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OPTIMUM DESIGN OF COMPOSITE WING SPAR SUBJECTED TO BENDING LOADS

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Title: OPTIMUM DESIGN OF COMPOSITE WING SPAR SUBJECTED TO BENDING LOADS  

 

Abstract: Composite materials offer high strength in the desired directions which is important in minimizing aircraft wing weight. Experimental and numerical results will be compared for static bending loads on a composite aircraft wing model. The wing model will consist of an outer skin, two spars, and flat ribs at the root and wing tip. The flat ribs will be cured using a heated press while the rest of the parts will be cured with an autoclave. All wing models will consist of a NACA 0016 airfoil with a chord length of 23.5 cm (9.25 inches), and wing length of 38.7 cm (15.25 inches). [1] The front and rear spars will be located at about 22 % and 72 % chord length respectively. A fixture will be built in order to hold the wing root fixed and also clamp down on the wing skin. Loads will be applied at the wing tip to simulate a bending load as if an aircraft were in flight. Strain values on the wing skin will be measured with the use of strain gages while it is being loaded. Experimental results from 3 wing models of the same design will be used to validate a finite element analysis (FEA) model. It was concluded that the FEA model matched the experimental strain results from the leading edge well. It was also seen that how well the FEA model correlated with the experimental testing was dependent on how well the fixture performed. 

 

Authors: Juan R. Lazarin, Faysal Kolkailah, Eltahry Elghandour  

 

Conference: CAMX 2017 –Orlando

 

SKU/Code: TP17-0046

 

Pages: 14


 

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