The aim of this work is to describe the structural analysis of a multifunctional aircraft fuselage panel. The structure of the panel has an embedded antenna tiles. The panel consists of UniDirectional (UD) carbon fibre reinforced composite skin stiffened with ortho-grid ribs, and a transparent skin window made using UD glass fibre reinforced composite. The orthogrid structure is a structural reinforcement but also the antenna tiles support. The presented work proposes a numerical multiscale strategy. The laminate is simulated with solid elements, in order to capture the real kinematics of the material, but several laminas are condensed in a single finite element. The performance of each lamina is obtained using the Serial-Parallel (SP) mixing theory. The specific formulations developed have been very useful to identify and study the mechanical performance of these new structures and the localization of unknown and un-predicted hot-spots in the structure.Postprint (published version

Martínez García, Javier

Otero Gruer, Fermín Enrique

Turon Pujol, Francesc

English

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Structural Analyses of Orthogrid Fuselage Panel for Integrated Ku-band SatCom Antenna.F. Turon, F. Otero and X. MartinezEuropean Conference on Multifunctional StructuresEMuS 2019X. Martinez and H. Schippers (Eds)STRUCTURAL ANALYSES OF ORTHOGRID FUSELAGE PANEL FORINTEGRATED KU-BAND SATCOM ANTENNA – EMUS 2019FRANCESC TURON∗, FERMIN OTERO∗ AND XAVIER MARTINEZ†∗∗ International Center for Numerical Methods in Engineering (CIMNE)Technical University of Catalonia (UPC)Gran Capitn s/n, 08034 Barcelona, Spaine-mail: fturon@cimne.upc.edu, fotero@cimne.upc.edu† Departamento de Ciencia e Ingeniera Nutica, FNBTechnical University of Catalonia (UPC)Pla de Palau 18, 08003 Barcelona, Spaine-mail: x.martinez@upc.eduKey words: Composites, Numerical Methods, Antenna, Panel, Multifunctional, Serial-ParallelMixing TheoryAbstract. The aim of this work is to describe the structural analysis of a multifunctional aircraftfuselage panel. The structure of the panel has an embedded antenna tiles. The panel consistsof UniDirectional (UD) carbon fibre reinforced composite skin stiffened with ortho-grid ribs,and a transparent skin window made using UD glass fibre reinforced composite. The ortho-grid structure is a structural reinforcement but also the antenna tiles support. The presentedwork proposes a numerical multiscale strategy. The laminate is simulated with solid elements,in order to capture the real kinematics of the material, but several laminas are condensed ina single finite element. The performance of each lamina is obtained using the Serial-Parallel(SP) mixing theory. The specific formulations developed have been very useful to identify andstudy the mechanical performance of these new structures and the localization of unknown andun-predicted hot-spots in the structure.1 INTRODUCTIONThe Fokker 100 is the target aircraft to integrate the ortho-grid stiffness fuselage panel withthe integrated antenna. The forward crown panel at the top of the fuselage is selected for theortho-grid panel location, as in this section the antenna will have the best performance with thelower loads. It should be noted that in this work no aerodynamic loads are included, and that avariation of temperature conditions is not foreseen.The panel is manufactured using thermoset prepreg materials, which are suitable for aerospacestructural applications. The panel skin and the ortho-grid ribs consist of carbon fibre reinforcedmaterial: Hexcel 8552 resin and AS4 fibre. On the other hand, the transparent panel windowskin consists of glass fibre reinforced material: Cytec FM906-27 resin and S2 glass 187-460fibre.The ortho-grid ribs are laid-up using automated fibre placement over the panel skin. The162Francesc Turon, Fermin Otero and Xavier Martinezpanel skin has a material transition in the glass window edges through interleave plies of carbonand glass. Finally, the whole fuselage panel is cured in an autoclave.Even with nowadays computational capacity it is almost impossible to analyse a structuralcomponent taking into account a detailed definition of the laminate layup. For this reason, inthis work a numerical multiscale strategy is proposed. The laminate is simulated with solid ele-ments, in order to capture the real kinematics of the material, but several laminas are condensedin a single finite element. The performance of each lamina is obtained using the serial-parallelmixing theory.2 SERIAL-PARALLEL MIXING THEORYThe serial-parallel mixing theory could be defined as a phenomenological homogenization,where the behavior of the composite is obtained from the constitutive response of their materialscomponents. This theory has been developed by Rastellini et al. [1], and is a natural evolutionof the parallel mixing theory developed by Car et al. [2]. The theory is based on the com-patibility conditions, but introduces a modification in the iso-strain hypothesis. The iso-straincondition is imposed in the reinforcement direction and a new iso-stress condition is imposed inthe transversal directions. Commonly, for UD composites, the formulation applies an iso-straincondition in fibre direction and a iso-stress condition in the other ones.Taking only two composite components, the equations that define the stress (σ) equilibriumand setting up the strain (ε) compatibility between the individual components follow the hy-pothesis previously described are:Parallel behavior: cεp = mεp = f εpcσp = mkmσp + f k f σp(1)Serial behavior: cεs = mkmεs + f k f εscσs = mσs = f σs(2)where the superscripts c, m and f stand for composite, matrix and fibre, respectively and ikis the volume-fraction coefficient of each constituent in the composite.This theory can predict the linear and non-linear behavior of structural elements made ofcomposite materials. Composite materials that can be modelled are those formed of long fibresembedded in a matrix. The theory predicts the different behavior of the composite, dependingon the load direction. The potential of this theory is to predict accurately the response of com-posites in the linear and non linear range as has been proved in several papers [3, 4]. Amongthe different failure modes that can be captured, in [5] it is shown that the serial-parallel mixingtheory is able to simulate the delamination problem naturally, without having to define specificelements or predefine the path of fracture.3 NUMERICAL MODEL DESCRIPTIONFinite element numerical models are used to study the structural response of the panel fordifferent