Titanium (Ti) alloys have been commonly used in the aerospace industry, not only because they have a high strength-to-weight ratio (comparing to the steels) but also their satisfactory corrosion resistance. Furthermore, they can be assembled with the carbon fibre composite parts. However, conventional manufacturing methods cause high material scrap rate and require lots of machining to obtain the final shape and size, which increases both the manufacturing time and cost. In order to improve the efficiency and reduce the cost of Ti parts, Additive Manufacturing (AM) has been developed. Rolled Wire and Arc Additive Manufacturing (rolled WAAM) is one of the AM processes. The main characteristics of this technology is the reduced β grain size to refine the alloy's microstructure. Both the ultimate tensile strength and yield strength of Ti alloy made by rolled WAAM are at least 10% higher than traditional wrought Ti. This project is to investigate the fatigue crack growth rates of the Ti-6Al-4V built by rolled WAAM process in both the longitudinal and transverse orientations to study the effect of rolling on fatigue crack growth rate of WAAM processed Ti. The project was carried out by testing the fatigue crack growth rates for 4 compact tension specimens. The test results of different orientations were compared with each other, and scatters in fatigue life and fatigue crack growth rate were found. Fatigue crack growth rate is lower in the longitudinal specimens. The results are also compared with those of the unrolled WAAM specimens tested in a previous project. It was found that rolling can significantly improve the fatigue crack growth behaviour in WAAM processed Ti, and can reduce the difference between the two orientations, i.e. achieving better isotropic material properties. Recorded scatters may be caused by the process induced residual stresses, error in measurement, and the test machine load range being much higher than the applied loads. More specimens can be tested to validate above observations further.

Cette thèse porte sur la fissuration par fatigue sous chargement variable en mode mixte I+II+III et sur les effets d'histoire induits par la plasticité confinée et les contraintes internes. Les essais réalisés montrent qu’il existe des effets antagonistes d’histoire du chargement à courte distance et à longue distance, que la forme de la séquence de chargement est cruciale et, par comparaison, que les effets de contact et de frottement sont d’importance moindre. Les contraintes internes jouent un rôle majeur sur la vitesse de fissuration par fatigue et sur le mode de fissuration. Une démarche a été mise en place pour étudier le comportement élasto-plastique d’une section représentative du front de la fissure par éléments finis. Pour ne retenir des calculs réalisés que le minimum d’information nécessaire, on se donne une approximation cinématique, le champ de vitesse est partitionné en composantes de modes I, II, III élastiques et plastiques, chaque composante étant caractérisée par un facteur d’intensité et une distribution spatiale fixe. Les calculs réalisés ont permis de sélectionner 7 trajets de chargement différents en mode I+II et en mode I+II+III, qui présentent les mêmes amplitudes pour chaque mode, les mêmes maxima, minima et valeurs moyennes. Ces trajets, censés être équivalents au sens des critères de rupture usuels, ne le sont pourtant pas lorsqu’on considère le comportement élasto-plastique du matériau et conduisent à des vitesses et des trajets de fissuration expérimentaux très différents les uns des autres. Les simulations numériques et la modélisation simplifiée sont en accord avec les résultats expérimentaux. Les calculs réalisés ont également permis de discuter le rôle des phases de chargement en mode III sur la fissuration. Le comportement du matériau étant non-linéaire, la direction du chargement nominal ne coïncide pas nécessairement avec celle de l’écoulement plastique. Ajouter une phase de chargement en mode III peut, dans certains cas, modifier très significativement le comportement de la fissure (direction de propagation, vitesse de fissuration, écoulement plastique). / This thesis deals with fatigue crack growth in non-proportional variable amplitude mixed mode I + II + III loading conditions and analyses the effects of internal stresses stemming from the confinement of the plastic zone in small scale yielding conditions. The tests showed that there are antagonistic long-distance and short-distance effects of the loading history on fatigue crack growth. The shape of loading path, and not only the maximum and minimum values in this path, is crucial and, by comparison, the effects of contact and friction are of lesser importance. Internal stresses play a major role on the fatigue crack growth rate and on the crack path. An approach was developed to analyze the elastic-plastic behavior of a representative section of the crack front using the FEA. A model reduction technic is used to extract the relevant information from the FE results. To do so, the velocity field is partitioned into mode I, II, III elastic and plastic components, each component being characterized by an intensity factor and a fixed spatial distribution. The calculations were used to select seven loading paths in I + II and I + II + III mixed mode conditions, which all have the same amplitudes for each mode, the same maximum, minimum and average values. These paths are supposed to be equivalent in the sense of common failure criteria, but differ significantly when the elastic-plastic behavior of the material is accounted for. The results of finite element simulations and of simulations using a simplified model proposed in this thesis are both in agreement with experimental results. The approach was also used to discuss the role of mode III loading steps. Since the material behavior is nonlinear, the nominal loading direction does not coïncide with the plastic flow direction. Adding a mode III loading step in a mode I+II fatigue cycle, may, in some cases, significantly modify the behaviour of the crack (crack growth rate, crack path and plasti flow).

