IMPACT BEHAVIOUR OF THIN CARBON FIBRE REINFORCED COMPOSITES COMPONENTS FOR AUTOMOTIVE APPLICATIONS

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1 IMPACT BEHAVIOUR OF THIN CARBON FIBRE REINFORCED COMPOSITES COMPONENTS FOR AUTOMOTIVE APPLICATIONS Loleï Khoun and Sylvain Bournival National Research Council Canada, Automotive and Surface Transportation Portfolio 75 bd de Mortagne, Boucherville, QC, J4B 6Y4, Canada Abstract In an effort of reducing vehicle weight, fibre reinforced polymer composite materials are more and more introduced in automotive industry to produce structural, semi-structural and body panel components. When used as automotive body panels, out-of-plane impact behaviour of composite becomes very important, as vehicle body panels can be subjected to different types of impact events, from small and medium road debris to pedestrian impacts. Understanding composite impact behaviour and being able to predict it become then very important, especially for part design and passenger and pedestrian safety. In this paper, out-ofplane impact behaviour of thin carbon fibre reinforced composite laminates was investigated under low velocity impact loading. Evolution of material elastic energy and absorbed energy with applied impact energy was determined using a vertical drop weight impact test setup. Damage type, size and location were then examined using a laser ultrasonic inspection system. An energy threshold was determined at 20 J, corresponding to a change in type of impact-induced damage. The evolution of the major damage mechanisms with an increase in impact energy was also established. Finally, modelling of impact behaviour of thin composite laminate was carried out using 2.5D non-linear dynamic numerical models. The material card was developed and calibrated using mechanical tests performed on the laminate. Single-layer and multi-layer approaches were evaluated to predict impact energy, impact damage and delamination. Hypotheses are presented to explain the discrepancy between the model and the experiments. Once validated, this numerical model will help automotive part designers to design carbon fibre reinforced composite components subjected to impact design requirements. Background Impact behavior of composite materials has been investigated for more than 25 years, and is still under investigation. Understanding the behaviour of composite under impact is crucial in term of design and structural requirements, as the presence of impact damage can significantly reduce the mechanical properties of the composite, such as strength and stiffness. Moreover, this damage may not be visible on the laminate surface (barely visible damage), while still reducing its mechanical properties. The vertical drop-weight test setup is used extensively to impact composite specimens. It consists of an impactor guided by a mechanism, free falling from a given height [1], as shown in Figure 1. A large range of weight and speed and impactor shapes can be tested which make it a versatile test, representative of different types of real-world impact events from tool drops to small and large roadway debris. Page 1

2 a) b) Figure 1: a) Impact equipment with double column impactor guide mechanism and b) detail of a hemispherical impactor [1] It has been demonstrated by various authors that impact behavior of composite materials is affected by a large number of parameters, intrinsic to the materials and the type of test carried out [2] [4]. The main parameters affecting impact damage can be classified under different categories such as: Projectile characteristics: impactor weight, impactor shape, impactor radius [5] [12]. Increase in damage area was usually observed with an increase in impactor mass and conical impactor, while impactor radius has an influence on the type of damage itself. Laminate characteristics: layup, thickness, sample size [3], [6], [12] [17]. Stacking sequence has a significant effect on the impact resistance of laminates such as shape of damage area, value of energy threshold and development of delamination. Laminate thickness has also an influence on the damage initiation threshold force. Material characteristics: type of fibre and resin, fibre volume fraction, porosity, fibre-matrix interface [5], [18]. As expected, glass fibre reinforcement and carbon fibre reinforcement behave very differently under impact loading due to their intrinsic properties. Environmental conditions: temperature, humidity, UV [19]. Few authors evaluate the effect of conditioning on low-velocity impact resistance and found that moisture has an effect on damage initiation energy and that combining high temperature and moisture level significantly reduce impact resistance of composites. Page 2

