ANALYSIS OF A GLASS-FIBRE SANDWICH PANEL FOR CARBODY CONSTRUCTIONS

Transcription

1 9th International Conference on Composite Structures (ICCS / 9) ANALYSIS OF A GLASS-FIBRE SANDWICH PANEL FOR CARBODY CONSTRUCTIONS by R.Barboni, P.Gaudenzi and A.Pascucci Università di Roma La Sapienza Dipartimento Aerospaziale, Roma, Italia A.Horoschenkoff Daimler-Benz AG, Research and Technology, Munich, Germany Abstract The analysis of a glass-fibre sandwich panel for railway applications is considered. The influence of a special fire-resistant treatment on the mechanical properties of the structure is analysed for static and cycling loadings. The main design and simulation issues are first highlighted. The failure loads of the composite panel are then evaluated from a numerical and an experimental point of view both for the laminated skin of the sandwich and for the entire structure. 1. INTRODUCTION The possible use of composite materials in secondary and primary structures has recently received more and more attention within the framework of railway applications [1],[2],[3]. In particular, some interest has been devoted to the adoption of sandwich panels made by composite laminated skins with an insulating core for the construction of railway carriages. In fact, several advantages of such a technology as compared to more traditional ones can be easily identified: Weight reduction Integrated acoustic insulation Integrated thermal insulation 1

2 Reduction of finishing work Reduction of the assembly phases While these advantages may offer cost savings, some drawbacks are expected due to: Joining difficulties Manufacturing process still in its "infancy " stage Difficulties of quality control Maintenance aspects (repair, recyclability) One of the most important requirements of the component is the fire resistance and the emission of toxic gases in the case of fire. In this paper, the design and testing of glass fibre sandwich panels for carbody constructions is analysed. The structure is first described in some detail. Then the main design and simulation issues are highlighted and some numerical predictions are presented. In particular, the loss of symmetry of the skin laminate and the non-linearity of its constitutive behaviour due to the treatment of the matrix for fire resistance are investigated. 2. DESCRIPTION OF THE PANEL The analysed sandwich is a segment of the side wall of a carriage. Each face of the sandwich panel is a fibre-reinforced laminate, whose thickness is t f =4mm; the thermosetting matrix is vinylester with the addition of aluminium trioxide powder (ATH), used to improve the fire resistance, while the composite reinforcement is glass fabric. The core is polymeric foam, with a thickness of t c =40mm; the skins are adhesively bonded to the core materials (epoxy-based adhesive). One skin essentially consists of two outer plies which are the fireresistant part of the skin and of three plies which are the carrying part of the skin (it is placed in proximity to the core). The two parts of the skin are distinguished by the amount of ATH in the matrix and the fibre-reinforcement. The thermosetting matrix mixed with ATH has better fire resistance, but lower mechanical properties. Hence the powder weight is three times greater in the outer plies than in the inner plies. The stacking sequence of the skin is illustrated in Fig.1, while the characteristics of the plies are described in Tab.1. In this table it can be noted that 2

3 the fibre content is higher in the inner laminate, in accordance with the defined task. Due to the presence of fire-resistant layers, the laminate is not symmetric. Furthermore, the lower fibre content and the lower toughness of the matrix, caused by ATH, results in other critical effects, such as the reduction of stiffness and strength of the laminate and the nonlinearity of the stress-strain relation. 3. INFLUENCE OF ATH ON THE MECHANICAL BEHAVIOUR The ATH powder is an inorganic substance Al 2 O 3, used to increase the fire resistance of the polymeric material in which it is inserted. The principle is based on an endothermic reaction produced at about 200 C, which generates water molecules and aluminium oxides. Besides, the ATH powder is relatively cheap compared to other modifiers and it does not cause any manufacturing problems, but on the other hand it causes some disadvantages: substantial reduction of strength and stiffness. Laminates (made of 9 layers having the same fibre directions) constituted with glass fabric, vinylester matrix and several amounts of ATH have been carried out: 0 ATH ATH is not used 60 ATH ATH weight = 60% matrix weight 180 ATH ATH weight = 180% matrix weight The laminates were subjected to fire resistance tests, whose results are illustrated in Tab 2. Tensile and shear tests have been performed; it has been noted that the Young s modulus is substantially constant in two intervals. The Young s moduli of both intervals and the knee strain (strain range between them) decrease with increasing ATH content. Engineering constants, ultimate strains and strengths of several laminates, owing to different ATH powder fractions, are illustrated in Tab.3 and Tab ANALYTICAL STUDY Structural characteristics of the laminate (skin) The structural characteristics of the laminate, which is the sandwich skin, have been computed by means of the classical lamination theory. The experimental data, illustrated in Tab.3, Tab.4, have been used to assign the 3

