The 7050 aluminum alloy is used in many aerospace structural applications. Previous studies have identified that fatigue cracks develop very rough crack-suiface profiles, which cause very high crack-closure levels due to a combination of plasticity, roughness and debris. Previotisly, tests were conducted on compact, C(T), specimens to generate crack-growth-rate data from threshold to near fracture over a wide range in stress ratios (R). New threshold testing methods, based on compression precracking, were used to generate the data in the near-threshold regime. The plasticity-induced crack-closure model, FASTRAN, was used to correlate the data over a wide range in stress ratios and crack-growth rates from threshold to near fracture. To account for the very high crackclosure levels, a very low constraint factor, like plane-stress conditions, had to be used in the model. In addition, the crack-opening loads were measured during these tests using a local strain-gauge method to generate another AK^,ff-rate curve. These two curves differed only in the near-threshold regime. Herein, fatigue-crack-growth tests were conducted on C(T) specimens under spike overloads and simulated aircraft spectrum loading. Fatigue tests were also conducted on single-edge-notch bend (SEN(Bj), specimens over a wide range in loading conditions (constant amplitude and three aircraft spectra). All specimens were machined from a single forged block of7050-T7451. However, no residual stresses were measured in both the SEN(B) and C(T) specimens. Two European standard spectra were used, but modified to have only tension-tension loading. The purpose of this paper was to evahtate the two different effective stress-intensity factor curves for making crack-growth and fatigue-life predictions. Small-crack theory was used to make fatigue-life predictions using inclusion-particle sizes from the literature. Fatigue predictions on the SEN(B) specimens agreed fairly well (±30%) using a 12micrometer semicircular initial flaw located at the semicircular-edge notch under all loading conditions, except the model was unconservative (factor of three) on one of the severe aircraft spectra (Mini-TWIST-^, Level 1). For the C(T) specimens subjected to single-spike overloads, the life-prediction code also produced much more retardation than observed in the tests. However, the predicted crack-length-against-cycles under the Mini-Falstaff-\-spectrum were only about 15% longer than the tests. The discrepancy under the single-spike overloads and the severe aircraft spectra was suspected to be caused by the low constraint factor and/or crack paths meandering around overload plastic zones. Ideally, a roughness-induced crack-closure model: in addition to the plasticity model, would be needed to obtain more reasonable results."plasticity-induced" crack closure, other closure mechanisms have been identified, such as roughness-, fretting-product-, and oxide-debris-induced closure. These mechanisms have greatly improved our understanding of the complex interactions that occur dur...
Plasticity effects and crack-closure modeling of small fatigue cracks were used on a Ti-6Al-4V alloy to calculate fatigue lives under various constant-amplitude loading conditions (negative to positive stress ratios, R) on notched and un-notched specimens. Fatigue test data came from a high-cycle-fatigue study by the U.S. Air Force and a metallic materials properties handbook. A crack-closure model with a cyclic-plastic-zone-corrected effective stress-intensity factor range and equivalent-initial-flaw-sizes (EIFS) were used to calculate fatigue lives using only crack-growth-rate data. For un-notched specimens, EIFS values were 25-μm; while for notched specimens, the EIFS values ranged from 6 to 12 μm for positive stress ratios and 25-μm for R = −1 loading. Calculated fatigue lives under a wide-range of constant-amplitude loading conditions agreed fairly well with the test data from low- to high-cycle fatigue conditions.
A technology for non-intrusive real-time structural health monitoring using piezoelectric active sensors is presented. The approach is based on monitoring variations of the coupled electromechanical impedance of piezoelectric patches bonded to metallic structures in high-frequency bands. In each of these applications, a single piezoelectric element is used as both an actuator and a sensor. The resulting electromechanical coupling makes the frequency-dependent electric impedance spectrum of the PZT sensor a good mapping of the underlying structure’s acoustic signature. Moreover, incipient structural damage can be indicated by deviations of this signature from its original baseline pattern. Unique features of this technology include its high sensitivity to structural damage, non-intrusiveness to the host structure, and low cost of implementation. These features have potential for enabling on-board damage monitoring of critical or inaccessible aerospace structures and components, such as aircraft wing joints, and both internal and external jet engine components. Several exploratory applications will be discussed.
Plasticity effects and crack-closure modeling of small fatigue cracks were used on a Ti-6Al-4V alloy to calculate fatigue lives under various constant-amplitude loading conditions (negative to positive stress ratios, R) on notched and un-notched specimens. Fatigue test data came from a high-cycle-fatigue study by the U.S. Air Force and a metallic materials properties handbook. A crack-closure model with a cyclic-plastic-zone-corrected effective stress-intensity factor range and equivalent-initial-flaw-sizes (EIFS) were used to calculate fatigue lives using only crack-growth-rate data. For un-notched specimens, EIFS values were 25-μm; while for notched specimens, the EIFS values ranged from 6 to 12 μm for positive stress ratios and 25-μm for R = −1 loading. Calculated fatigue lives under a wide-range of constant-amplitude loading conditions agreed fairly well with the test data from low- to high-cycle fatigue conditions.
The 7050 aluminum alloy is used in many aerospace structural applications. Previous studies have identified that fatigue cracks develop very rough crack-surface profiles, which cause very high crack-closure levels due to a combination of plasticity, roughness and debris. Previously, tests were conducted on compact, C(T), specimens to generate crack-growth data from threshold to near fracture over a wide range in stress ratios (R). New threshold testing methods, based on compression precracking, were used to generate the data in the near-threshold regime. The plasticity-induced crack-closure model, FASTRAN, was used to correlate the data over a wide range in stress ratios and crack-growth rates from threshold to near fracture. To account for the very high crack-closure levels, a very low constraint factor, like plane-stress conditions, had to be used in the model. In addition, the crack-opening loads were measured during these tests using a local strain-gage method to generate another ΔKeff-rate curve. These two curves differed only in the near-threshold regime. Herein, fatigue-crack-growth tests were conducted on C(T) specimens under spike overloads and simulated aircraft spectrum loading. Also, fatigue tests were conducted on single-edge-notch bend, SEN(B), specimens over a wide range in loading conditions (constant amplitude and three aircraft spectra). All specimens were machined from a single forged block of 7050-T7451. However, no residual stresses were measured in both the SEN(B) and C(T) specimens. Two European standard spectra were used, but modified to have only tension-tension loading. The purpose of this paper was to evaluate the two different effective stress-intensity factor curves for making crack-growth and fatigue-life predictions. Small-crack theory was used to make fatigue-life predictions using inclusion-particle sizes from the literature. Fatigue predictions on the SEN(B) specimens agreed fairly well (± 30%) using a 12-micrometer semi-circular initial flaw located at the semicircular-edge notch under all loading conditions, except the model was unconservative (factor of 3) on one of the severe aircraft spectra (Mini-TWIST+, Level 1). For the C(T) specimens subjected to single-spike overloads, the life-prediction code produced much more retardation than observed in the tests. However, the predicted crack-length-against-cycles under the Mini-Falstaff+ spectrum were only about 15% longer than the tests. The discrepancy under the single-spike overloads and the severe aircraft spectra was suspected to be caused by the low constraint factor and/or crack paths meandering around overload plastic zones. Ideally, a roughness-induced crack-closure model, in addition to the plasticity model, would be needed to obtain more reasonable results.
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