3 Due to the large number of parameters influencing composite impact behaviour, predictive methods such as numerical simulations are then needed to reduce the cost associated with experimental testing, given a specific set of material properties. Impact behaviour and damage evolution in composite materials are usually modelled using 3D elements and interface cohesive elements for the composite laminate [9], [20] [22]. 3D finite element models give good prediction of the laminate impact response and the involved damage mechanisms, especially through-thickness damages like delamination. However, these 3D models needs large computational resources and are mainly demonstrated on small and simple academic geometries (small circular or rectangular plates) and would not be suitable to model large and complex shape laminate, such as real automotive components (hood, roof, A-B pillars, etc.). 2D and 2.5D finite element models are better suited to model large and thin components: they are more precise for such geometry type and require less computational resources and computational time. However such models cannot evaluate through-thickness stresses and therefore cannot predict damage mechanisms related to through-thickness stresses, such as delamination. Common methods to model delamination evaluate crack propagation caused by delamination [23], [24], delamination in a 2D cross section [25] and 3D methods [26]. Only few authors have used 2.5D approaches to model the behaviour of composites structures under impact with delamination and progressive damage [27] [29]. The current analysis however attempts to compare two different 2.5D approaches and evaluate their strengths and weaknesses in a design and optimization context. In this paper, out-of-plane behaviour of thin carbon fibre reinforced composite laminates was experimentally investigated under low velocity impact loading conditions. Evolution of material elastic energy and absorbed energy with applied impact energy was determined using a vertical drop weight impact test setup. Damage type, size and location were then examined using a laser ultrasonic inspection system and microscopy. Two 2.5D non-linear dynamic numerical approaches were developed, using the composite mechanical properties, experimentally measured. The models were correlated to the experimental results obtained from the vertical drop weight impact tests to validate the use of a 2.5D model to predict the laminate impact behaviour. Materials Experiments Unidirectional carbon fibre reinforced epoxy prepreg with 60wt% of carbon fibre and 40wt% of epoxy resin content was used to manufactured [0/90/0] s composite panels with a thickness of 1.2 mm. These composite panels are representative of automotive component that can be subjected to out-of-plane impact. Impact specimens were cut from those panels with the following dimensions: x mm (6 x4 ), where the long direction corresponded to the 0º direction of the panel. Impact tests Vertical drop weight tests, based on ASTM D7136, were used to evaluate the behaviour of thin laminates under impact event. Impact tests were performed with a hemispherical impactor, with a mass of 5.83kg and a diameter of 38 mm. Impact behaviour of the thin CF/epoxy laminate was evaluated for impact energy between 5 J and 25 J, as shown in Table 1. The range of impact energy was chosen so that different types of damage were induced to the composite specimens from barely visible defects to indentation and perforation. Chosen impact energies correspond to impact speed from 4.9 km/h to 10.8 km/h (1.36 m/s to 3 m/s). Page 3

4 For each impact test, the following variables, schematized in Figure 2, were measured: F max : maximum recorded contact force F i : recorded contact force at which the force versus time curve has a discontinuity, which can correspond to a first failure within the specimen t T : contact duration (total duration of the impact event) E impact : actual impact energy (incident kinetic energy) E a : energy absorbed by the specimen during the impact event E elas : elastic energy (E impact = E elas + E a ) E max = absorbed energy at the time of maximum recorded contact force F max a) b) Figure 2: Representative data obtained during an impact test: a) Contact force versus time and b) Energy versus time Page 4

5 Nb of samples Impactor geometry Impactor mass (kg) Table 1: Impact test plan Impactor diameter (mm) Impact speed (m/s) Impact energy 3 hemispherical hemispherical hemispherical hemispherical hemispherical (J) Damage Analysis Three methods were used to observe impact damage. First, samples were visually inspected to evaluate damages on the composite outer layers. Then, Laser Ultrasonic Inspection System (LUIS) was used to scan the specimens before and after the impact event. This non-destructive method is similar to classic ultrasonic inspection method, however, in laser-ultrasonics, the generation and detection of ultrasound is done by two lasers. When the two lasers, one for generation and the other one for detection, impinge at the same location on the surface of the specimen, a train of ultrasonic echoes is generated and then detected. Any change in the structure of the specimen, such as porosity, delamination or contaminant, can be detected by this method. Finally, in order to have more information of the damage location through the thickness of the specimen, one specimen per energy level was cut at different sections and observed under optical microscope. Sections were cut every centimeter or centimeter and a half along the damage length. Three to seven cuts were done, depending of the damage dimension. Impact tests Impact behaviour of thin carbon fibre / epoxy laminate Figure 3 and Figure 4 present the measured contact force and energy with time, respectively, at different impact energies. Results are very reproducible, especially below 20 J. From those graphs, it can be noticed that the contact duration remain the same at about 12 ms. The total absorbed energy, the elastic energy, the energy at maximum load and the maximum load were extracted from those graphs (as shown in Figure 2) and plotted in function of the impact energy in Figure 5, Figure 6, Figure 7 and Figure 8, respectively. Results from preliminary tests performed at higher impact energies (between 25 J and 60 J) are also plotted on the graphs. As expected, the total absorbed energy increases with the applied energy (Figure 5). Two regimes can be observed: a linear increase up to J where the total absorbed energy is almost equal to the impact energy, followed by a slower increase of the total absorbed energy with the impact energy after 25 J and larger value scatter. Those two regimes, with a change of regime around 20 J, are also observed for the elastic energy, the maximum load and the energy to maximum load. For the elastic energy (see Figure 6), it remains constant at a value around 5 J and then increase linearly for impact energy of 20J and above. The maximum load (see Figure 8) and the energy to maximum load (see Figure 7) follow the same trend with a linear increase Page 5

