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a0.5in.spherical hardened-steelindenterhead.Dentdepths rangedfrom0.03to0.0325in.,measuredontheconvex side of the specimen. A starter notch of 0.3 in. was produced at the edge of the specimen with a jeweler’s saw blade. The specimen was fatigue-loaded under constant-amplitude loading to produce an initial crack length of 0.37 in., at whichtimereadingsofcracklengthvscyclesbegan.Thesameconstant-amplitudecyclicloadingusedtoproducethe initialcracklengthwasusedduringthetesting.Thecracklengthsweremeasuredwithanopticalmicroscopeat160 magnification. Nine specimens were tested, including three replications for each of the three conditions. Crack growth data are given in both tabular and graphical forms for all specimens. Crack growth rate data are also presented in graphical form. The overall crack growth in the dented specimens was significantly greater than in the pristine specimens. It was also faster, on average, in the reworked specimens; reworking, in general, did not recapture the life displayed by the pristine specimens.

Using Hi-Lok fasteners and Universal rivets, we experimentally investigated the fatigue behavior of fastener joints under constant-amplitude loading. The study generated fatigue-based stress severity factors for fastener joints at two different load transfer and mean stress levels. Fatigue tests on open-hole, filled-hole, and load transfer specimens were conducted at mean stress levels of 20.68 and 41.37 MPa, with varying stress ratios to achieve lives ranging between 10 3 and 10 6 cycles. The regression-based mean life curves were used to generate the fatigue-based correction factors for hole filling and bearing load transfer as a function of the maximum stress level. The fatigue life of load transfer specimens fell between those of the open-hole and filled-hole configurations. With increasing load transfer and mean stress levels, the fatigue life approached that of the open-hole case. In addition, at stress levels corresponding to fatigue lives in excess of 10 5 cycles, the failure modes of load transfer specimens transitioned from a net-section failure to a clamp-up-induced fretting failure occurring away from the fastener hole. The stress severity factor associated with the mean life of fastener joints and Hi-Lok filled-hole coupons increased with the maximum fatigue stress level. However, different trends were observed with respect to load transfer at the two mean stress levels used in the study.

Effects of plasticity induced crack closure on the calculated fatigue crack-growth behaviour of middle-crack tension specimens made of 2024-T351 alloy subjected to variable amplitude fatigue loading were examined. The specified loadings were four distinct aircraft spectra: (1) “modified” center-wing spectrum performed as part of this study; (2) fighter aircraft maneuver spectrum; (3) lateral gust spectrum; and (4) pressurized fuselage gust and maneuver spectrum. The experimentally observed crack-growth lives were compared with crack-growth predictions obtained using the crack closure concept in FASTRAN. Fatigue crack-growth material data extracted from the NASGRO materials database and the literature follows two distinct paths possibly due to different transitions from flat-to-slant (single or double shear) fracture. The material baselines for the FASTRAN code were selected to fit the two different paths. The two baseline curves were used to predict crack-growth life under four spectra to assess how the observed differences in the crack-growth rate trajectories affect the results. Using the lower bound baseline curve, which is suspected to be characterized by the double shear fracture mode, the FASTRAN model was able to predict crack- growth behaviour within ±25% of the test data for all four spectra.

The residual strength of a panel with a crack is dependent on the mode of failure. Generally, the panel will exhibit a brittle-fracture mode of failure (based on linear elastic fracture mechanics) or a mode of failure based on yielding. A common practice is to assume an elastic, perfectly plastic, material behavior, along with a net-section-yield criterion for panels that fail by yielding. However, the net-section-yield criterion is not appropriate for configurations such as panels with asymmetric geometry or loading, because equilibrium is violated. For asymmetric cases such as panels with edge cracks or off-center cracks, a plastic moment or hinge must exist, the details of which may be determined by satisfying equilibrium. The purpose of this study is to compare the experimental residual strengths of panels with edge cracks and off-center cracks with the residual strengths determined from the assumption of elastic, perfectly plastic, behavior and a plastic-moment (hinge) criterion. The material chosen for this study is 2024-T3 aluminum alloy. This material is commonly used for aircraft skin and it has a high fracture toughness, and so the mode of failure will less likely be brittle fracture. An elastic, perfectly plastic, material model was assumed, and expressions were developed for a plastic-hinge failure for panels with edge cracks as well as off-center cracks. A large number of panels with either edge cracks or off-center cracks were tested to failure, and the results were compared with the results obtained from the equations developed from the plastic-hinge assumption. The results from the plastic-hinge assumption compared favorably with the experimental results.

