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The American standard ASTM E837 presents a standard procedure to determine residual stresses in isotropic materials using the incremental hole drilling technique (IHD). The standard, however, presents limitations regarding its applicability, such as those related with the thin thickness of the samples. According to this standard, in depth non uniform residual stresses can only be determined, roughly, in plates where the thickness is greater than the mean diameter of the strain gage rosette used. This limitation excludes important experimental cases and, therefore, deserves to be investigated. In this work this limitation is numerically and experimentally investigated in detail, considering the case of residual stresses induced by laser shock peening (LSP) in aluminum alloy 7075-T651 plates. The obtained results using the incremental hole drilling technique (IHD), based on the integral method, are benchmarked against the results of several diffraction techniques, used as reference, and a procedure to correct the experimentally determined strain-depth relaxation curves, to accurately still apply the ASTM E837 standard procedure is discussed and validated.

The focus of this paper is on the use of the Laser Shock Peening (LSP) process, and the strategy with which it is employed, to retard fatigue cracks in thin metallic airframe structural panels. This approach allows for the extension of component life in both design and maintenance applications. LSP was carried out without an ablative coating. Various compressive residual stress fields were created by using different laser parameter sets, through the variation of power intensity and coverage. These were quantified using X-Ray Diffraction and Incremental Hole Drilling techniques. The size and position of the peened region was considered in order to balance induced compressive residual stresses and the resulting tensile residual stresses to obtain improved fatigue life and damage tolerance. The investigation was carried out using analytical fracture mechanics code AFGROW and validated through experimentation. Applying LSP to a pre-cracked aluminium panel provided clear retardation of the crack front as it propagated through the processed region. Decreasing the offset of the LSP region from the centre of the panel and increasing the width of the LSP region improved the results. The best configuration, a 15 mm wide LSP region with a 15 mm offset from the crack tip, yielded a fatigue life of more than 6 times the baseline fatigue value.

Fatigue and crack propagation tests are carried out on 7075-T73 open hole aluminium alloy specimens. A remarkable increment of fatigue life due to the Split Sleeve Cold ExpansionTM process is observed, further improved by applying the process twice on the same hole. Comparable increments are obtained by using the StressWaveTM method. Specific tests are performed in order to point out a possible relaxation of the residual stresses due to fatigue loading. The results obtained reveal no relaxation phenomena in the tested specimens. The compressive residual stress around cold expanded holes is numerically evaluated. Dedicated experiments based on the Sachs method are conducted to confirm the residual stress prediction.

The current paper describes a destructive, sectioning method for measuring residual stress in welded panels, using electric strain gauges. Since this method is carried out rather simply, it has a wide range of possible applications. For this reason, a more thorough analysis of the method is performed. The potential limits of its application are investigated, together with the parameters that can influence the obtained results. The sectioning method is verified by determining residual stress in welded panels numerically. The results show it is possible to estimate the residual stress field in a welded structure rapidly and accurately, by using reasonably priced equipment.

Fatigue tests were carried out on 4 mm thick spot welded joints; the material was stainless steel AISI 301, quarter hard. Some specimens were instrumented with a strain gauge bonded in correspondence with one of the edges of the spot weld. Strain gauge output was demonstrated to be a reliable instrument to monitor the nucleation and propagation of fatigue cracks. A good correlation was found between strain gauge output and spent fatigue life. Some fatigue tests were suspended when the strain gauge output was equal to pre-fixed values, corresponding to fatigue life in the range from 15 to 85%. Subsequently, the specimens were dissected to observe fatigue cracks. The same correlation existed between crack depth and fatigue life. Small cracks were observed in specimens fatigue tested up to 15% of the mean fatigue life; fatigue cracks in the joints under examination would be nucleated between 5 and 10% of fatigue life. Finite Element calculations were carried out, introducing in the models cracks similar to those observed in the fatigue tests. Calculated strain at the external surface compared well with the measured strain as a function of crack depth. Calculations demonstrated that small errors in strain gauge position can be tolerated without appreciable deterioration in crack dimension prediction.

The series version of the M346 military trainer aircraft is currently under construction at Aermacchi (Venegono Superiore, Italy). The design target life of the aircraft, which will be certified for Damage Tolerance, is 12 000 flight hours (FH), with the possible extension to 16 000 FH after specific inspections. Fatigue tests were performed on critical elements at the Department of Aerospace Engineering at University of Pisa in order to verify crack propagation calculations. The wing to fuselage connection is one of the most interesting elements from the fatigue point of view. Spars and frames, both integrally machined, are connected by two lug-fork joints; the base material is aluminium alloy 7050-T7451 for both the elements. High interference bushings, ForceMate®, produced by FTI (Fatigue Technology Inc., Seattle, WA) were used in the lug/fork connections. Experimental activity was carried out on two different specimens. The first, a Compact Tension specimen, was tested under constant amplitude loading to verify the fatigue crack growth rate data contained in NASGRO 4, the software used for Damage Tolerance evaluations. Experimental results were fully comparable with the NASGRO 4 material database. Additional variable amplitude loading tests were carried out in order to calibrate crack growth prediction models used in the analyses. The second specimen was a lug-fork joint designed as the actual joints present on the aircraft. Both constant and variable amplitude loading fatigue tests were carried out in this case too. Results obtained clearly indicated the beneficial effect of ForceMate bushings, compared to shrink fit bushings.