6 of their value with impact energy up to 20 J and a plateau of 4500 N for the maximum load and 15 J for the energy to maximum load for impact energy above 20 J. a) Impact 5J 1.36 m/s b) Impact 8J 1.7 m/s c) Impact 14J 2.2 m/s d)impact 20 J 2.65 m/s d) Impact 25J 3 m/s Figure 3: Load versus time for the different impact energy: a) 5 J, b) 8 J, c) 14 J, d) 20 J and e) 25 J Page 6

7 a) Impact 5J 1.6 m/s b) Impact 8J 1.7 m/s c) Impact 14 J 2.2 m/s e) Impact 20J 2.65 m/s f) Impact 25J 3 m/s Figure 4: Energy versus time for the different impact energy: a) 5 J, b) 8 J, c) 14 J, d) 20 J and e) 25 J Page 7

8 Figure 5: Total absorbed energy versus impact energy Figure 6: Elastic energy versus impact energy Figure 7: Energy to maximum load versus impact energy Page 8

9 Figure 8: Maximum load versus impact energy Damage analysis Table 2 shows pictures of the front and back faces of impacted laminates at the five studied impact energies, with the corresponding LUIS scans (in-plane and through-thickness scans) and the dimension of the damage area, isolated using open source ImageJ image analysis software. The in-plane LUIS scans shows the damage size for the laminate. Change in colours can be related to damages that occur at different plies or ply interfaces. The through-thickness LUIS scans shows were the first change in the material occurs within the thickness of the laminate. The red arrows representing the direction of the generated laser ultrasound, the damage or delamination in our case can be detected at the position of the first fringe of colour. The other fringes of colour correspond to the echoing of the ultrasound and are not representative of damage. As expected, damage size increases with the applied impact energy, and type of observed damage also changes with the applied impact energy. It can be noticed that LUIS method enables to visualize the extent of damage due to delamination, which was not possible with visual observations. At 5 J, damages are not visible on the outer layer of the specimen; no cracks or indents were observed on the impacted face of the laminate. However, as the laminate is very thin, delamination can be guessed under the laminate surface. This is confirmed by LUIS technique, which shows 25 mm x 20 mm damage located below the impacted surface of the laminate. At 8 J, damages are not visible or barely visible at the laminate surface. Small vertical cracks along the 0º fibre direction were observed and delamination can also be guessed under the laminate surface. LUIS scan shows larger damage with typical peanut shape meaning that the delamination occurs at the interface between the 0º and 90º plies. Again, defects seem to be located closed to the surface layers (top and bottom). At 14 J, damages are clearly noticeable on the surfaces of the laminate. Vertical matrix cracks, along the 0º fibre direction, were observed on the front and back face of the laminate, as well as delamination area in a shape of a peanut. Small fibre breakage along the 90º direction was also noticeable. LUIS scans shows larger peanut shape delamination areas, oriented along the 0º direction. Page 9

