PSI - Issue 60

2

M Mohan Kumar et al. / Procedia Structural Integrity 60 (2024) 177–184 M Mohan Kumar et al. / Structural Integrity Procedia 00 (2024) 000 – 000

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1. Introduction Panels in aircraft components are formed from skin and internal stiffeners by riveting, screwing and welding. The fabrication difficulties of the assembled panel entities are very less, and the damage resistant performance is good [1]. However, in the recent aircraft design and development, the rigidity, strength, fatigue, sealing performance and other aspects are posing some issues in complying with the aircraft usage requirements. These issues may be eliminated to a large extent by usage of integral stiffeners. Integral stiffened structures have few advantages compared to the riveted panels as they have very few or no external projections on their surfaces as the number of connected parts are greatly reduced and due to absence of any buckling effects of skin. Also, greater allowable compression loads in the stiffeners is observed due to absence of attached members and improvised joint efficiencies under tension loads because of integral doublers, etc [2]. The integral panel has obvious advantages not only in these aspects, but can also prevent any kind of defects and damage initiation caused by riveting structure [3]. Therefore, in the recent aircraft development programs, the integrated design and manufacturing of panels is gradually used to replace the built up panels and structures. Usage of integral structure in the place of conventional riveted assembly structure in large aircrafts has become a noticeable feature of aircraft structural design, and the research on the integral structure is gaining greater momentum in the aircraft structural design and development [4]. Poe [5] has established the crack growth studies of a panel with stiffeners based on unstiffened panel a nalysis data, and his predictions has a good agreement with the actual test data acquired from panels with stiffeners of varied geometries. Poe's work [5] depicts how a light stringer causes a slow crack growth, since it brings about less reduction of the stress intensity factor. In the view of this, the research on the fracture resistance of the entire structure with cracks has become an important part of the research on the damage tolerance of the integral structure.

Nomenclature P

Applied load, N

K 1

Stress intensity factor, MPa √ m

a

Crack length, mm Number of cycles

N

K IC Fracture toughness, MPa√m da/dN Crack growth rate, mm/cycle a c Critical crack length, mm K C

Critical stress intensity factor, MPa√m Stress intensity factor range, MPa√m Threshold Stress intensity factor, MPa√m

∆K ∆ K th

R Stress ratio 2. Description of stiffenedpanel geometry

The present work deals with metallic integrally and riveted stiffened aircraft fuselage crown panels made up of aluminum alloy 2024-T351. Since the crown panels (top portion of the fuselage) are the region of maximum tensile stress locations under pressurized loading conditions a typical crown panel is considered for the present study. The geometrical details of the panel considered is shown in Fig. 1. The panel dimensions consist of 1.2 m length and a 0.8 m width skin of thickness 1.5 mm. The skin is stiffened by 8-longitudinal Z section stringers. In case of integrally stiffened panel, stringers are integral with the skin, whereas a line of rivets is used to fasten the stringer to the skin in riveted stiffened panel. The riveted stiffened panel considered in this study has a rivet spacing of 25 mm as shown in the Figure1: