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Damages such as a crack/notch and delamination in aerospace, mechanical and offshore structures due to fatigue, corrosion or accidence are inevitable during services. Such damages will grow at an alarming rate due to the stress/strain concentration around the damage locations and cause possible failures of structures. Thus, the structural repair has become an important and practical research topic since the last several decades and attracted much attention in academy and industry. A key objective in a repair design is to lessen the stress/ strain concentration at the damaged part of a structure, e.g. the tips of the notch/crack and delamination, to reinforce the damaged structure. Structural repair with bonded materials has been the most traditionally used technology to increase the service life of damaged structures. The traditional method was to meld or mount additional high stiffness patches onto the damaged area to improve the mechanical function of a damaged structure. In 2002, a patch repair of cracks in the longeron of an F16 aircraft was reported by Hart and Boogers. The main cause for cracking was the occurrence of assembly stresses due to an improperly repaired access panel 2408. Since a conventional mechanical repair could not be done, a cost effective bonded patch repair was evaluated and applied to the cracked longeron. Titanium 6A14 V sheet was used for repair of the 2 mm thick 2024-T62 aluminium longeron flange. A symmetric bonded repair was done with a room temperature curing acrylate based adhesive. The feasibility of the proposed repair geometry was determined by periodic inspections to check on de-bonding and fatigue crack propagation. However, it was noted that the repair process was designed for an air craft with a limited service life (400 flight hours). A major problem is that additional stress concentration may possibly be induced at the bonding area. Moreover, the repair method using normal additional patch cannot adjust itself to newly induced damage due to unexpected external loadings.
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