Advanced Materials Research Vols. 891-892

Abstract: Plane bending fatigue testing was performed to study the fatigue properties of friction stir welded (FSW) 3 mm thick AA6061-T6 aluminum alloy plates. Fatigue cracks propagated with bends and curves on the specimens, showing large deviation from a linear line. This might be reflecting the material flow and microstructure in the weld zone. The fatigue strength of the unwelded base material (BM) was 110 MPa at 10⁷ cycles and FSW deteriorated it to 90 MPa. However, laser peening (LP) restored the degraded fatigue strength up to 120 MPa which is higher than that of the BM.

Abstract: Laser Shock Peening (LSP) is a material enhancement process used to introduce compressive residual stresses in metallic components. This investigation explored the effects of different combinations of LSP parameters, such as irradiance (GW/cm²) and laser pulse density (spots/mm²), on 3.2 mm thick AA6056-T4 samples, for integral airframe applications. The most significant effects that are introduced by LSP without a protective coating include residual stress and surface roughness, since each laser pulse vaporizes the surface layer of the target. Each of these effects was quantified, whereby residual stress analysis was performed using X-ray diffraction with synchrotron radiation. A series of fully reversed bending fatigue tests was conducted, in order to evaluate fatigue performance enhancements with the aim of identifying LSP parameter influence. Improvement in fatigue life was demonstrated, and failure of samples at the boundary of the LSP treatment was attributed to a balancing tensile residual stress.

Abstract: This paper reports the effectiveness of residual stress fields induced by laser shock peening (LSP) to recover pristine fatigue life. Scratches 50 and 150 μm deep with 5 μm root radii were introduced into samples of 2024-T351 aluminium sheet 2 mm thick using a diamond tipped tool. LSP was applied along the scratch in a band 5 mm wide. Residual stress fields induced were measured using incremental hole drilling. Compressive residual stress at the surface was-78 MPa increasing to-204 MPa at a depth of 220 μm. Fatigue tests were performed on peened, unpeened, pristine and scribed samples. Scratches reduced fatigue lives by factors up to 22 and LSP restored 74% of pristine life. Unpeened samples fractured at the scratches however peened samples did not fracture at the scratches but instead on the untreated rear face of the samples. Crack initiation still occurred at the root of the scribes on or close to the first load cycle in both peened and unpeened samples. In peened samples the crack at the root of the scribe did not progress to failure, suggesting that residual stress did not affect initiation behaviour but instead FCGR. A residual stress model is presented to predict crack behaviour in peened samples.

Abstract: Highly loaded aircraft components have to fulfill strict fatigue and damage tolerance requirements. For some components besides the crack initiation mainly the fatigue crack propagation behavior is the main design criteria. To improve the crack propagation behavior of a component several methods are known or have been described in literature. For thin aircraft panels i.e. the application of crenellations [1] or bonded doublers [2, 3] can be a solution. For thick structures mainly the introduction of compressive residual stresses is beneficial. In this paper the potential of compressive residual stresses obtained by Laser Shock Peening (LSP) and Shot Peening (SP) is investigated. By means of Laser Shock Peening the residual compressive stress field can extend much deeper below the treated surface than that produced by conventional Shot Peening (i.e. with steel or ceramic balls) [4, 5]. The effect of such deep compressive stress profile results in a significantly higher benefit in fatigue behavior after Laser Shock Peening or after the combination of Laser Shock Peening and Shot Peening on top. The measurement of residual stresses as a depth profile has been performed by incremental hole drilling (ICHD) and contour method. Finally crack propagation tests have been carried out to validate the process technology approach.

