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dc.contributor.advisorYang, Chihdar Charlesen_US
dc.contributor.authorMeka, Uday Sankar
dc.descriptionThesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineeringen
dc.description.abstractLarge composite structures have been increasingly used in the aviation industry. In order to achieve higher fuel efficiency, the use of light-weight, high-strength composite materials, such as carbon/epoxy, needs to be fully explored. New applications of composite materials include primary structures such as aircraft fuselages. This study dealt with thermal stresses induced in a composite aircraft fuselage, in which the fuselage skin was made of carbon/epoxy composite and was fastened to aluminum beams. These stresses resulted from the large coefficient of thermal expansion (CTE) difference and also the large temperature difference between the time of assembly, which was 75ºF and the actual flight condition, which was -65ºF). This temperature difference of around 140ºF induced high thermal stresses, not only in the fasteners but also in the aluminum beams and composite panels. The two main objectives of the study are as follows: To investigate the thermally induced stresses in the aluminum beams. To investigate the feasibility of thermally isolating the aluminum beams from the composite fuselage skins. An experimental program was conducted to measure the strains on the top surface of an aluminum beam, which was fastened to the composite panel from thermal loads due to temperature difference and CTE mismatch. An approach was also designed to study the effects of the length of the aluminum beam on stresses. An analytical model was developed to evaluate the fastener load transfer and the thermally induced stress within the fastened aluminum/composite assemblies. Five parameters were used to develop an analytical model to calculate the load transfer between the aluminum/composite hybrid structures: equivalent area of the aluminum beam and composite panel, equivalent temperatures of the aluminum beam and composite panel, and equivalent fastener stiffness were determined using three-dimensional finite element analysis. An attempt has been made to study the effect of fastener diameter, fastener spacing, material of the metallic beam, size of the metallic beam, thickness of the composite panel on the five parameters required to find the load transfer so that a relation could be established for a working engineer to determine these parameters without doing any finite element work. Equations correlating the five parameters with geometric and material properties were provided.en
dc.format.extentxxi, 151 leaves, ill.en
dc.format.extent2958660 bytes
dc.rightsCopyright Uday Sankar Meka, 2007. All rights reserved.en
dc.subject.lcshElectronic dissertationsen
dc.titleFinite element and analytical models for load transfer calculations for structures utilizing metal and composites with large CTE differencesen

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  • AE Theses and Dissertations
    Electronic copies of theses and dissertations defended in the Department of Aerospace Engineering
  • CE Theses and Dissertations
    Doctoral and Master's theses authored by the College of Engineering graduate students
  • Master's Theses
    This collection includes Master's theses completed at the Wichita State University Graduate School (Fall 2005 --)

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