Stiffened Panels of the FAA / NASA Wide

This paper presents the fracture analyses conducted on the FAA/NASA stiffened and unstiffened panels using the STAGS (STructural Analysis of General Shells) code with the critical crack-tip-opening angle (CTOA) fracture criterion. The STAGS code with the "plane-strain" core option was used in all analyses. Previous analyses of wide, flat panels have shown that the high-constraint conditions around a crack front, like plane strain, has to be modeled in order for the critical CTOA fracture criterion to predict wide panel failures from small laboratory tests. In the present study, the critical CTOA value was determined from a wide (unstiffened) panel with anti-buckling guides. The plane-strain core size was estimated from previous fracture analyses and was equal to about the sheet thickness. Rivet flexibility and stiffener failure was based on methods and criteria, like that currently used in industry. STAGS and the CTOA criterion were used to predict load-against-crack extension for the wide panels with a single crack and multiplesite damage cracking at many adjacent rivet holes. Analyses were able to predict stable crack growth and residual strength within a few percent (5%) of stiffened panel tests results but over predicted the buckling failure load on an unstiffened panel with a single crack by 10%. INTRODUCTION Widespread fatigue damage is of great concern to the aging commercial transport fleets because the residual strength of a stiffened structure with a single long crack may be significantly reduced by the existence of adjacent smaller cracks as postulated by Swift [1]. Whereas a single long crack in a fuselage structure may produce flapping, a process by which a cracked fuselage would peel open in a small local region and lead to safe decompression, a fuselage with a long lead crack and multiple-site or multiple-element damage (MSD or MED) cracking may not flap. Tests on panels with long lead cracks and MSD have shown that the presence of an array of small adjacent cracks strongly degrade residual strengths [2,3]. One of the objectives in the NASA Aging Aircraft Research Program [4] is to develop the methodology to predict flapping or failure in damaged fuselage structures in the presence of widespread fatigue damage. The approach is to use a finite-element shell code with global-local, adaptive mesh capabilities and appropriate local fracture criteria to predict progressive failure in complex structures. In the future, fuselage structures may be designed by analysis, and verified by tests, to produce flapping or improved crack arresting capability under MSD or MED conditions. In an effort to develop the methodologies required to predict the residual strength of complex fuselage structures with MSD, a series of tests and analyses have been performed from the coupon level to subscale fuselage simulation tests [5,6]. These series of tests were used to verify the residual strength methodology based on the critical crack-tip-opening angle (CTOA) failure criterion. NASA and the FAA have designed a series of wide, stiffened panels with single and multiple-site damage cracking at many simulated rivet holes [7,8]. These panels had "fatigue" MSD cracks instead of sawcuts. The panels were made of 2024-T3 sheet material (thickness of 1.6 ram) with five 7075-T6 stiffeners (thickness of 2.2 mm). Stiffeners were placed symmetrical about the thickness of the sheet. The central stiffeners were cut and the panels were allowed to buckle during the stable tearing and crack linkup process. Load-crack extension, load-strain records and out-of-plane displacement measurements were made during all tests. Stable crack growth in metallic materials has been studied extensively using elastic-plastic finite-element methods, see reference 9. These studies were performed to investigate various fracture criteria. Of these, the crack-tip-opening angle (CTOA) or displacement (CTOD) criterionwasshownto bethemostsuitedfor modelingstablecrackgrowth andinstabilityduring the fracture process. Numerousinvestigatorshave also experimentallymeasuredCTOA (or CTOD) duringfracture. Dawickeet al. [10], usingahigh-resolutionphotographiccamerawith a videosystem, hasshownthatthecriticalanglesduringstablecrackgrowth in thin-sheetaluminum alloyswere nearly constantafter a small amountof tearing. The non-constantCTOA region (measuredat the free surface)was shownto be associatedwith severetunnelingduring the initiation of stabletearing. NewmanandDawicke [11] andDawickeet al. [12] haveextensively appliedtheCTOA fracturecriterionandthe finite-element methodto analyzestablecrackgrowth and fracture of manycomplexcrack configurationsmadeof aluminumalloys. Seshadriand Newman[13] havealsousedthe finite-element methodandthe CTOA criterionto predict stable tearingin thepresenceof severeout-of-planebucklingfor bothaluminumalloysandsteels. This paperpresentsthe fractureanalysesconductedon the FAA/NASA stiffenedpanels usingtheSTAGS(STructuralAnalysisof General Shells)code[14,15]with thecritical crack-tipopeningangle(CTOA) fracturecriterion. TheSTAGScode,with the"plane-strain"coreoption, was used in all analyses. Previousanalysesof wide, flat panelshave shownthat the highconstraintconditionsarounda crackfront, like planestrain,hasto be modeledin order for the critical CTOAfracturecriterionto predictwidepanelfailuresfrom smalllaboratorytests[ 16]. In thepresentstudy,thecritical CTOA valuewasdeterminedfrom a wide (unstiffened)panelwith anti-bucklingguides. Theplane-straincoresizewasestimatedfrom previousfractureanalyses andwasaboutequalto the sheetthickness.Rivet flexibility andstringerfailure werebasedon methodsandcriteria,like that currentlyusedin industry [1]. Comparisonswere madebetween load-crackextensionon stiffenedand unstiffenedpanelswith singlecracks and MSD. An assessment of the capabilityof the STAGS code with the critical CTOA failure criterion to predictresidualstrengthwasmade.