Friction Stir Welding (FSW) is a new solid-state welding process that shows great promise for use in the aerospace and transportation industries. One of the primary benefits of this process is that mechanical properties of the base material are not as severely degraded as they are with conventional fusion welding. However, fatigue crack initiation and growth properties of the resulting weld nugget are not fully understood at this time. The primary goal of this project is to characterize the fatigue crack growth properties of friction stir welds in 6061-T6 aluminum as relates to the microstructural evolution of the weld. This was accomplished by producing friction stir welds and testing fatigue crack growth response in different crack orientations with respect to the weld. In addition, residual stress measurements were conducted for all cases, using both the crack compliance and contour methods. The results from the methods were compared in order to evaluate the accuracy of each method. Being an immature technology, the potential for discovery of new applications for the FSW process exist. With this in mind, novel applications of the FSW process, including the addition of particles during welding were explored. The first step was the investigation of property changes that occur when secondary cast phases are refined using the FSW process. The FSW process successfully refined all secondary phases in A380 and A356, producing an increase in hardness. Next, methods for the creation of particle metal matrix composites using FSW will be investigated. Nano-scale alumina particles were successfully added to the matrix and homogenously distributed. Using multiple weld passes through the composite was found to increase the uniformity of particle distribution. However, the alumina particle composite failed to provide any statistically significant hardness increase over the base material. The FSW process was also evaluated for weldability of traditionally difficult alloy systems. FSW was found to show very good weldability for dissimilar cast and wrought alloys, as well as for high-pressure die castings. Lastly, the feasibility of friction stir welding/processing in repairing crack defects in complex structural members in combination with cold-spray technology was determined. Friction Stir processing was used on a cold spray 6061-T6 block, resulting in significant increases in hardness over the base material, as well as a reduction in porosity. In addition, FSP was shown to eliminate crack-type defects in cold spray materials, a finding that has important applications in part repair. The deliverables of this work include an understanding of the fatigue crack growth response of FSW/FSP 6061-T6, as well as a feasibility study exploring novel uses for the FSW/FSP process. In addition, the deliverables include CNC code, fixtures, procedures, and analytical code for the creation and analysis of FSW/FSP joints. This will be important for the continuation of FSW/FSP work at WPI.

In an attempt to find a way of improving the damage tolerance of composite bonded repairs to metallic aircraft structures, the effect of using conventional crack growth retardation techniques in conjunction with bonded repairs was experimentally investigated. Hence, an experimental test program was set up to determine whether fatigue crack growth under bonded repairs is retarded further by giving the crack to be repaired a crack growth retardation treatment prior to repair patch application. In addition, it was set up to determine the influence of a bonded repair on the effectiveness of a crack growth retardation method. Centrally cracked aluminium plates were used. Stop drilling followed by cold hole expansion and the application of single overloads were selected as retardation treatments. Two patch materials were considered; boron/epoxy and Glare 2. Further test variables were the aluminium alloy and the plate thickness. Fatigue testing was carried out under constant amplitude loading and baseline results were determined first. In addition to optically monitoring the crack growth, local and global out-of-plane deformations were visualised with holographic interferometry and shadow moire??. Furthermore, the stress intensity factors under the repair patch were examined with strain gauges and measurement of the central crack opening displacement. Disbonds and fracture surfaces were studied after residual strength tests. The crack growth results obtained showed that retardation treatments decrease crack growth rates under a repair patch and that the effectiveness of a retardation treatment is increased by the patch. Although identical crack growth rates were observed under boron/epoxy and Glare 2 patches, the reinitiation period after the retardation treatment lasted longer when Glare 2 patches were applied. Analytical predictions of the extent of retardation based on existing models showed that the conjunctive effect of retardation treatments and bonded repairs was underestimated. A sustained reduction in crack growth rates was observed under bonded repairs with a prior overload retardation treatment. It was concluded that the damage tolerance of bonded repairs is increased by the application of a crack growth retardation treatment because the crack growth is retarded further. These findings indicate that the range of cracks in aircraft for which bonded repairs can be considered is expanded and that economic benefits can be obtained.

