybrid laminates showed superior fatigue crack propagation resistance in the longitudinal direction, which may be attributed to the bridging effect imposed by the intact fibres in the crack wake. It has also been shown that the effectiveness of fatigue crack growth reduction increases with the thickness of the carbon fibre/epoxy layer. The resistance to fatigue crack propagation can be further improved by introducing compressive residual stresses in the aluminium layer by postcure stretching the laminate in the plastic region of the aluminium alloy.     

Application of the energy release rate approach for delamination growth in Glare

Fatigue crack growth in a fibre metal laminate such as Glare is accompanied by delamination growth at the interface between the aluminium and glass fibre/adhesive layers. To incorporate this delamination growth in crack growth prediction methods, the energy release rate approach is applied to describe the delamination growth rate. Tests were performed to determine the relationship between the delamination growth rate and the calculated energy release rate.     

Glare Design Aspects and Philosophies

The very nature of Glare is its crack bridging mechanism, which provides superior damage tolerance properties. Depending on the property, Glare shows either monolithic metal or composite behaviour, which challenges the definition of strength justification and certification procedures. Airworthiness regulations have to be interpreted for Glare in order to guarantee the same level of safety as obtained for aircraft structures made of other materials and to take at the same time benefit of its particular properties.Cut-outs are highly fatigue sensitive due to the stress concentrations they cause. In aircraft fuselages these cut-outs are quite large in the case of the windows and doors. The stress level may be increased through the application of Glare in the doubler packages, due to the improved fatigue behaviour compared to conventional aluminium. Glare also presents the possibility of tailoring the material to the load, i.e. fibres aligned with the load, e.g., a 45 degree orientation. FE analysis defined the total doubler package and a test programme was run to confirm the behaviour of the material and to predict the crack behaviour of the Glare door corner.Some aspects of the detailed design of aircraft structures in Glare, the design of splices and riveted joints are discussed. In order to apply Glare in very large fuselage panels, a splice concept was developed, which allows a number of longitudinal splices to be cured in the same curing cycle as the basic material. Through the introduction of this splicing concept, the width of a panel is no longer limited to the maximum width of the aluminium sheet. Internal local reinforcements (doublers) can be integrated into the panel during lay-up. A discussion on the design of riveted joints in Glare is held.     

Response of fibre reinforced aluminium–lithium laminates to different fatigue conditions

Under fatigue conditions fibre reinforced aluminium–lithium laminates do not respond in the same manner as monolithic aluminium alloys. The variation of fatigue crack growth rates with initial loading condition has been examined for both carbon and glass fibre reinforced laminates, and compared with the behaviour of unreinforced 8090 aluminium–lithium alloy for a range of conditions (different initial nominal stress intensity factor range, load range and reversed loading). During fatigue, cracks grow in the metal layers of these laminates whilst the fibres in the crack wake remain intact, bridging the crack faces. The fibre bridging mechanism, inherent in this laminate system, reduces the fatigue crack growth rate. The magnitude of the bridging effect appears to be inversely related to the applied load range. This relationship can account for the behaviour observed in the performed experiments. This revised version was published online in November 2006 with corrections to the Cover Date.     

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.     

Bridging effect on mode I fatigue delamination behavior in composite laminates

This paper discusses the bridging effect of fibres on mode I fatigue delamination growth in unidirectional and multidirectional polymer composite laminates based on a series of double cantilever beam (DCB) tests. From the results, there is sufficient evidence that fibre bridging can decrease the crack growth rate da/dN significantly, and using only one fatigue resistance curve to determine the delamination behavior in composite materials with large-scale fibre bridging may be inadequate. The bridging created in fatigue delamination is different from that of quasi-static delamination at the same crack length. So it is incorrect to use the resistance curve (R-curve) from quasi-static delamination tests to normalize fatigue delamination results.     

