Advanced Materials and Innovative Structural Concepts I
Chair: Ravinder Chona
DESIGN AND TESTING OF ADVANCED COMPOSITE LOAD INTRODUCTION STRUCTURE FOR AIRCRAFT HIGH LIFT DEVICES|
Tamas Havar, Eckart Stuible
Abstract: Innovative composite structures are increasingly being used in the aircraft industry. A critical point of these composite structures is their attachment to the surrounding aircraft structure. In cooperation between different EADS Business Units, new advanced composite load introduction structures are developed to minimize weight and manufacturing costs.
For the development of new composite high lift elements, the load introductions are of most interest and therefore design rules for these composite attachments are investigated. Hereby, the design of single pin load introduction can be divided into composite lugs (e.g. Airbus vertical stabilizer attachment) and composite load introduction loops (e.g. helicopter blade attachment). Both designs are analyzed with the finite element method using non-linear boundary conditions (e.g. gap elements, deformable body contact definitions) and are afterwards verified by static and dynamic tests. The investigations demonstrate that for pure tension loading, the load introduction loops show the best performance, since the fibers are aligned with the load direction. For bending loadings that are perpendicular to the tension loads, the performance of the load introduction loop drastically decreases, because the load is internally carried primarily by the matrix and not the fibers. On the contrary, the lug shows a better performance, since part of the fibers is aligned with the loading. For a mixed mode loading both design show similar performance. According to the design rules, composite load introduction loops should therefore only be used at structures, which are primarily loaded in tension. For structures with various loadings, lug designs are a better compromise.
In addition bearing friction is considered in the analysis of the load introduction. Based on non-linear calculations, design rules for single pin load introduction structures are developed. The design rules show, that the friction moments for different rotation axes vary up to 40%. The highest moment occurs around the rotation axis of the bearing. For swiveling of the bearing, which is the most critical loading for linkages, the friction moment is 27% lower than for rotation. For tilting of the bearing, the friction moment is smallest with 60% of the rotation moment. This investigation shows that the former approaches, where the rotation friction moment was applied for all 3 directions, are far too conservative.
A High Lift Support System for future aircrafts is developed and tested based on the former investigations. It consists of a hybrid flap track functioning as load carrying fairing with an aluminum frame and a composite skin as well as an innovative composite kinematics system. The kinematics system has several composite linkages and fittings, which connect the flap to the flap track. The linkages with two attachment points have primarily tension and compression loads. Therefore, these linkages are designed as load introduction loops. Linkages with three attachments points and fittings have mostly multiple loadings and are consequently designed as composite lugs. Each linkage element is analyzed using non-linear finite element methods showing sufficient strength for the occurring loads. Innovative manufacturing concepts for the composite Linkages and the flap track are developed to minimize manufacturing costs. The complete composite flap support system is manufactured and verified in a full scale fatigue test in accordance with Airbus specifications. The fatigue test proved that the composite flap support system is able to withstand all fatigue as well as static loads even with predefined defects.
HYBRID STRUCTURE SOLUTION FOR THE A400M WING ATTACHMENT FRAMES|
Matthijs Plokker, Derk Daverschot, Thomas Beumler
Abstract: The mainframes in the centre fuselage of A400M have the vital function of introducing the wing loads into the fuselage. The lift distribution among the wing causing a torque and the frame cur-vature causing a bending moment, result in high running stress and limit loads in the frames below the rear wing attachment. These frames are made of 7000 aluminium alloy and are there-fore primarily dimensioned by fatigue reasons, especially crack growth. Consequently, three different options for the A400M mainframes have been investigated to fulfil the fatigue, mainte-nance, and weight targets.
The investigation was focused to reinforce the inner flanges of the frames. The inner flange is the highest loaded and critical part of the frame. Three reinforcement options were investigated:
Thickening of the aluminium inner flange to reach an acceptable stress level
Application of riveted Titanium strap attached to inner flange
Application of bonded Glare® strap attached to inner flange
In the option 2 and 3, the strap acts as damage containment feature that redistribute the load in case of a damaged frame. The first part of the investigation was the comparison of the three designs on weight, fatigue behaviour, and cost.
