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Mechanical Behavior and Failure Mechanisms of Carbon Fiber Composite Pyramidal Core Sandwich Panel after Thermal Exposure
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J. Mater. Sci. Technol., 2013, 29(9), 846e854
Mechanical Behavior and Failure Mechanisms of Carbon Fiber Composite Pyramidal Core Sandwich Panel after Thermal Exposure Jiayi Liu, Zhengong Zhou*, Linzhi Wu, Li Ma Center for Composite Materials and Structures, Harbin Institute of Technology, Harbin 150001, China [Manuscript received May 12, 2012, in revised form November 1, 2012, Available online 18 April 2013]
An attempt has been made here to evaluate the effect of thermal exposure on the mechanical behavior and failure mechanisms of carbon fiber composite sandwich panel with pyramidal truss core under axial compression. Analytical formulae for the collapse strength of composite sandwich panel after thermal exposure were derived. Axial compression tests of composite laminates and sandwich panels after thermal exposure were conducted at room temperature to assess the degradation caused by the thermal exposure. Experimental results showed that the failure of sandwich panel are not only temperature dependent, but are time dependent as well. The decrease in residual compressive strength is mainly attributed to the degradation of the matrix and the degradation of fiberematrix interface, as well as the formation of cracks and pores when specimens are exposed to high temperature. The measured failure loads obtained in the experiments showed reasonable agreement with the analytical predictions. KEY WORDS: Composite; Thermal exposure; Mechanical properties; Sandwich panel
1. Introduction The interest for carbon fiber composite sandwich panel (CFCSP) has grown rapidly over the last decade due to their lightweight attributes and potential multifunctional applications[1e5]. The superior specific strength and stiffness of carbon fiber reinforced polymer composite (FRPC) sandwich panels relative to metallic sandwich panels have been revealed at room temperature[6]. However, FRPC sandwich panel is quite sensitive to temperature variation. The FRPC is combustible, and susceptible to deterioration of mechanical properties when they are exposed to high temperature[7e9]. The matrix dominated properties are expected to be more affected for FRPC when they are exposed to high temperature. Currently, the mechanical behavior and failure mechanisms of CFCSP at room temperature have been studied in detail[10e15], while less is known about the structural integrity of CFCSP after exposure to high temperature. In aircraft structures, some composite parts adjacent to the hot source are exposed to a high temperature environment in presence of oxygen in service, leading to the creation of an oxidized layer on the material surfaces and to cracking of the composite materials[16e18]. The * Corresponding author. Prof., Ph.D.; Tel.: þ86 451 86402396; Fax: þ86 451 86402386; E-mail address: [email protected] (Z. Zhou). 1005-0302/$ e see front matter Copyright Ó 2013, The editorial office of Journal of Materials Science & Technology. Published by Elsevier Limited. All rights reserved. http://dx.doi.org/10.1016/j.jmst.2013.04.013
degradation of matrix and the fiberematrix interface have been reported when carbon fiber composite materials were exposed to high temperature[19]. The thermal distortion of panels made by FRPC under transverse thermal-loading conditions has been discussed[20]. The severe thermal damage by cracks and delamination, and high mass losses have been revealed for 8552/ IM7 when the material was exposed to 340e450 C for 30 min[21]. Foster and Bisby[8] studied the residual tensile strength of carbon FRPC after exposure to elevated temperatures of 100, 200, 300 and 400 C for 3 h. They found that the tensile strength for the carbon FRPC showed no obvious reduction with temperature up to 300 C, but there were severe reductions at 400 C. Cleary et al.[22] tested glass fiber reinforced polymer (GFRP)-wrapped concrete cylinders in axial compression after exposure to elevated temperatures. The results showed the GFRP-confined concrete cylinders lost about 2%, 4%, 13% and 18% of their initial room temperature ultimate strength after exposure for 90 min at temperatures of 120, 135, 150 and 180 C, respectively. The glass transition temperature of the resin system used in this study was quoted as 121 C. With increasing use of CFCSP in aircraft structures where high temperature can occur, understanding the effect of high temperature exposure on mechanical behavior of CFCSP is an important safety issue. Thus, the present work is an attempt to fill this information needed. The effect of high temperature exposure on mechanical properties and damage mechanisms of CFCSP with pyramidal truss cores under axial compression was investigated in this study. The outline of the present paper is as follows. Firstly,
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analytical expressions are derived for the collapse load of composite sandwich panels after thermal exposure. Secondly, the modulus and strength of composite laminates after thermal exposure were studied in order to determine the mechanical properties of composite sandwich panel. Finally, the in-plane compressive responses of pyramidal truss core sandwich panels were investigated and compared with the analytical predictions.
(bending) or core shear buckling, (ii) face sheet wrinkling, and (iii) face sheet crushing. The difference between the core shear buckling and Euler buckling is subtle, since both modes are always simultaneously active and the final deformed shapes are somewhat similar. The differences between the modes can be found through observing the deformation of the core[23].
2. Analytical Predictions for Competing Failure Modes
Euler buckling and core shear buckling are two possible modes of macro elastic buckling in a sandwich panel under axial compression. The Euler buckling load PB(T, t) can be estimated from classical buckling theory[24] as
The schematic of the procedure used to investigate the effect of thermal exposure on mechanical behavior of composite sandwich panel is shown in Fig. 1. For truss core sandwich panel, the width of face sheet may be wider than the width of core as shown in Fig. 1(b). The width of face sheet and core is denoted by symbol B and B0 , respectively. Fig. 1(c) shows the schematic of a pyramidal truss core sandwich panel with length L, face sheet thickness hf, and core height Hc, subjected to in-plane quasi-static compression. A Cartesian coordinate is also established in order to facilitate the analysis as shown in Fig. 1(c): 1-axis is along the length direction, 2-axis is along the width direction and 3-axis is along the thickness direction. Fig. 2 shows the schematic of the pyramidal unit cell with the diameter of truss d, length of truss l, inclined angle u, and the distance k between two closest struts. The relative density r of the pyramidal core is given by
r¼
pd 2 pffiffiffi 2 sin u 2cos u þ 2k
(1)
Three different failure modes exits for sandwich panels under axial compression suggested by Xiong et al.[12]: (i) Euler
2.1. Macro elastic buckling
PB ðT ; tÞ ¼
k12 p2 DðT; tÞ L2
(2)
where k1 ¼ 2 for a column with built-in ends, and DðT ; tÞ is the equivalent flexural rigidity of the composite column. The value of PB ðT ; tÞ and DðT; tÞ is dependent on the thermal exposure temperature T and time t. The flexural rigidity can be expressed as follows: DðT ; tÞ ¼ D0 ðT ; tÞ þ 2Df ðT ; tÞ
(3)
where 2 Ef ðT; tÞBhf hf þ Hc D0 ðT ; tÞ ¼ 2
Df ðT ; tÞ ¼
Ef ðT ; tÞBh3f 12
(4)
(5)
Fig. 1 Schematic of the procedure used to investigate the effect of thermal exposure on mechanical behavior of composite sandwich panel: (a) composite sandwich panel was exposed to high temperature; (b) composite sandwich panel was cooled at room temperature; (c) composite sandwich panel was tested at room temperature under axial compression.
