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Introduction. Preventing failure of composite material systems has been an important issue in engineering design. There are two types of physical failures that occur in laminated composite structures and interact in complex manner are intralaminar and interlaminar failures. Intralaminar failure is manifest in micromechanical components of the lamina such as fiber breakage, matrix cracking, and debonding of the fiber-matrix interface. Generally, aircraft structures made of fiber reinforces composite materials are designed such that the fibers carry the bulk of the applied load. Interlaminar failure such as delamination refers to debonding of adjacent lamina. The possibility that intralaminar and interlaminar failure occur in structural components is considered a design limit, and establishes restrictions on the usage of full potential of composites. Due to the lack of through-the-thickness reinforcement, structures made from laminated composite materials and adhesively bonded joints are highly susceptible to failure caused by interfacial crack initiation and growth. The delamination phenomenon in a laminated composite structure may reduce the structural stiffness and strength, redistribute the load in a way that the structural failure is delayed, or may lead to structural collapse. Therefore, delamination is not necessarily the ultimate structural failure, but rather it is the part of the failure process which may ultimately lead to loss of structural integrity.Most of the components on the aircraft are increasingly being replaced with composite materials. The main attraction is the effective reduction in mass with a comparative increase in stiffness, strength, fatigue and impact resistance, thermal conductivity and corrosion resistance. Through these replacements, the structural weight can be reduced, which will in turn lead to a more economical commercial aircraft [1]. The major structural applications for fiber-reinforced composites are in the field of military and commercial aircrafts, for which weight reduction is critical for higher speeds and increased payloads. Ever since the production application of boron fiber-reinforced epoxy skins for F-14 horizontal stabilizers, the use of fiber-reinforced polymers has experienced a steady growth in the aircraft industry. Carbon fiber-reinforced epoxy has become the primary material in many wings, fuselage, and empennage components. The structural integrity and durability of these early components have built up confidence in their performance and prompted developments of other structural aircraft components, resulting in an increasing amount of composites being used in military aircrafts. The F-22 fighter aircraft also contains nearly 25% by weight of carbon fiber-reinforced polymers. The outer skin of B-2 and other stealth aircrafts is almost all made of carbon fiber-reinforced polymers. The stealth characteristics of these aircrafts are due to the use of carbon fibers, special coatings, and other design features that reduce radar reflection and heat radiation [2].A...

This paper examines critical load and corresponding displacement of double cantilever beam (DCB) composite specimens made of glass/epoxy of three different layups. Experiments were conducted on these laminates, and the fracture energy, Ic , was evaluated considering the root rotation at the crack tip. The present model requires the applied load-displacement history and crack extension to estimate fracture energy. Reduction schemes based on cubic and power law are also proposed to determine Young's modulus and energy release rate and found good agreement with the published and test results.

This article examines the fracture toughness of end notched flexure (ENF) composite specimens of three different lay-ups. Experiments were conducted on these glass/epoxy specimens and the critical fracture energy,, was evaluated based on compliance based beam method (CBBM). Classical methods require crack length measurements, which are not easy to obtain as propagation occurs rapidly without a clear opening. The CBBM is based on crack equivalent concept, which does not require crack length monitoring during propagation and hence the crack growth resistance curve (R-curve) can be generated in a much easier way. Moreover, the CBBM accounts the non-negligible energy dissipation in the fracture process zone (FPZ) in addition to stress concentrations near crack tip, contact between specimen arms at the pre-crack region and root rotation effects. Hence, the complete R-curves of ENF specimens of different lay-ups were obtained using the CBBM with higher degree of accuracy. It was observed that the unidirectional specimen did show higher propagation toughness value than the angle-ply and cross-ply specimens.

Large deviations have been observed while analysing composite double cantilever beam (DCB) specimens assuming each cracked half as a simple cantilever beam. This paper examines the effect of rotational spring stiffness(K)on the critical fracture energy(GIC)considering nonzero slope at the crack-tip of the DCB specimen by modelling each cracked half as the spring-hinged cantilever beam. The critical load estimates of DCB specimens fromGICare found to be in good agreement with in-house and existing test results of different composite material systems.