A coupon test program was performed to investigate the behaviour of aluminium reinforced with bonded Glare® strap. This shows a constant crack growth rate for long ranges of crack lengths (in contrary to aluminium or titanium reinforced inner flange)
The inspection and cost requirements of the bonded Glare® strap reinforcement are well within the programme boundaries. Additionally, the Glare® reinforced aluminium inner flange offers a weight opportunity of almost 20 kg’s.
The certification of this application on A400M will be achieved by a series of coupon, compo-nent and the full-scale fatigue test. The local metal bonding process on a large part with com-plex contour is considered as challenging. Therefore a series of experiments is carried out to investigate the effect of artificial delaminations and defects below the detection size in the frame/strap bondline. A detailed finite element investigation is being done to determine the maximum size of delamination that sustains limit load condition. This analysis is validated by panel component testing. Completing the design philosophy, manufacturing process and certifi-cation of this particular application, anti peeling fastener will be installed locally, i.e. in rivet pitches that equal the determined maximum artificial delamination length.
CONTROL OF CRACK GROWTH RATES AND CRACK TRAJECTORIES FOR ENHANCED FAIL SAFETY AND DAMAGE TOLERANCE IN WELDED AIRCRAFT STRUCTURES|
Phil Irving, Xiang Zhang, Yu-E Ma, Guido Servetti, Stewart Williams, Gary Moore, Jorge dos Santos, Marco Pacchione
Abstract: Use of welds in aircraft structures creates significant local inhomogeneities in material microstructure, in strength and toughness, and in residual stress fields – all located in and around the line of the weld. Initiation and growth of fatigue cracks in the vicinity of the weld line is markedly affected by these inhomogeneities. Welding in addition creates large integral structures which require additional design considerations and analysis techniques in order to improve and enhance the fail safety and damage tolerance capability of this kind of structures. Frequently weld residual stresses and microstructure changes can act to accelerate fatigue crack growth rates. However, weld stresses can be modified via cold work and by post weld machining operations, and could result in reduced tension or even compression stress fields which will reduce crack growth rates and enhance fatigue and damage tolerance behaviour. This attractive prospect is contingent on two key capabilities.
• The ability to control development of microstructure and residual stress during and after welding so that desirable distributions with reduced crack growth rates are achieved.
• The ability to predict crack fatigue growth rates and crack trajectories in response to material property distributions and to internal as well as externally imposed stresses.
Recent research has made significant progress in both control of residual stress and prediction of fatigue crack growth rates in the presence of local residual stress fields. The paper will describe work that has been undertaken in controlling residual stress in welded structures, together with recent work in predicting fatigue crack growth rates and crack trajectories under the influence of weld residual stress fields. The predictions have been validated against experimental data.
Experiments to explore the ability of cold work applied during and after welding to modify the residual stress fields associated with weld operations will be described. Residual stresses measured on weld samples with and without cold work demonstrate that the stresses can be successfully modified, and that this in turn has modified the fatigue crack growth rates through the welded samples.
To validate fatigue crack growth rate predictions, crack growth samples in CT, M(T) and ELSEN geometries of three different sizes were machined from friction stir welded 2198 and 2195 sheet and plate in a range of thicknesses and weld orientations. Residual stress distributions in the samples were measured using the neutron diffraction and synchrotron X-ray techniques. The different sample size and crack orientations resulted in a wide range of residual stress distributions. Fatigue crack growth rates were measured in the samples to provide an experimental reference data set for the predictions. Further tests were conducted to establish the influence of residual stresses acting together with local stress concentrations on fatigue crack trajectories and the stability of the crack path within and outside the weld.
To predict crack growth rates in the welded samples, values of stress intensity factors associated with the residual stress field (Kresid) acting on the crack tip throughout the crack trajectory were calculated using residual stress field data. Kresid values were used to modify the values of Kmax and Kmin acting on the crack tip due to the applied loads. Fatigue crack growth rate changes arising from the changed mean stress were calculated from parent plate data using the Walker and NASGROW equations. Good agreement of model predictions with experiments for all levels of residual stresses and crack growth paths was obtained.
The implications of the results on capability to control and predict fatigue crack development in welded aircraft structures will be discussed.