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Table 2 Experimental results for compressive modulus of unidirectional carbon fiber struts after exposure at 200 and 300 C for different time Exposure temperature ( C) 200 300
Fig. 2 Schematic of the unit cell of the pyramidal core.
where Ef ðT ; tÞ is the equivalent elastic modulus along 1-axis of the face sheet. The value of Ef ðT ; tÞ is dependent on the thermal exposure temperature T and time t. The equivalent elastic modulus of the face sheet can be obtained by experiment. The core shear buckling load Ps ðT ; tÞ is set by the shear stiffness of the core, which can be formulated as Ps ðT; tÞ ¼ Gc ðT ; tÞB0 Hc
(6)
where Gc ðT ; tÞ is the effective shear modulus of the pyramidal truss core, and it can be given by Gc ðT; tÞ ¼
1 r Ec ðT ; tÞsin2 2u 8
(7)
where Ec ðT ; tÞ is the effective elastic modulus of the core material in compression[25]. In present paper, the measured values for the effective elastic modulus of the core material after thermal exposure are shown in Tables 1 and 2. According to Allen’s expressions[24], the critical buckling load Pcr ðT; tÞ is given by 1 1 1 ¼ þ Pcr ðT ; tÞ PB ðT; tÞ Ps ðT; tÞ
3h
6h
9h
12 h
15 h
19.66 4.62
19.18 4.10
18.91 3.76
18.62 3.35
18.28 2.62
face sheet is connected with the aspect ratio d/l of the core member: the higher rotation restraining effect will be produced by the larger aspect ratio. It is difficult to compute k2 theoretically for the complex boundary conditions of the intercellular face sheets. While the value of k2 can be obtained numerically, which is about 1.3 for the sandwich panels in this paper. 2.3. Face sheet crushing The face sheet of sandwich panel may crush under axial compression, since the pyramidal truss core is much more compliant as compared to the face sheet. The matrix crushing, fiberematrix shear failure and delamination will be the possible micro failure modes for the composite laminates in compression. Here, we ignore the specific micro failure modes and take them as face sheet crushing macroscopically. The face sheet crushing load can be expressed as Pfc ðT ; tÞ ¼ 2Bhf sfc ðT ; tÞ
(10)
where sfc ðT ; tÞ is the crushing strength of face sheet. The crushing strength of face sheet can be obtained by experiment. 3. Materials and Experimental Methods
(8) 3.1. Material
2.2. Face sheet wrinkling The face sheet will wrinkle under axial compressive load when they are relatively thin. Wrinkling is a local buckling behavior of the face sheet and occurs between the adjacent nodes of attachment. The force Pfw ðT ; tÞ associated with the wrinkling of the face sheet can be estimated from the following expression[23] 2k 2 p2 Df ðT ; tÞ Pfw ðT; tÞ ¼ 2 pffiffiffi 2 2 cosðu þ tÞ
Exposure time
(9)
Fig. 3 shows an example of the manufactured CFCSP with pyramidal truss core, which was fabricated from unidirectional carbon/epoxy prepreg by the molding hot-press method. The sandwich face sheets were fabricated from 14-plies unidirectional carbon/epoxy prepreg stacked in the sequence of [0/90/0/ 90/0/90/0]s, and the truss cores consisted of unidirectional carbon fiber struts. Top and bottom face sheets were interconnected with truss cores. The sandwich face sheets and truss cores were manufactured in one manufacturing process without bonding. The similar in fabricate procedure was introduced in detail by Wang et al[10]. The pyramidal core presented in this paper has dimensions as d ¼ 2.5 mm, k ¼ 7 mm, l ¼ 21.2 mm, Hc ¼ 15 mm, u ¼ 45 and relative density r ¼ 2:24%. For the sandwich face sheets, the thickness of face sheet hf ¼ 1.68 mm, the length of face sheet L ¼ 225 mm, and the width of face sheet
The factor k2 depends on the end-constraints set by the pyramidal truss cores. The rotation restraining effect of truss core to Table 1 Experimental results for compressive modulus of unidirectional carbon fiber struts after exposure at different temperatures for 6h 20 C
100 C
150 C
200 C
250 C
300 C
21.56
21.18
20.10
19.18
16.01
4.10
Fig. 3 The sandwich panel with carbon fiber composite pyramidal truss cores.
