Design and Optimization of Defense Hole System for Uniaxially Loaded Laminates

Introduction

A composite material consists of two or more materials mixed together to give a material with good properties. A typical composite material consists of a material with high mechanical strength and stiffness (reinforcement), for example unidirectional or woven fibers, embedded in a material with lower mechanical strength and stiffness (matrix). To tailor the properties of the composite material, a laminate is formed by stacking on top of each other layers of reinforcement oriented in different directions.

Composite materials, if properly used, offer many advantages over metals. Examples of such advantages are: high strength and high stiffness-to-weight ratio, good fatigue strength, corrosion resistance and low thermal expansion. Nevertheless, conventional composites made of pre-impregnated tape or fabric also have some disadvantages, such as poor transverse properties, inability to yield and sensitivity to moisture and high temperatures, which must be accounted for in the design.

Composite materials such as glass fiber, aramide fiber, boron fiber and carbon-fiber-reinforced plastics have been used for a few decades, especially in the aircraft industry. Aircraft structures also include a large number of open holes and cut-outs e.g. holes for electric wires and hydraulic pipes or holes required for assembly or maintenance where a laminate containing open holes is subjected to different types of loading.

Reducing stresses in structures and optimizing their weight are the main goals for designers and engineers which improve the structural efficiency, performance and durability. Most engineering structures are assembly of different parts. Parts and components are assembled to the main structure by bolts, rivets, etc. Joining by mechanical fasteners is one of the common practices in the assembly of structural components. Among the most important elements in aircraft structures in general and in composite structures in particular are mechanically fastened joints. Improper design of the joints may lead to structural problems or conservative design leading indirectly to overweight structures and high life-cycle cost of the aircraft. Typical examples of mechanically fastened joints in composite aircraft structures are: the skin-to-spar/rib connections, a wing structure, the wing-to-fuselage connection and the attachment of fittings etc. Since the failure of the joints can lead to the catastrophic failure of the structures, an accurate design methodology is essential for an adequate design of the joints. Because of the complex failure modes of composite materials, the mechanical joining of structures made of composite materials demands much more rigorous design knowledge and techniques than those currently available to the traditional methodology for metallic joints. The holes that are needed for such joints induce stress concentration. These high stress concentration spots are likely places for crack initiation. There have been number of incidents of aircraft fuselage failure resulting from crack initiation in the vicinity of riveted holes. The best known example of this was the Aloha airlines incident in 1988 where it was found that many short cracks exist in the row of the riveted lap splice joints. Cracks propagate under fatigue loading and can eventually link up and cause failure for the whole structure. To prevent such scenario from taking place stress relief system usually introduced within the vicinity of these riveted hole to reduce the stress i.e., increasing the load carrying capacity of the structure and reducing the weight. Up to this moment, most Airframe structures (on commercial level) are made of aluminum and some other alloys except the new Boeing Aircraft B787. This new aircraft is made of laminated composite material with many layers of different thicknesses.