Fibre Metal Laminates

In this chapter a brief overview of the history of Fibre Metal Laminates Arall and Glare is given as background information for the other, technical chapters of this book. This is a summary of a more complete description that can be found in [1].

This chapter gives a definition of Glare as a material-type. Following discussion of the various Glare grades and special material features, a step is made to the value of this material for use in aircraft. Several established yet still unexplored fields of application are mentioned and briefly discussed.

As the name Fibre Metal Laminates (FML) already suggests, they are a mixture of sheet material and fibres, bonded by an adhesive. Arall and Glare are the first and second generation FMLs respectively However, other potential variants have been developed with success as well. In the early nineties a carbon titanium laminate was developed for operating temperatures up to 300 °C, while at this moment, a glass aluminium laminate is under development for elevated temperatures up to 180 °C. As well as developing other laminates, different manufacturing processes have also been investigated. For example, a special manufacturing process for seamless tubes was successfully developed.

Aircraft are operated in a wide variety of environments, for periods of up to 30 years. In order to maintain safe and economical travel, it is necessary to have a good knowledge of the long-term behaviour of the aircraft materials. For longer periods of time, one of the main threats to materials is the influence of the environment, especially that of moisture combined with temperature. The combination of materials used in Glare complicates the investigation of the influence of moisture. However, by investigating environmental influences on each of the different components of Glare, a good understanding of the effects on the laminate as a whole can be obtained. In most cases, the laminated build-up has a positive effect on the environmental durability of Glare, but care needs to be taken in areas of combined moisture absorption and stress concentrations.

The material design allowable is a statistically determined minimum value of a material strength property. These values are necessary for the design of aircraft. In the case of Glare, material design values can be obtained through the use of the Metal Volume Fraction. The Metal Volume Fraction (MVF) is the amount of metal present in a Glare sheet. For the different material design allowables there is a (linear) relationship between Glare with a low MVF and a high MVF. The second subject concerns qualification of Glare. Qualification of basic Glare includes qualification of the underlying processes and materials. The qualification of spliced Glare is focussed on the qualification of the details in splice. Using the qualified details any splice layup can be made without having to qualify each splice configuration separately.

This chapter tries to answer the question: “How do we improve the Glare properties by searching for the most relevant aluminium alloy?” First, a quick overview of the available improvements made in the aluminium industry is given. A brief summarisation of the influence of the aluminium alloy properties upon the residual strength of the laminate follows, and finally the new aluminium alloys most promising for Glare are presented.

This section describes several methods that can be used to model the stress-strain curve of Glare.

Shear loads in an aircraft fuselage will occur as a result of bending and torsion. Yield values under shear are necessary during the design phase since the material should not deform plastically below the limit load. The Metal Volume Fraction (MVF) method for shear yield strength prediction of Glare laminates was investigated. The maximum deviation of the Iosipescu shear test data from the prediction line is 6.5%. The results indicate a good applicability of the MVF method.

This chapter describes the work performed to investigate the feasibility of using the Norris failure criterion for blunt notch strength prediction in Glare under arbitrary in-plane loading conditions (especially tension and shear). The described Norris method for Glare blunt notch is based on the experimentally determined blunt notch strengths in L-direction and LT-direction and the blunt notch strength under a 45° fibre off-axis angle.The Norris failure model was found to be very accurate in predicting multiaxial (biaxial and shear component) blunt notch strengths of Glare based on the Metal Volume Fraction approach. Moreover, a set of formulae is given to calculate the failure strength of an arbitrary laminate.

This chapter presents a concise overview of the current design methodologies employed for the static strength of stiffened Glare panels, regarding stability aspects (buckling / post-buckling) under compressive and shearing loads. In the light of the current developments in potential applications of Glare, the focus is on fuselage structures that are representative for Ultra-High Capacity Aircraft (UHCA). A number of experiments on such types of panels have been conducted for shear and compression loading. The results are discussed in this chapter and are compared to predictions from the current design methodologies and from analyses using the Finite Element Method (FEM).