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B ¼ 100 mm. The width of lattice truss core B0 ¼ 70.4 mm, since there are two unit cells along width direction. The mechanical properties of composite laminates were also studied in this paper, since the modulus and strength of composite laminate will be used to determine the mechanical properties of composite sandwich panel. In the present fabricating method, the ends of truss strut were embedded in the face sheets, which would cause a bad influence on the continuity of the carbon fiber in composite face sheets. Thus, the mechanical properties of the face sheet may decrease somewhat. In order to determine the mechanical properties exactly, the composite laminates tested in this paper were cut from sandwich panels and obtained by removing the pyramidal truss cores. 3.2. Test methods The composite laminates and sandwich panels were exposed to different temperatures for different time when they had been fabricated. Thermal exposures were accomplished by placing the specimens in a programmable insulated temperature-controlled air oven and heating to required temperature. In order to investigate the effect of thermal exposure temperature on mechanical properties, some of the composite laminates and sandwich panels were exposed to temperatures of 20 C (room temperature), 100, 150, 200, 250, and 300 C for 6 h, whereas other composite laminates and sandwich panels were exposed to temperatures of 200 and 300 C for 3, 6, 9, 12, or 15 h, in order to study the effect of thermal exposure time on mechanical properties. This temperature regime was selected to simulate temperatures that might be experienced by the fiber reinforced polymer system in an actual fire situation[8]. After thermal exposure, the specimens were cooled down to room temperature and kept at room temperature for a period of more than 24 h. Then specimens were tested at room temperature under axial compression till failure. For the composite laminates, compression test was performed on INSTRON 5569 according to ASTM 3410-75[26]. Prior to testing, the strain gages were bonded in axial direction along the centerline of the laminates in order to measure the axial strains. For truss core sandwich panels, the response of truss core sandwich panels was measured in accordance with ASTM C365/ C 364M-05[27]. The axial compression tests were conducted in the quasi-static regime with a displacement rate of 0.5 mm/min. Five repeated tests were conducted in each case to gauge the variability in the test measurements. 4. Results and Discussion 4.1. Thermal analysis of epoxy resin The glass transition temperature (Tg) of the resin used in materials is about 140 C, which is provided by the manufacturer (Beijing Institute of Aeronautical Materials, China). Thermogravimetric analysis (TGA) has been carried out to determine the temperature, at which the epoxy resin matrix began to burn off, signifying severe and irreversible chemical decomposition and loss of the matrices’ mechanical properties. The TGA test was performed in air at a heating rate of 20 C/min by a TA Instruments TGA Q50 thermogravimetric analyzer. Fig. 4 shows the retained massetemperature curve for the epoxy resin matrix measured using TGA. It can be found that the epoxy resin was decomposed in two stages, a first rapid degradation starting from 250 C and ended by a pseudo-plateau. The second
Fig. 4 TGA curves for epoxy resin.
stage leads to complete degradation of epoxy resin. Thermal analysis of polymer matrix by Regnier and Mortaigne[28] indicated that decomposition reaction mainly involves random scission of the polymer chain. Most of the polymer is fragmented into low molecular weight hydrocarbon volatiles and nonflammable gases, and only a small percentage of the original mass is retained as solid char. 4.2. Compressive responses of composite laminates The compressive modulus and strength of composite laminates after thermal exposure are studied here, which will be used to determine mechanical properties of composite sandwich panels. The change of residual compressive modulus and strength vs the exposure temperature is illustrated in Fig. 5. The measured modulus and strength data of composite laminates were based on the average of five tested specimens. It can be seen that the residual compressive modulus and strength decreased with exposure temperature, the severe reductions in modulus and strength were observed for exposure temperatures more than 250 C. The modulus and strength decreased by 82.55% and 96.84%, respectively, when the specimens were exposed to 300 C for 6 h. The thermal degradation due to the thermo-oxidative degradation of the epoxy resin matrix caused the loss of mechanical properties[29]. The effect of thermal exposure time on mechanical properties of composite laminates was also revealed as shown in Fig. 6. It can be found that the residual compressive modulus and strength of composite laminates decreased as exposure time increased at a specific high temperature (200 or 300 C). The decrease in modulus and strength with exposure time can be explained by more deterioration in mechanical properties with longer exposure to high temperature. The effect of thermal exposure on failure modes of composite laminates was observed in Fig. 7. The delamination failure was observed under axial compression for the specimens exposed to temperature in a range of 20e200 C as shown in Fig. 7(a), while the delamination and buckling failure modes were observed for the specimens exposed to temperatures of 250 and 300 C as shown in Fig. 7(b). This behavior is attributed to the degradation of the matrix property when specimens were exposed to higher temperatures, which makes the ability that resins protect and unify the fibers, as well as transfer stress between fibers be reduced. As a result, the fibers, which are not in perfect alignment but rather in a continually wavy state cannot evenly carry the load[30]. Thus the fiber loading difference due to this continually wavy state resulted in the delamination and buckling failure mode.
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Fig. 5 Effect of exposure temperature on the residual compressive modulus (a) and the residual compressive strength (b) of composite laminates
4.3. Compressive responses of sandwich panels An attempt has been made here to evaluate the effect of high temperature exposure on the mechanical behavior and failure mechanisms of CFCSP with pyramidal truss core. The composite sandwich panels were exposed to temperatures of 20, 100, 150, 200, 250, and 300 C for 6 h in order to investigate the effect of thermal exposure temperature on mechanical properties of sandwich panels. During heating, the distinct odor of burning epoxy was evident for the specimens heated at 150, 200, 250, and 300 C, indicating some decomposition of the polymer matrix. After thermal exposure, color changes appeared in the specimens as shown in Fig. 8. The initial black coloration changed successively to light red, then brown and finally black. This observation is indicative of chemical modification occurring at the surface of the epoxy resin due to thermal-oxidation[31]. Small residual deformations due to differential shrinkage between the fibers and the polymer matrix were also observed for the specimens exposed to 150 C. Deformations became more severe as exposure temperature increased. The severe thermal deformations have been revealed for the specimen exposed to 300 C for 6 h as shown in Fig. 8(f). After thermal exposure and before conducting compression tests, Olympus microscope was used to observe the surfaces of composite face sheet. The surfaces of composite face sheet after exposure at different temperatures for 6 h are shown in Fig. 9. It can be seen that the surface was relatively sound and did not show any remarkable degradation for the specimen exposed to temperature in the range of 20e150 C. While the micro crack was observed on the surface of face sheet, which was exposed to 200 C as shown in Fig. 9(d), and the pore and micro cracking were observed on the
Fig. 6 Effect of exposure time on the residual compressive modulus (a) and the residual compressive strength (b) of composite laminates.
surface of face sheet for the specimens exposed to temperatures of 250 and 300 C as shown in Fig. 9(e and f). Some of the cracks have developed into delaminations. This phenomenon indicated that the epoxy matrix is susceptible to the harsh thermal condition. The effect of high temperature exposure on failure mode of composite sandwich panels under axial compression has been also revealed. The sandwich panel displayed a linear elastic
Fig. 7 Compressive failure modes for the composite laminates after thermal exposure: (a) delamination failure for specimens exposed to temperature ranging from 20 to 200 C; (b) delamination and buckling failure for specimens exposed to 250 and 300 C.