The Influence of the Constituent Properties on the Residual Strength of Glare

Aircraft manufacturers like Boeing and Airbus are currently designing new, high capacity aircraft, e.g., A3XX. To make this aircraft cost effective for the next 30 years, a strong impulse is given to the development of new technologies like the application of new aircraft materials. One of these studies investigates the feasibility of using the Fibre Metal Laminate Glare in the damage tolerance critical upper part region of the aircraft in order to reduce weight and increase safety. The present study investigates the crack resistance of Glare in case of foreign object damage as a function of the constituent's characteristics. An experimental program has been performed on Fibre Metal Laminates built up from several combinations of aluminium alloys and fibres. Especially the fracture mechanism was studied. It was found that a larger strain hardening region and a lower yield stress of the aluminium layers had a positive influence on the residual strength due to the capability of transferring high loads away from the cracked area. Increasing the stiffness and lowering the ultimate strain of the fibres reduces the residual strength, since stiffer fibres attract more load and because the final fracture is dominated by fibre failure.     

Fatigue resistance and residual strength of riveted joints in FML

Static and fatigue tests have been carried out on 190-mm-wide fibre metal laminate (FML) riveted lap joints. The specimens were made with a Glare 3-3/2-0.3 material, i.e. a laminate composed of three layers of 2024-T3 aluminium alloy, thickness 0.3 mm and two double layers (0/90) of pre-preg FM-94-27%-S2 glass fibre. Seven configurations were tested which differed in the number of rivet rows (two or three), in the rolling direction of metal layers (perpendicular or parallel to the load direction), in the rivet type (solid or special, such as Hi-Lok or Lock-Bolt), in the rivet material for solid rivets (7050-T73 or 2017-T3), in the rivet diameter (4.8 or 5.5 mm) and in the presence of interlaminar doublers in the overlap area (titanium, aluminium, glass fibre). An additional difference was in the pre-formed rivet head: solid rivets had countersunk head, while Hi-Loks and Lock-Bolts had protruding head. The fatigue tests demonstrated the efficiency of a selective local reinforcement in the overlap area; in some cases, the fatigue resistance was so high that fatigue cracks nucleated in the laminates, rather than in the overlap area, as commonly expected.     

Fatigue crack growth in fibre metal laminates with multiple open holes

The fatigue crack growth behaviour of hybrid S2‐glass reinforced aluminium laminates (Glare) with multiple open holes was investigated experimentally and analytically. It was observed that the presence of multiple‐site fatigue damage would increase crack growth rates in the metal layers as two propagating cracks converged. An analytical crack growth model was established for predicting crack growth rates based on empirical Paris equation. The effective stress intensity factor at crack tips is a function of mode I far‐field stress intensity factor, crack opening stress intensity factor and effective non‐dimensional stress intensity factor that incorporated the crack‐bridging effect in fibre metal laminates. The predicted results under different applied stress can capture the trend of averaged crack growth rates in experiments, although deviation exists in the predictions.     

Comparison of Interlaminar Fracture Toughness in Unidirectional and Woven Roving Marine Composites

This paper investigates the effect of fibre lay-up and matrix toughness on mode I and mode II interlaminar fracture toughness (G_Ic and G_IIc) of marine composites. Unidirectional and woven roving fibres were used as reinforcements. Two vinyl ester resins with different toughness were used as matrices. Results from both modes showed toughness variation that is consistent with matrix toughness. Values of G_Ic were not significantly influenced by fibre lay-up except at peak load points in the woven roving/brittle-matrix composite. Each peak load point, caused by interlocked bridging fibres, signified the onset of unstable crack growth. For unidirectional specimens, crack growth was stable and G_Ic statistically more reliable than woven roving specimens, which gave fewer G_Ic values due to frequent unstable crack growth. Mode II tests revealed that, except for crack initiation, G_IIc was higher in woven roving composites. This was due to fibre bridging, perpendicular to the crack growth direction, which encouraged stable crack growth and increased energy absorption. Mode II R-curves were obtained for the woven roving specimens. These R-curves provide additional information useful for characterising delamination resistance. The paper concludes that composites with woven roving fibres show similar mode I delamination characteristics to the unidirectional composites; but their mode II delamination characteristics, after crack initiation, are quite different.     