10 At 20 J, indentation of the laminate was observed, creating large matrix and fibre cracks at the front and back faces of the laminate. In-plane LUIS scan shows that delamination extends to almost the entire length of the sample. Finally, at 25 J, perforation occurred for most of the impacted laminates. Large fibre breakage, fibre peeling and matrix cracks were observed. Figure 9 plots the evolution of the damage area with the impact energy. A linear increase is observed. The size of the damage area is multiplied by 10 between 5 J and 25 J impact energies. The observation of the through-thickness section with optical microscope (see Figure 10) gave a better understanding of the damage formation with impact energy: At 5 J impact energy, transverse matrix cracks (vertical bending cracks and shear cracks oriented at 45º) arose under the impact load point, at laminate mid-plane, within the two plies oriented at 0º, and in the bottom ply oriented at 0º. Delamination also arose at the 0º/90º ply interface of the last two plies. At 8 J impact energy, more transverse matrix cracks were observed within the 0º plies, at the laminate mid-plane, and the top and bottom plies. Delamination appeared at most of the 0º/90º ply interface with larger delamination at the 0º/90º ply interface of the last two plies Similar types of damage were observed at 14 J impact energy, but with a larger extent. More transverse matrix cracks are created and delaminations were visible at all 0º/90º ply interfaces. Also, fibre breakages arose as well in 90º ply at the impact load point. Finally, at 20 J, created damages were such that most of the samples were split in two. Multiple transverse matrix cracks were located within all 0º plies and delamination areas were very large. Fibre breakage also arose within all 90º plies. This evolution of impact damage with impact energy at the impact load point is schematized in Figure 10-b. Therefore, the change in composite response to an impact event between 20 J and 25 J, as observed with the change in regime (slope) for the total absorbed energy, the elastic energy, the maximum load and the energy to maximum load, corresponds to a change in induced damage types: below 20 J, the main type of damage consists in delamination and transverse matrix cracks in the plies oriented at 0º. Those transverse matrix cracks are due to the bending stresses created during the impact test and the bending of the specimen, and shear stresses. Increase in impact energy up to 20 J leads to an increase of those damages. above 20 J, indentation and then perforation (at 25 J) occurred, which are associated with large fibre breakage. 20 J impact energy then corresponds to the penetration threshold of this CF/epoxy laminate, in case of out-of-plane impact event. Page 10

11 Table 2: Damage visualization at different impact energy by visual inspection and LUIS scans Visual inspection LUIS scans ImageJ In-plane Through-Thickness Front Back a) 5J impact energy Front Back b) 8J impact energy Front Back c) 14J impact energy Page 11

12 Table 2: Damage visualization at different impact energy by visual inspection and LUIS scans (con t) Visual inspection LUIS scans ImageJ In-plane Through-Thickness Front Back d) 20J impact energy Front Back e) 25J impact energy Page 12

13 Figure 9: Damage area versus impact energy 5J 8J 14J 20J a) Through-thickness micrographs of impact damage at impact loading point with increase in impact energy, obtained from optical microscope (x20) b) Schematic of impact damage at impact loading point Figure 10: Impact damage evolution with increase in impact energy at impact loading point: a) micrographs obtained with optical microscope, b) schematic representation of the impact damage Page 13

14 MAT54 material model Impact modelling Impact modelling was carried out with LS-Dyna software. MAT54 material model was used to model the composite laminate. This material model supports laminated shell theory and is therefore capable of modelling several plies of different material or fibre orientation. This material model takes into account four modes of failure: tensile failure in the longitudinal direction (fibre), compressive failure in the longitudinal direction (fibre), tensile failure in the transverse direction (matrix) and compressive failure in the transverse direction (matrix), and uses the Chang-Chang failure criterion [30] [32], which corresponds to a strength-based ply failure. MAT54 material model requires more than 30 parameters, some related to measurable intrinsic properties of the material, and other computational parameters that must be estimated (see Table 3). In Table 3, we easily recognize the density (ρ), Young s modulus (E i ) and shear modulus (G ij ) and Poisson s ratio (ν ij ). The limits of the material are given by the tensile (subscript t) and compressive (subscript c) maximum strengths in the longitudinal direction (Xt and Xc), transverse direction (Yt and Yc) and shear (Sc) and the associated maximum strains (the tensile and compressive maximum strains in the longitudinal direction, DFAILT and DFAILC, the transverse maximum strain DFAILM and the maximum shear strain DFAILS, see Figure 11). The material constitutive parameters shown in green in Table 3 were measured experimentally by performing five standard mechanical tests: tensile tests in the ply longitudinal and transverse directions (ASTM D3039), compressive tests in ply longitudinal and transverse directions (ASTM D3410) and in-plane shear tests (ASTM D3518). The remaining MAT54 parameters were then estimated by implementing the measured material properties in MAT54 material card and modelling all the experimental tests. Figure 12 compares experimental and predicted strain-stress curves obtained after some iteration. A good correlation can be observed between the prediction and the experiment using the measured and estimated MAT54 parameters. This material data card was then used to model the composite impact tests. Page 14