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Fig. 8 Specimens after exposure to different temperatures for 6 h at: (a) 20 C; (b) 100 C; (c) 150 C; (d) 200 C; (e) 250 C; (f) 300 C.
response until a sudden unloading of the specimen when the exposure temperature was 20 C. After the linear elastic response, the sandwich panel failed in core shear buckling, and the node rupture occurred at the same time as shown Fig. 10(a). The core shear buckling was dominant failure mode for the specimens exposed to 20 C. No practical laboratory scale column can be designed to probe the Euler buckling mode. The same failure modes were also observed for the specimens exposed to temperature of 100e200 C. However, face sheet crushing was observed for the specimens exposed to temperatures of 250 and 300 C as shown in Fig. 10(b and c). The behavior is attributed to severe heat damage to face sheet, which was caused by the degradation of the matrix, as well as the formation of delamination cracking and pores when specimens were exposed to higher temperatures. Thus, the face sheet crushing occurred for the specimens exposed to temperatures of 250 and 300 C. In order to provide insight into the effect of high temperature exposure on strengthening mechanism, the fiberematrix
interfaces of specimens were examined by scanning electron microscopy (SEM). Fig. 11 shows the fiberematrix interfaces of specimens after exposure to different temperatures for 6 h. The fibers are covered with matrix representing a good fiber to matrix adhesion on the fracture surface of specimens exposed to room temperature as shown in Fig. 11(a), and the fiberematrix interfaces of specimen exposed to 150 C did not exhibit severe change as compared to that of specimen exposed to room temperature as shown in Fig. 11(b). Whereas the fibers are separated from the matrix for the specimen exposed to temperatures of 250 and 300 C, indicating a low fiber to matrix adhesion as shown in Fig. 11(c and d). A summary of the predicted and measured failure loads and failure modes is given in Tables 3 and 4. It can be seen that the predicted failure loads are close to the measured failure loads, despite the predicted values are higher than the experimental values. It can be seen that the residual compressive strength of sandwich panels was dependent on thermal exposure temperatures from the experimental results in Table 3. Severe reductions
Fig. 9 Typical pictures of surfaces of composite face sheet after exposure to different temperatures for 6 h at: (a) 20 C; (b) 100 C; (c) 150 C; (d) 200 C; (e) 250 C; (f) 300 C.
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Fig. 10 The failure modes for composite sandwich panel after exposure to different temperatures for 6 h at: (a) 20 C; (b) 250 C; (b) 300 C.
in residual compressive strength were observed for sandwich panels exposed to 300 C for 6 h. The residual compressive strength of sandwich panels exposed to 300 C for 6 h decreased by 93.30% as compared to that of sandwich panels exposed to
room temperature for 6 h. Table 4 shows the evolution of residual compressive strength with exposure time. It is clearly shown that at a specific high temperature (200 or 300 C), the residual compressive strength decreased as exposure time
Fig. 11 SEM images of fiberematrix interfaces of specimens: (a) interface with thermal exposure to 20 C for 6 h; (b) interface with thermal exposure to 150 C for 6 h; (c) interface with thermal exposure to 250 C for 6 h; (d) interface with thermal exposure to 300 C for 6 h.
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Table 3 Summary of the predicted and measured failure loads and collapse modes after exposure at different temperatures for 6 h. The failure modes in this table are abbreviated as: CSB (core shear buckling) and FC (face sheet crushing) Exposure temperature ( C)
Exposure time (h)
Ppred (kN)
Predicted failure mode
Pmeasured (kN)
Observed failure mode
20 100 150 200 250 300
6 6 6 6 6 6
59.14 58.09 54.96 52.21 30.60 4.16
CSB CSB CSB CSB FC FC
48.96 47.68 45.12 40.30 24.99 3.28
CSB CSB CSB CSB FC FC
Table 4 Summary of the predicted and measured failure loads and collapse modes after exposure at 200 and 300 C for different time. The failure modes in this table are abbreviated as: CSB (core shear buckling) and FC (face sheet crushing) Exposure temperature ( C)
Exposure time (h)
Ppred (kN)
Predicted failure mode
Pmeasured (kN)
Observed failure mode
200
3 6 9 12 15 3 6 9 12 15
53.56 52.21 51.45 50.60 49.68 4.79 4.16 3.67 3.18 2.46
CSB CSB CSB CSB CSB FC FC FC FC FC
40.61 40.30 39.82 39.31 38.91 3.62 3.28 2.80 2.49 2.08
CSB CSB CSB CSB CSB FC FC FC FC FC
300
increased. This can be attributed to more deterioration in residual compressive strength with longer exposure to high temperature. In addition, it is also indicated that the rate of loss in residual compressive strength with time is more pronounced for 300 C than 200 C. The results presented above indicated that both the thermal exposure temperature and time are important factors affecting the failure of composite sandwich panels. 5. Conclusion The aim of the present investigation is to study the effect of high temperature exposure on the mechanical behavior and failure mechanisms of CFCSP with pyramidal truss core. The effect of thermal exposure on mechanical behavior of composite sandwich panels has been revealed. Delamination cracking, pores, and the degradation of fiberematrix interface have been shown to occur after thermal exposure. Severe reduction in residual compressive strength was observed when specimens were exposed to 300 C. High temperature exposure had an important effect on failure modes of composite sandwich panels. The core shear buckling was observed for the specimens exposed to temperature in a range of 20e200 C, while failure was characterized by face sheet crushing for the specimens exposed to temperatures of 250 and 300 C. The residual compressive strength of composite sandwich panels after thermal exposure was predicted using the proposed prediction method in this paper and compared with measured results. Experimental results showed that the measured failure loads are in reasonable agreement with the predictions. Acknowledgments The authors gratefully acknowledge the financial support of the project from the Major State Basic Research Development Program of China (No. 2011CB610303), and the National Natural Science Foundation of China (Nos. 90816024 and 11272105), the
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J. Mater. Sci. Technol., 2013, 29(9), 846e854
Mechanical Behavior and Failure Mechanisms of Carbon Fiber Composite Pyramidal Core Sandwich Panel after Thermal Exposure Jiayi Liu, Zhengong Zhou*, Linzhi Wu, Li Ma Center for Composite Materials and Structures, Harbin Institute of Technology, Harbin 150001, China [Manuscript received May 12, 2012, in revised form November 1, 2012, Available online 18 April 2013]
