国电联合动力技术有限公司】前程无忧官方招聘.doc

These loads can cause damage or weakening of the structure especially for aging military and civilian aircraft thereby affecting its load carrying capabilities. Hence, a repair or reinforcement of the damaged or weakened part of the structure to restore the structural efficiency and thus assure the continued airworthiness of the aircraft has become an important issue in recent years to military and civilian aircraft operators. The US Air Force in recent years has shown considerable interest in the use of advanced composites to repair cracked metallic aircraft structures to enhance their life. One issue preventing using bonded composite patches, as a standard means of repairing damaged metallic aircraft structures is the fact that the integrity of the repairs is unknown. In this paper the design, analysis and durability of adhesively bonded composite patch repairs of cracked aircraft aluminum panels is reported. Pre-cracked 2024-T3 clad aluminum panels of 381 89 1.6 mm (15 3.5 0.063 in.) repaired with octagonal single sided boron/epoxy composite patch were used as test specimen. Two different composite ply configurations, 5- and 6-ply were investigated. Linear and non- linear finite element analyses were performed on the test specimen using 8-noded 24 degree of freedom (DOF) hexagonal elements for the aluminum panel, boron/epoxy patch and adhesive material subjected to uni-axial tensile loading. The stress distributions obtained were used to predict the increase in strength and durability of the repaired structure. A comparison of the stress values at critical points was made. The analysis also was used to validate various assumptions made in the design of the composite patch. Experimental investigations were conducted on the cracked aluminum panel repaired with a 5-ply composite patch as well as on two baseline-unpatched panels (one with a crack and one with no crack) by uni-axial tensile testing to validate the analytical results. The experiment was conducted on the Instron tension-testing machine. It was found that the maximum skin stress decreases significantly after the application of the patch and the region of maximum skin stress shifts from the crack front for an unpatched panel to the patch edges for a patched one.Article Outline1. Introduction2. Problem definition3. Design of patch3.1. Overlap length3.2. Patch thickness3.3. Adhesive3.4. Peel stresses3.5. Stress intensity factor3.6. Strain in patch fibers4. Design variables input to CRAS5. Modeling using PatchGen5.1. Geometry5.2. Finite element analysis using Nastran6. Results and discussion7. Experimental verification7.1. Machining of aluminum panels7.2. Fabrication of patches7.3. 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