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Fatigue damage evolution in thick composite laminates: Combination of X-ray tomography, acoustic emission and digital image correlation
Composites Science and Technology 183 (2019) 107815
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Fatigue damage evolution in thick composite laminates: Combination of Xray tomography, acoustic emission and digital image correlation
T
Abderrahmane Djabalia,b,c, Lotfi Toubalb,∗, Redouane Zitounec, Saïd Rechaka a
Mechanical Engineering and Development Laboratory, Ecole Nationale Polytechnique, 10 Av. Hassen Badi, El-Harrach, 16200, Algiers, Algeria Mechanical Engineering Department, Université Du Québec à Trois-Rivières, 3351 Boul. Des Forges, Trois-Rivières, G9A 5H7, Québec, Canada c Institut Clément Ader (ICA), CNRS UMR 5312, IUT-A of Toulouse Université, 133c, Av. de Rangueil, 31077, Toulouse Cedex, France b
A R T I C LE I N FO
A B S T R A C T
Keywords: Thick composite laminates Fatigue damage evolution X-ray computed tomography Acoustic emission Digital image correlation
The main purpose of this study is to provide a thorough experimental investigation of fatigue damage mechanisms and evolution in thick carbon/epoxy laminate subjected to bending load. The use of X-ray computed tomography (CT) in this study has allowed the visualization of all damage present in the studied laminates, which made it possible to identify, quantify and locate them precisely and therefore, to identify the physical origin of residual strength decrease and acoustic emissions (AE). Furthermore, the results of the AE analysis have provided very valuable information about the nature and evolution of damage. However, the determination of the depth and size of internal damage was not possible with this technique. The displacement field measured by digital image correlation (DIC) made it possible to determine and monitor the strain field evolution during the experiments. The combination of the results of the three non-destructive techniques used in this work has allowed better characterization of fatigue damage evolution in the studied laminates, and provide a complete and accurate description of the different mechanisms involved during their damage process.
1. Introduction Faced with an annual growth rate of more than 4% in the air traffic according to the International Civil Aviation Organization (ICAO), and a global community increasingly aware of the amount of greenhouse gases emitted by aircraft (carbon dioxide, nitrogen oxides) and their detrimental effects on our planet as well as the depletion of fossil fuels, the reduction of fuel consumption in future aircraft has become one of the main challenges of the aviation industry. These new environmental requirements are in line with the economic interest of the actors, since the cost of fuel represents more than 25% of aircraft operating cost and its price is highly unstable. Consequently, aircraft manufacturers have targeted significant reductions in fuel consumption and pollutant emissions for future aircraft, which can only be achieved with simultaneous work on increasing the engines energy efficiency, improving the aircraft aerodynamics and lightning their structure. As the lifetime of new aircraft is between 35 and 40 years, a reduction of 1 kg of their weight allows us, during this period, to save several thousand kilograms of fuel and avoid the emission of tons of CO2 into the atmosphere. Due to their lightness, high strength and stiffness, and their good resistance to corrosion and fatigue, composite materials have been the great alternative to metallic materials in the aeronautical and
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aerospace industry for a few decades. In recent years, the proportion of composite materials used in aircraft structures has exceeded 50% and the substitution of metal materials by composites in aeronautics is no longer reserved for secondary structures, but it was extended to some primary structures, such as the fuselage and wings of the Airbus A350 and Boeing 787, and the fan blades of the LEAP engine. This progress is considered as a major step towards the aim of building a whole aircraft structure from composite materials. However, during this development, the thickness of the composite laminates used in this area has greatly increased and led to several questions about the mechanical and failure behavior of these materials which are little studied in the literature. In the literature, there is no agreement on the definitions or thickness of what can be called a thick composite laminate. Nevertheless, most previous work in this area considers a composite laminate of more than 6 mm thick as a thick composite [1–4]. The study of the mechanical and failure behavior of thick composites has been the subject of several research papers [5–15]. However, most of these papers have been focused on the effect of specimen thickness on the mechanical properties of the composites, and more particularly on their compressive strength. In this context, an extensive literature review of size and scale effects in composites is given by Wisnom [16]. In most of these works, a significant decrease in the strength of the thick specimens
Corresponding author. E-mail address: lotfi[email protected] (L. Toubal).
https://doi.org/10.1016/j.compscitech.2019.107815 Received 30 May 2019; Received in revised form 30 August 2019; Accepted 3 September 2019 Available online 04 September 2019 0266-3538/ © 2019 Elsevier Ltd. All rights reserved.
