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.

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.

No description available.

<div>An important challenge in predicting fatigue and creep crack growth is describing crack growth rates under transient conditions. Transient conditions occur when similitude is violated at the crack tip due to the applied loads or material behavior. Crack growth models like the Paris law, valid for homogeneous materials under constant-amplitude cyclic loading or sustained loading, no longer apply. Transient crack growth rates are strongly influenced by changes in plastic deformation at the crack tip. Activation of time-dependent damage and viscoplastic deformation at high temperatures further complicates the problem.</div><div><br></div><div>This thesis advances knowledge and predictive capabilities for transient creep and fatigue crack growth in metals, with specific applications to two technologically-relevant nickel-base superalloys. Finite element computations of crack growth following overloads and in multilayered materials are conducted. Crack extension is an outcome of the boundary value problem through an irreversible cohesive zone model and its interaction with plasticity and viscoplasticity in the bulk material.</div><div><br></div><div>First, fatigue crack growth in rate-independent materials is analyzed. The plasticity formulation considers both plastic strain and gradients of plastic strain, which produce hardening beyond that predicted by classical plasticity models. The computations demonstrate that hardening due to plastic strain gradients plays a significant role in transient fatigue crack growth following overloads. Fatigue crack growth transients associated with material inhomogeneity are studied through the case of a crack growing toward interfaces between plastically dissimilar materials. Interactions between the interface strength and the yield strength mismatch are found to govern crack growth rates near the interface. Hardening due to plastic strain gradients is important for finding the critical conditions associated with crack bifurcation at an interface and penetration through an interlayer.</div><div><br></div><div>Subsequently, crack growth in rate-dependent materials is analyzed. For materials characterized by power-law viscoplasticity, fatigue crack growth rates following overloads are found to depend strongly on the material rate sensitivity. The computations predict a transition from acceleration- to retardation-dominated post-overload crack growth as the rate sensitivity decreases. The predicted post-overload crack growth rates show good agreement with high-temperature experimentally-measured trends for Alloy 617, a solid solution strengthened nickel-base superalloy proposed for use in next-generation nuclear power plants. The results demonstrate why Alloy 617 behaves in a relatively brittle manner following overloads despite being characterized as a creep-ductile material. Crack growth is also studied in materials where rate dependence is captured through time-dependent damage and dislocation storage and dynamic recovery processes. This approach is relevant for high-strength creep-brittle materials, in which the viscoplastic zone grows with the advancing crack. The computations predict crack growth retardation for several loading waveforms containing overloads. The amount of retardation depends strongly on the overload ratio and subsequent unloading ahead of the crack tip. The predicted post-overload crack extension shows good agreement with high-temperature experimentally-measured trends for Alloy 718, a precipitation-hardened nickel-base superalloy used in turbine engines and power generation applications. The results demonstrate why Alloy 718 behaves in a ductile manner following overloads, despite being characterized as a creep-brittle material.</div>

This thesis investigates fatigue crack growth rate behavior in the threshold and near-threshold regimes for a Ti-6Al-4V (STOA) alloy using two proposed ASTM procedures- (1) load-shedding (LS) using a larger load-shed rate than the current ASTM Standard E647 load-reduction (LR) test procedure, and (2) compression pre-cracking constant-amplitude (CPCA) or load-increasing (CPLI) and load-shedding (CPLS). Tests were conducted at a low stress ratio (R = 0.1) on compact C(T) specimens of two different widths (W = 51 and 76 mm) and threshold fatigue crack growth rates were generated. These test data were compared to previous test data produced from the same batch of material using the current LR and the CPCA test procedure. While no test procedure provided an exact representation of the threshold value (?Kth), the compression pre-cracking (CP) procedures were the most promising. The LR, LS, and CPLS test procedures were influenced by prior loading-history and various crack-closure mechanisms, leading to higher ?Kth values and slower crack growths in the threshold regime. The LS tests (at shed-rates of -0.08,-0.32, and -0.95 mm-1) generated ?Kth values that were 15% to 32% higher than the estimated threshold stress-intensity factor range (?*Kth)R=0.1. The CP test procedures are a more accurate alternative for developing near-threshold and threshold fatigue crack growth rates. The CPLS test procedure produced a ?Kth value that was 10% higher than (?*Kth)R=0.1. LR and LS tests produced different ?Kth values as a function of the specimen width for the given load ratio. The CP test procedures produced consistent crack growth rates over the same range of ?K values examined, independent of the specimen width. Further research is required for developing test procedure(s) capable of providing a more definitive representation of the ?Kth value and closureree fatigue crack growth rates in the threshold regime.

Fatigue and crack-growth tests were conducted on 7050-T7451 aluminum alloy under a wide range of loading conditions. Crack-growth tests were conducted on compact, C(T), specimens under constant-amplitude loading, single-spike overloads, and a simulated aircraft spectrum loading. Fatigue tests were also conducted on single-edge-notch bend, SEN(B), specimens under constant-amplitude loading and three aircraft load spectra. The FASTRAN, life-prediction code, was used to make crack-growth predictions on the C(T) specimens; and to make fatigue-life calculations using a 12-micrometer initial flaw size at the center of the edge-notch on the SEN(B) specimens. The predictions agreed fairly well with most of the tests, except the model was unconservative on the single-spike overload tests and the severe spectrum Mini-TWIST+ Level 1 tests. The discrepancy was suspected to be caused by a low constraint factor and/or crack paths meandering around overload plastic zones. A roughness- and plasticity-induced crack-closure model would be needed to improve the model.

No description available.

No description available.