11:00   Advanced Materials and Innovative Structural Concepts II
Chair: Luigi Lazzeri
30 mins
John Bakuckas, Keith McIver, Ching Hsu
Abstract: Adhesive bonding technology using composite and metallic patches offers an efficient and cost-effective approach to airplane structural repairs. Compared to conventional, mechanically fastened metallic repairs, bonded repairs have no stress concentrations due to holes, are less damaging to the parent material since no drilling or machining are required, and are more aerodynamically and structurally efficient. The application of bonded repairs has been studied primarily in the military sector where durability and damage tolerance aspects have been demonstrated. However, several technical challenges need to be addressed before bonded repair technology will be generally accepted and implemented in both military and commercial primary structural applications. Currently, credit is typically not provided in certification programs of bonded repairs for slowing crack growth or restoring residual strength. Of primary concern is the ability to predict the fatigue behavior and ensuring the durability of bonded repairs. In an effort to gain a better understanding of the durability and damage tolerance aspects of bonded repair technology, the Federal Aviation Administration (FAA) and the Boeing Company have partnered in a 3-year Cooperative Research and Development Agreement. A phased approach is being undertaken where the initial study will focus on the test and analysis of bonded repairs on a B727 fuselage structure using the FAA Full-Scale Aircraft Structural Test Evaluation and Research facility. The program objectives are to characterize the long-term durability of bonded repairs under simulated flight load conditions up to one typical design service goal and then determine if the repair patches meet damage tolerance requirements in a residual strength test. Bonded repairs to two damage scenarios are being considered, namely, a midbay through-the-thickness crack (fatigue presharpened) and a lap-joint scribe. Both boron/epoxy and aluminum patches were used to repair this damage. A photogrammetry method is being used to obtain full-field displacement and strain measurements at the patch regions. The acoustic emission method is being used to monitor for damage growth in real time and serve as an early warning for imminent failure. Several nondestructive inspection (NDI) methods are being used, including flash thermography and computer-aided tap techniques to scan for patch disbonds and eddy current to monitor crack growth. An overview of the results obtained from this study will be provided in terms of data to validate analytical predictions of bonded repair patch durability and residual strength, an assessment of several NDI in detecting disbonds and fatigue cracking, and lessons learned that will help identify safety and structural integrity issues of bonded repairs.
30 mins
Ivan Meneghin, Pascal Vermeer, Marco Pacchione
Abstract: Adhesive bonding provides solutions to realize cost effective and low weight aircraft fuselage structures in particular where the Damage Tolerance (DT) is the design criteria. Bonded structures that combine Metal Laminates (MLs) and eventually Selective Reinforcements can guarantee slow crack propagation, crack arrest and large damage capability. MLs are produced by adhesive bonding of a number of thin sheets to obtain the required thickness; the laminate may incorporate bonded local reinforcements. Selective Reinforcements expand the local reinforcement’s concept using fatigue insensitive materials to further improve the crack growth and residual strength performances. To optimise the design exploiting the benefit of bonded structures incorporating selective reinforcement requires reliable analysis tools. The effects of bonded doublers / selective reinforcements is very difficult to be predicted numerically or analytically due to the complexity of the underlying mechanisms and failures modes acting: debonding at the interface between the doubler and the skin around the crack tip, load redistribution between damaged part and intact reinforcements, fatigue damage of the metallic reinforcement eventually causing its premature failure, crack bridging of long crack in case of fatigue insensitive doublers. In addition, secondary effects like tensile residual stresses from the bonding process and secondary bending due to the eccentric doublers increase the complexity of the phenomena. Reliable predictions of crack growth can only be based on sound empirical and phenomenological consideration strictly related to the specific structural concept. With the purpose of investigating solutions applicable to pressurized fuselages large flat stiffened panels that combine MLs and selective reinforcements have been tested in frame of the EU Project (6th Framework Programme, No. 502846) DIALFAST - Development of Innovative and Advanced Laminates for Future Aircraft Structures. The large test campaign (for a total of 35 stiffened panels) has quantitatively investigated the role of: • different metallic skin concepts (monolithic vs. MLs) • aluminium, titanium, glass-fibre reinforcements • stringers material and cross sections • geometry and location of doublers / selective reinforcements. The performances of the different structural concepts are compared in terms of fatigue crack propagation over broken stringer and residual strength (2 bays crack). Bonded doublers and selective reinforcements confirmed to be outstanding tools to improve the DT properties of structural elements with a minor weight increase. • The effectiveness of the metallic doublers is dominated by its fatigue crack nucleation period • Aluminium doubler geometry/weight can be optimised to retard crack nucleation (thick and narrow instead of thin and wide) • Bonding aluminium doublers with Glass fibre reinforcements prevent fatigue nucleation creating a fatigue insensitive reinforcements • Titanium doublers provide the advantage of high stiffness and long fatigue nucleation period. These advantages are to a certain extent reduced by the internal stresses created during hot bonding process • Selective reinforcements (Titanium or glass-fibre) may significantly increase residual strength The tests confirmed the considerable role of the skin / stringer material, however ML skin provides negligible advantages in terms of through crack propagation. The role of stiffener section as determined by tests is in good agreement with prediction based on displacement compatibility analysis. The crack retardation mechanisms, based on local stiffening and crack-bridging effects will be discussed for the different reinforcement concepts.
