13:30   Session 7: Fatigue crack growth and life prediction methods  III Chair: Carlos E. Chaves 13:30 20 mins Fracture mechanics-based approach for anomaly size acceptability of additively manufactured metals Simone Romano, Andrew Perry, Francesco Sausto, Apostolos Karafillis Abstract: The traditional approach for quantifying material capability generally assumes that material properties in components are fully represented by test bar results. However, this assumption might not always be true for complex additively manufactured components, which may contain anomaly populations different than those expected based on relatively simple-geometry mechanical test coupons. In this work, we present a possible fracture mechanics-based approach for definition of anomaly size acceptability limits for AM materials, to augment the traditional building block of fatigue test-based material capability assessments. The fracture mechanics analysis is validated against fractography observations of tested bars and with fatigue tests on bars containing artificially notched coupons. 13:50 20 mins Fatigue crack growth on several materials under single-spike overloads and aircraft spectra James Newman, Kevin Walker Abstract: In the mid-1960’s, the phenomenon of flat-to slant crack growth was studied by many in the aircraft industry. At low stress-intensity factors, a crack surface is flat, and the behavior is referred to as the tensile mode. The stress state in the crack-front region is under plane-strain conditions (high constraint). As the crack grows with higher stress-intensity factors, a 45-degree shear lip starts to develop at the intersection of the crack front and free surfaces. With further crack extension, a complete shear failure occurs through the thickness of the sheet or plate. This behavior is the shear mode, which is under low constraint or plane-stress conditions. In 1965, Schijve found that the transition from flat-to-slant crack growth on a 2024-T3 aluminum alloy over a wide range in stress ratios (R) occurred at a constant crack-growth rate. Hudson and Newman also showed the same behavior on 7075-T6 and Ti-6Al-4V alloys. The materials considered herein are 2024-T3, 7075-T6 and 9310 steel. Four areas of research are presented: (1) constraint loss during plane-strain to plane-stress crack-growth behavior, (2) fracture behavior, (3) single-spike overload-underload behavior, and (4) simulated aircraft spectrum loading. The FASTRAN crack-closure based life-prediction code was used to correlate the constant-amplitude crack-growth-rate data over a wide range in stress ratios (R) and rates from threshold to fracture, and to calculate or predict the crack-growth behavior on the single-spike overload-underload tests. Crack-closure behavior is strongly dependent upon the level of constraint. The main objectives were to see if the constraint-loss region can be experimentally measured and whether constraint-loss behavior is the primary reason for crack-growth delays after single-spike overloads. If the current crack-growth models cannot predict the delay cycles from a simple spike overload, then questions arise about their accuracy under more complex aircraft spectrum loading. Tests were also conducted on the two aluminum alloys under the Mini-TWIST (standard European) transport wing spectrum. Crack-growth analyses using crack-closure theory without constraint loss was unable to predict crack growth under spike overloads or simulated aircraft spectra. However, predicted crack length against cycles with constraint-loss behavior compared well with all tests. 14:10 20 mins Effects of the secondary aging heat treatment T6I4 on fracture toughness and fatigue crack growth resistance of AA7050 alloy Carlos Antonio Baptista, Ana Marcia Antunes, Luis Rogério de Oliveira Hein, Samara Cristiny de Paula, Gabriel Cotrim de Cesare Peinado Abstract: The tensile response and the fatigue behavior of aircraft age-hardenable aluminum alloys are strongly influenced by factors like the size, spacing and volume fraction of strengthening precipitates. Recent papers have pointed out that the secondary aging T6I4, by producing a higher density of fine strengthening precipitates as compared to the conventional T7451 temper, can improve the ability of AA 7050 alloy to accommodate plastic strain during loading and promote a more uniform strain distribution. In this sense, the present work had as main goal to evaluate the effects of T6I4 aging on the fracture toughness and fatigue crack growth behavior of AA 7050 plate samples. The alloy was previously hot rolled at approximately 400°C to 75 mm thick plate and received in the commercial T7451 temper condition. Its chemical composition was determined by optical emission spectrometry. The T6I4 condition was obtained from the as-received material by means of a solution treatment (486C / 4h) followed by two-step ageing (145C / 30 min and 65C / 24h). The fracture toughness of the material in both T7451 and T6I4 conditions was determined using the Chevron notch methodology according to ASTM E1304-97(2020) standard. The fatigue crack growth tests were conducted under constant amplitude (R = 0.1 and 0.5) and variable amplitude loadings. In the former case, C(T) notched specimens in accordance to ASTM E647-15e1 standard were employed and the crack length was measured by the compliance method. In the latter case, the Mini FALSTAFF spectrum loading of dog-bone test specimens was adopted and the crack length was measured by quantitative fractography according to the method developed by DSTG (Australia). The obtained results showed an expressive increase of the materials’ KQvM values for the T6I4 condition, as well as a significant decrease on the crack growth rate, especially during the early stages of fatigue at lower constant amplitude ratio, possibly due to a more pronounced crack closure. The specimens treated to T6I4 condition and tested under spectrum loading also presented fatigue lives greater than twice the fatigue life presented by the T7451 material condition. 