Authors :
Hemanth.Kumar.G, Praveen D N, Lohitesh Jaga Kumar
Volume/Issue :
Volume 2 - 2017, Issue 6 - June
Google Scholar :
https://goo.gl/b3teKW
Scribd :
https://goo.gl/xpyTNg
Thomson Reuters ResearcherID :
https://goo.gl/3bkzwv
Abstract :
Damage tolerance design philosophy is followed in the airframe design to achieve the minimum weight of the structure without compromising on the safety of the structure. This philosophy includes fail-safe design of the structure. Fuselage structure of the aircraft is made up of stressed skin, longitudinal longenors, and circumferential. The most common cause of structural failures is fatigue under service loading. Fatigue c racks can initiate and propagate to critic al dimensions leading to a catastrophic failure of the air frame. The current airframe design concepts permit a fatigue crack to initiate from a manufacturing flaw early in service life and propagate. But it should not lead to a catastrophic fail u re of the aircraft.The current investigation includes the evaluation of a fusel age-stiffened panel for its damage tolerance capability with one of its frames in the broken condition. The cracking location is idealized as a flat stiffened panel with a skin crack subjected to uniaxial tensile loading analysis will be carried out with a frame broken condition. A Finite element analysis approach will be followed in this investigation. Geometrical dimensions representative of actual aircraft in service will be considered. The material used for the stiffened panel will be taken as 2024-T3 aluminum alloy.A panelstrength diagram will be derived from the stress analysis of this cracked stiffened panel with an frame broken conditionIn this analysis only the frame is considered in the design of the fuselage and analysis is done considering only this conditions with a broken frame.
Keywords :
Stiffened panel, Stress intensity factor, Fatigue crack, Finite element analysis, Fail-safety, Catastrophic failure.
Damage tolerance design philosophy is followed in the airframe design to achieve the minimum weight of the structure without compromising on the safety of the structure. This philosophy includes fail-safe design of the structure. Fuselage structure of the aircraft is made up of stressed skin, longitudinal longenors, and circumferential. The most common cause of structural failures is fatigue under service loading. Fatigue c racks can initiate and propagate to critic al dimensions leading to a catastrophic failure of the air frame. The current airframe design concepts permit a fatigue crack to initiate from a manufacturing flaw early in service life and propagate. But it should not lead to a catastrophic fail u re of the aircraft.The current investigation includes the evaluation of a fusel age-stiffened panel for its damage tolerance capability with one of its frames in the broken condition. The cracking location is idealized as a flat stiffened panel with a skin crack subjected to uniaxial tensile loading analysis will be carried out with a frame broken condition. A Finite element analysis approach will be followed in this investigation. Geometrical dimensions representative of actual aircraft in service will be considered. The material used for the stiffened panel will be taken as 2024-T3 aluminum alloy.A panelstrength diagram will be derived from the stress analysis of this cracked stiffened panel with an frame broken conditionIn this analysis only the frame is considered in the design of the fuselage and analysis is done considering only this conditions with a broken frame.
Keywords :
Stiffened panel, Stress intensity factor, Fatigue crack, Finite element analysis, Fail-safety, Catastrophic failure.