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compared to the thin specimens was reported. According to these works, the observed decrease is due to the influence of test configuration that are not designed for thick composites [5–7], the thermal gradients that develop during the curing process [9–12], the interlaminar shear, the manufacturing defects and matrix dominated failures [13–15]. Furthermore, the effect of impact damage on the mechanical properties and delamination behavior of thick composite laminates, can be found in Refs. [4,17,18]. It was found that a near edge impact affects the compressive strength of specimens and gives a large area of delamination, whereas a central impact affects the tensile strength of specimens and gives a great number of fiber breaks. Furthermore, it has been found that when a preload is applied, the cumulative surface area of delamination increases, and the projected damaged area decreases. On the other hand, some authors have investigated the damage behavior of thick composites under compressive loading [19–21], they reported that the first failure mechanism appears is shear failure in the matrix precipitated by pre-existing fiber misalignment, and that the kink bands and delamination are predominant failure mechanisms. Nevertheless, only very few papers have reported on the behavior of thick composites under fatigue loading. These papers have investigated the effects of thickness [22–24], impact damage [23], curing cycles [11], and heat generation [25] on the fatigue-behavior of these materials. It was mentioned that the fatigue life of laminates under compression loading decrease with increasing thickness [22]. However, under tension-compression loading, the thicker composites perform more reliably in fatigue, and that even when they are subjected to a low-energy impact [23]. In the last few decades, the non-destructive techniques (NDT) have almost become indispensable to the study and in-service control of composite materials and structures. Among the various NDT used in the study of composite materials, the acoustic emission (AE) is the most common and effective technique for damage detection, identification, location and real-time monitoring [26–33]. Furthermore, with good penetrating ability in most materials and high spatial resolution, the Xray computed tomography (CT) has recently become a powerful and potential tool for materials characterization, it was used in several studies on composite materials as a non-destructive technique and efficient tool for damage, voids and fiber orientation characterization [34–43]. However, a literature review on non-destructive evaluation (NDE) of thick-section composites and sandwich structures [44] shows that the NDE of thick composite materials appears to be relatively immature compared to that of thin composites, and that there are no references in the literature on inspection reliability or acceptance criteria for thick composite structures. The increased use of the thick composite laminates in so-called vital structures of the aircraft requires a full characterization of their static and especially dynamic mechanical and failure behavior. Nonetheless, to date, the modeling, sizing and non-destructive evaluation and control of thick composite materials and structures are mainly based on the results of studies conducted on thin composites. In this regard, the aim of the present study is to provide a thorough experimental investigation of fatigue damage evolution in thick carbon/epoxy laminate subjected to bending load using three non-destructive evaluation and monitoring techniques, as well as to evaluate the potentialities and the limits of each of these techniques. The choice of bending load was motivated by the fact that most primary aircraft structures, such as the fuselage and wings are mainly subjected to bending loads. Furthermore, these tests allow to have a complex stress field, which enrich the study and make it closer to reality.
Table 1 Mechanical properties of HexPly T700-M21-GC. Mechanical properties of material (T700-M21) Young's modulus in the L direction (GPa) Young's modulus in the T direction (GPa) Shear modulus (GPa) Poisson's ratio Ply thickness (mm) Fiber content (%)
E1 = 142 E2 = 8.4 G12 = 3.8 ν12 = ν13 = 0.33 ν21 = 0.023 h = 0.26 Vf = 59
T700GC) supplied by Hexcel Composites. This prepreg is already used in the aeronautical industry to manufacture primary and secondary structures of some aircraft, such as the A380 and A400M. The mechanical properties of the used prepreg are listed in Table 1 [45]. The laminate was prepared in a controlled atmosphere (white room) and compaction was carried out using vacuum pump. A mould for the laminate was prepared and placed in a vacuum bag and evacuated to 0.7 bars. Curing was then conducted at 180 °C (raising at the rate of 2 °C/min) for 120 min during which the pressure was maintained at 7 bar in an autoclave and 0.7 bar of vacuum inside the mould. The composite plate of 300 × 250 × 9.4 mm3 is made up of 36 layers. For the stacking sequence, a typical aeronautical sequence was adopted [45°/135°/90°/0°/45°/90°/90°/135°/90°/90°/135°/90°/90°/45°/0°/ 90°/135°/45°]s. Using the abrasive water jet process, 17 longitudinal (L) and 3 transverse (T) specimens were trimmed from this plate into a rectangular shape with dimensions of 182 × 15.2 × 9.4 mm3 according to ASTM D790-03 (Fig. 1). According to X-ray diffraction measurements and acid digestion tests, the void and fiber contents of the specimens are about 1.6% and 58%, respectively. However, as shown in Fig. 2, the void content is higher in the middle plies than near-surface plies, and the ply thickness is not constant and depends on several factors, such as fiber orientation and position in the specimen thickness. 2.2. Instrumentation 2.2.1. Acoustic emission A two-channel MISTRAS data acquisition system from Physical Acoustics Corporation with a sampling rate of 4 MHz and a 40 dB preamplification was used to record AE data. AE measurements were achieved using two resonant MICRO-80 sensors with an operating frequency range of 100–1000 kHz. The sensors are coupled on the faces of the specimens with a thin layer of silicon grease and grasped with two small clamps. A threshold of 35 dB was used to suppress background noise. The peak definition time (PDT), hit definition time (HDT) and hit lockout time (HLT) were set at 50, 150 and 300 μs, respectively. A three pencil lead break tests were performed before each test out to calibrate the source location of AE events and check the coupling state between the specimen and the sensors. 2.2.2. X-ray tomography The X-ray CT measurements were carried out using micro-computed tomography scanner EasyTom 130 (Fig. 3). The specimens tested were exposed to radiation produced by X-ray tube and rotated through 360° to obtain a sequence of 2D tomograms for third-dimension reconstruction. The X-ray voltage and current were set to 130 kV and 300 mA, respectively. The scanner has sealed micro-focus X-ray tube that has a spot size of 3 μm and high-resolution flat panel detector with 1920 × 1536 pixels of 127 μm. The maximum possible resolution, under the measurement conditions of this study, was around 18 μm.