4 properties of plies in the principal directions; the layer properties in the weft direction were reduced by 10%, compared to the ones in warp direction. Hence, checking of normal strain in local references with increasing load has been used to consider the two intervals in which the Young s moduli are constant. Since the constitutive behaviour of each fabric layer is nonlinear, the local inplane stresses σ 1, σ 2, τ 12 of every ply have a nonlinear dependence upon the external tensile load (Fig.2, Fig.3, Fig.4). The diagram is arrested at the first ply failure load of the laminate. The nomenclature of the laminate used in such graphs is such that the plies are numbered from inside to outside (see Fig.1). For each of the considered cases (traction, bending, shear and torsion) the responsible of the first ply failure is always the shear stress component. For tensile and bending load cases the first broken layer is the fire-resistant lamina at (±45), while for the shear and torsion load cases the first ply to fail is the fireresistant lamina at (0,90). Calculated first ply failure and ultimate loads of the laminate are illustrated in Tab.5. The first ply failure load is about 50-60% of the ultimate load of the laminate; for the tensile load it appears as normal strain of the middle plane ε x =0.79%. In the tensile case, a strong decrease of the Young s modulus for a mean stress σ x =40N/mm 2 and a strain ε x =0.2% can be observed. On the contrary, the Young s modulus is reduced constantly with increasing load for the bending case and it negligibly decreases for shear and torsion loads. In Tab.6, the decrease of the engineering coefficients of the laminate, loaded up to first ply failure, are summarised. Failure modes of the sandwich The loads, causing fracture of the face for the cases traction, constant bending and in-plane shear, have been calculated by means of the classical lamination theory. In addition, the core shear failure and wrinkling loads have been determined, too. The numerical results for a test sandwich panel (110x80cm) are summarised in Tab.7. The critical compressive loads on the test sandwich panel have been performed by means of the finite element program ADINA [7], by using the composite shell element; the skin (whose properties were experimentally determined) has been considered an orthotropic layer while the core has been 4

5 treated as an isotropic medium. The mesh used for the sandwich panel has 6 elements for each edge, every element having 8 nodes and each node with 6 degrees of freedom, which total degrees of freedom. In Tab.8, critical loads for several cases of boundary conditions are shown. The critical shear load for the simply supported plane, calculated with the F.E. method, has the value N xy = N. It can be noted that for each load (N x, M x, N xy, M xy ) the sandwich failure occurs at levels which are lower than the ones due to local or global instability. Finally, the four-point bending results have been determined analytically: the value of the distance between the inner and outer supports, above which first ply failure occurred and below which core shear failure occurred, is L=93.5cm, which corresponds to a load P of 4 610N. 5. EXPERIMENTS Structural characteristics of the laminate (skin) Firstly, tensile, three-point bending and shear tests on the laminate considered have been performed; Fig. 5 and Fig. 6 show the graphs (normal stress versus strain) of tensile and bending experiments, respectively, while Tab. 9 and Tab 10 describe the results. Besides, the ultimate shear stress has also been determined: τ xy(ultimate) =99.2N/mm 2. In Fig. 7, experimental forms of Young s modulus with increasing load for the tensile and the bending cases are illustrated. The form of Young s modulus for the tensile case presents a strong discontinuity compared to the bending case, for which the Young s modulus decreases rather constantly. For both tensile and bending cases, specimens have been subjected to increasing loads up to a load level causing first ply failure, with the objective of identifing the reason for the nonlinear behaviour of the composite; Fig. 8, Fig. 9, and Fig. 10, Fig.11 show both surfaces of the laminate for traction and bending respectively. The formation of microcracks in the matrix, which are arranged in transversal direction to the load, could be observed on the surfaces of the laminates. This structural damage clearly reduces the mechanical properties. In 5