loading cases. To reduce the computation time required for the simulations, and tobe able to include all details in the models developed, only a reduced section of the fuselage263Francesc Turon, Fermin Otero and Xavier MartinezFigure 1: Reduce numerical model for the ortho-grid fuselage panel.is analysed. In order to obtain the same results that would be reached with the whole panelconsistent boundary conditions are imposed.In the Figure 1, the central green window represents the UD glass fibre reinforced compositetransparent window skin. While the grey part of the skin uses UD carbon fibre prepreg material.Figure 2 shows in detail the different geometries in which every part of the model is discretizedto model the connection between skin/ribs, rib/rib and the glass/carbon skin transition.To represent the transition regions, it is necessary to define models with a high level of detailin order to account for changes in layer orientation, composite intersections (glass to carbon),and all different specificities shown in Figure 3. Consequently the quantity of elements requiredby the model is high, which increases the computational cost of the numerical analysis, in termsof time and memory. In order to improve the machine time and simplify the numerical model anumerical multiscale approach is proposed.In the numerical model, the laminate is simulated with solid finite elements, to capture thereal kinematics of the material, but several layers are condensed in a single finite element. Theperformance of each lamina is obtained using the SP theory. On the right side of the Figure3 the homogenized transition zone is shown. As can be seen, there are three element throughthe thickness, and each one of these elements contains internally several layers. The upper partof the figure shows the numerical model developed while in the lower part are represented thedifferent layers assigned to each region of the model.364Francesc Turon, Fermin Otero and Xavier MartinezFigure 2: Detailed view of the glass/carbon skin transition, and rib connections.Figure 3: Detailed view of the glass/carbon and homogenized model skin transition.4 RESULTSIn the following it is presented the numerical results obtained for a tension load case. Atension flow (261.5 N/mm) in the top side and a fixed support in the bottom side of the panelis imposed. It is also applied an the internal pressure (0.08 N/mm2) of the fuselage. Therefore,specific boundary conditions have to be imposed on both lateral side of the panel.The total displacement field obtained is shown in Figure 4. The left side shows a inside viewof the displacements obtained and the right side a external view of the panel. The maximumvalues of displacements are presented in the top of the panel because of the tension load, andthey are extended to the middle of the panel for the effect of the internal pressure. In Table 1the most relevant displacement results are presented.The longitudinal ribs of the panel take most of the load applied. A mean stress value ofaround 320 MPa is presented in the longitudinal ribs. The maximum value of the principalstress obtained is located on the longitudinal rib close to the Glass/Carbon skin transition in therib connection. The rib connections have half of the UD carbon plies cut to allow the continuity465Francesc Turon, Fermin Otero and Xavier MartinezTable 1: Relevant displacement results obtainedMax. Long. Displ. (Z dir.) [mm] 2.35Max. Radial Displ. [mm] 1.85Mean Radial Displ. [mm] 1.1Max. Relative Radial Displ. [mm] 0.75Figure 4: Total displacement field obtained for the tension load case.of the plies in the perpendicular direction. The maximum stress obtained is around 700 MPaand it is found in the continue layers of the rib connection. The left side of Figure 5 shows thecritical zone regarding the stresses level. Another critical area is the region in the connectionwhere the cut layers end, as this region contains only matrix and the stresses are close to 60MPa. This stress level is close to the maximum ultimate strength of the matrix component.The skin and the material transition zones do not reach the maximum design strain value (i.e.3600 μ%). However, in the rib connections and the rib/skin connection the principal strain fieldpresents higher values. An amplified view of the rib connection is shown in the right side ofFigure 5. For the additive material in the ribs connection, a higher design value of maximumstrain could be accepted. The mentioned figure shows a strain value in this material around6000 μ%.5 CONCLUSIONSA multiscale strategy that combines the finite element method and the SP formulation isproposed. The numerical approach presented allows the analysis of complex composite struc-566Francesc Turon, Fermin Otero and Xavier MartinezFigure 5: Zone with the maximum value of the principal stresses in the fuselage panel.tures with an affordable computational cost, which is basic for the development of new mul-tifunctional composite structures. The formulation developed has been very useful to studythe mechanical performance of the ortho-gid fuselage panel, and to identify the localization ofunknown and un-predicted hot-spots in the structure.REFERENCES[1] Rastellini F. and Oller S. and Salomón O. and Oñate, E. Composite materials non-linearmodelling for long fibre-reinforced laminates continuum basis, computational aspect andvalidations. Computers and Structures. (2008) 86:879-896.[2] Car E. and Oller S. and Oñate, E. An anisotropic elastoplastic constitutive model for largestrain analysis of fiber reinforced composite materials. Computer Methods in Applied Me-chanics and Engineering. (2000) 185:245-277.[3] Martinez X. and Oller S. Numerical simulation of matrix reinforced composite materialssubjected to compression loads. Archives of computational methods in engineering. (2009)16:357-397.[4] Otero F. and Oller S. and Martı́nez X. and Salomón O. Numerical homogenization forcomposite materials analysis. Comparison with other micro mechanical formulations.Composite Structures. (2015) 122:405-416.[5] Martinez X. and Oller S. and Barbero E. Study of Delamination in Composites by Usingthe Serial/Parallel Mixing Theory and a Damage Formulation. In: Mechanical Responseof Composites. Computational Methods in Applied Sciences, vol 10. Springer, Dordrecht(2008)667

Structural analyses of orthogrid fuselage panel for integrated Ku-band SatCom antenna

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Structural analyses of orthogrid fuselage panel for integrated Ku-band SatCom antenna

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