An attempt has been made here to evaluate the effect of thermal exposure on the mechanical behavior and failure mechanisms of carbon fiber composite sandwich panel with pyramidal truss core under axial compression. Analytical formulae for the collapse strength of composite sandwich panel after thermal exposure were derived. Axial compression tests of composite laminates and sandwich panels after thermal exposure were conducted at room temperature to assess the degradation caused by the thermal exposure. Experimental results showed that the failure of sandwich panel are not only temperature dependent, but are time dependent as well. The decrease in residual compressive strength is mainly attributed to the degradation of the matrix and the degradation of fiberematrix interface, as well as the formation of cracks and pores when specimens are exposed to high temperature. The measured failure loads obtained in the experiments showed reasonable agreement with the analytical predictions. KEY WORDS: Composite; Thermal exposure; Mechanical properties; Sandwich panel
1. Introduction The interest for carbon fiber composite sandwich panel (CFCSP) has grown rapidly over the last decade due to their lightweight attributes and potential multifunctional applications[1e5]. The superior specific strength and stiffness of carbon fiber reinforced polymer composite (FRPC) sandwich panels relative to metallic sandwich panels have been revealed at room temperature[6]. However, FRPC sandwich panel is quite sensitive to temperature variation. The FRPC is combustible, and susceptible to deterioration of mechanical properties when they are exposed to high temperature[7e9]. The matrix dominated properties are expected to be more affected for FRPC when they are exposed to high temperature. Currently, the mechanical behavior and failure mechanisms of CFCSP at room temperature have been studied in detail[10e15], while less is known about the structural integrity of CFCSP after exposure to high temperature. In aircraft structures, some composite parts adjacent to the hot source are exposed to a high temperature environment in presence of oxygen in service, leading to the creation of an oxidized layer on the material surfaces and to cracking of the composite materials[16e18]. The * Corresponding author. Prof., Ph.D.; Tel.: þ86 451 86402396; Fax: þ86 451 86402386; E-mail address: [email protected] (Z. Zhou). 1005-0302/$ e see front matter Copyright Ó 2013, The editorial office of Journal of Materials Science & Technology. Published by Elsevier Limited. All rights reserved. http://dx.doi.org/10.1016/j.jmst.2013.04.013
degradation of matrix and the fiberematrix interface have been reported when carbon fiber composite materials were exposed to high temperature[19]. The thermal distortion of panels made by FRPC under transverse thermal-loading conditions has been discussed[20]. The severe thermal damage by cracks and delamination, and high mass losses have been revealed for 8552/ IM7 when the material was exposed to 340e450 C for 30 min[21]. Foster and Bisby[8] studied the residual tensile strength of carbon FRPC after exposure to elevated temperatures of 100, 200, 300 and 400 C for 3 h. They found that the tensile strength for the carbon FRPC showed no obvious reduction with temperature up to 300 C, but there were severe reductions at 400 C. Cleary et al.[22] tested glass fiber reinforced polymer (GFRP)-wrapped concrete cylinders in axial compression after exposure to elevated temperatures. The results showed the GFRP-confined concrete cylinders lost about 2%, 4%, 13% and 18% of their initial room temperature ultimate strength after exposure for 90 min at temperatures of 120, 135, 150 and 180 C, respectively. The glass transition temperature of the resin system used in this study was quoted as 121 C. With increasing use of CFCSP in aircraft structures where high temperature can occur, understanding the effect of high temperature exposure on mechanical behavior of CFCSP is an important safety issue. Thus, the present work is an attempt to fill this information needed. The effect of high temperature exposure on mechanical properties and damage mechanisms of CFCSP with pyramidal truss cores under axial compression was investigated in this study. The outline of the present paper is as follows. Firstly,
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analytical expressions are derived for the collapse load of composite sandwich panels after thermal exposure. Secondly, the modulus and strength of composite laminates after thermal exposure were studied in order to determine the mechanical properties of composite sandwich panel. Finally, the in-plane compressive responses of pyramidal truss core sandwich panels were investigated and compared with the analytical predictions.
(bending) or core shear buckling, (ii) face sheet wrinkling, and (iii) face sheet crushing. The difference between the core shear buckling and Euler buckling is subtle, since both modes are always simultaneously active and the final deformed shapes are somewhat similar. The differences between the modes can be found through observing the deformation of the core[23].
2. Analytical Predictions for Competing Failure Modes
Euler buckling and core shear buckling are two possible modes of macro elastic buckling in a sandwich panel under axial compression. The Euler buckling load PB(T, t) can be estimated from classical buckling theory[24] as
The schematic of the procedure used to investigate the effect of thermal exposure on mechanical behavior of composite sandwich panel is shown in Fig. 1. For truss core sandwich panel, the width of face sheet may be wider than the width of core as shown in Fig. 1(b). The width of face sheet and core is denoted by symbol B and B0 , respectively. Fig. 1(c) shows the schematic of a pyramidal truss core sandwich panel with length L, face sheet thickness hf, and core height Hc, subjected to in-plane quasi-static compression. A Cartesian coordinate is also established in order to facilitate the analysis as shown in Fig. 1(c): 1-axis is along the length direction, 2-axis is along the width direction and 3-axis is along the thickness direction. Fig. 2 shows the schematic of the pyramidal unit cell with the diameter of truss d, length of truss l, inclined angle u, and the distance k between two closest struts. The relative density r of the pyramidal core is given by
r¼
pd 2 pffiffiffi 2 sin u 2cos u þ 2k
(1)
Three different failure modes exits for sandwich panels under axial compression suggested by Xiong et al.[12]: (i) Euler
2.1. Macro elastic buckling
PB ðT ; tÞ ¼
k12 p2 DðT; tÞ L2
(2)
where k1 ¼ 2 for a column with built-in ends, and DðT ; tÞ is the equivalent flexural rigidity of the composite column. The value of PB ðT ; tÞ and DðT; tÞ is dependent on the thermal exposure temperature T and time t. The flexural rigidity can be expressed as follows: DðT ; tÞ ¼ D0 ðT ; tÞ þ 2Df ðT ; tÞ
(3)
where 2 Ef ðT; tÞBhf hf þ Hc D0 ðT ; tÞ ¼ 2
Df ðT ; tÞ ¼
Ef ðT ; tÞBh3f 12
(4)
(5)
Fig. 1 Schematic of the procedure used to investigate the effect of thermal exposure on mechanical behavior of composite sandwich panel: (a) composite sandwich panel was exposed to high temperature; (b) composite sandwich panel was cooled at room temperature; (c) composite sandwich panel was tested at room temperature under axial compression.
