International Research Journal of Engineering and Technology (IRJET)
e-ISSN: 2395-0056
Volume: 09 Issue: 05 | May 2022
p-ISSN: 2395-0072
www.irjet.net
Fail Safe Design of an Aircraft Stiffened Panel by Stress Intensity Factor Determination Through Modified Virtual Crack Closure Integral Method Shivanandappa N D1, Guru Kiran E2 1Assistant 2PG
Professor, Dept. of Mechanical Engineering, J N N College of Engineering, Shivamogga, Karnataka, India. Student, Dept. of Mechanical Engineering, J N N College of Engineering, Shivamogga, Karnataka, India.
------------------------------------------------------------------------***----------------------------------------------------------------------designed to a satisfy a particular value of static and dynamic loading conditions, deformation and functional importance during the design of various interconnected criterion. Service loads during the operation of an aircraft components. During service, the component parts are prone for design and verification of damage tolerance and to crack initiation due to fluctuating loads and are going to durability are also very important. Fatigue and the grow over a period. Stiffened panel is one such component resulting crack growth are a major challenge during the part which is prone to crack initiation. The panel must be design of aircrafts. For the continued airworthiness of an able to withstand the loads even in the presence of crack and aircraft during its entire economic service life, fatigue and not fail catastrophically without giving any warning. A faildamage tolerance design, analysis, testing and service safe design approach is employed for evaluating the design experience correlation play a pivotal role. of the stiffened panel. The method involves the finite element analysis of the panel with crack to determine the stress The design of an aircraft considers finding an optimal intensity factor by modified virtual crack closure integral proportions of payload and weight of the vehicle. It needs method. The stress intensity factor for three different skin to be stiff and strong enough to fly under exceptional thicknesses and varying crack lengths are determined and circumstances. Also, the aircraft has to fly even when one compared with the fracture toughness value of the material of the parts fail during the flight. of the stiffened panel. The results show that the skin offers more resistance to crack propagation as the thickness The skin is a load carrying member in the modern increases. Also, the result shows that the stress intensity aircrafts. Folded sheet metals can carry compressive loads factor of the panel of thickness 2.2 mm goes beyond the unlike the flat sheets that carry only tension. Stiffeners fracture toughness value after it reaches a crack length of combined with a section of skin are analysed as thin walled 635 mm and comes below the fracture toughness value as structures, known by the name stringers. the crack length reaches 1016 mm indicating that the failsafe design is achieved. In the current case, a part of stiffened panel from the fuselage segment is considered for the analysis and then Keywords โ Aircraft Fuselage, Fail-safe Design subjected to tensile loading which is equal to the hoop Approach, Fluctuating Load, Fracture Toughness, MVCCI stress developed in the fuselage. In case of the damage method, Stiffened Panel, Stress Intensity Factor. existing in the fuselage the damage should not exceed beyond the design limit and the structure should not 1 INTRODUCTION undergo failure leading to catastrophic failure of the aircraft structure. So the design of structure to be made in For the airplanes to fly free of danger, they have to terms of damage tolerance to avoid the failure of structure. satisfy damage tolerance requirements and follow In this context the damage existing in the skin can be airworthiness regulations. A structural component is said tolerated by increasing the skin thickness of the stiffened to be damage tolerant if it remains in operation after an panel. initial damage is detected. Analysis of fatigue crack growth
Abstract - The safe flying of an aircraft is of paramount
is the main focus of damage tolerance assessment. It involves determining how cracks propagate during service life.
The geometric model of the stiffened panel with fuselage segment is been created in CATIA modeling software and then imported into MSC.PATRAN for finite element modeling. The finite element model is solved using MSC.NASTRAN for solving stiffened panel subjected to the tensile loading with a center crack.
Modern airplanes operate in a complex environment, loading conditions, human resource and economic requirements. The major components of the aircraft are
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