PSI - Issue 21

Hande Yavuz / Procedia Structural Integrity 21 (2019) 112–119 H. Yavuz/Structural Integrity Procedia 00 (2019) 000 – 000

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4. Conclusions

Materials selection plays a crucial role for the determination of candidate materials for a specific engineering application. CES Edupack software, which ensures prompt visualization of such materials, was used along with material selection charts. To construct light and safe panels, one of the best candidate materials was identified as carbon fiber reinforced epoxy composites in the frame of Ashby’s multiple constraints design methodology. The mechanical damage analysis of various composite laminates were then explored via MATLAB and Abaqus simulations. Considering Hashin model for damage analysis in unnotched laminates, the highest matrix tensile failure load was attributed to [+45/-45/0/90] s laminate which indicates ~17% difference with the [+45/+30/0/90] s laminate. Among the studied notched laminates, which have the same ply orientations as of unnotched ones, the highest matrix tensile failure load was again determined for [+45/-45/0/90] s laminate which possesses ~14% difference with the [+45/+30/0/90] s laminate. In all notched laminates, the effect of stress concentration on damage evolution was clearly detected from hole edge to hole boundary. Moreover, damage initiates and propagates in QI laminate due to stresses developed near free edges. These results show that either unnotched or notched laminates, [+45/-45/0/90] s configuration would likely to be prefferred over [0/+45/-45/90] and [+45/+30/0/90] s configuration with respect to uniaxial tensile loading in the frame of Hashin failure criteria. In further studies, skin panels may be considered as twisting plates along with multiple objectives such as minimum mass and minimum cost. In this way, mass and cost objectives can be aggregated into single solution by defining a locally linear value (or utility) function called penalty function. This function should be used to develop exchange constant(s) as long as cost and mass objectives considered together. Besides, computational structural analysis may also be performed on stiffened composite plates having various stacking sequence including mechanical or chemical fasteners instead of single notched laminates by adjusting various natural and essential boundary conditions. However, in case the quasi-static behavior of such laminated structures is intended to be researched, determination of the ratio of kinetic energy to internal energy and evaluations of two energies individually should be realized in order to conclude whether results are reasonable. These aspects would also be associated with kinetic energy history which indicates quasi-static behavior. Furthermore, experimental studies would also be embedded in such work in order to validate finite element analysis for each configuration. Interlaminar failure analysis (e.g., delamination in skin panel) which may induce decrease of the local moment of inertia of a skin panel due to ply separation would be conducted using cohesive zone method. References ASTM D3039/D3039-14, 2014. Standard Test Method for Tensile Properties of Polymer Matrix Composite Materials, ASTM International, USA. ASTM D5766/D5766M-11, 2018. Standard Test Method for Open-Hole Tensile Strength of Polymer Matrix Composite Laminate, ASTM International, USA. Bristow, J.W., Irving, P.E., 2007. Safety Factors in Civil Aircraft Design Requirements. Engineering Failure Analysis 14, 459 – 470. CES Edupack, 2018. Granta Design Limited, Cambridge, UK. Dababneh, O., Kayran, A., 2014. Design, Analysis and Optimization of Thin Walled Semi-Monocoque Wing Structures Using Different Structural Idealization in the Preliminary Design Phase, International Journal of Structural Integrity 5, 214 – 226. De Florio, F., 2016. Airworthiness: An Introduction to Aircraft Certification and Operations. Butterworth-Heinemann, Cambridge, MA, USA, pp. 37 – 83. Degenhardt, R., Castro, S.G.P., Arbelo, M.A., Zimmerman, R., Khakimova, R., Kling, A., 2014. Future Structural Stability Design for Composite Space and Airframe Structures. Thin-Walled Structures 81, 29 – 38. Duman, A., Çelensü, B., Öçal, B., 2019. Failure Analysis in Laminated Composites for Aerospace Applications, Senior Design Project, University of Turkish Aeronautical Association, Ankara, Turkey. Gay, D., 2015. Composite Materials: Design and Applications. CRC Press, Taylor & Francis Group, Boca Raton, FL, USA, pp. 161. Gou, S., 2007. Aeroelastic Optimization of an Aerobatic Aircraft Wing Structure, Aerospace Science and Technology 11, 396 – 404. Hashin, Z., 1980. Failure Criteria for Unidirectional Fiber Composites, Journal of Applied Mechanics 47, 1083 – 1094. Hinrichsen, J., Bautista, C., 2001. The Challenge of Reducing Both Airframe Weight and Manufacturing Cost, Air & Space Europe 3, 119 – 121. Johnson, A.F., Thomson, R. S., David, M., Joosten, M.W., 2015. Design and Testing of Crashworthy Aerospace Composite Components, in “ Polymer Composites in the Aerospace Industry ”. Irving P., Soutis C. (Eds.). Woodhead Publishing, Waltham, MA, USA, pp. 261 – 293. Jorgensen, O., 1991. Optimization of the Flutter Load by Material Orientation, Mechanics of Structures and Machines 19, 411 – 436. Kaplan, S.S., Çetin, A., Yavuz, H., 2017. Design of Internal Rods in Central Wing-Box of an Aircraft, 8 th North American Materials Education Symposium. Cambridge, MA, USA, paper #2. ABAQUS v6.14, Dassault Systemes Simulia Corp, Providence, RI, USA. Ashby, M.., 2011. Materials Selection in Mechanical Design, Butterworth-Heinemann, MA, USA, pp. 197 – 216.