6 the tensile case the fire-resistant layer shows a higher density of cracks compared to the load-carrying ply; it means that the ATH makes the matrix brittle. Besides, on the surface of the load-carrying ply one observes a first phase of the formation of damage; the formation of cracks has focussed in the zone of the intersection among warp and weft, particularly among the weft fibres in the proximity of the external surface. Then, for the tensile and bending cases, the first ply failure has been studied in detail; in fact specimens, subjected to first ply failure load, have been analysed with C-scan and in addition micrographs of cross sections have been performed, as is illustrated in Fig.12 and Fig.13. The micrographs show cracks in the fire-resistant ply, with the principal direction orientated (±45) with respect to the global reference, which complies with analytical results. Additionally, specimens have been subjected to a cycling tensile load with decreasing maximum normal strain, with footstep ε x =0.02%, beginning from ε x =0.18% (extreme strain limit of the first phase, in which the stiffness behaviour is linear). Thus, as shown in Fig. 14, the elastic limit of the laminate has been defined as a deformation ε x of 0.09% 0.1%; such a specimen does not exhibit any damage on the surface. Since the fatigue limit has been determined for a strain ε x of 0.4%, the increase in damage of the laminate, subjected to few load cycles, has been qualitatively studied. In Fig. 15, three graphs (stress versus strain) concerning tensile tests with 1,5 and 10 cycles are illustrated; it can be noticed that the highest deformation ε x is 0.35%, which is within the strain range of damage tolerance. The laminate absorbs most of the energy in the first cycle compared to the following ones; besides, the formation of microcracks is still not saturated after 10 cycles. In addition, the stress-strain relation becomes linear with rising cycles; the stiff contribution of the matrix quickly vanishes and after 10 cycles the Young s modulus E x is N/mm 2. Reduction of residual deformations of the laminate Since the laminate is not symmetric, residual strains are present. Several square laminated plates (with a 30cm side) have been manufactured and the 6

7 curvatures have been measured; the mean values of which are: k x =0.215m -1, k y =0.197m -1, k xy =0.278m -1. Fire resistance layers could be added in the proximity of the outer side of the load-carrying plies to obtain symmetry, but other disadvantages occur: strong increase of weight and formation of cracks for weak loads on the surface adjacent to the adhesive ply. Hence a concept has been elaborated to introduce an interlayer (thin ply with casually orientated glass fibres) which is able to reduce coupling between in-plane and out-of-plane action. First the stiffness properties of the interlayer were determined experimentally. Thus, the coefficients of the coupling matrix as a function of position and thickness of the interlayer have been analysed. A substantial reduction of coupling coefficients has been obtained for the laminate with an interlayer (whose thickness is t=0.85mm) placed between the load-carrying layers (between the outer one with (0,90) orientation and the one with (±45) orientation). Finally, the laminate was built with the new stacking sequence; a strong reduction of curvatures has been measured: k x =41.7%, k y =57.3%, k xy =93.7%. However, no interlayer was used for further investigations since the reduction of curvature is obtained by means of an excessive increase of the skin thickness and of the corresponding weight. Four-point bending test Sandwich beams have been designed and realised; the main manufacturing steps are summarised: 1) Construction of two laminates (with dimensions 850x400 mm): The Bag Molding technique has been used; hence, after preparing the tool plate, glass fabrics (6 pieces with (0,90) orientation and 4 with (±45) orientation of fibres), peel plies, bleeder ply and vacuum bag were cut and the matrixes were prepared, according to the weight contents of Tab.11. Hence, the lay up according to the stacking sequence in Fig.1 was implemented; subsequently, vacuum bag assembly processing was realised with the following temperature and pressure cycles: T=100 t=1h p=7bar t 1.5h (from the beginning, until T=40 C) 7

8 2) Using an electric saw, the edges of the laminates (which constitute impurity) were removed and the core panel cut. 3) Adhesion between faces and core The laminates were warmed at a low temperature T C, then the adhesive paste was smeared on the side of the load-carrying layer of the laminate. Subsequently, faces and core were fixed together with no adhesive tapes; in addition, the sandwich panel was vacuum-bagged, as shown in Fig.16, and the adhesive resin was polymerized by means of the following curing process: T=60 C t=2h p=1bar 4) Using an electric saw, four sandwich beams were cut having the dimensions: L=800mm length of beam t f =4mm thickness of face b=80mm width t c 40mm thickness of core h=48mm thickness t a 0.5mm thickness of adhesive layer 5) Finally, two strain rosettes were placed on the skins in the middle cross section to measure the deformation ε x. The equipment of the four-point bending test was regulated so that the distance between the outer supports was L out =700mm and the distance between the inner supports was L in =300mm. Three beams were tested; as expected, failure occurred in the core. The separation surface is oblique and it is placed between the inner and outer supports; Fig.17 shows the beam crack at the instant of failure; it is shown clearly how the faces slide over each other (shear deformation). The experimental graphs, load versus normal strain of the skin in the middle cross section, give information about the flexural rigidity, whose value is D=40 000N/mm 2 ; the failure load is P=4 570N and the failure normal strain at the edge of the middle section is ε x =0.03%. The maximum normal strain of the skins is lower than the elastic limit of the laminate; in fact, the specimens tested do not show damage to the skins, and the flexural rigidity remains constant. Acknowledgements The authors wish to express their appreciation to Mr.Wenzl for his contribution to fire resistance tests on the laminate. In addition, special thanks are 8

