STRESS INTENSITY FACTOR FOR CRACKS IN STIFFENED PANELS USING MODIFIED CRACK CLOSURE INTEGRAL METHOD

Abstract

The airframe of the aircraft fuselage is predominantly made up of stretch cladding, circumferential frames and longitudinal spars. During the aircraft's landing, the skin made by metal of the fuselage expands resulting in metal fatigue. Fatigue damage accumulates during each load cycle on the fuselage structure during operation. When the cumulated damage reaches a critical value; the cracks start from riveted holes. In addition to fatigue damage, cracks may occur because of corrosion or accidental damage. These cracks propagate to critical sizes leading to catastrophic failure of the structure. Damage tolerance analysis focuses on this phase called crack propagation. Damage tolerance is the ability of the structure to support the anticipated structure. Loads in the presence of fatigue cracks, corrosion or accidental damage until such damage is detected by inspection or breakdown and repaired. Large transport aircraft are designed to tolerate large cracks and are therefore tolerant of damage. An aircraft damage tolerance analysis involves determining the crack growth life from either an initial size or a detectable size up to the critical crack size. A typical crack growth analysis is generally performed by idealizing the geometry and loading of a single component to enable the use of a library of solutions in code to analyze crack growth. One of the important parameters is the stress intensity factor, which is a function of stress, crack length and the geometry parameter. This paper focuses on the derivation of the stress intensity factor for different crack lengths and for different curved, stiffened plate configurations, hence the geometry factor data using the FEA (NASTRAN) and Modified Crack Closure Integral (MCCI) method. The geometry parameter data is used in the damage tolerance analysis of longitudinal cracks for the mission spectrum.