Issue 35

R. Sepe et alii, Frattura ed Integrità Strutturale, 35 (2015) 534-550; DOI: 10.3221/IGF-ESIS.35.59 535 loads of varying amplitude and complexity for a specified period of testing in order to simulate the fatigue loads that are predicted to occur over the life of the airframe. The main aim of the full scale test is to produce fatigue damage, that may be expected during service, in order to discover where fatigue cracking may occur and to assess its behaviour. Together with the experimental activity, in the last 20 years, the finite element method has been further developed and now represents a useful tool in the numerical evaluation of local stress and strain; the results obtained can be used to carry out both fatigue strength and a damage tolerant assessment. Cordes [6] predicts fatigue cycles from stress/strain data, a calculated crack-tip energy, and a calculated fracture toughness. Grbovic and Rasuo [7] show that it is possible, by using finite element analysis (FEA), to obtain not only the good estimation of the fatigue life of the assembly such as the spar of the light aircraft, but also a good prediction of a number of load cycles which will propagate a crack on the spar to a certain length. Šedek et al. [8] evaluate the crack growth in an unstiffened and integrally stiffened wing panel without and with retardation effect evaluation. Crack growth predictions are performed by using both the linear damage accumulation principle and the FASTRAN retardation model. Furukawa et al. [9] a new methodology for modeling fatigue crack propagation in pressurized cylindrical shell structures. The methodology involves the evaluation of four stress intensity factors that are used to characterize cracked shell structures, the determination of the crack propagation trajectory as an integral part of the simulation process, and the prediction of the load cycles needed for the crack to propagate from its initial length until the final crack length. In addition, the numerical analyses make it possible to study the fatigue and damage tolerant behaviour of complex structures using fracture mechanics approach [10-12]. Moreover, in order to determine an acceptable fatigue life for the joint structures or to develop an effective riveted reinforcement methodology an extensive research has been conducted in the area of riveted patch joint performance: the majority of numerical analyses have been performed using the Finite Element Method (FEM), but some work has also been done using the Dual Boundary Element Method (DBEM) [13-15]. This paper reports the results of extended fatigue testing of a fuselage panel using a full-scale aircraft structural testing machine. The test panel was instrumented with strain gages, and previously quasi-static tests were conducted to ensure a proper load transfer to the panel. To support these tests, geometric nonlinear shell finite element analyses were conducted to predict strain and stress distributions and other parameters governing the crack initiation. The FE model allows to reproduce the effective shape and conditions of the specimen tested; moreover, it gives a complete description of the state of strain and stress in the whole panel. Crack formation during the fatigue test was monitored at every 10˙000 cycles using high-magnification visual inspection. After 164˙292 cycles a crack appeared on the panel frame, successively growing up and leading to several failures. After 12˙708 cycles more the panel broke down definitively at the middle-lower bay, near the third frame. Subsequently, a finite element analysis was carried out to correlate failure events; due to the actual biaxial nature of the fatigue loads, Sines criterion was used. The analysis was performed taking into account the different materials of which the panel is composed. The output shows a good correlation between experimental data and numerical results, that successfully may predict the failure locations on the panel. E XPERIMENTAL TEST Test description he full scale aeronautical test panel is shown in Fig. 1. The panel, whose sizes are 2181x2181 mm 2 (excluding the aluminum gripping plates), consists of two panels with the thickness of 1.2 mm, four frames and thirteen stringers. Stringers are connected to the frames with stringer clips. The frames are connected to the skins by shear ties. The lower and upper panel skins are joined by a lap transversal joint that is parallel to the stringer direction. The upper and lower skins are made of Al 2024-T3, the stringers, frames, shear ties and stringer clips are made of Al 7075-T6 51. The tested panel has been instrumented by strain gages, that are located on the internal side of the panel. Seven strain gages type CEA-13-250UW-350 of MM Vishay were bonded on the specimen. The strain gauges S1, S2, S4, S6, S7 (Fig. 2) measure the strains on the skin, while those S3 and S5 measure strains on lower pad of central frames. S2, S3, S4, S5, S6 provide strains in longitudinal direction of the panel, while S1 and S7 provide strains in transversal direction. The strain gages were bonded by a two-component epoxy adhesive in order to ensure good performance in case of large strains and each strain gage was attached to an acquisition system through a quarter bridge linkage. The positioning coordinates (X, Y, Z) of the strain gages are reported in Tab. 1. Some interesting machines were designed and manufactured in the past [16-19], but they seem to be able to apply only limited loading conditions on reduced scale specimens. Three loading axes, one diagonal, and two normal with respect to the side of the panel, have to be necessary to simulate correctly the structural behaviour of an aircraft fuselage, where the