Fatigue is an important design aspect, especially for the large number of geometrical notches in aircraft structures. The amount of flight cycles needed for a fatigue crack to grow from a detectable length to the critical length determines the maintenance inspection intervals, whereas the initiation life determines the threshold and start of frequent inspections. The longer the fatigue life of a cracked structure, the larger the inspection intervals can be, which reduces maintenance costs. Glare has excellent fatigue crack growth characteristics due to its laminated composition. In this chapter the fatigue mechanisms in Glare will be highlighted.

Mechanically fastened joints have proven to be weak links in an aircraft structure. Therefore, these joints require high quality design and manufacturing processes. This chapter focuses on the joint design optimisation. Glare provides the means to improve the joint design using advanced material design. The enhanced fatigue properties will demonstrate some design advantages. The following key characteristics need to be highlighted: load transfer, secondary bending, influence of biaxiality and curvature and influences of fastener installation. Secondary bending induced by the inherent eccentricities of joints introduces a non-linear bending stress in the joints. A fast and simple means to calculate the stresses in joints is provided in this chapter and it is shown by tests that the neutral line model for this purpose is accurate. A complex three-dimensional stress system exists around rivets, which is affected by the installation procedures. Improved riveting installation procedures, which control interference fits, can delay crack initiation by a factor of ten. The joint fatigue behaviour for Glare is described in three parts, i.e. crack initiation, crack growth and residual strength. Fatigue crack initiation can be predicted using the Fokker severity concept adapted for Glare. Stress levels in equally thick aluminium and Glare sheets show higher stresses for the aluminium layers in Glare. Fibre-bridging and bridging by the intact aluminium layers stop the crack created by the earlier crack initiation in Glare. Crack growth is significantly slower in Glare than in aluminium joints.Finally two failure modes, i.e. fastener failure and net section failure, dominate the residual strength behaviour of Glare joints. The first failure mode in Fibre Metal Laminates is a somewhat different failure behaviour than for monolithic aluminium, because of the build up of FML. This allows for delamination of the layers during the installation process. Countersunk rivets are most likely to cause delamination in the outer layers, since the contact area is smaller and thus provides less support for the rivet. The fastener will therefore be more easily pulled through the sheet. The net section failure mode phenomenon in Glare is characterised by intact fibres bridging the crack. Experiments and closed form solutions have shown that also rivet pull through can occur. The residual strength calculation of joints with fatigue cracks is based on the remaining net section strength of the aluminium. Residual strength is up to 1.5 times higher, dependent on the metal volume fraction for Glare, than for aluminium joints.The higher fatigue crack growth resistance makes lower thickness or higher stresses in Glare possible. Many tests indicated that a reduction of 30% in thickness still leads to higher fatigue lives compared to aluminium 2024 joints.

The residual strength of fatigue-cracked Glare specimens is virtually independent of the relative initial crack length. The number of cut fibres and the amount of delamination determine the residual strength instead.To obtain fracture, the energy available for crack growth (energy release rate G) must be larger than the crack resistance R of the material. With this knowledge a K_R-curve can be determined experimentally. The K_R-curve is better for Glare 2 than for Alclad-2524-T3 as regards specific weight. From experimental data it follows that for Glare the K_R-curve is a function of: the Metal Volume Fraction, the properties of the metal and fibre layers, the interfaces between these layers, the rolling direction in comparison with load direction and the quantity of fibres in de load direction. A prediction model for the residual strength of Glare was developed, but it can only be used for of Glare types that are based on the 2024-T3 alloy and loaded in either L-T or T-L direction. Additional research is needed to extend the model.To fulfil the two-bay crack criterion after FOD (necessary for newly designed aircraft), the residual strength of the Glare skin can be improved. This can be done by adding extra “crack stopping” fibre layers in the laminate. A test on a skin panel showed that this idea works and the crack tended to flap to the centre of the stiffener bay.

The spirit of Glare is its crack bridging mechanism, which provides superior damage tolerance properties. Depending on the specific property, Glare shows either monolithic metal or composite behaviour. This 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.

To assess the potential of Glare as a skin material for ultra large aircraft, several Glare skin panels were designed and manufactured for the Megaliner Barrel Programme. This programme is being executed as a joint effort between Airbus, Fokker Aerostructures and the Dutch laboratory for Aeronautical Engineering. This chapter gives an overview of the Glare development process for the Megaliner Barrel and focuses on requirement definition and the iterative sizing and design process. Some attention is also given to the complications that occurred due to a simultaneous development of the Glare shells, the Fabrication Technology for Glare components and the tools and data to design using Glare. The article starts with a short description of the structural layout of the Megaliner Barrel.