15 Table 3: MAT54 material data card Materials parameters Density ρ Modulus E 1 E 2 E 3 G 12 G 23 G 13 Poisson s ratio ν 12 Max. strength Xc Xt Yc Yt Sc Max. strain DFAILM DFAILS DFAILT DFAILC EPS Damage factors SOFT FBRT YCFAC PEL SOFT2 SOFTG Shear weighting ALPH BETA Element deletion TFAIL Shear damage EPSF EPSR TSMD Minimum stress limit SLIMT1 SLIMC1 SLIMT2 SLIMS Figure 11: Stress strain behavior of MAT54 material model Page 15

16 a) b) c) d) e) Figure 12: Model correlation with standard mechanical tests: a) Tensile test in fibre direction, b) Tensile test in transverse (matrix) direction, c) Compression test in fibre direction, d) Compression test in transverse direction and e) In-plane shear test Page 16

17 Impact modelling The impact test model consists in a composite rectangular plaque modelled of MAT54 2.5D shell elements and an impactor, modelled as a rigid body with motion constrained in the perpendicular direction. The plaque was modelled with a structured mesh of variable density (see Figure 13), the elements at the point of impact having a size of 1 mm. The nodes on the 4 edges of the plaque were fixed in the three directions. A contact of type surface to surface was implemented between the laminate and the impactor in a manner similar to the one described in [27]. Two approaches were considered to model the composite laminate: The first approach consisted in modelling the composite laminate and its stacking sequence with a single layer of shell elements. LS-Dyna can, with the help of the *PART_COMPOSITE keyword and the activation of the laminated shell theory, model a part with several plies. In the element, an integration point is used to model each ply; each can be defined with a different material and fibre orientation. This approach enables to reduce significantly the number of elements and the associated computational time. For this specific model, 6561 elements were used (4335 shell elements and 2226 elements for the impactor) for a computational time between 18 and 27 minutes. However, one major drawback of the approach is that delamination cannot be predicted. The second approach consisted in modelling each ply of the composite laminate with a layer of shell elements. A contact of type *CONTACT_AUTOMATIC_ONE_WAY_SURFACE_TO_ SURFACE_TIEBREAK is applied between each layer of elements. This type of contact is based on the cohesive zone model: the elements will not penetrate each other and they will stick together until a certain de-bonding criteria is met, at which point the force keeping the elements together will be removed (with some energy release) and the contact is converted to a standard surface to surface contact. While having more elements, this approach enables the modelling of delamination. For this specific model, 6 layers were modelled with a total of shell elements for an average computational time of 68 minutes. 0 fibre orientation Figure 13: Numerical model of the plaque and impactor Page 17

18 The predicted load and predicted impact energy (two types of black dashed line for the models) were compared with the experimental results on Figure 3 and Figure 4 respectively: For the single layer approach, the evolution of the load was fairly well predicted from the onset to first failure (see Figure 3). After the first failure, the load remained low and almost constant as if the specimen failed with no further material resistance. The predicted absorbed energy was then too high in most of the evaluated cases with only a little elastic energy returned to the impactor (see Figure 4). This predicted behaviour led to excessive failure with larger damage area in the transverse direction than observed experimentally. This approach predicted damages corresponding to matrix failure and fibre failure, even at low impact energy. Results obtained with the multi-layered approach remained closer to the experimental data. The evolution of the load with time was predicted after the first failure (see Figure 3) and good correlation was observed for the absorbed energy up to impact energy of 14 J (see Figure 4). More discrepancy was observed for impact energy of 20J and up with an excessive level of absorbed energy and almost no elastic energy. By analyzing the model time frame for the single layer approach, it was observed that the model first predicted failure in tension at the center of the bottom ply in the transverse direction. This failure then reached the neutral axis, at which point, compressive failure happened on the top ply in the longitudinal direction. Shortly after, all the failure modes in the central element were reached, which caused the deletion of this element. This caused a redistribution of the stresses to the other elements in the laminate, and a cascading effect of failure and deletion of elements until failure of the specimen. This is represented by the sudden jump in the curve seen in Figure 3c. The same sequence was observed for the five impact energy tested. This predicted failure sequence is in agreement with the impact damage evolution observed by microscopy; however it starts to happen at lower impact energy. The time frame for the multi-layered approach begins in the same fashion, only this time delamination caused a redistribution of stresses which delayed the impact of the cascading effect. The delamination is controlled by the tiebreak contact condition applied between the layers: a linear spring for tensile and shear forces is applied between both surfaces. When the strain in that spring reaches a certain value delamination happens, the contact spring is removed and the contact condition is modified to compression-only contact. However, if the strain remains below the threshold, tiebreak contact conditions are maintained and elastic energy is restored to the laminate. In the 5J multi-layer model for example, the strain threshold was not reach, preventing delamination and keeping the element from reaching its failure limits and be deleted (as it was the case for the single layer model). The predicted delamination area of the multi-layer model is shown in Figure 14. The predicted delamination zone is larger than the one observed with the non-destructive technique (Table 2c), however, this could be caused by the model boundary conditions that may be too restrictive and not exactly representative of the experimental conditions. In the model the edges of the plaque are fixed in the three directions, while slight in-plane displacement can occur at the edge of the specimen during the experiment. This artificial restriction created larger stresses towards the edges of the laminate, leading to an extent of delamination area in those zones (circled in Figure 14). The model correctly predicted little to no delamination between layers 3 and 4, both with fibres oriented in the same direction. Page 18