30 mins
Malgorzata Skorupa, Tomasz Machniewicz, Andrzej Skorupa, Adam Korbel
Abstract: Some results of an experimental investigation on the influence of various factors on the fatigue behaviour of riveted lap joints are presented. The variables considered are the squeeze force and the fastener type and material. Also, the effect of a staggered sheet thickness in the overlap region is studied. The constant amplitude fatigue tests were carried out on lap joint specimens with three rows of rivets installed under load control. The 2mm thick sheet material was a Russian Al alloy D16Cz in the clad condition which, as shown elsewhere [1], is similar to the 2024-T3 Alclad regarding its chemical composition, mechanical properties and the fatigue crack growth behaviour. The round head rivets and the so called rivets with compensators according to a Russian standard were used. The compensator is a small protrusion on the mushroom rivet head. For both type rivets the ratio of the machined head diameter to the rivet shank diameter is near 2. Two Al alloy rivet materials used in the Polish aircraft industry were considered, namely PA24 and PA25 with the minimum shear strength of 185 MPa and 240 MPa respectively. The rivets can be installed in the as received condition. The tensile and compressive tests on the rivet wire have revealed that PA24 is similar to the Western AD rivet material. This was also confirmed by the measurement results on the driven head dimensions (D - diameter, H - height) versus the rivet squeeze force (Fsq) which coincided with the data reported for the AD rivets [2]. For PA25, the measured D vs. Fsq and H vs. Fsq data points fell below and above respectively the results for the Western rivets. In the present investigation, the round head rivets were of the PA24 alloy whilst both alloys were used for the rivets with compensators. The rivet shank diameter for all rivets was Do=5 mm. The effect of the rivet type and material was studied in the fatigue test series performed at the stress ratio R=0.1 and the maximum nominal stress Smax=120 MPa for D/Do ranging from 1.3 to 1.6. Consistent with the literature data, e.g. [3], the fatigue lives (Nf) of the specimens joined using the rivets of the PA24 alloy considerably increased with increasing the D/Do ratio, i.e. with elevating the Fsq level. The endurances for the rivets with compensators of PA24 were always longer than for the round head rivets. The joints with the PA25 rivets are superior to the others only at D/Do values below 1.5. High Fsq values required to obtain larger head diameters for the PA25 material caused cracking of the driven rivet heads which adversely affects the joint fatigue properties. In all cases the fatigue fracture occurred at one of the outer rivet rows and the cracks initiated on the faying surface from or near the rivet holes. At a given D/Do value, the crack nucleation site was always found to be more shifted from the rivet hole for a rivet with compensator than for a round head rivet. The latter behaviour can be explained by the larger hole expansion promoted by the compensator. In order to improve the fatigue quality of a lap joint Schijve proposed to stagger the sheet thickness within the overlap region for a joint with three rivet rows [4]. The concept is intended to provide for high stresses in the outer rivet rows by both reducing the secondary bending and lowering the load transmission through these rows. To verify the utility of the staggered thickness configuration the constant amplitude fatigue tests at R=0.1 have been carried out on three specimen series with the PA24 round head rivets shown in Fig. 1a and b. For the staggered thickness geometry with ps=1.5p the weight of the joint is the same as for the standard configuration. In order that the ”pure” effect of the joint geometry should be extracted in the tests, the D/Do ratio of 1.3 has been chosen. It was the highest value at which the crack initiation occurred still at the rivet hole edge. The fatigue test results shown in Fig. 1c indicate the beneficial influence of the thickness staggering, especially for the larger rivet row spacing. Though the percentage gain in the fatigue life decreases at the lower stress levels, still a considerable life increase of nearly 125 kcycles compared to the standard joint is observed at Smax=80 MPa for the staggered thickness configuration with ps=1.5p. References [1] Schijve J., Skorupa M., Skorupa A., Machniewicz T. and Gruszczyński P. Int. J. Fatigue 26 (2006), 1-15. [2] Rijck de J.J.M. PhD thesis, Delft University of Technology, Delft, 2005. [3] Müller R.P.G. PhD thesis, Delft University of Technology, Delft, 1995. [4] Schijve J. Doc. B2-06-02. Delft University of Technology, Delft, 2006.