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...
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.
Eccentrically-loaded single-edge crack tension, ESE(T), specimens made of A36 structural steel were tested over a wide range in stress ratios (R=0.1 and 0.7) in laboratory air. Two test methods were used: (1) ASTM Standard E647 load-reduction method and (2) compression precracking. After compression precracking (CP), three different loading sequences were used: (1) constant amplitude (CPCA), (2) load reduction (CPLR), and (3) constant stress-intensity factor (CPCK). The crack-compliance method was used to determine that the specimens had no residual stresses; and that the effects of tensile residual stresses from compression precracking dissipated in about 2 compressive plastic-zone sizes. Agreement was found between the A36 and TC-128B steel ΔK-rate data tested at both low and high stress ratio (R) conditions. At R=0.1 loading, the CPCA and CPLR tests generated lower thresholds and faster rates than using the standard ASTM load-reduction method. All load-reduction tests exhibited an accumulation of debris at the crack front near threshold conditions. A crack-closure analysis was preformed to calculate the effective stress-intensity factor range (ΔKeff) against rate using measured 1 % offset (OP1) values for all R=0.1 tests. The ΔKeff-rate data correlated well with the high-R results.
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