Damage Tolerance Evaluation for Wing Structure with Large Cutout

Abstract

Wings are the lift generating components in the airframe structure. Wings are also used as fuel tanks in the transport aircraft. Cutouts are provided in the bottom skin of the wing to permit entry into the airplane fuel tanks for inspection or component repair. Bottom skin is under tension stress field during flight. Cutouts in the bottom skin will act as stress risers due to stress concentration effect. The high tensile stress locations are the most probable locations of fatigue cracking in the structure. The damage tolerance design philosophy says that cracks are allowed in the structure, but the cracks should not lead to catastrophic failure of the structure. The damage-tolerance evaluation of structure is intended to ensure that should fatigue, corrosion, or accidental damage occur within the LOV (Limit of validity) of the airplane, the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the damage is detected [1]. In the damage tolerance design philosophy the safety is ensured by inspection. Identification of the critical locations in the structure is most important to ensure the safety of the structure throughout the service life of the structure. Aircraft designer needs to ensure the structural integrity of the airframe without compromising on the safety of the structure. This would be possible only by adopting the damage tolerance design principles. The current project includes the stress analysis of a wing box of a medium size transport aircraft having large cutout in the bottom skin to identify the critical location for fatigue crack initiation. The aircraft under consideration is a conceptual Light Transport Aircraft. A local analysis is followed to obtain more accurate stress value and the distribution of stress. Aluminium alloy 2024-T351 Material is used for the wing box. Finite element method is adopted for stress analysis of the structural components. MSC NASTRAN and MSC PATRAN FEM packages are used to carry out the analysis. The maximum stress location is found from the wing box FE model (Global analysis). Damage tolerance evaluation includes the stress intensity factor calculations at crack tip. This is carried out by simulation of the crack in the finite element model of the bottom skin of the wing (Local analysis). Stress intensity factor (SIF) at different crack lengths is calculated using Modified Virtual Crack Closure Integral (MVCCI) method. The SIF calculated at every crack length is compared with the fracture toughness of the material. Variation of SIF as a function of crack length is plotted. II. LITERATURE SURVEY Damage tolerance philosophy is a refinement of the fail-safe philosophy. It assumes that cracks will exist, caused either by processing or by fatigue, and uses fracture mechanics analyses and tests to determine whether such cracks will grow large enough to produce failures before they are detected by periodic inspection. Three key items are needed for successful damage-tolerant design: residual strength, fatigue crack growth behaviour, and crack detection involving non-destructive inspection. Of course, environmental conditions, load history, statistical aspects, and safety factors must be incorporated in this methodology.[2] The recent Air Force requirement to apply linear elastic fracture mechanics approach in damage tolerance design of aircraft structures, warrants the critical review of various approaches. Pir M. Toor [3] has critically reviewed some damage tolerance design approaches and their application to aircraft structures. The