Confront with the serious aging problem in aircraft structure field, the profession was tasked to unveil the mysterious in the mechanism of aging. In decades, many endeavours were put into different subjects such as, fatigue and crack calculation, corrosion analysis, reliability evaluation, life prediction, structure monitor and protection, structure repair, etc. In an effort of developing a reasonable model for life prediction and reliability evaluation, a wide range of topics in the field of aging structure reliability are reviewed. Many existing methods and tools are carefully studied to distinguish the advantages, disadvantages and the special application. With consideration of corrosion fatigue life, and based on the data obtained through investigating service status of the aging aircraft, a fuzzy reliability approach is proposed and presented. Initially, the thesis presents the literature review in the field, introducing the well-established theories and analysis tools of reliability and points out how such these methods can be used to assess the life and reliability of aging structure. Meanwhile, some characteristic parameters and distributions, as well as some crucial calculation formulations, procedures for aging aircraft reliability/risk analysis are given. Secondly, mathematical models are established to evaluate the initial crack size and to assess both randomness and fuzziness of the variables, which also successfully work out the probability of survival of existing structures over a time period and predict the operation time under specific reliability requirement. As a practical approach to the reliability of aging aircraft structure, example is presented and evaluated. While conduct the calculation, a few programs based on FORTRAN code are developed to solve the none-linear equation, to work out the multi dimension integration and to simulate the survival probability. The crack life prediction software AFGROW is selected for comparison of the calculation results, which also shows the appropriate accuracy of the established model. As conclusion, the effects of some variables including fuzzy factors on reliability and life of aging aircraft structure are finally discussed. It is apparent that the confines of the model are existing as fact because of the huge assumption of the parameters input and model uncertainties. Suggestions on further prospective research are proposed respectively.

It is generally accepted that the fatigue crack growth depends on the stress intensity factor range (ΔK) and the maximum stress intensity factor (K_max). Numerous driving forces were introduced to analyze fatigue crack growth for a wide range of stress ratios. However, it appears that the effect of the crack tip stresses and strains need to be included into the fatigue crack growth analysis as well. Such an approach can be successful as long as the stress intensity factors are correlated with the actual elastic-plastic crack tip stress-strain field. Unfortunately, the correlation between the stress intensity factors and the crack tip stress-strain field is often altered by residual stresses induced by reversed plastic deformations. A two-parameter model (ΔK_tot, K_max,tot) based on the elastic-plastic crack tip stress-strain history has been proposed. The applied stress intensity factors (ΔK_appl, K_max,appl) were modified and converted into the total stress intensity factors (ΔK_tot, K_max,tot) in order to account for the effect of local crack tip stresses and strains on the fatigue crack growth. The fatigue crack growth was regarded as a process of successive crack re-initiations in the crack tip region and predicted by simulating the stress-strain response in the material volume adjacent to the crack tip and estimating the accumulated fatigue damage. The model was developed to predict the mean stress effect for steady-state fatigue crack growth and to determine the fatigue crack growth under simple variable amplitude loading histories. Moreover, the influence of the applied compressive stress on fatigue crack growth can be explained with the proposed two-parameter model. A two-parameter driving force in the form of: Δκ = K_max,tot^p ΔK_tot^(1-p) was derived based on the local stresses and strains at the crack tip using the Smith-Watson-Topper (SWT) fatigue damage parameter: D = σ_maxΔε/2. The parameter p is a function of material cyclic stress-strain properties and varies from 0 to 0.5 depending on the fatigue crack growth rate. The effects of the internal (residual) stress induced by the reversed cyclic plasticity manifested themselves in the change of the resultant (total) stress intensity factors driving the crack. Experimental fatigue crack growth data sets for two aluminum alloys (7075-T6 and 2024-T351), two steel alloys (4340 and 4140), and one titanium alloy (Ti-6Al-4V) were used for the verification of the model under constant amplitude loading. This model was also capable of predicting variable-amplitude fatigue crack growth. Experimental fatigue crack growth data sets after single overloads for the aluminum alloy 7075-T6, steel alloy 4140, and titanium alloy Ti-6Al-4V were also used for the verification of the model. The results indicate that the driving force Δκ can successfully predict the stress ratio R effect and also the load-interaction effect on fatigue crack growth.