Cut-outs are highly fatigue sensitive due to the large stress concentrations. In aircraft fuselages these cut-outs are quite large in the case of the windows and doors. Much effort is usually needed to limit the stresses and displacements around these cut-outs. Displacements are mainly controlled by edge members and stresses are controlled by doubler packages. 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. For the door corners this may result in the use of Glare with a 45 degrees 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.

This chapter handles some aspects of the detailed design of aircraft structures in Glare, and the design of splices, internal doublers and riveted joints in particular. 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. In addition, internal local reinforcements (doublers) can be integrated into the panel during layup. A brief review of the development from splice concept to splice design is given. The use of doublers is discussed. The chapter concludes with a discussion on the design of riveted joints in Glare.

Laminates can be sensitive to delamination buckling, which occurs when a partially delaminated panel is subjected to a compressive load. The interaction of local buckling and extension of the delaminated zone typically results in a decrease of the residual strength and, eventually, in a collapse of the structure. Fortunately, this phenomenon has never been observed in experimental tests with Glare. Although the delaminated layers buckle locally, the delamination front does not propagate within the range of compressive stresses that can be expected in typical aerospace structures.Numerical analyses can give more insight into the mechanisms that prevent Glare panels from collapsing. In this paper, some experimental observations regarding delamination in Glare are discussed and, based on these observations, a numerical model is constructed at a meso-mechanical level. In this approach, solid-like shell elements are used to model the individual layers. They are connected by interface elements, which are capable of modelling delamination between the layers. These numerical techniques are used to simulate two classical delamination-buckling tests of a Glare laminate.

In recent years, significant progress has been made in understanding the behaviour of Fibre Metal Laminates, especially Glare. Now, this material has been chosen to serve as skin material in fuselage sections of the Airbus A380. However, a very challenging task must still be solved. All the available knowledge and research results must be transferred into quality standards and procedures, design and sizing methods in order to meet the stringent requirements set up by clients, airworthiness authorities and Airbus quality standards. Therefore, the following paper describes major activities that are carried out to transfer Glare from an invention into a real innovation in the aerospace industry.

Advanced Fibre Metal Laminates, Arall and Glare, were developed at TU Delft, The Netherlands. These patented, engineered materials were initially developed for application in the construction of fatigue prone aircraft components such as lower wing skins and fuselage skins. While ideal for these applications, commercial application in these primary structures was slow since new aircraft models were required in order to take full advantage of these materials and to avoid high additional costs of certification. Material characteristics of Fibre Metal Laminates are such that they also provide significant advantages over monolithic metals (aluminium, magnesium and titanium) and composites when used in the manufacture of secondary structural aircraft components. The materials are high-strength and lightweight. They are highly resistant to fatigue-crack propagation and impact damage. They are also highly resistant to flame penetration. Such characteristics make these materials ideally suited for flap skins, cargo bay liners, floors and firewalls. The materials are also ideal for use in the manufacture of speciality airline containers. This paper will review the characteristics of Fibre Metal Laminates and will give examples of significant improvements made in the service lives of aircraft components manufactured with these materials.

Different machining processes have been tested for their applicability when machining Glare, including drilling, milling, water-jet cutting and laser-jet cutting. For machining processes in general, three phenomena are important when a particular process is applied: the wear of the cutting tool, the delamination of the laminate, and the introduction of heat into the cutting tool and the laminate. Some changes in cutting tool geometry, tool material and machining procedures may be required for successful application of a particular machining process for Glare.This chapter describes the milling and drilling process in more detail. The influence of lubrication, laminate thickness and other parameters is also discussed.

Despite the limited formability of Fibre Metal Laminates (FMLs), most structural components for a stiffened wing or fuselage structure can be manufactured. In general the applicable manufacturing processes are a combination of conventional processes used for sheet metal parts and laminating principles used for composites. The formability of Glare is related to the different failure modes that may occur. These failure modes are related to the constituents of the laminates and the layup of the laminates. Despite the limited formability, the manufacture of stringers does not pose a problem, as has been demonstrated by extensive research. On the other hand, in-plane deformations of Glare laminates and the related forming processes, such as stretch forming, are affected by the limited failure strains of Glare.