Since Friction Stir Welding has been identified as a key technology for primary aerospace structures, fatigue crack growth testing and modelling is required. However, there is still a complete lack of information regarding the fatigue crack propagation in friction stir welds under variable amplitude loading and flight loading conditions. Experimental investigations have been performed on centre cracked 4 mm thick AA2024-T3 base metal and FSW specimens: regarding the FSW samples, the crack was placed 5 mm out of the weld centre, in the most critical part of the joint: the coupons have been tested under simple variable amplitude load sequences, and under a standardized flight-simulation load history, Falstaff. The fatigue crack propagation was then predicted using widespread aerospace fracture mechanics software packages. Interaction effects and internal residual stresses were firstly separately simulated and than combined in order to evaluate the ability to predict the fatigue crack propagation on FSW welded structures under service loading conditions.

Fatigue tests were carried out on 2024-T351, thickness 1.6 mm, central hole specimens containing pins installed with five different interference-fit levels. Tests clearly demonstrated the beneficial effect of interference fit on fatigue resistance, up to the maximum value examined, 2.5%. A three-dimensional (3D) finite-element model was used in order to characterize the stress field around the hole. A large specimen, with a 40-mm-diameter hole filled with interference-fit pin, was instrumented by strain gauges and statically tested in order to check FEM results. A very good correlation existed between measured and numerically evaluated strains. FEM results demonstrated the well-known effect of interference-fit fasteners on reducing stress ranges. By increasing the interference level, the stress range was practically unchanged, while the mean stress decreased. Interference-fit produces a biaxial stress state, which must be taken into account for fatigue evaluation. In the present case, a simple criterion, based on hoop strain, predicted the fatigue results quite well with the exception of open hole fatigue test results, which were overestimated.

An experimental study was performed in order to determine the influence of the sequence of operations on the effectiveness of Laser Shock Peening (LSP) treatment in increasing the fatigue performances of open-hole aluminium specimens. Residual stress measurements, fractographic analysis and FEM analysis were performed, indicating the presence of compressive residual stresses on the surface of the treated specimens and tensile residual stresses in the mid-section along the thickness of the specimens. Negative effects on fatigue lives were encountered on the specimens with the hole already present, while positive effect were observed in specimens in which the hole was drilled after LSP treatment. These results indicate that LSP can be a good solution for “in production” application, in which open holes are to be drilled after the LSP treatment. The application in which LSP is used “in service” on structures with pre-existing cut-outs, has proven to be impracticable in the investigated configuration.

Fatigue crack propagation tests of through cracks in 3 mm thick aluminium alloy 6013-T6 plates, under combined membrane and bending stresses were carried out. Five different values for the ratio of bending to membrane stresses, SB, were examined: 0, 0.55, 1.25, 1.8 and 2.23. Firstly, the results were elaborated by taking into account only the membrane stress and the front dimension of the crack for the evaluation of the stress intensity factor range. The results relevant to the lower SB values, evaluated in this mode, show good agreement with the results obtained when only the membrane stress was applied, while the results obtained at the higher SB values exhibit a remarkable increase in the crack propagation rate. The results were successively evaluated on the basis of tabular stress intensity factor solutions available in the literature; the agreement between the predicted and the experimental data occurs when mean values of stress intensity factor, rather than local values, were used at the back face of the plate, i.e., the face where the bending produces a compressive stress, to counteract the high gradient of stress intensity factor present in this area.

Comparative fatigue tests were carried out on Friction Stir Welded specimens of a 2195-T8 aluminum–lithium alloy that differed significantly in width. The width of the larger specimens was over thirteen times greater than that of the small specimens. Fatigue results showed a clear “size effect”, i.e. fatigue life of large specimens was about 40% of the corresponding value of small specimens. The Equivalent Initial Flaw Size methodology was adopted to correlate the two sets of results. Fatigue crack initiation life was disregarded with respect to crack propagation life, and fatigue life was evaluated only as propagation of a small pre-existing defect. Following this methodology, test results of small specimens were used to evaluate the initial equivalent flaw contained in each specimen. It was assumed that this data followed a normal distribution. The equivalent initial flaw in larger specimens was evaluated by simple geometrical considerations. A very good assessment of mean fatigue life and scatter in the fatigue results of large specimens was obtained by simulating the propagation of these defects. Calculations were carried out by taking also welding residual stresses into account, but the results demonstrated that this effect was not significant.

The results of fatigue and crack propagation tests carried out on dog-bone specimens made of 2024-T3 are described. Two types of specimens were investigated: the first was machined from a 1.27-mm-thick sheet, while the second was machined from a bonded metal-laminated sheet, made of four 0.3-mm-thick layers. Crack propagation tests confirmed the high resistance of metal-laminated sheets to the propagation of fatigue cracks, compared to monolithic sheets, once again. At the same time, standard fatigue tests, carried out up to the final failure of the specimens, demonstrated a comparable fatigue resistance of monolithic and laminated specimens. As a consequence, it can be concluded that fatigue cracks nucleated earlier in the metal-laminated specimens, compared to the monolithic ones, but propagated more slowly. This behaviour was attributed to the presence of sharp edges in the inner laminas of metal-laminated materials which cannot be eliminated by deburring. Additional tests were carried out on monolithic specimens containing burrs and sharp edges at the holes. These specimens were drilled and reamed after stacking and pressing them to form a package. The specimens were fatigue tested without deburring the holes. A decrease in the fatigue resistance was observed. The formation of burrs and sharp edges was additionally promoted by inserting plastic foils between the specimens during the machining operations. Fatigue resistance of these specimens is progressively lower.