It is generally accepted that the fatigue crack growth depends on the stress intensity factor range (ΔK) and the maximum stress intensity factor (K_max). Numerous driving forces were introduced to analyze fatigue crack growth for a wide range of stress ratios. However, it appears that the effect of the crack tip stresses and strains need to be included into the fatigue crack growth analysis as well. Such an approach can be successful as long as the stress intensity factors are correlated with the actual elastic-plastic crack tip stress-strain field. Unfortunately, the correlation between the stress intensity factors and the crack tip stress-strain field is often altered by residual stresses induced by reversed plastic deformations. A two-parameter model (ΔK_tot, K_max,tot) based on the elastic-plastic crack tip stress-strain history has been proposed. The applied stress intensity factors (ΔK_appl, K_max,appl) were modified and converted into the total stress intensity factors (ΔK_tot, K_max,tot) in order to account for the effect of local crack tip stresses and strains on the fatigue crack growth. The fatigue crack growth was regarded as a process of successive crack re-initiations in the crack tip region and predicted by simulating the stress-strain response in the material volume adjacent to the crack tip and estimating the accumulated fatigue damage. The model was developed to predict the mean stress effect for steady-state fatigue crack growth and to determine the fatigue crack growth under simple variable amplitude loading histories. Moreover, the influence of the applied compressive stress on fatigue crack growth can be explained with the proposed two-parameter model. A two-parameter driving force in the form of: Δκ = K_max,tot^p ΔK_tot^(1-p) was derived based on the local stresses and strains at the crack tip using the Smith-Watson-Topper (SWT) fatigue damage parameter: D = σ_maxΔε/2. The parameter p is a function of material cyclic stress-strain properties and varies from 0 to 0.5 depending on the fatigue crack growth rate. The effects of the internal (residual) stress induced by the reversed cyclic plasticity manifested themselves in the change of the resultant (total) stress intensity factors driving the crack. Experimental fatigue crack growth data sets for two aluminum alloys (7075-T6 and 2024-T351), two steel alloys (4340 and 4140), and one titanium alloy (Ti-6Al-4V) were used for the verification of the model under constant amplitude loading. This model was also capable of predicting variable-amplitude fatigue crack growth. Experimental fatigue crack growth data sets after single overloads for the aluminum alloy 7075-T6, steel alloy 4140, and titanium alloy Ti-6Al-4V were also used for the verification of the model. The results indicate that the driving force Δκ can successfully predict the stress ratio R effect and also the load-interaction effect on fatigue crack growth.

Thermally cycling a thin plate of nickel-based superalloy with an intense in-plane thermal gradient, or hot spot, produces thermally induced crack growth not represented by classic thermo-mechanical fatigue (TMF). With the max hot spot temperature at 1093 C (2000 F) of a 1.5 mm thick, 82.55 mm diameter circular plate of B-1900+Hf, annular buckling and bending stresses result during each thermal cycle which drive the crack initiation and propagation. A finite element analysis (FEA) model, using ANSYS 7.1, has been developed which models the buckling and as well as represents the stress intensity at simulated crack lengths upon cool down of each thermal cycle. The model approximates the out-of-plane response at heat-up within 5% error and a difference in the final displacement of 0.185 mm after twelve thermal cycles. Using published da/dN vs. Keff data, the number of cycles needed to grow the crack to the experimental arrest distance is modeled within 1 mm. The number of cycles to this point is within 5 out of 462 in comparison to the experimental test.

In an attempt to find a way of improving the damage tolerance of composite bonded repairs to metallic aircraft structures, the effect of using conventional crack growth retardation techniques in conjunction with bonded repairs was experimentally investigated. Hence, an experimental test program was set up to determine whether fatigue crack growth under bonded repairs is retarded further by giving the crack to be repaired a crack growth retardation treatment prior to repair patch application. In addition, it was set up to determine the influence of a bonded repair on the effectiveness of a crack growth retardation method. Centrally cracked aluminium plates were used. Stop drilling followed by cold hole expansion and the application of single overloads were selected as retardation treatments. Two patch materials were considered; boron/epoxy and Glare 2. Further test variables were the aluminium alloy and the plate thickness. Fatigue testing was carried out under constant amplitude loading and baseline results were determined first. In addition to optically monitoring the crack growth, local and global out-of-plane deformations were visualised with holographic interferometry and shadow moire??. Furthermore, the stress intensity factors under the repair patch were examined with strain gauges and measurement of the central crack opening displacement. Disbonds and fracture surfaces were studied after residual strength tests. The crack growth results obtained showed that retardation treatments decrease crack growth rates under a repair patch and that the effectiveness of a retardation treatment is increased by the patch. Although identical crack growth rates were observed under boron/epoxy and Glare 2 patches, the reinitiation period after the retardation treatment lasted longer when Glare 2 patches were applied. Analytical predictions of the extent of retardation based on existing models showed that the conjunctive effect of retardation treatments and bonded repairs was underestimated. A sustained reduction in crack growth rates was observed under bonded repairs with a prior overload retardation treatment. It was concluded that the damage tolerance of bonded repairs is increased by the application of a crack growth retardation treatment because the crack growth is retarded further. These findings indicate that the range of cracks in aircraft for which bonded repairs can be considered is expanded and that economic benefits can be obtained.

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