The Self Forming Technique and the splice concept were developed for the manufacture of large single or double curved skin panels. This technique is very similar to the layup processes used for full composite structures. The application of the spliced concept enables the designer to design large skin panels, the only limitation being the size of the autoclave used for curing. Other details such as doublers, stringers, and thickness steps can be made during the same or subsequent autoclave cycle. The layup tool for the manufacture of those large skin panels can be a single curved or a double curved tool. The layup tools, which should have a high accuracy, have been designed as welded aluminium structures. For both single and double curved skin panels a few demonstrator panels have been made, which have proven the concepts for the large skin panel manufacture.

The non-destructive ultrasonic C-scan method is an effective tool to check the quality of Glare panels and components. This automated method measures the ultrasonic attenuation through the panel and images it over its entire surface. Local defects and porosity give a significant increase in attenuation and can therefore be located within the component. To determine the significance of these indications, the scope of the inspection method has to be defined and a reference system developed. An “effect of defects” programme is needed to provide suitable rejection levels.

This chapter shows the after-sales related activities for Glare structures within the Airbus consortium. For the Glare demonstrator panel on the MRT A310 MSN 484 of the German Air Force a supplemental SRM has been developed. This SRM is the starting point for the activities related to the A380 design and development. Allowable damage limits and repairs on Glare parts will be tested as part of the “Megaliner Barrel” research project.

Burn-through tests showed that Glare is very fire resistant. Depending on the Glare grade and thickness, only the outer aluminium layer will melt, whereas the other layers will remain intact. Carbonisation of the fibre layers and delamination of the laminate will insulate the non-exposed side of the material from the flames, which results in relatively low temperatures at the non-exposed side. Burn-through tests meant for baggage compartment liners were performed for more than 10 minutes, which did not result in burn-through or flame leakage of the Glare sheet.Lightning strike testing revealed a similar behaviour. Due to arc attachment of the lightning strike, heat was created and consequently the outer aluminium layer locally melted and/or vaporised. Also in this case the underlying fibre layers carbonised and were partially damaged, depending on the severity of the strike. In all cases, all layers underneath the first prepreg layer remained intact, whereas in a monolithic aluminium sheet the material melted through the thickness. In contrast with aluminium, Glare consequently showed a much better fatigue behaviour after a lighting strike, since most of the layers remained intact.

An aircraft structure should be able to tolerate a certain amount of impact energy and damage, dependent on the location and probability of impact. The structural design and material choice must guarantee the structural integrity after a realistic impact event. For Glare the impact damage resistance is related to the aluminium and glass fibre/epoxy properties. Glare is stronger than aluminium at higher impact velocities due to the strain rate dependent behaviour of the fibres. The dent depth after impact is comparable to aluminium with the same thickness. The residual properties for Glare are comparable to (e.g. compressive strength), or better than (e.g. fatigue) monolithic aluminium. The inspection of a Glare structure is no more complicated than for aluminium, since Glare has the same indentation and because the damage size due to impact is smaller than the visible dent size.

Corrosion is a problem that is hard to avoid in an aircraft structure. The different types of environments in which an aircraft operates and the large variety of materials used create a spectrum of different types of corrosion. Aluminium alloys are the most commonly used materials in aircraft structures. Aluminium itself is not so sensitive to corrosion, but the addition of other elements to the alloy makes it much more prone to corrode. In Glare the outer layers always consist of a sheet of an aluminium alloy, therefore the same type of corrosion-resistance can be expected as for a monolithic aluminium sheet. A difference may exist due to thickness effects. The thin sheets of aluminium that are used in Glare have a different microstructure compared to thicker sheets. This may have a positive or negative influence on the occurrence of different types of corrosion. If corrosion does occur, it will be limited to the outer aluminium layers, since the prepreg layers act as a corrosion barrier. Therefore Glare has better corrosion resistance than monolithic aluminium sheets.

