AIAA-31776-769 - JOURNAL OF AIRCRAFT Vol. 45, No. 2,...

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Thermally Induced Loads of Fastened Hybrid Composite/Aluminum Structures Chihdar Yang, Wenjun Sun, Waruna Seneviratne, and Ananthram K. Shashidhar § Wichita State University, Wichita, Kansas 67260-0044 DOI: 10.2514/1.31776 Large composite structures have been increasingly used in the aviation industry. New applications of composite materials include primary structures such as aircraft fuselages. These large composite parts are sometimes attached, either by the fasteners or adhesive bonding, to metallic structures. Because of the large coef cient of thermal expansion mismatch between the metallic and composite structures, the temperature change from the aircraft assembly line to the actual ight condition induces high thermal stresses during ight in both the composite fuselage and aluminum frames. An experimental program was executed to determine the interaction between the fastened Z- shaped aluminum beams and the solid composite laminate. An analytical model was also developed to simulate the thermal/mechanical behavior of the hybrid composite/metal structure. Finite element analysis was conducted to determine the parameters necessary for the analytical model. The results from the developed analytical model were found to correlate well with experimental results. Nomenclature A a = equivalent cross-sectional area of the aluminum beam A c = equivalent cross-sectional area of the composite panel E a = Young s modulus of the aluminum beam E c = Young s modulus of the composite panel F = load transferred by the fasteners L a = length of a unit of the aluminum beam L c = length of a unit of the composite panel P = bypass force T = present temperature T a = equivalent temperature of the aluminum beam T c = equivalent temperature of the composite panel T o = initial temperature u = displacement in the x direction ± a = coef cient of thermal expansion of the aluminum beam ± c = coef cient of thermal expansion of the composite panel ± u M = change in length due to mechanical load ± u T = change in length due to temperature change " = original strain-gauge reading " M = strain due to tensile/compressive stress " T = strain due to thermal expansion/contraction " T 0 = strain-gauge output due to change in temperature I. Introduction G LOBAL competitiveness has become vital for airplane makers in recent years. To achieve higher fuel ef ciency, the use of lightweight, high-strength composite materials, such as carbon/ epoxy, needs to be fully explored. Large composite structures have been increasingly used in the aviation industry. New applications of composite materials include primary structures such as aircraft fuselages.These largecomposite partsaresometimes attached,either by the fasteners or adhesive bonding, to metallic structures.
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This note was uploaded on 02/24/2011 for the course AERO 520 taught by Professor Rs during the Spring '11 term at Istanbul Technical University.

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AIAA-31776-769 - JOURNAL OF AIRCRAFT Vol. 45, No. 2,...

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