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Table 2 Experimental results for compressive modulus of unidirectional carbon fiber struts after exposure at 200 and 300 C for different time Exposure temperature ( C) 200 300
Fig. 2 Schematic of the unit cell of the pyramidal core.
where Ef ðT ; tÞ is the equivalent elastic modulus along 1-axis of the face sheet. The value of Ef ðT ; tÞ is dependent on the thermal exposure temperature T and time t. The equivalent elastic modulus of the face sheet can be obtained by experiment. The core shear buckling load Ps ðT ; tÞ is set by the shear stiffness of the core, which can be formulated as Ps ðT; tÞ ¼ Gc ðT ; tÞB0 Hc
(6)
where Gc ðT ; tÞ is the effective shear modulus of the pyramidal truss core, and it can be given by Gc ðT; tÞ ¼
1 r Ec ðT ; tÞsin2 2u 8
(7)
where Ec ðT ; tÞ is the effective elastic modulus of the core material in compression[25]. In present paper, the measured values for the effective elastic modulus of the core material after thermal exposure are shown in Tables 1 and 2. According to Allen’s expressions[24], the critical buckling load Pcr ðT; tÞ is given by 1 1 1 ¼ þ Pcr ðT ; tÞ PB ðT; tÞ Ps ðT; tÞ
3h
6h
9h
12 h
15 h
19.66 4.62
19.18 4.10
18.91 3.76
18.62 3.35
18.28 2.62
face sheet is connected with the aspect ratio d/l of the core member: the higher rotation restraining effect will be produced by the larger aspect ratio. It is difficult to compute k2 theoretically for the complex boundary conditions of the intercellular face sheets. While the value of k2 can be obtained numerically, which is about 1.3 for the sandwich panels in this paper. 2.3. Face sheet crushing The face sheet of sandwich panel may crush under axial compression, since the pyramidal truss core is much more compliant as compared to the face sheet. The matrix crushing, fiberematrix shear failure and delamination will be the possible micro failure modes for the composite laminates in compression. Here, we ignore the specific micro failure modes and take them as face sheet crushing macroscopically. The face sheet crushing load can be expressed as Pfc ðT ; tÞ ¼ 2Bhf sfc ðT ; tÞ
(10)
where sfc ðT ; tÞ is the crushing strength of face sheet. The crushing strength of face sheet can be obtained by experiment. 3. Materials and Experimental Methods
(8) 3.1. Material
2.2. Face sheet wrinkling The face sheet will wrinkle under axial compressive load when they are relatively thin. Wrinkling is a local buckling behavior of the face sheet and occurs between the adjacent nodes of attachment. The force Pfw ðT ; tÞ associated with the wrinkling of the face sheet can be estimated from the following expression[23] 2k 2 p2 Df ðT ; tÞ Pfw ðT; tÞ ¼ 2 pffiffiffi 2 2 cosðu þ tÞ
Exposure time
(9)
Fig. 3 shows an example of the manufactured CFCSP with pyramidal truss core, which was fabricated from unidirectional carbon/epoxy prepreg by the molding hot-press method. The sandwich face sheets were fabricated from 14-plies unidirectional carbon/epoxy prepreg stacked in the sequence of [0/90/0/ 90/0/90/0]s, and the truss cores consisted of unidirectional carbon fiber struts. Top and bottom face sheets were interconnected with truss cores. The sandwich face sheets and truss cores were manufactured in one manufacturing process without bonding. The similar in fabricate procedure was introduced in detail by Wang et al[10]. The pyramidal core presented in this paper has dimensions as d ¼ 2.5 mm, k ¼ 7 mm, l ¼ 21.2 mm, Hc ¼ 15 mm, u ¼ 45 and relative density r ¼ 2:24%. For the sandwich face sheets, the thickness of face sheet hf ¼ 1.68 mm, the length of face sheet L ¼ 225 mm, and the width of face sheet
The factor k2 depends on the end-constraints set by the pyramidal truss cores. The rotation restraining effect of truss core to Table 1 Experimental results for compressive modulus of unidirectional carbon fiber struts after exposure at different temperatures for 6h 20 C
100 C
150 C
200 C
250 C
300 C
21.56
21.18
20.10
19.18
16.01
4.10
Fig. 3 The sandwich panel with carbon fiber composite pyramidal truss cores.
