Open AccessThe Impact Behaviour of Carbon Fiber-Epoxy Composite Leading Edge using Finite Element Method
Author(s) -
Nurihan Omar,
Yulfian Aminanda,
Jaffar Syed Mohamed Ali,
Zainudin A. Rasid,
Hafizal Yahaya
Publication year - 2019
Publication title -
international journal of recent technology and engineering
Language(s) - English
Resource type - Journals
ISSN - 2277-3878
DOI - 10.35940/ijrte.b3013.078219
Subject(s) - materials science , projectile , composite material , composite number , enhanced data rates for gsm evolution , finite element method , epoxy , wing , structural engineering , toughness , stacking , fracture toughness , damage tolerance , engineering , telecommunications , physics , nuclear magnetic resonance , metallurgy
Composite material has been widely used in aircrafts due to its high strength to weight ratio that leads to weight saving of the aircrafts. Equally important, aircraft material should be tough i.e. it should have the ability to absorb high energy and thus resist fracture. The aircraft’s wing design requires the material to have high toughness as parts of the wing especially its leading edge is subjected to impact loadings. Using finite element software of LS-DYNA, this research focuses on studying the impact behaviour of composite panels that represent the leading edges of wings when the panels are subjected to rigid sphere projectile. Three shapes of panels are used: flat, semi-circular and semi ellipse while panels can be of 2, 4 and 8 layers to vary its thickness. The panels are made of laminated composites with woven carbon fibres and the angle of orientations are [0/90] n, [0/45]n and [45/-45]n where n will give the number of layer for the composite. The Mat-58 material type suitable for woven type fibre is used where failure criteria of Hashin is applied. It was found that the simulation results are in a very close agreement with the finding from experiments conducted earlier. Furthermore, the optimum stacking sequence was found to be the [0/45]2 stacking sequences