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13:30   Fatigue Crack Growth and Life Prediction Methods: III Chair: John E. Moon 13:30 30 mins SOME ELEMENTARY ENVIRONMENTAL INVESTIGATIONS ON STANDARD-GLARE® Thomas Beumler, Jos Sinke, Bob Borgonje Abstract: The application of GLARE® panels on the A380 fuselage and tailplane demanded, among other efforts, the development of a fatigue strength justification philosophy. Although certified as a metal structure in general, some key properties of Fiber Metal Laminates had to considered for more detailed investigations, e.g. the multiple load path character of the laminate itself – allowing for the application of the fatigue approach, the internal stresses which lead to relatively early fatigue crack initiation in the FML and the obligatory Multiple Site Damages (MSD) in FML riveted joints. Associated to these topics the transition phase from crack initiation to fatigue crack propagation (start of crack bridging), the relatively long fatigue crack propagation phase, the relatively low crack propagation rates in FML’s and the almost linear dependency of residual strength on the net aluminium section demanded attention. The hybrid character of any FML and in particular the presence of the fiber/epoxy prepreg demanded an extended investigation of the material behaviour, beyond the usual investigations from elementary to full scale level performed for monolithic aluminium. In a six years program, the Aerospace Faculty of Delft University of Technology and Airbus shared and extensive investigation of Standard-GLARE® specimens and -structures under different environmental conditions as well as under different temperature conditions. Involved have been several static specimens and tests (e.g. bearing, blunt notch, rivet pull through) and a significant number of specimens dealing with fatigue crack initiation, fatigue crack propagation and residual strength. However, even a part of the static specimens, for example compression filled hole, have been tested with fatigue cracks in order to complement the design- and justification philosophy agreed for the A380 FML structures. While Delft University provided most of the “as-received”-, accelerated aged- and outdoor exposed test data, Airbus focused mainly on temperature related investigations, panel tests and on tests which required special NDI techniques. The combined program included tests from elementary scale to full scale panel level, with more than 200 specimens up to flat panel size (riveted repairs and bonded repairs) exposed outdoors in Queensland/Australia. This paper will discuss some of the challenges which had to be mastered or still have to be mastered in order to monitor the material behaviour at the vicinity of a bore hole under various conditions. A selection of test results on elementary scale and from coupons will be presented. Prior to any discussion of results it must be noticed that, for the purpose of the investigation, none of the involved specimens has been treated with the complete Airbus surface protection system. Specimens involved in outdoor exposure have been primed and painted in dark blue color (U.S. Airways paint system at the time) in order to increase the specimens temperature and thereby to increase diffusion rates. Drilling and fastener installation has been done after painting and without application of anti fretting compound in order to investigate a worst case. Thanks to the barrier function of the external metal sheets, diffusion in FML’s is limited to either holes or sheet edges. The edge is an item to investigate with relatively simple specimens. However, it was necessary to create a set of new specimen types in order to investigate the potential influence of moisture on the properties of a hole drilled into the laminate. The outdoor exposure weight gain campaign demonstrated again, that diffusion is a reversible process, the weight of the specimens followed the seasons of the tropic ambient air (specimens under glass cover). It is therefore concluded that the artificial ageing process applied earlier on the Standard-GLARE® qualificaton specimens is a rather conservative one. As well it could be expected that diffusion into the matrix systen under outdoor conditions would not significantly influence the mechanical properties of the laminate, different to experiences with specimens exposed extremely long in an environmental chamber (e.g. 20000 hours). The monitoring of crack initiation on the mating surfaces of riveted FML joints remains another challenge, because the detection of a 1mm crack is demanded. NDI and test specialists have been pushed to the limits by the FML community. More or less sophisticated methods are available, some associated with significant costs, which all influence the test result in the one or the other way. This information is of importance if one tries to compare or judge test results from different sources. Combining results obtained from similar specimens but different inspection methods might suggest the impression of high scatter, which is for the individual result not necessarily the case. For the practical application of the result, i.e. the strength justification of the aircraft, it is then the minimum requirement that the developed analytical method for strength prediction covers all test results conservatively. A particular item of interest is the investigation of the freqency influence on the transition point crack initiation / crack propagation. Even for monolithic aluminium the cause of the frequency influence remains a speculation to some extend. Vogelesang and Schijve performed intensive investigations on aluminium and identified the transition of tensile mode to shear mode of small cracks as at least one of the reasons. However, in frame of this FML project frequency tests with elementary specimens made of both, aluminium and GLARE® have been repeated, in order to verify that they behave similar. Interesting is, that the frequency influence could be demonstrated for both materials for fatigue up to 1mm crack length, i.e. for a very low crack propagation portion of the fatigue life. The relevance for the frequency investigation was the sub-project which dealed with variable temperature tests. Both the cooling and the heating of a specimen demands time and leads automatically to a low frequency. The challenge was to separate potential temperature-, frequency- and environmental influences