The aluminum alloy 2524-T3 is replacing 2024-T3 for many applications because 2524-T3 has a greater fracture toughness while retaining the same strength. Much attention has been given to the effect of multiple site damage on 2024-T3 aluminum; however, very little has been reported about 2524-T3. Twenty-two panels of 2524-T3, each with a different crack configuration, were tested for critical (linkup) strength, and the results were compared with an identical set of previously tested 2024-T3 panels with MSD. The panels were 24 in. wide with a midspan row of 0.25-in.-diam holes at 1-in. pitch. Each panel had a central lead crack with collinear MSD cracks emerging from both sides of the adjacent holes. A comparison of the results showed the 2524-T3 panels to average approximately 27% greater strength than the 2024-T3 panels. The linkup or plastic-zone-touch model used to predict the critical (link-up) strength of the panels was found to be highly conservative. Consequently the test data were used for a semi-empirical analysis to develop a modified link-up model for 2524-T3, similar to the one previously developed for 2024-T3. The average error between the critical strengths from testing and those predicted by the link-up model was approximately 20%, whereas that for the modified link-up model was approximately 3%.

Fracture analysis of cracks along mechanically fastened joints is a central issue in the damage tolerance assessment of semimonoque aircraft structures. Nonlinear elastic-plastic finite element analysis, employing the cracktip-opening-angle criterion, was used to evaluate residual strength of bolted lap-joint 2024-T3 aluminum panels with multiple site damage subjected to Mode-I loading. A total of 36 different crack configurations were analyzed and compared to the experimental results. The effect of different aspects of the finite element modeling procedure on the accuracy of the residual strength predictions was investigated. Modeling of the geometric details of the test panels, such as fasteners and antibuckling restraints, was found to have a considerable effect on the predicted residual strengths. The nonlinear analyses provided reasonable estimates of the residual strength. The numerical results were compared with residual strength predictions obtained from a semi-empirical link-up model. Both approaches were comparable in terms of accuracy when compared to the test data.

A modified linkup model has been developed for determining the residual strength of 2024-T3 aluminum panels with multiple-site damage. This model was developed by semi-empirical analysis of test data from flat unstiffened open-hole panels. The model was later validated with test data from 36 open-hole stiffened panels. During the investigation, 36 bolted lap joint panels with different crack configurations were tested to further validate the previously developed modified linkup model. Stress intensity factors for the modified linkup model for the different crack configurations were determined from finite element analysis with the FRANC2D/L code. The residual strengths predicted by the modified linkup model correlate well with the test data, as was the case with the earlier studies of the stiffened panels. The results for the bolted lap joint panels show a sudden reduction in the residual strength at the early stages of multiple-site damage followed by a gradual linear decrease with increasing multiple-site damage crack size. The results also demonstrate that a stress-intensity-factor-based solution can be formulated with empirical analysis of test data from a simple configuration and then used to analyze more complex configurations.

The increased use of composite materials and their relatively high cost and limited availability make it essential to develop low cost, effective nondestructive testing and inspection techniques (NDT/NDI). One of the oldest and widely used NDT/NDI methods is the coin tap test. The objective of this research was to determine if the sound signals generated by tapping a composite sandwich panel could be classified by an artificial neural network (ANN) as originating from damaged or non-damaged areas on the panel and if possible, to make accurate damage level assessments. Tap sound signals were recorded from several test panels using an ordinary condenser microphone and related equipment. Two separate signal-preprocessing techniques were employed, one using Fourier transforms and one using Wavelet transforms. Wavelet transformation of the signals tended to produce the best results. Artificial neural network configurations were developed using the backpropagation-learning algorithm that correctly classified da...

Much attention has been given to the development of technology for the purpose of determining the strength of aluminum panels that have multiple-site damage. The linkup model has been investigated because of its simplicity, and a number of modified linkup models have been presented. However, most of the attention has been given to 2024-T3 aluminum. Little attention has been given to 7075-T6 because it is a more brittle material with lower fracture toughness, making it more suitable to be analyzed by conventional linear elastic (brittle) fracture mechanics technology. The work presented here involves the study of 7075-T6 panels with multiple-site damage. Both the classical linear elastic fracture model and the linkup model are shown to be highly inaccurate. A modified linkup model and a modified brittle fracture model have both been developed by empirical analysis. Both of these models appear to have a high degree of accuracy over a wide range of crack geometry. The modified linkup model for 7075-T6 is then compared with a previously developed modified linkup model for 2024-T3. This comparison shows that 2024-T3 panels with multiple-site damage have greater strength than 7075-T6 panels, especially for panels with small ligament lengths.

