Damage tolerant evaluation of cracked stiffened panels under fatigue loading

Reliable damage tolerant design of airframe structures relies on the accurate life prediction of fatigue cracks propagating from a detectable size to the critical size, which is challenging due to the complex load sequence effect under spectrum loading. This paper aims at gaining a further understanding of the complex influence of the loading history on fatigue damage through a detailed numerical simulation of the near-tip crack behaviour using the crack closure concept. The spectrum loading is broken down into a number of simple yet representative loading scenarios with overload/underload superimposed onto the baseline constant amplitude fatigue loading. Detailed finite element (FE) simulation of the plasticity-induced crack closure (PICC) has been carried out to catch the transient behaviour of PICC and link it to the fatigue damage. The load interaction effect has been analysed with the aim to identify the possible dominant loading cycle which could simplify the fatigue life prediction process in the industry. It is concluded that more reliable damage tolerant design can be achieved if the load sequence effect on fatigue damage can be taken into account more accurately for a structure under spectrum loadings.

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Failure simulations of open-hole IM7/977-3 coupons subjected to fatigue loading using Autodesk Helius PFA

Journal of Composite Materials ◽

10.1177/0021998316669579 ◽

2016 ◽

Vol 51 (15) ◽

pp. 2119-2129 ◽

Cited By ~ 7

Author(s):

Richard W Dalgarno ◽

Jason E Action ◽

Donald H Robbins ◽

Stephen P Engelstad

Keyword(s):

Air Force ◽

Carbon Fibers ◽

Fatigue Loading ◽

X Rays ◽

Hole Configuration ◽

Open Hole ◽

Modeling Strategy ◽

Improve Correlation ◽

Damage Tolerant ◽

Air Force Research Laboratory

Finite element simulations of three laminates in open-hole configuration subjected to constant amplitude tension–tension fatigue loading are investigated as part of the Damage Tolerant Design Principles program organized by the Air Force Research Laboratory. All coupons were made from unidirectional IM7/977-3 plies, which are composed of intermediate modulus carbon fibers and a toughened epoxy matrix. Government furnished experimental data from an assortment of fatigue loaded unnotched coupons were used to characterize the behavior of the composite material in the simulations. The commercial software Autodesk Helius PFA was used to model the non-linear response of the material. Blind simulations of coupon stiffness and damage at several cycle numbers and residual coupon tensile and compressive strengths are benchmarked against experimental measurements and X-rays. Upon review of the experimental results, a second round of simulations was performed where the modeling strategy was updated to improve correlation to experiment.

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Characterization of Fatigue Crack Growth Rate of AA6056 T651 and T6: Application to Predict Fatigue Behaviour of Stiffened Panels

Materials Science Forum ◽

10.4028/www.scientific.net/msf.636-637.1511 ◽

2010 ◽

Vol 636-637 ◽

pp. 1511-1517 ◽

Cited By ~ 6

Author(s):

Pedro Miguel Guimarães Pires Moreira ◽

Valentin Richter-Trummer ◽

Sérgio M.O. Tavares ◽

Paulo Manuel Salgado Tavares de Castro

Keyword(s):

Growth Rate ◽

Fatigue Crack ◽

Crack Growth Rate ◽

Fatigue Crack Growth ◽

Crack Growth ◽

High Speed ◽

Thick Plate ◽

Fatigue Loading ◽

Fatigue Behaviour ◽

Stiffened Panels

Tensile and fatigue crack growth tests of the 6056 T651 and T6 aluminium alloys were carried out. The fatigue crack propagation tests were performed on compact tension 4mm thick (CT) specimens, under cyclic loading with R ratios 0.1 and 0.5. The resulting data was used to predict the fatigue behaviour of stiffened panels subjected to fatigue loading under similar R ratios. The AA6056-T651 panels were fabricated using High Speed Machining (HSM) starting with 30mm thick plates. AA6056-T651 CT specimens were cut from the panels mentioned above, whereas AA6056-T6 CT specimens were machined from 5mm thick material. It was found that the AA6056-T651 (HSM material) specimens, machined from a 30mm thick plate presented higher rupture and yield stress than the AA6056-T6 material extracted from a 5mm thick plate. When tested at the same R value the AA6056-T6 specimens present higher crack growth rate than the AA6056-T651 specimens.

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