Fatigue crack growth in fibre metal laminates under selective variable-amplitude loading

To extend the predictive capability of existing crack growth models for fibre metal laminates under constant amplitude fatigue loading to variable-amplitude loading, further research on variable-amplitude fatigue mechanisms in fibre metal laminates is necessary. In response to this need, an experimental study into the effects of multiple overloads, underloads and various block-loading sequences on crack growth in the fibre metal laminate Glare was investigated. Crack growth retardation effects were observed in the tests; however, the magnitude of these effects was lower than seen in monolithic aluminium because of fibre bridging. As a result, predictions of the observed behaviour were attempted using an existing constant-amplitude fatigue crack growth model for Glare in combination with a linear damage accumulation law.     

Fatigue behaviour and life prediction of fibre reinforced metal laminates under constant and variable amplitude loading

ABSTRACT Fatigue crack growth of fibre reinforced metal laminates (FRMLs) under constant and variable amplitude loading was studied through analysis and experiments. The distribution of the bridging stress along the crackline in centre‐cracked tension (CCT) specimen of FRMLs was modelled numerically, and the main factors affecting the bridging stress were identified. A test method for determining the delamination growth rates in a modified double cracked lap shear (DCLS) specimen was presented. Two models, one being fatigue‐mechanism‐based and the other phenomenological, were developed for predicting the fatigue life under constant amplitude loading. The fatigue behaviour, including crack growth and delamination growth, of glass fibre reinforced aluminium laminates (GLARE) under constant amplitude loading following a single overload was investigated experimentally, and the mechanisms for the effect of a single overload on the crack growth rates and the delamination growth rates were identified. An equivalent closure model for predicting crack‐growth in FRMLs under variable amplitude loading and spectrum loading was presented. All the models presented in this paper were verified by applying to GLARE under constant amplitude loading and Mini‐transport aircraft wing structures (TWIST) load sequence. The predicted crack growth rates are in good agreement with test results.     

FATIGUE BEHAVIOUR OF FIBRE ALUMINIUM LAMINATES WITH DIFFERENT RESIDUAL STRESSES

In this study two kinds of fibre aluminium laminates (aramid aluminium laminates, ARALL and glass aluminium laminates, GLARE) with different residual stresses in the aluminium layers were prepared. Fatigue crack propagation tests were performed. It is found that the residual stress condition plays an important role in the fatigue behaviour of fibre aluminium laminates. With a decrease of the tensile residual stress in the aluminium layers, the fatigue crack growth rate of the laminates is greatly reduced, and the shape of the curves of fatigue crack propagation rate as a function of the stress intensity factor changed. Compared to GLARE, the ARALL is more sensitive to the residual stress condition. The fatigue properties of non-prestressed GLARE are better than those of ARALL. The influence of the residual stress is discussed in detail.     

Modelling cyclic shear deformation of fibre/epoxy layers in fibre metal laminates

Relations for a theoretical model were derived describing the shear deformation of uni-directional fibre/epoxy layers in Fibre Metal Laminates. The shear deformation is induced by the cyclic shear stresses at the interface that result from the cyclic load transfer from aluminium to the bridging fibre/epoxy layers during fatigue loading of the laminate. With the presented relations the crack opening contribution as result of shear deformation can be calculated. As the shear deformation is hard to measure experimentally, the theoretical model has been validated by comparing the model with finite element analysis. A good correlation has been obtained between the presented theoretical model and the FE-model representing the complex reality.     

