R. Jones

Damage tolerance of graphite/epoxy composites

Abstract A major long-term objective in the evolution of advanced fibre composites as aircraft structural materials must be the establishment of a damage tolerance methodology. The availability of this methodology would greatly aid the design of safe efficient composite structures as well as their management in service. There are, however, many difficulties in meeting this objective, including the multiplicity of failure modes in the composites, the numerous types of potentially significant defects, which may arise during manufacture or in service, and sensitivity to moisture and temperature. This paper briefly considers the present status of the damage tolerance approach for composites and then highlights the critical issues as they relate to the problem of delamination damage, an aspect of primary concern. Selected theoretical and experimental studies in the literature, providing at least partial answers to some of these issues, are briefly discussed and a unified theoretical approach is proposed.

Repair of mirage III aircraft using the BFRP crack-patching technique

Abstract The BFRP crack-patching technique has been applied to the field repair of fatigue cracks in the aluminium alloy wing skins of Mirage III fighter aircraft. Finite-element procedures were used in patch design. The repair was qualified using fatigue-crack propagation studies on panels simulating the cracked and repaired area. A field support unit was designed to allow repairs to be carried out by air force personnel during routine maintenance of the aircraft. To date over 150 patches have been applied and nearly three years of operational history gained. While some crack growth was observed after repair of a few wings, the patch stopped further growth and no wing skin has required further repair.

Crack Patching: Analysis and Design

ABSTRACT Bonded composite patches are frequently used to retard crack growth. This repair procedure is usually referred to as crack patching. The present paper outlines the various methods for the analysis and design of fiber composite patches in thin and thick structures. As illustrative examples the repair of fatigue cracks in the wing skins of Mirage III aircraft, of surface flaws in Macchi landing wheels, and of cracks in a truss is considered.

Designing for damage tolerant bonded joints

Abstract This paper discusses the damage tolerant design methodology for adhesively bonded joints and bonded repairs. Attention is focused on the effects of rate dependence, creep, and strain/load holds on the load carrying performance of structural adhesives.

Characterisation of pure and mixed mode fracture in composite laminates

Abstract The extensive use of advanced fibre composite materials for aircraft construction has necessitated the development of a damage tolerance methodology for aircraft components. Such a methodology would facilitate the design of more efficient and reliable composite structures and their maintenance in service. Therefore, it is necessary to characterize and understand the complex failure modes of fibre composites, including the influence of temperature, moisture and various defects arising from manufacture or service conditions. This paper briefly discusses the present status of some approaches to the experimental characterization of pure and mixed-mode fracture of composite laminates.

Assessment of the effect of impact damage in composites: Some problems and answers

In order to provide through-life support for structural components, fabricated from advanced composite materials, it is desirable to establish a damage tolerance methodology. The availability of this methodology would greatly aid the design of safe, efficient composite structures as well as their management in service. There are, however, many difficulties in meeting this objective, including the multiplicity of failure modes in the composites, the numerous types of potentially significant defects which may arise during manufacture or in service and the sensitivity to moisture and temperature. This paper discusses the analysis and testing of impact damage composite laminates. Selected experimental studies in the literature are briefly discussed. n nIt is shown that uniaxial S−N curves for damaged laminates have a generic shape with a pronounced threshold level. Indeed, it is clear that further research is needed to understand the physical reasons for such a threshold value.

Application of the T∗-integral and S criteria in finite element analysis of impact damage at fastener holes in graphite/epoxy laminates under compression

This paper illustrates the application of two fracture criteria, Atluris T∗-integral and Sihs strain energy density factor S for estimating the residual strength of an impact damaged fastener hole in a composite laminate. Finite element analyses are performed, and the magnitude and distribution of T∗ and S are determined around the delamination. It is found that the profiles of these distributions are extremely complex and sensitive to the modelling of the stress fields close to the delamination. The effects of local closure, shear moduli and cracktip singularity are investigated. It is also shown that the representation of the stress singularity in the finite element model has a strong effect on the distribution of T∗ and S. The distribution of both T∗ and S is such that three local maxima occur and are situated at the same locations around the delamination. These locations approximately coincide with the points of maximum growth as revealed by ultrasonic C-scan of the damage growth of several specimens. The T∗-integral is shown to be more sensitive to the values of interlaminar shear moduli than S. The results of the analyses suggest that both the T∗-integral and S criteria may be successfully employed in the prediction of residual strength. However, S is better at predicting the direction of damage growth provided the stress fields near the delamination can be accurately modelled.

Is fatigue testing of impact damaged laminates necessary

Abstract A major long-term objective in the evolution of advanced fibre composites as aircraft structural materials must be the establishment of a damage tolerance methodology. The availability of this methodology would greatly aid the design of safe, efficient composite structures as well as their management in service. There are, however, many difficulties in meeting this objective, including the multiplicity of failure modes in the composites, the numerous types of potentially significant defects which may arise during manufacture or in service and sensitivity to moisture and temperature. This paper discusses the necessity of fatigue testing of impact damaged composite laminates. Selected experimental studies in the literature are briefly discussed. It is shown that uniaxial S-N curves for damaged laminates have a generic shape with a pronounced threshold level. Indeed, it is clear that further research is needed to understand the physical reasons for such a threshold value.

Effects of plasticity on several fracture parameters for impact damage and related problems

Abstract Apart from temperature and moisture effects, matrix plasticity is another factor that may influence the behaviour of delamination damage in composite materials. The present paper illustrates how, for a range of problems, matrix plasticity does not affect the governing fracture parameters.

A Hybrid Contour Method

ABSTRACT This paper discusses the use of hybrid and isoparametric elements in elastic stress analysis and elastic-plastic fracture mechanics, A new approach is developed that involves the use of a hybrid stress recovery method together with isoparametric elements. This procedure, which we will call the hybrid contour method, is an efficient alternative method for calculating the J integral. One advantage of this approach is that it accesses only the displacement field and, like J, appears to be path independent.