At the end of the second millennium did the aircraft industry decide for the first time to apply the fiber metal laminate GLARE in a large quantity on a civil transport aircraft. It was focused on an application of the material on the pressurised fuselage, the decision driven by the demand for weight saving at an affordable cost level. GLARE material properties are linked to both, monolithic aluminium properties as well as fiber composite properties, with the preference depending on the particular strength feature under investigation. Between others, three major structural mechanic advantages compared with monolithic aluminium can be identified for fiber metal laminates, i.e. the lower density, the crack bridging capability of the fibers in presence of a fatigue crack in the particular aluminium sheets and the possibility to tailor the material according to structural mechanic requirements by appropriate orientation of the fibers. Essential for an economic application of the material is the acceptance of 'flying with undetectable fatigue damages'. This rule and it's implications on structural sizing and justification is discussed in depth in this report, leading to the particular subject of the strength behaviour of riveted joints in a fatigued condition. Because GLARE contains an epoxy resin prone to moisture absorption, strength degradations due to environmental influences have to be taken into account. However, as for other materials, it is searched for a realistic exposure in order to design a structural component to the limits of the material strength and to avoid unnecessary high reserves. An outdoor exposure program is under evaluation as part of the structural investigation, which extends the time frame of this thesis. Structural components are supposed to remain for up to 6 years on a tropical exposure site. However, evaluated weight gain measurements allow extrapolations for 30 years aircraft operation and the definition of a representative accelerating ageing process. Extensive results on particular GLARE related investigations are reported since the end of the 1980's. This thesis is compiling the available information on fatigue issues and environmental related material issues, it is extending the previous research and it interprets the material properties within the framework of the mandatory airworthiness rules. For that purpose, two structural items which are tested by Airbus under full scale conditions are investigated concerning all particular strength properties which are required for certification or not are investigated. The application of airworthiness regulations which are defined for monolithic aluminium are discussed for GLARE in chapter 1. While fatigue sensitive monolithic aluminium structures have a relatively long crack initiation life but a relatively short crack propagation life, GLARE shows the opposite behaviour. What does that mean for the certification of a GLARE structure? Which fatigue methods should be used for the certification and is GLARE a single load path or a multiple load path structure? Some basic rules are discussed and proposals concerning the structural certification are provided. No type certification of a commercial passenger aircraft is accepted without test substantiation. The test pyramid is starting with elementary specimens (can be tested early in a project) and finishes with a full scale fatigue test. Chapter 2 discusses the particular influences which have to be considered for structural certification, e.g. crack initiation scatter and temperature sensitivities to crack initiation, and which have to be reconsidered for GLARE. Chapter 3 presents all specimen types which belong to the outdoor exposure program, i.e. test series from which some specimens are shipped to a tropic exposure site and some are tested under laboratory conditions, for comparison. The relevance of the particular specimen types and their history for GLARE is briefly discussed. A review of accelerating ageing experiences of previous researches with Fiber Metal Laminates is done in chapter 4. Diffusion coefficients for GLARE3 made of 2024T3/FM94 are determined. First weight gain measurements which are available from the outdoor exposure test site are investigated and interpreted in chapter 5. Predictions of the weight gain of GLARE around bore holes for both, long range aircraft and short range aircraft, are performed. Chapter 6 presents elementary specimen test results for the two tested materials, related to different ageing conditions. First experiments with specimens which have been exposed for one year outdoors are included. The crack initiation and crack propagation scatter for a row of bore holes is determined. Chapter 7 is dealing with coupon specimens, which are representative for the full scale structure tested by Airbus. A review and calibration of crack initiation, crack propagation and residual strength methods under development by Airbus Deutschland and Delft University is performed for a prediction of the full scale behaviour. The calculated results are discussed in view of a fatigue & damage tolerance certification. Full scale test results are related to the elementary- and the coupon experiments performed in this thesis. The outdoor exposure investigation is extended by a few panel tests, i.e. non-stiffened riveted repair panels, non-stiffened bonded repair panels and door corner cut out specimens. This analyses are collected in chapters 9 to 11. All experiments and analysis are performed with/for GLARE composed of aluminium 2024T3 and prepreg FM94/S-glass, cured at 120°C. Due to the wide range of specimen types each of them couldn'd be provided in a high quantity, for economic reasons. Therefore just property trends can be provided. The investigations performed in frame of this thesis contribute to the verification of the TU Delft / Airbus computer program 'FML F&DT Toolbox', which is developed at the same time.