Self-healing of delamination fatigue cracks in carbon fibre–epoxy laminate using mendable thermoplastic

This article examines the self-healing repair of delamination damage in mendable carbon fibre–epoxy laminates under static or fatigue interlaminar loading. The healing of delamination cracks in laminates containing particles or fibres of the mendable thermoplastic poly[ethylene-co-(methacrylic acid)] (EMAA) was investigated. The results showed that the formation of large-scale bridging zone of EMAA ligaments along the crack upon healing yielded a large increase (~300%) in the static mode I interlaminar fracture toughness, exceeding the requirement of full restoration. The mendable laminates retained high healing efficiency with multiple repair cycles because of the capability of EMAA to reform the bridging zone under static delamination crack growth conditions. Under fatigue loading, healing by the EMAA was found to restore the mode I fatigue crack growth resistance, with the rates of growth being slightly less than that pertinent to the unmodified laminate. The EMAA bridging zone, which generated high toughness under static loading conditions, does not develop under fatigue loading because of rapid fatigue failure of the crack bridging ligaments. Similar to the multiple healing capability of EMAA under static loading, multiple healing of delamination fatigue cracks is confirmed, with the fatigue crack growth rates remaining approximately unchanged. This study shows that EMAA was capable of full recovery of fatigue crack growth resistance and superior healing efficiency for static loading.     

Implications of interface friction for crack growth in fibre-reinforced metal matrix composites by three-dimensional finite element modelling

A three-dimensional finite element model of a compact tension specimen consisting of a Ti-6Al-4V matrix reinforced with unidirectional, continuous SiC fibres under monotonic and cyclic loading has been developed. This has enabled true Coulomb frictional interface sliding resulting from thermal residual stresses to be modelled. The results, which include the action of individual bridging fibres close to the crack-tip, are compared to results from a two-dimensional weight function method which uses fibre-induced bridging tractions on the crack face based on a constant interface strength. Reasonable agreement was found between the two methods used. An investigation of the fibre stresses showed that together with normal crack bridging tractions a strong bending component is present in the fibres which also affects crack opening and could affect the mode of fibre failure. The influence of processing induced thermal residual stresses and friction at the fibre-matrix interface on the crack growth behaviour during monotonic and cyclic loading has been assessed. It was found that the bridging fibres strongly reduce the crack-tip stress intensity factor. The thermal residual stresses produce a crack-tip opening load in the absence of an external load and have an influence on the crack-tip load ratio. The effect of the crack-tip load ratio on the fatigue threshold has a significant impact on the likelihood of crack arrest.     

Damage tolerance assessment by bend and shear tests of two multilayer composites: Glass fibre reinforced metal laminate and aluminium roll-bonded laminate

The damage tolerance of an aluminium roll-bonded laminate (ALH19) and a glass fibre reinforced laminate (GLARE) (both based on Al 2024-T3) has been studied. The composite laminates have been tested under 3-point bend and shear tests on the interfaces to analyze their fracture behaviour. During the bend tests different fracture mechanisms were activated for both laminates, which depend on the constituent materials and their interfaces. The high intrinsic toughness of the pure Al 1050 layers present in the aluminium roll-bonded laminate (ALH19), together with extrinsic toughening mechanisms such as crack bridging and interface delamination were responsible for the enhanced toughness of this composite laminate. On the other hand, crack deflection by debonding between the glass fibres and the plastic resin in GLARE was the main extrinsic toughening mechanism present in this composite laminate.     

Damage Mechanisms in Composite/Composite Bonded Joints Under Static Tensile Loading

This paper presents an experimental study of the initiation and development of damage in composite/composite joints. The materials studied are glass fibre reinforced polyester laminates and these are bonded together using an epoxy adhesive. Two types of joint interface are examined, assemblies between woven roving layers and between mat layers. The experimental techniques employed are dye impregnation, extensometry, visual observation and photoelasticimetry. It is shown that the nature of the surface layer does not influence the damage initiation load but does affect subsequent crack propagation. This revised version was published online in June 2006 with corrections to the Cover Date.