19 a) Layers 1 and 2 interface b) Layers 2 and 3 interface c) Layers 3 and 4 internface d) Layers 4 and 5 interface e) Layers 5 and 6 interface Figure 14: Delamination area predicted with 14J impact mode Discussion The single layer model predicted a sequence of laminate failure corresponding to the overall evolution of impact damage with impact energy, observed by optical microscopy and represented in Figure 10. However, while major laminate failure (fibre failure, indentation, penetration) occurred for impact energy above 14J in the experiment, the numerical model predicts matrix and fibre failure at impact energy as low as 8J. It is assumed that, in the single layer model, as failure by delamination couldn t happen, failure modes of the central elements were reached too early, causing premature deletion of those elements. The redistribution of the stresses to the other elements in the laminate then created a cascading effect of failure and deletion of elements until failure of the specimen. The multi-layered model predicted well the variation of load and energy with time for the tested impact energies. The use of tiebreak contact between the plies of the laminate allowed the prediction of delamination and eliminated the premature element deletion at the center of the laminate. The predicted of modes of failure with applied impact energy were then closer to the experiments. Page 19

20 To explain the differences, it must be noted that, in the single layer model, the throughthickness strains remained the same, while, in the multi layered model, the nodes of one layer were free to slide on neighboring layers (especially after delamination). This effect is illustrated in Figure 15, where the displacement in the Y-direction for nodes initially at the same X and Y coordinates are plotted. It can be noted that, after delamination at 3 ms, the second and fifth layers (B and E) oriented at 90º, undergo more displacement in the Y-direction than the other layers, and therefore different strains. The different strains allow for a redistribution of stresses in the elements of each layer. On the contrary, in the single layer model, as the plies of the laminate are modeled with a single element, the strains must remain the same for all the plies. This caused some plies to reach the maximum strain prematurely, starting a chain reaction leading to the entire failure of the element and the elimination of energy. As for the accuracy of both models, a few hypotheses can be proposed to explain the discrepancies with the experimental results. 1) Material parameters not directly measured: Table 3 lists many material parameters, the most important of which were determined with a set of in-plane tests. However there are parameters that can only be estimated as there are no experiments that can isolate them. Some of those parameters have a significant effect on the failure prediction. 2) Strain rate influence: All in-plane tests were conducted at a relatively slow strain rate compared to the impact tests. It is possible some properties may be strain rate dependent and vary when tested at higher rate. 3) Out-of-plane parameters: Due to the in-plane geometry of the sample (thin laminate), no out-of-plane properties like G 23 and G 13 were measured for the composite laminate. However those out-of-plane parameters could have some effect on the impact analyses. 4) Boundary conditions: Model boundary conditions might be too restrictive compared to the experiment. For example, in the experiment, there could be some slippage between the plaque and the clamping frame, or a certain rotation at the extremities that are not happening in the model. Page 20

