Fiber-metal laminates were developed as a family of hybrid materials that consist of bonded thin metal sheets and fibers embedded in adhesive, i.e. rubber-thoughened epoxy (Fig 1). This laminated layout creates a material with excellent fatigue, impact, and damage tolerance characteristics, and a low density (Vogelesang and Gunnink 1986, Vlot et al. 1999). The bondlines act as barriers against corosion of the metallic sheets, and the metal layers protect the fiber/epoxy layers from picking up moisture. The laminate has an inherent high burn-through resistance as well as good damping and insulation properties. The material; like monolithic aluminium, these sheets can be machined and formed into products. The laminate can also be cured in an autoclave into a complete structure: large, curved panels with co-cured doublers and stiffening elements. The so-called "splicing concept" makes a larger panel size possible compared to conventional aluminium structures, with consequent reduced assembly costs. The possible 20% weight reduction was the prime reason behind the development of these materials for aircraft structures.@en

Fibre metal laminates were developed at Delft University of Technology during the last two decades as a family of new hybrid materials consisting of bonded thin metal sheets and fibre/adhesive layers. This laminated structure provides the material with excellent fatigue, impact and damage tolerance characteristics and low density. While the 20 per cent weight reduction was the prime driver behind the development of this new family of materials, it turns out that additional benefits like cost reduction and an improved safety level has become more and more important. The combination of these aspects in one material makes fibre metal laminates a strong candidate material for fuselage skin structures of the new generation of high capacity aircraft. The focus on this application currently leads to industrialization and qualification that makes this material available to the aircraft designer.@en

The number of ageing aircraft is increasing and will further increase considerably in the near future. The control of damage by inspection and repair becomes therefore more and more important. The bonded repair method proved to be a cost-effective way to repair damage. However, this method has some disadvantages. One of them is the introduction of thermal residual stresses in the repair, due to the mismatch in coefficients of thermal expansion between the skin material and the repair material. Two analytical models are available to determine these residual stresses in the repairs, i.e. the Wang-Rose model and the Van Barneveld-Fredell model. A new computer program CalcuRep2000 [1] is being developed. This computer program will be used to calculate the stresses in bonded repairs, and is developed by the US Air Force in co-operation with the faculty of Aerospace Engineering of Delft University of Technology in The Netherlands. For this program an analytical model, which is able to calculate the thermal stresses in a bonded rapair accurately, is essential. This paper describes how the most adequate analytical model was identified by comparison with Finite Element (FE) analysis. Two repair configuratios were distinguished: The test specimen, i.e., bonded repair specimen used for fatigue tests in laboratories, and the in-field specimen, i.e., a realistic repair. The difference between both repair configurations is that a test specimen is free to expand during curing while in a realistic repair configuration expansion is partly prevented by the cooler surroundings of the heated area. The analytical model of Van Barneveld and Fredell proved to be the most accurate one. The results of this model agree with the results of the FE model of the test specimen and the in-field specimen. The Wang-Rose model calculates the thermal residual stresses accurately in the case of a test specimen. but the stresses calculated in the case of an in-field specimen are less accurate. No useful experimental data are available in literature from measurements of the thermal residual strains on bonded repairs. Therefore experiments were conducted to measure thermal residual strains in bonded repairs. The same repair configurations were used for these experimentss. The results obtained with the experiments are compared with FE results. Finally, the temperature distribution calculated by the analytical models is compared with the temperature distribution measured during the experiments of the in-field specimen. The influence on the thermal residual stresses of a different temperature field is discussed here, using the FE results@en

Fibre Metal Laminates (e.g. Glare) are composed of a bonded structure of metalic layers and fibre reinforced adhesive layers and were developed by laboratory testing and design during the last fifteen years. The Glare Technology Development (GTO) has been aiming at final technology readiness for application in aircraft design, production and service. The current paper gives an overview of the work that was done in this large project.@en