Abstract: This paper will provide an overview on potential applications for the aerospace industry for repairing aircraft as well as to ensure salvage for identified hot spots in terms of fatigue and crack growth performance. Residual stress engineering is a field of engineering aiming to improve the economic and ecological impact of future aircraft structures by controlling the residual stresses induced by Laser Shock Peening (LSP). Managing the residual stresses for designing structures represents an innovative approach for next generation aircraft. Predicting crack turning induced via a LSP treatment and the optimization of the LSP treatment itself for reaching the crack growth design stress for the targeted weight benefit will be discussed. Advanced forming processes in aircraft manufacturing represent another potential area of interest and the benefits and challenges of applying laser peen forming in this context will be presented. The aeronautical industry requirements for future developments of the laser shock process will also be included for applications ranging from the repair environment to design and manufacturing of aircraft structures.

Abstract: Fatigue crack growth and propagation analysis in welded joints have to deal with the complexity of modeling multiple weld toe surface cracks originating from weld toes. Fitness-For-Service (FFS) assessments for weld toe surface cracks employ a fracture mechanics and Paris Law approach to predict the fatigue crack propagation life of a semi-elliptical surface crack (SESC) to failure. A fatigue crack growth algorithm for assessing multiple surface crack growth, coalescence and propagation life was initially validated with previuously report crack growth data for a fillet shoulder specimen. Next a parametric study for single, double, and triple SESCs located along the weld toe line of a fillet weld was investigated with three starting crack depth sizes (0.1mm, 0.5mm, 1.0mm) coupled with three different crack aspect ratios (a/c = 1.0, a/c = 0.5 and 0.25) giving a total of 27 cases studied.

Abstract: Recently a new methodology was developed for automated fatigue crack growth (FCG) life analysis of components based on finite element stress models, weight function stress intensity factor solutions, and algorithms to define idealized fracture geometry models. This paper describes how the new methodology is being used to integrate FCG analysis into highly automated design assessments of component life and reliability. In one application, the FCG model automation is supporting automated calculation of fracture risk due to inherent material anomalies that can occur anywhere in the volume of the component. Automated schemes were developed to divide the component into a computationally optimum number of sub-volumes with similar life and risk values to determine total component reliability accurately and efficiently. In another application, the FCG model automation is supporting integration of FCG life calculations with manufacturing process simulation to perform integrated computational materials engineering. Calculation of full-field, location-specific residual stresses or microstructure is being linked directly with automated life analysis to determine the impact of manufacturing parameters on component reliability.

Abstract: Arbitrary, non-planar progressive fracture analysis is of critical importance to the integrity of structures. While significant progress has been made in the last 25 years, there are still technical issues regarding computational expense, robustness, and accuracy. Based upon the recent significant enhancements available through the use of multicomplex finite element methods, a new high-order progressive strain energy based progressive crack growth algorithm is proposed.

Abstract: In cyclic stress and strain controlled tests, the fatigue life was determined for twin roll cast AM50 magnesium alloy sheet metals. Different stress and strain ratios were applied to cover the technical relevant load range. The investigated sheet metal shows a strong basal texture which leads to twinning in compression parallel to the sheet plane direction. Accordingly, at low cycles to failure N_f the cyclic plastic deformation is dominated by twinning and detwinning. In this study an energy based approach for lifetime estimation is presented and compared with the Smith-Watson-Topper damage model as well as with experimental test results. The energy based approach shows a comparatively high correlation at all investigated stress and strain ratios in a wide cycle range (20< N_f <1.000.000) and is adequate for the lifetime prediction at various mean stresses.

Abstract: This paper discusses an evaluation method of creep-fatigue lives of YH61 single crystal superalloy under multiaxial loading at high temperature. Three types of creep-fatigue tests were performed using three types of the single crystalsuperalloy specimens at 1173K. They were push-pull tests using solid bar specimens, tension-torsion tests using hollow cylinder specimens and biaxial tension-compression tests using cruciform specimens. Anisotropic strain and Mises stress in combination with frequency modified fatigue equation were applied for evaluating the creep-fatigue lives in the three types of tests. The former parameter gave a relatively large scatter but the latter parameter a small scatter in the correlation.