Conclusion
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Conclusion The response of composite sandwich fuselage side panels with window cutouts has been evaluated for internal pressure and axial tension. The panels have been tested with combined loading conditions that are representative of the design limit load and design ultimate load conditions. The strain magnitudes around the cutouts on the outer surfaces of the test panels for these loading conditions for the material, suggesting that the structure satisfies the design requirements. The finite element analytical results compare very well with the experimental results. For a panel made of sandwich pattern frame configuration, the finite element analysis and experimental results correlate well and the strain results are less than the ultimate strain\ allowable for the material. The damage tolerance of them panel is also demonstrated by testing the panel at design limit load conditions with a notch at a window cutout region that is in the location of the highest value of axial stress. For this case, the maximum value for the axial strain obtained from the test and from the analysis is also compared and shown such that the panel made of composite sandwich have minimum strain than panel made of aluminum and having minimum weight which is an important parameter in the aircraft design. Future Work We are planning to do investigation on the same panel with honeycomb structure as a core for the sandwich pattern and expecting a better result.
description
composite
Transcript of Conclusion
ConclusionThe response of composite sandwich fuselage side panels with window cutouts has been evaluated for internal pressure and axial tension. The panels have been tested with combined loading conditions that are representative of the design limit load and design ultimate load conditions.The strain magnitudes around the cutouts on the outer surfaces of the test panels for these loading conditions for the material, suggesting that the structure satisfies the design requirements. The finite element analytical results compare very well with the experimental results. For a panel made of sandwich pattern frame configuration, the finite element analysis and experimental results correlate well and the strain results are less than the ultimate strain\ allowable for the material. The damage tolerance of them panel is also demonstrated by testing the panel at design limit load conditions with a notch at a window cutout region that is in the location of the highest value of axial stress. For this case, the maximum value for the axial strain obtained from the test and from the analysis is also compared and shown such that the panel made of composite sandwich have minimum strain than panel made of aluminum and having minimum weight which is an important parameter in the aircraft design.
Future WorkWe are planning to do investigation on the same panel with honeycomb structure as a core for the sandwich pattern and expecting a better result.
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