For the repair of Glare skins it is important that the current repair methods and tooling for aluminium structures can be used in order to avoid an increase of the operational costs for airliners. It has been shown that Glare can be repaired with normal riveted repair procedures, which are described in the Structural Repair Manual (SRM) of existing aircraft. Workshop operations like drilling, cutting and deburring could be performed on Glare without any problems. The excellent fatigue properties of Glare provide the riveted repair on a Glare skin with very good damage tolerance properties. Fatigue cracks in the first rivet rows of a repair will only occur in one or two layers of the skin. Stresses can be transferred to other layers in the laminate. Depending on the Glare grade, the residual strength results were equal or better to identical repairs on aluminium structures, even when the fatigue loading was performed at a much higher stress level.The repair of a Glare stringer could also be performed according to an existing SRM. Again, all normal operations, which were also performed on the Glare skin repair, did not lead to problems when performed on a Glare stringer.

The growing number of ageing aircraft around the world, both military and civil, has led to an increasing demand for safe, damage tolerant, and cost-effective repairs to solve the inherent fatigue damage and corrosion damage problems. Adhesive bonding of Glare repair patches can provide such a solution. While Glare can be used for conventional repair methods like riveting and bolting, bonded patches provide a uniform and efficient load transfer into the patch and reduce the risk of high stress concentrations caused by mechanically fastened repairs. The advantage of Glare repair patches compared to other patch materials lies in the small mismatch in the coefficient of thermal expansion (CTE) with the aluminium skin, the excellent fatigue properties, the high strength, the moderate extensional stiffness and high bending stiffness. This makes Glare a promising candidate for safe and cost-effective bonded repairs for thin skin structures, not only for ageing aircraft, but also for repair of in-service damage of relatively new aircraft (high request of repair performance). Extensive research has been performed in a joint research programme of the Faculty of Aerospace Engineering of Delft University of Technology and the Center for Aircraft Structural Life Extension (CAStLE) of the United States Air Force Academy. An overview of this research will be presented here.

This chapter describes the application of Glare repairs to a USAF C-5A transport. One of the problems in the ageing C-5A fleet is the multiple small cracks in the upper aft-crown section of the fuselage skin. These cracks are possibly caused due to the usage of stress-corrosion sensitive 7079-T6 aluminium. The crown section experiences significant longitudinal tensile bending in addition to biaxial tension due to internal pressurisation. Multiple short cracks (25 to 50 mm long) were found, see Figure 31.1, in the crown of several aircraft. In the autumn of 1995 two bonded Glare repairs were applied at fuselage stations 1700 and 1784 by a team of Wright Laboratory, San Antonio-Air Logistics Center, and United States Air Force Academy/Center for Structural Life Extension (CAStLE) personnel [1].The cracks are believed to have nucleated in the rivet holes of the 7079-T6 aluminium skins due to high fit-up stresses induced in manufacturing and the bending moment caused by the T-tail. Figure 31.1Crack Iocations in crown section of C-5A.

The eddy current non-destructive inspection technique can be used effectively on fibre metal laminates. It uses the principle of electromagnetic induction, i.e. the interaction between a magnetic field with an electrically conductive object. With this method, the different Fibre Metal Laminate types can be easily discriminated and fatigue cracks in the metal layers of the laminate can reliably be detected in, e.g. plain material, lap joints and spliced laminates.

The difference between the terms “Environmental Friendly” and “Sustainable Development” is their space and time frame. Sustainable Development has a much wider scope in terms of time and space.Glare may not be completely environmental friendly (although it is completely harmless for humans when proper precautions are taken) during its life cycle, but, more important, it is a very promising material with respect to Sustainable Development. Recycling of Glare can not create a completely closed loop of energy and material without any landfill. But, this limited environmental friendliness does not make a contribution to Sustainable Development impossible.The IPAT theory shows that for a Sustainable Society, an enhancement of the efficiency of technology by a factor of 5 to 50 is needed in the next 50 years. The promising aspect of Glare is that it contributes to this required factor to make a 20% higher efficiency possible, of the new generation of aircraft, because of reduced fuel consumption.Weight savings, for instance due to the use of Glare, are usually used to study how much energy, i.e. fuel can be saved. However, the question should be raised whether this saving will create more transport as a result of lower prices. This is called the ‘rebound effect’. But it is also possible that more payload is transported on each flight and thus the number of flights is reduced. The effect of the application of Glare is therefore not certain.