if present in various fatigue tests for a better understanding and more optimized application of the advanced technology. Keywords: GLARE, fatigue crack initiation, fatigue crack propagation, residual strength, exposure, temperature, frequency 14:00 30 mins ADVANCES IN CRACK GROWTH MODELLING OF 3D AIRCRAFT STRUCTURES Sharon Mellings, John Baynham, Robert Adey Abstract: This work is focused on the 3D simulation of crack growth in metal structures exposed to complex multi-axial loadings as regularly occur in aircraft. Simulation of these cracks under the applied loading can be used by the design engineers to investigate changes in the design, loading or materials. Given an initial crack, the aims of the simulation are to identify the stress intensity factors occurring at any stage of the loading cycle, and to predict the time taken for the crack to grow through the structure. This means that vulnerability to fast crack growth is assessed, and that fatigue growth calculations of growth direction and distance are performed and accumulated. The result is that crack life predictions are made of the crack size variation with number of cycles. In additional multiple cracks can be simulated in the model, and the complex stress behaviour of the interaction of the cracks can be investigated under the applied loading conditions. The use of multi-scale techniques which allow extremely detailed representation of the crack(s) is discussed. These techniques readily allow the designer to perform what-if investigations using a variety of scenarios which can be selected from any part of the model, to determine the rate of crack growth, and the path which the crack will follow. The resulting large scale model may use the sometimes approximate geometry of the small scale model, or more accurate geometry can be constructed to take better account of the real geometry. These approaches are discussed and demonstrated. The effects of residual stress may be significant, and the way in which such effects can be included in the multi-axial crack growth investigations is discussed. This allows for built-in stresses from either the manufacturing process or from surface treatment effects to be simulated. If surface treatments are being used to generate a residual stress field to slow any crack growth, then the effectiveness of different residual stress fields can be investigated analytically without the need to use expensive test programs. In the analysis the stress intensity factor results are computed in all modes (modes I, II and III) and this allows the turning of the crack to be simulated. Multiple load cases can be superimposed to give multi-axial loading and various techniques to calculate the growth direction under multi-axial loading are available. The differences between the methods are evaluated for a range of different loading scenarios. Under some loadings, a crack may tend to close, so that the mode I stress intensity factor is zero but mode II and mode III stress intensity factors are non-zero. The use of contact simulation on crack surfaces to prevent interference and allow tangential slip is considered and examples are presented. The simulation tool is aimed at simplifying the modelling process, as cracks can be added into the model simply by selecting the crack location, orientation and size. It is not necessary to manually modify the mesh in advance. Therefore this is especially ideal for parameter studies aimed at determining worst-case scenarios for crack location, and improving the accuracy of the growth predictions. 14:30 30 mins DEVELOPMENT OF ADVANCED RISK ASSESSMENT METHODOLOGY FOR AIRCRAFT STRUCTURES CONTAINING MSD/WF Min Liao Abstract: A risk based management approach/tool has been adopted by many military air fleets. In the past few years, Canadian Forces (CF) has been introducing and revising a Record of Airworthiness Risk Management (RARM) process to manage technical and operational airworthiness for all CF aircraft. Today, the RARM has become the single most critical decision making tool in the CF air force. In RARM, both qualitative (defining hazard probability as ‘frequent’, ‘remote’, ‘extreme improbable’, etc.) and quantitative (defining hazard probability as ‘10-3’, ‘10-5’, ‘10-8’, etc.) risks are defined for all CF aircraft platforms including UAVs and helicopters. When there is sufficient data available, a quantitative risk assessment (QRA) can be performed to substantiate the assignment of a risk index number. However, some challenges are remaining for a QRA, such as how to establish methodologies to get sufficient/meaningful input data, particularly for a complicated damage scenario such as MSD (multi-site damage)/MED(multi-element fatigue damage)/WFD(widespread fatigue damage). This paper presents the latest results of IAR/NRC research on risk assessment methodologies /tools for aircraft structures containing MSD/WFD, in order to support Canadian Forces air fleet life cycle management. To carry out MSD/WFD risk analysis, a methodology for determining the initial crack size distribution (ICSD) was developed to describe the random distributed MSD cracks in a wing panel. A NRC crack growth analysis program was developed and used to simultaneously grow the MSD cracks, considering the interaction effects of MSD cracks and adjacent structure failure (MED). In this program, numerous beta factors for MSD and MED cracks were developed, and verified with p-version finite element analysis results (StressCheck) and other sources. The crack growth program was also designed to allow an efficient Monte Carlo Simulation for risk analysis purpose. With the aid of finite element modeling, a global-local technique was developed to determine the residual strength curve for the MSD/MED wing panel. Finally, the coupled Monte Carlo simulation with probability integration techniques was developed and included in the NRC risk analysis code, ProDTA, to accurately calculate the PoF (probability of failure) of the MSD/MED structures. Through several case studies on different aircraft structures, sensitivity studies were performed to show the significance of initial crack size distribution, maximum stress distribution, residual strength curve, crack growth curve, and probability of detection on the risk analysis. The best practices and the gap between theory and practice are discussed in order to establish the generic risk assessment procedures for the fleet management.