IJSR - INTERNATIONAL JOURNAL OF SCIENTIFIC RESEARCH 157 Volume : 2 | Issue : 8 | August 2013 • ISSN No 2277 - 8179 Research Paper Engineering Dr. P. Ravinder Reddy Professor and Head, Mechanical Engineering Department, Chaitanya Bharathi Institute of Technology, Gandipet, Hyderabad – 500 075 P. RamaLakshmi Assistant Professor, Mechanical Engineering Department, Chaitanya Bharathi Institute of Technology, Gandipet, Hyderabad – 500 075 P. Anjani Devi Assistant Professor, Mechanical Engineering Department, Chaitanya Bharathi Institute of Technology, Gandipet, Hyderabad – 500 075 ABSTRACT In this investigation, J-integral for the cracked panel with stiffeners is studied considering the finite ele- ment models. In the first case, the panel with two stiffeners is examined where in position of the stiffeners is increased. The crack length is varied keeping the stiffener at one inch from the crack tip is presented as case two. Isotropic (steel and aluminum) and orthotropic materials (glass epoxy and graphite epoxy) are selected for the whole stiffened panel. The finite element models for different stiffener and crack configurations are generated using ANSYS11 and the non dimensional value of the J- integral is calculated and is plotted against position of the stiffener for case one and crack length for case two. The presence of stiffeners made significant contribution in the reduction of J-integral when compared with the sample without stiffeners. Influence of Stiffeners on Fracture Parameters in Isotropic and Orthotropic Materials KEYWORDS : J- integral, FEA, stiffener, crack length, orthotropic 1. INTRODUCTION AND LITERATURE REVIEW Due to light weight and high operating stresses aircraft struc- ture is a prominent example where structural efficiency is high. There are few important parameters that need to be considered – capability to generate at a reasonable cost, adequate service life [1] and perform the required function. Sheets and stringers which include fuselage skin panels, wings, spar webs, stiffener form the important part in the making of aircraft structures. Cracks arise in these structural elements thereby reducing the total load carrying capacity and the stiffness of the structure. The need for light weight, large scale metallic structure has brought new set of problems related to fracture. The weight of the aircraft is addresses by thin skin and stringers. By adhesive bonding or riveting stringers are joined to the skin. S Habeeb [2] et al examined the load bearing characteristics and crack arrest- ing capabilities in a stiffened and unstiffened panel subjected to uniform remote displacement field. From linear elastic analy- sis there is a decrease in stress intensity factor when the crack approached the stiffener. Franc2dl code was used to perform fracture analyses on adhesively bonded stiffened panel. When fuselage is pressurized and depressurized during each take off and landing cycle of aircraft, the metal skin of fuselage expands and contracts, resulting in metal fatigue. Due to the presence of large number of rivet holes, the fuselage skin has large num- ber of high stress locations and these are locations of potential crack initiation. From the principles of displacement compat- ibility and superposition, Rans et al [3], presented an analytical model for estimating fracture parameters in cracked skin panels containing bonded stiffening elements. Results were validated with the experimental data available in literature. Superposi- tion, fracture mechanics and displacement compatibility ap- proach used by Poe [4] was adopted by Alderliesten [5-11] to understand the crack growth behavior of fiber metal laminates. Venkatesha et al [1] investigated crack initiation, crack growth, fast fracture and crack arrest features in the stiffened panel. MSC Patran software was used for modeling and MSC Nastran - Solver for linear static stress analysis of stiffened panel of fuselage. Fracture mechanics provides a tool for assessing the criticality of flaws in structures that can be used directly accord- ing to Barsom and Rolfe [12]. They also state that the science of fracture mechanics can be used to describe quantitatively the trade-offs among stresses, material toughness, and flaw size. 2. PROBLEM FORMULATION In the present work, J-integral is estimated with the help of ANSYS package. Two problems are selected. In the first prob- lem shown in figure 1, aircraft panel having crack is examined to evaluate J-integral. As an extension of the same, stiffeners is added and J-integral is determined with variation in the posi- tion of the stiffener is shown in figure 2. In case two, the position of the stiffener is kept constant (one inch from the crack tip) and the crack length is varied to find out the non-dimensionalised J- integral values. In the second case one stiffener is used while in the first case, two stiffeners are placed uniformly on the aircraft panel. Figure 1: Plate with stiffener Figure 2: Plate without stiffener 3. METHODOLOGY