14:30 20 mins Low-cycle fatigue property and life prediction model of TC18 titanium alloy material Shaopu Su, Binwen Wang, Xianmin Chen, Lei Li, Dengke Dong Abstract: As alpha+ beta type titanium alloy, TC18 is widely applied in the landing gear and flap slides with the advantages of high strength, excellent plasticity and weldability. Yet more service datum show that low-cycle fatigue failure sometimes occurs in the blocking hook of carrier-based aircraft due to its serious servicing environment. Consequently low-cycle fatigue performance plays an important role in the fatigue life evaluation of aircraft load-bearing structural parts. Through the low-cycle fatigue experiments of standard fatigue coupons, the static and low-cycle fatigue behaviours of TC18 titanium alloy were obtained, such as the stress-strain constitutive relationship, the cyclic stress-strain curve and the strain-life curve. In order to meet the engineering fatigue life evaluation for TC18, on the one hand, based on Manson-Coffin and reliability analysis theory, the strain life model was built and the low-cycle fatigue life curve under a certain survival rate was observed to investigate the life at certain confidence and reliability; on the other hand, combined with the damage mechanical theory, the gradual degradation process of the mechanical properties of TC18 titanium alloy with the increase of load times was described by the Lemaitre plastic damage model, and the fatigue damage accumulation parameters were determined based on experimental results. We obtain the conclusion as following:1) the cyclic behaviour of TC18 titanium alloy material is a strain cycle softening relationship under different strain-rate loading;2) its low-cycle fatigue fracture form is described that: the crack initiated in the form of quasi-cleavage fracture, and propagated by quasi-cleavage fracture or mixed-type fracture of quasi-cleavage and dimple, finally sheared at the edge of the specimen to form a shear dimple fracture ;3) Based on the experimental results, the engineering Manson-Coffin strain life model and macro-cyclic plastic damage theory model in the paper can be used for the low-cycle fatigue life evaluation of TC18 titanium alloy material. 14:50 20 mins Measuring small fatigue crack growth with the aid of marker bands in recrystallized annealed TI6AL4V Ingrid Kongshavn, Simon Barter, Larissa Sorensen Abstract: Recrystallized annealed Ti6Al4V (RA) has a small grain size typically with equiaxed alpha surrounded by remnant beta. This material and microstructure is used in a large number of structural parts for aircraft, and now, additively manufactured parts that may have similar microstructures. In all cases, such structures need to be durable and resist fatigue damage. One aspect of this type of microstructure is that fatigue cracks can remain very small for a long part of a component's life. This may not be mostly due to nucleation, rather, under variable amplitude spectra, loaded cracks tend to nucleate quickly and grow very slowly for much of the time. For this and other reasons, it is important to understand the growth of small natural fatigue cracks that may occur in aircraft critical structures. A known method to measure crack growth is through the addition of loading cycles to variable amplitude spectra for quantitative fractographic analysis. To some extent, the marker loads chosen depend on the requirements of the test and material being tested. While aluminium alloys are relatively easy to mark, the fine-grained Ti6Al4V alloys are difficult to mark at small crack sizes. Typical markers may be over- and/or underloads, high loads, higher than the largest spectrum load, or constant amplitude blocks with varying mean and/or maximum stresses, etc. Many test programs containing such markers for aluminium alloys can be found in the literature. However, examples for Ti6Al4V RA under spectrum loading are scarce, due to the fine mixed-phase microstructure that does not lend itself well to marking from nucleation to fracture. It is with this intent that specific markers have been developed for a coupon truncation study using Ti6Al4V RA, with a focus on understanding small to long crack growth in this material. In this paper, the process used to develop these markers and their effectiveness; the results of crack growth curves measured for a wing root bending moment spectrum; and, some observations on the nature of cracking over the first 250 microns (‘crack initiation life’) will be discussed, with comparisons to cracks produced in another common Ti6Al4V form: beta annealed material.

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