9 given to all the laboratory staff at the Daimler-Benz AG, Research and Technology, Department F2K-F, in Ottobrunn (Munich, Germany), where the present work was developed during a stay of A.Pascucci. As to the part of the work developed in Rome and concerning the numerical calculations and, partly, the interpretation of the experimental results, the support of CNR contract n ct07 (responsible Dr. P. Gaudenzi) is gratefully acknowledged. References 1. Batchelor, J., "Use of Fibre Reinforced Composites in Modern Railway Vehicles", Materials in Engineering, 1981,Vol.2, pp Ruhmann, D.C., "The design, fabrication and testing of the Glasshopper prototype covered hopper rail cars", Composite Structures, 1994, pp Takao, K., Yoshimura, M.,Tagawa, N., Matsudaira, Y., Nagano, K., "Development of the Superconducting Maglev Vehicles for the Yamanashi Test Line", 4th Japan International SAMPE, 1995, Japan 4. Jones, R. M., Mechanics of Composite Materials, McGraw-Hill Edition, Vinson, J.R., Sierakowski, R. L., The Behaviour of Composed Structures of Composite Materials, Kluwer Academic Publisher, Zenkert, D., An Introduction to Sandwich Construction, KTH Department of Lightweight Structures, Sweden, ADINA Theory and Modelling Guide. ADINA R&D Inc Report ARD 87-8,

10 Position Number Main characteristic Kind of fibre Name Matrix Weight fibre content 3 2 Separate layer - - VE-ATH Fire-resistant layer Glass fabric 800L VE-ATH % 1 3 Load-carrying layer Glass fabric W1480 VE-ATH 60 70% Tab.1 Material characteristics of the laminate Time interval [sec] 0 ATH 60 ATH 180 ATH First flame formation Failure (Gas formation) Tab.2 Fire test results for laminates with different ATH contents 0 ATH 60 ATH 180 ATH ε x (phase1) [%] ε x (knee) [%] ε x (phase2) [%] E x (phase1) [N/mm 2 ] E x (phase2) [N/mm 2 ] ε x(ultimate) [%] σ x(ultimate) [N/mm 2 ] Tab.3 Tensile tests results for laminates with different ATH contents 0 ATH 60 ATH 180 ATH G xy [N/mm 2 ] τ xy(ultimate) [N/mm 2 ] Tab.4 Shear tests results for laminates with different ATH contents First ply failure load Ultimate load Traction N x [N/mm] Bending M x [N] In-plain shear N xy [N/mm] Torsion M xy [N] Tab.5 Ultimate and first ply failure loads on the laminate E x [%] E y [%] G xy [%] ν xy [%] N x (0-453N/mm) M x (0-619N) N xy (0-187N/mm) M xy (0-140N) Tab.6 Decrease of engineering coefficients between initial conditions and loads causing first ply failure 10

11 Face fracture First ply failure Shear core Wrinkling failure Traction N x [N] Bending M x [Nm] Transversal shear N xz [N] In-plain shear N xy [N] Tab.7 Failure loads on the test sandwich panel [N] The others free The others simply The others clamped supported Loaded edges clamped Loaded edges simply supported Loaded edges clamped and simply supported Load edges clamped and free Tab.8 Critical compressive load on the test sandwich panel with several boundary conditions [%] N/mm 2 N/mm 2 ε x(ultimate) 2.29 σ x(ultimate) E x(din) (10-50%) ε x (phase1) σ x (phase1) 0-40 E x (phase1) ε x (knee) σ x (knee) ε x (phase2) σ x (phase2) E x (phase2) Tab.9 Tensile test on the laminate, main strains, stresses and Young s moduli [%] N/mm 2 N/mm 2 ε x(ultimate) 2.53 σ x(ultimate) E x(din) (10-50%) Tab.10 Three-point bending test on the laminate, ultimate strain, stress of the edge and secant Young s modulus Position Fibre [g] Matrix [g] TBPB Hardening [g] TBEH Hardening [g] ATH [g] Tab.11 Weight content of the layers Pos. 3 0/90 Pos. 2 ± 45 Pos. 2 Pos. 3 0/90 Pos. 1 ± 45 Pos. 1 0/90 Pos. 1 OUTSIDE INSIDE Fig.1 Stacking sequence scheme of the laminate 11