Two modie edlinkup modelsweredevelopedfordetermining thecritical stressbased on linkup (ligamentfailure ) of2024-T3aluminumpanelswithmultiplesitedamage.ThesemodelsweredevelopedforusewithstandardMilitary Handbook for Metallic Materials and Elements for Aerospace Vehicle Structures (MIL-HDBK-5G )yield strength values. For this investigation, ligament failure stresses predicted by these models are compared with test stresses determined from a variety of stiffened panels including single-bay panels with the lead crack centered between stiffeners and two-bay panels with the lead crack centered beneath a severed stiffener. The stresses predicted by the modie ed linkup models correlate well with the test data. The results of this investigation should add to the understanding of the extent to which nonlinear behavior can be modeled with simplie ed engineering models. Nomenclature a = lead crack half-length an = nominal lead crack half-length c = multiple site damage (MSD) crack length D = hole diameter Fcol = collapse stress ` = half-length for MSD crack and hole, cC D/2 L = ligament length t = panel thickness tS = stiffener thickness W = panel width WS = stiffener width

An aging aircraft accumulates fatigue cracks commonly referred to as multiple site damage (MSD). For ductile materialssuchas2024-T3aluminum,MSDcracksmaylowerthestrengthsignie cantlybelowthatwhichispredicted by conventional fracture mechanics or net section yield failure methods. An analytical model generally referred to as the linkup model (or the plastic-zone-touch model )has previously been used to describe the MSD phenomenon. However, the linkup model is only accurate for limited geometric cone gurations. Two modie cations to the linkup model were developed through regression analysis of test data obtained from the literature and from experimental results conducted in this investigation. The modie ed models show signie cantly improved correlation with the test data over a wide range of cone gurations for e at 2024-T3 aluminum panels with MSD at open holes. Nomenclature a = lead crack half-length an = nominal lead crack half-length c = MSD crack length D = hole diameter L = ligament length ` = half-length for MSD crack and hole, c + D/2 t = panel thickness W = panel width b a = correction to stress intensity of the lead crack, b a/`b W b a/` = correction to stress intensity of the lead crack for the effect of the adjacent MSD crack b b = correction to stress intensity of the adjacent MSD crack for the effect of an open hole b ` = correction to stress intensity of the adjacent MSD crack, b `/ab b I (c/`) b `/a = correction to stress intensity of the adjacent MSD crack for the effect of the lead crack b W = e nite-width correction to the stress intensity of the lead crack, I [sec(p a/w)] r c = critical stress for ligament failure based on

The purpose of this stud y was to determine the effect of dents on the stable crack growth in 0.04 inch thick 2024 -T3 bare aluminum sheet. The test specimens were either pristine dented or reworked. The edge -cracked pin -loaded specimens of 8 inches in width were tested at consta nt amplitude loading with a stress ratio of 0.2 producing stable crack growth of close to 4 inches completely through two dents on the crack line. Dents were produced with a drop tower having a one inch spherical hardened steel indenter head. Dent depths ranged from 0.03 inch to 0.0325 inch measured on the convex side of the specimen. A starter notch of 0.3 inch was produced at the edge of the specimen with a jewelers saw blade. The specimen was fatigue loaded under constant amplitude loading to produce a n initial crack length of 0.37 inch at which time readings of crack length vs. cycles began. The same constant amplitude cyclic loading used to produce the initial crack length was used during the testing. The crack lengths were measured with an optical microscope at 160X magnification. Nine specimens were tested including three replications for each of the three conditions. Crack growth data is given in both tabular and graphical form for all specimens. Crack growth rate data is also presented in grap hical form. The overall crack growth in the dented specimens was significantly greater than in the pristine specimens. It was also, on the average, faster in the reworked specimens; reworking, in general, did not recapture the life displayed by the pristi ne specimens.

Combined simultaneous radiotherapy and multidrug chemotherapy have been utilized in an attempt to eradicate or shrink tumors of the head and neck area in advanced stages to allow subsequent surgical extirpation. Thirty-six patients (1 stage II; 11 stage III; 24 stage IV; 1 unknown primary) were treated with simultaneous radiotherapy and fractionalized doses of bleomycin, adriamycin, and 5-fluorouracil. Thirty of 36 patients completed the treatment schedule. Twenty-nine of 30 (97%) had a 75% or greater response to treatment; 20/30 (66%) had 100% response; 9/30 (30%) had 75% response; and 1/30 had 50% response. Of the 30 patients completing the regimen, 19/30 (63%) are alive three to 27 months later, and 12 (40%) of these have no evidence of disease. Six patients died within four months of initiation of therapy, mostly from complications of their disease. This treatment regimen appears effective in the control of locoregional disease or to produce enough shrinkage of tumor to allow subsequent surgical extirpation.