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B ¼ 100 mm. The width of lattice truss core B0 ¼ 70.4 mm, since there are two unit cells along width direction. The mechanical properties of composite laminates were also studied in this paper, since the modulus and strength of composite laminate will be used to determine the mechanical properties of composite sandwich panel. In the present fabricating method, the ends of truss strut were embedded in the face sheets, which would cause a bad influence on the continuity of the carbon fiber in composite face sheets. Thus, the mechanical properties of the face sheet may decrease somewhat. In order to determine the mechanical properties exactly, the composite laminates tested in this paper were cut from sandwich panels and obtained by removing the pyramidal truss cores. 3.2. Test methods The composite laminates and sandwich panels were exposed to different temperatures for different time when they had been fabricated. Thermal exposures were accomplished by placing the specimens in a programmable insulated temperature-controlled air oven and heating to required temperature. In order to investigate the effect of thermal exposure temperature on mechanical properties, some of the composite laminates and sandwich panels were exposed to temperatures of 20 C (room temperature), 100, 150, 200, 250, and 300 C for 6 h, whereas other composite laminates and sandwich panels were exposed to temperatures of 200 and 300 C for 3, 6, 9, 12, or 15 h, in order to study the effect of thermal exposure time on mechanical properties. This temperature regime was selected to simulate temperatures that might be experienced by the fiber reinforced polymer system in an actual fire situation[8]. After thermal exposure, the specimens were cooled down to room temperature and kept at room temperature for a period of more than 24 h. Then specimens were tested at room temperature under axial compression till failure. For the composite laminates, compression test was performed on INSTRON 5569 according to ASTM 3410-75[26]. Prior to testing, the strain gages were bonded in axial direction along the centerline of the laminates in order to measure the axial strains. For truss core sandwich panels, the response of truss core sandwich panels was measured in accordance with ASTM C365/ C 364M-05[27]. The axial compression tests were conducted in the quasi-static regime with a displacement rate of 0.5 mm/min. Five repeated tests were conducted in each case to gauge the variability in the test measurements. 4. Results and Discussion 4.1. Thermal analysis of epoxy resin The glass transition temperature (Tg) of the resin used in materials is about 140 C, which is provided by the manufacturer (Beijing Institute of Aeronautical Materials, China). Thermogravimetric analysis (TGA) has been carried out to determine the temperature, at which the epoxy resin matrix began to burn off, signifying severe and irreversible chemical decomposition and loss of the matrices’ mechanical properties. The TGA test was performed in air at a heating rate of 20 C/min by a TA Instruments TGA Q50 thermogravimetric analyzer. Fig. 4 shows the retained massetemperature curve for the epoxy resin matrix measured using TGA. It can be found that the epoxy resin was decomposed in two stages, a first rapid degradation starting from 250 C and ended by a pseudo-plateau. The second
Fig. 4 TGA curves for epoxy resin.
stage leads to complete degradation of epoxy resin. Thermal analysis of polymer matrix by Regnier and Mortaigne[28] indicated that decomposition reaction mainly involves random scission of the polymer chain. Most of the polymer is fragmented into low molecular weight hydrocarbon volatiles and nonflammable gases, and only a small percentage of the original mass is retained as solid char. 4.2. Compressive responses of composite laminates The compressive modulus and strength of composite laminates after thermal exposure are studied here, which will be used to determine mechanical properties of composite sandwich panels. The change of residual compressive modulus and strength vs the exposure temperature is illustrated in Fig. 5. The measured modulus and strength data of composite laminates were based on the average of five tested specimens. It can be seen that the residual compressive modulus and strength decreased with exposure temperature, the severe reductions in modulus and strength were observed for exposure temperatures more than 250 C. The modulus and strength decreased by 82.55% and 96.84%, respectively, when the specimens were exposed to 300 C for 6 h. The thermal degradation due to the thermo-oxidative degradation of the epoxy resin matrix caused the loss of mechanical properties[29]. The effect of thermal exposure time on mechanical properties of composite laminates was also revealed as shown in Fig. 6. It can be found that the residual compressive modulus and strength of composite laminates decreased as exposure time increased at a specific high temperature (200 or 300 C). The decrease in modulus and strength with exposure time can be explained by more deterioration in mechanical properties with longer exposure to high temperature. The effect of thermal exposure on failure modes of composite laminates was observed in Fig. 7. The delamination failure was observed under axial compression for the specimens exposed to temperature in a range of 20e200 C as shown in Fig. 7(a), while the delamination and buckling failure modes were observed for the specimens exposed to temperatures of 250 and 300 C as shown in Fig. 7(b). This behavior is attributed to the degradation of the matrix property when specimens were exposed to higher temperatures, which makes the ability that resins protect and unify the fibers, as well as transfer stress between fibers be reduced. As a result, the fibers, which are not in perfect alignment but rather in a continually wavy state cannot evenly carry the load[30]. Thus the fiber loading difference due to this continually wavy state resulted in the delamination and buckling failure mode.
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Fig. 5 Effect of exposure temperature on the residual compressive modulus (a) and the residual compressive strength (b) of composite laminates
4.3. Compressive responses of sandwich panels An attempt has been made here to evaluate the effect of high temperature exposure on the mechanical behavior and failure mechanisms of CFCSP with pyramidal truss core. The composite sandwich panels were exposed to temperatures of 20, 100, 150, 200, 250, and 300 C for 6 h in order to investigate the effect of thermal exposure temperature on mechanical properties of sandwich panels. During heating, the distinct odor of burning epoxy was evident for the specimens heated at 150, 200, 250, and 300 C, indicating some decomposition of the polymer matrix. After thermal exposure, color changes appeared in the specimens as shown in Fig. 8. The initial black coloration changed successively to light red, then brown and finally black. This observation is indicative of chemical modification occurring at the surface of the epoxy resin due to thermal-oxidation[31]. Small residual deformations due to differential shrinkage between the fibers and the polymer matrix were also observed for the specimens exposed to 150 C. Deformations became more severe as exposure temperature increased. The severe thermal deformations have been revealed for the specimen exposed to 300 C for 6 h as shown in Fig. 8(f). After thermal exposure and before conducting compression tests, Olympus microscope was used to observe the surfaces of composite face sheet. The surfaces of composite face sheet after exposure at different temperatures for 6 h are shown in Fig. 9. It can be seen that the surface was relatively sound and did not show any remarkable degradation for the specimen exposed to temperature in the range of 20e150 C. While the micro crack was observed on the surface of face sheet, which was exposed to 200 C as shown in Fig. 9(d), and the pore and micro cracking were observed on the
Fig. 6 Effect of exposure time on the residual compressive modulus (a) and the residual compressive strength (b) of composite laminates.
surface of face sheet for the specimens exposed to temperatures of 250 and 300 C as shown in Fig. 9(e and f). Some of the cracks have developed into delaminations. This phenomenon indicated that the epoxy matrix is susceptible to the harsh thermal condition. The effect of high temperature exposure on failure mode of composite sandwich panels under axial compression has been also revealed. The sandwich panel displayed a linear elastic
Fig. 7 Compressive failure modes for the composite laminates after thermal exposure: (a) delamination failure for specimens exposed to temperature ranging from 20 to 200 C; (b) delamination and buckling failure for specimens exposed to 250 and 300 C.