21 Figure 15: Y-displacement with time of nodes initially superposed in the multi-layer model (14J model) Conclusion In an effort to better understand out-of-plane impact behaviour of thin carbon fibre reinforced composite laminates and predict this behaviour, vertical drop weight impact tests were performed on thin laminates representative of automotive body panel components. The evolution of material absorbed energy and damage size with applied impact energy was determined experimentally. An energy threshold was determined at 20 J, corresponding to a change in type of impact-induced damage from delamination and transverse matrix cracks to fibre breakage with specimen indentation and perforation. Evolution of the major damage mechanisms with an increase in impact energy was also established using laser ultrasonic inspection system and optical microscopy observations. In an industrial design and optimization context, it is sometimes appealing to make simplifications to a model in order to obtain preliminary results at a lesser cost. This paper presented two different shell based approaches, and identified their limits. The first approach, with a faster solving time, used a single layer of shell elements with as many through thickness integration points as there are plies, while the second approach used a layer of shell elements for each ply with contact conditions between each ply. A specific material card was developed for the tested material and calibrated using mechanical tests performed on the laminate. Both models behaved well up to the first failure. The single layer approach predicted an earlier failure than what was observed during the experiments due to its inability to take into account ply delamination. The multi-layer approach showed better correlation with the experiment at 5 J, 8 J and 14J impact energy, but discrepancy remained for higher impact energy. Hypotheses were presented to explain the discrepancy between the models and the experiments. Verification and validation of theses hypotheses are part of the ongoing work of this two-year project. Once validated, those numerical models will help automotive part designers to design carbon fibre reinforced composite components subjected to impact design requirements. Page 21

22 Acknowledgements The authors thank Manon Plourde and Christian Néron from NRC-Boucherville for their precious help for the sample impact testing and damage analysis. The authors also thank Zongxun Wang, Eric Langley, Steve Roddy and John Ingram from Magna Exteriors for providing technical advices, and the people at the Magna-NRC Composite Centre of Excellence for supplying the materials and manufacturing the composite laminates. References [1] ASTM, ASTM D Standard Test Method for Measuring the Damage Resistance of a Fiber- Reinforced Polymer Matrix Composite to a Drop-Weight Impact Event 1, ASTM Int., [2] M. O. W. Richardson and M. J. Wisheart, Review of low-velocity impact properties of composite materials, Compos. Part A Appl. Sci. Manuf., vol. 27, pp , [3] S. Agrawal, K. K. Singh, and P. K. Sarkar, Impact damage on fibre-reinforced polymer matrix composite - A review, J. Compos. Mater., vol. 48, no. January 2013, pp , [4] S. N. A. Safri, M. T. H. Sultan, N. Yidris, and F. Mustapha, Low Velocity and High Velocity Impact Test on Composite Materials A review, Int. J. Eng. Sci., pp , [5] R. Hosseinzadeh, M. M. Shokrieh, and L. Lessard, Damage behavior of fiber reinforced composite plates subjected to drop weight impacts, Compos. Sci. Technol., vol. 66, pp , [6] Z. Aslan, R. Karakuzu, and B. Okutan, The response of laminated composite plates under lowvelocity impact loading, Compos. Struct., vol. 59, pp , [7] R. Karakuzu, E. Erbil, and M. Aktas, Impact characterization of glass/epoxy composite plates: An experimental and numerical study, Compos. Part B Eng., vol. 41, no. 5, pp , [8] T. Mitrevski, I. H. Marshall, R. Thomson, R. Jones, and B. Whittingham, The effect of impactor shape on the impact response of composite laminates, Compos. Struct., vol. 67, pp , [9] A. Qiu, K. Fu, W. Lin, C. Zhao, and Y. Tang, Modelling low-speed drop-weight impact on composite laminates, Mater. Des., vol. 60, pp , [10] Y. P. Siow and V. P. W. Shim, An Experimental Study of Low Velocity Impact Damage in Woven Fibre Composites, J. Compos. Mater., vol. 32, no. 12, pp , [11] S. Wakayama, S. Kobayashi, T. Imai, and T. Matsumoto, Evaluation of burst strength of FW-FRP composite pipes after impact using pitch-based low-modulus carbon fiber, Compos. Part A Appl. Sci. Manuf., vol. 37, no. 2006, pp , [12] H. Y. Choi, H.-Y. T. Wu, and F.-K. Chang, A New Approach toward Understanding Damage Mechanisms and Mechanics of Laminated Composites due to Low-Velocity Impact: Part II - Analysis, J. Compos. Mater., vol. 25, [13] E. V. González, P. Maimí, P. P. Camanho, C. S. Lopes, and N. Blanco, Effects of ply clustering in laminated composite plates under low-velocity impact loading, Compos. Sci. Technol., vol. 71, pp , Page 22

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