12 σ 1 [N/mm²] Nx [N/mm] lamina1 (0,90) lamina2 (+45,-45) lamina3 (0,90) lamina4 (+45,-45) lamina5 (0,90) Fig.2 Normal stress (in warp direction) of each layer versus tensile load σ 2 [N/mm²] Nx [N/mm] lamina1 (0,90) lamina2 (+45,-45) lamina3 (0,90) lamina4 (+45,-45) lamina5 (0,90) Fig.3 Normal stress (in weft direction) of each layer versus tensile load 12

13 τ 12 [N/mm²] Nx [N/mm] lamina1 (0,90) lamina2 (+45,-45) lamina3 (0,90) lamina4 (+45,-45) lamina5 (0,90) Fig.4 In-plane shear stress of each layer versus tensile load Fig.5 Tensile test on the laminate, normal stress versus normal strain 13

14 Fig.6 Three-point bending test on the laminate, normal stress of the edge versus correspondent normal strain E tensile E bending EX [N/mm 2 ] σ x [N/mm 2 ] Fig.7 Secant Young s moduli for the tensile and bending cases versus the mean stress through the cross section and the normal stress at the edge, respectively 14

15 Fig.8 Tensile test, surface of the fire resistance layer, whose calculated strain ε x =0.895% Fig.9 Tensile test, surface of the load carrying layer, whose calculated strain ε x =0.681% 15

16 Fig.10 FPB test, surface of the fire resistance layer, whose calculated strain ε x =1.418% Fig.11 FPB test, surface of the load carrying layer, whose calculated strain ε x =-1.128% 16

17 [µm] Fig.12 Tensile test, micrograph of cross section of the laminate [µm] Fig.13 TPB test, micrograph of cross section of the laminate 17

18 Fig.14 Normal stress versus strain of laminates subjected to cycling tensile load Fig.15 Normal stress versus strain diagrams of laminates subjected to cycling tensile load with 1, 5, 10 cycles, respectively 18

19 Fig.16 Vacuum bag process during the bonding phase Fig.17 Failure of the sandwich beam 19

20 Table headings Tab.1 Material characteristics of the laminate Tab.2 Fire test results for laminates with different ATH contents Tab.3 Tensile tests results for laminates with different ATH contents Tab.4 Shear tests results for laminates with different ATH contents Tab.4 Shear tests results for laminates with different ATH contents Tab.5 Ultimate and first ply failure loads on the laminate Tab.6 Decrease of engineering coefficients between initial conditions and loads causing first ply failure Tab.7 Failure loads on the test sandwich panel Tab.8 Critical compressive load on the test sandwich panel with several boundary conditions Tab.9 Tensile test on the laminate, main strains, stresses and Young s moduli Tab.10 Three-point bending test on the laminate, ultimate strain, stress of the edge and secant Young s modulus Tab.11 Weight content of the layers Fig.1 Fig.2 Fig.3 Fig.4 Fig.5 Fig.6 Fig.7 Figure captions Stacking sequence scheme of the laminate Normal stress (in warp direction) of each layer versus tensile load Normal stress (in weft direction) of each layer versus tensile load In-plane shear stress of each layer versus tensile load Tensile test on the laminate, normal stress versus normal strain Three-point bending test on the laminate, normal stress of the edge versus correspondent normal strain Secant Young s moduli for the tensile and bending cases versus the mean stress through the cross section and the normal stress at the edge, respectively Fig.8 Tensile test, surface of the fire resistance layer, whose calculated strain ε x =0.895% Fig.9 Tensile test, surface of the load carrying layer, whose calculated strain ε x =0.681% Fig.10 FPB test, surface of the fire resistance layer, whose calculated strain ε x =1.418% Fig.11 FPB test, surface of the load carrying layer, whose calculated strain ε x =-1.128% Fig.12 Tensile test, micrograph of cross section of the laminate Fig.13 TPB test, micrograph of cross section of the laminate Fig.14 Normal stress versus strain of laminates subjected to cycling tensile load Fig.15 Normal stress versus strain diagrams of laminates subjected to cycling tensile load with 1, 5, 10 cycles, respectively Fig.16 Vacuum bag process during the bonding phase Fig.17 Failure of the sandwich beam 20