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Fig. 8 Specimens after exposure to different temperatures for 6 h at: (a) 20 C; (b) 100 C; (c) 150 C; (d) 200 C; (e) 250 C; (f) 300 C.
response until a sudden unloading of the specimen when the exposure temperature was 20 C. After the linear elastic response, the sandwich panel failed in core shear buckling, and the node rupture occurred at the same time as shown Fig. 10(a). The core shear buckling was dominant failure mode for the specimens exposed to 20 C. No practical laboratory scale column can be designed to probe the Euler buckling mode. The same failure modes were also observed for the specimens exposed to temperature of 100e200 C. However, face sheet crushing was observed for the specimens exposed to temperatures of 250 and 300 C as shown in Fig. 10(b and c). The behavior is attributed to severe heat damage to face sheet, which was caused by the degradation of the matrix, as well as the formation of delamination cracking and pores when specimens were exposed to higher temperatures. Thus, the face sheet crushing occurred for the specimens exposed to temperatures of 250 and 300 C. In order to provide insight into the effect of high temperature exposure on strengthening mechanism, the fiberematrix
interfaces of specimens were examined by scanning electron microscopy (SEM). Fig. 11 shows the fiberematrix interfaces of specimens after exposure to different temperatures for 6 h. The fibers are covered with matrix representing a good fiber to matrix adhesion on the fracture surface of specimens exposed to room temperature as shown in Fig. 11(a), and the fiberematrix interfaces of specimen exposed to 150 C did not exhibit severe change as compared to that of specimen exposed to room temperature as shown in Fig. 11(b). Whereas the fibers are separated from the matrix for the specimen exposed to temperatures of 250 and 300 C, indicating a low fiber to matrix adhesion as shown in Fig. 11(c and d). A summary of the predicted and measured failure loads and failure modes is given in Tables 3 and 4. It can be seen that the predicted failure loads are close to the measured failure loads, despite the predicted values are higher than the experimental values. It can be seen that the residual compressive strength of sandwich panels was dependent on thermal exposure temperatures from the experimental results in Table 3. Severe reductions
Fig. 9 Typical pictures of surfaces of composite face sheet after exposure to different temperatures for 6 h at: (a) 20 C; (b) 100 C; (c) 150 C; (d) 200 C; (e) 250 C; (f) 300 C.
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Fig. 10 The failure modes for composite sandwich panel after exposure to different temperatures for 6 h at: (a) 20 C; (b) 250 C; (b) 300 C.
in residual compressive strength were observed for sandwich panels exposed to 300 C for 6 h. The residual compressive strength of sandwich panels exposed to 300 C for 6 h decreased by 93.30% as compared to that of sandwich panels exposed to
room temperature for 6 h. Table 4 shows the evolution of residual compressive strength with exposure time. It is clearly shown that at a specific high temperature (200 or 300 C), the residual compressive strength decreased as exposure time
Fig. 11 SEM images of fiberematrix interfaces of specimens: (a) interface with thermal exposure to 20 C for 6 h; (b) interface with thermal exposure to 150 C for 6 h; (c) interface with thermal exposure to 250 C for 6 h; (d) interface with thermal exposure to 300 C for 6 h.
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Table 3 Summary of the predicted and measured failure loads and collapse modes after exposure at different temperatures for 6 h. The failure modes in this table are abbreviated as: CSB (core shear buckling) and FC (face sheet crushing) Exposure temperature ( C)
Exposure time (h)
Ppred (kN)
Predicted failure mode
Pmeasured (kN)
Observed failure mode
20 100 150 200 250 300
6 6 6 6 6 6
59.14 58.09 54.96 52.21 30.60 4.16
CSB CSB CSB CSB FC FC
48.96 47.68 45.12 40.30 24.99 3.28
CSB CSB CSB CSB FC FC
Table 4 Summary of the predicted and measured failure loads and collapse modes after exposure at 200 and 300 C for different time. The failure modes in this table are abbreviated as: CSB (core shear buckling) and FC (face sheet crushing) Exposure temperature ( C)
Exposure time (h)
Ppred (kN)
Predicted failure mode
Pmeasured (kN)
Observed failure mode
200
3 6 9 12 15 3 6 9 12 15
53.56 52.21 51.45 50.60 49.68 4.79 4.16 3.67 3.18 2.46
CSB CSB CSB CSB CSB FC FC FC FC FC
40.61 40.30 39.82 39.31 38.91 3.62 3.28 2.80 2.49 2.08
CSB CSB CSB CSB CSB FC FC FC FC FC
300
increased. This can be attributed to more deterioration in residual compressive strength with longer exposure to high temperature. In addition, it is also indicated that the rate of loss in residual compressive strength with time is more pronounced for 300 C than 200 C. The results presented above indicated that both the thermal exposure temperature and time are important factors affecting the failure of composite sandwich panels. 5. Conclusion The aim of the present investigation is to study the effect of high temperature exposure on the mechanical behavior and failure mechanisms of CFCSP with pyramidal truss core. The effect of thermal exposure on mechanical behavior of composite sandwich panels has been revealed. Delamination cracking, pores, and the degradation of fiberematrix interface have been shown to occur after thermal exposure. Severe reduction in residual compressive strength was observed when specimens were exposed to 300 C. High temperature exposure had an important effect on failure modes of composite sandwich panels. The core shear buckling was observed for the specimens exposed to temperature in a range of 20e200 C, while failure was characterized by face sheet crushing for the specimens exposed to temperatures of 250 and 300 C. The residual compressive strength of composite sandwich panels after thermal exposure was predicted using the proposed prediction method in this paper and compared with measured results. Experimental results showed that the measured failure loads are in reasonable agreement with the predictions. Acknowledgments The authors gratefully acknowledge the financial support of the project from the Major State Basic Research Development Program of China (No. 2011CB610303), and the National Natural Science Foundation of China (Nos. 90816024 and 11272105), the
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