2. Experimentation
2.2.3. Digital image correlation A two-dimensional (2D) digital image correlation system (StrainMaster, LaVision Inc.) was used for in situ measurement of displacement and surface strain fields in the thickness of the specimens
2.1. Materials and specimens All specimens of the present study were made from a unidirectional T700 carbon-fiber/M21 epoxy-resin system prepreg (M21/35%/268/ 2
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Fig. 1. Dimensions and stacking sequence of the composite plate and specimens.
the ASTM D790-03 standard. A three longitudinal and three transverse specimens were tested. The testing machine was an Instron model LMU150, equipped with a 150 kN load cell. The fatigue bending tests were conducted in an MTS-370.10 servohydraulic testing machine equipped with a 100 kN load cell and hydraulic clamping jaws. All fatigue tests were performed in sinusoidal displacement control at room temperature and constant displacement ratio (R) of 0. The displacement ratio is defined as δmin/δmax, where δmin and δmax are the minimum and maximum applied displacement, respectively. The specimens were subjected to three-point bending tests with displacement ranging from 35% δf to 75% δf (δf is the static failure displacement), with a step size of 5% and each for blocks of 10 000 cycles. To prevent thermal effects, the tests frequency was set at 5 Hz [46] and the load blocks were interspersed with rest period of 10 min to let the specimens returned to its initial temperature (Fig. 4). The quasistatic and dynamic tests were instrumented with two acoustic emission sensors and CCD camera (Fig. 5). The elastic modulus of specimens
during testing. The CCD cameras of this system have a sensor of 7.2 × 5.4 mm2, with a resolution of 1628 × 1236 pixels and rate of 14 frames per second. A speckle pattern provided by the system manufacturer was applied on the side surface of the test specimens. The post processing of the results was performed by means of the commercial software Davis 8.2.3 (LaVision). The correlation parameters used for the measurement of the displacement field are 9 pixels for the scanning step and 16 pixels for the digital gauge. The calculation of the strain field is carried out after the measurement of the displacement field. 3. Experimental setup and procedure A quasi-static three-point bending tests were initially performed to determine the elastic properties of laminate, the fatigue tests parameters and the loading type and ply orientation influence on the laminate failure. The tests were performed at room temperature with a loading rate of 1 mm/min and a span length of 150.4 mm, according to
Fig. 2. (a) 3D reconstructed CT scan showing the distribution of the segmented voids in the specimen. (b) Optical micrograph showing the variation of ply thickness through the specimen thickness. 3
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Fig. 3. Experimental set-up for X-ray CT scanning.
displacement of specimens. The stress is calculated by the following equation, which yield the apparent strength based on homogeneous beam theory: σf = 3PL/2bd 2 . Where P, L, b and d are load, support span, width and depth of the specimen, respectively. The strain is calculated by the following equation: εf = 6Dd/ L2 . Where εf and D are strain in the outer surface and maximum deflection of the center of the specimen measured by DIC, respectively. The elastic moduli were determined by the Chord-modulus Method. The first part of the curves shows a linear elastic behavior. At a strain of approximately 1% and 1.35% for L and T specimens, respectively, the stress–strain curves start to deviate from the elastic line and the composite behaves as a quasi-elastic material until failure where the composite has a quasi-brittle behavior. The final failure of the first specimen (L) occurred when several layers in the tensile and compressive zones of the specimen were suddenly and simultaneously broken, while for the second specimen (T) the final failure is initiated by the sudden breaking of several layers in the compressive zone followed by the breaking of almost the same number of layers in the tensile zone. Although damage initiation in the specimen L is earlier than that in the specimen T, the strain at failure of the specimen L is much greater than that of the specimen T. Since the total number of
were non-destructively measured before testing using the impulse excitation technique (IET) according to ASTM E1876-09, and all specimens were scanned after testing with a micro-computed tomography scanner (EasyTom 130). Fig. 1 illustrates the orientation of specimens during the X‐ray scan. 4. Results and discussions 4.1. Quasi-static bending tests The mechanical behavior and damage of thick carbon/epoxy laminates subjected to static and dynamic bending loads were already investigated in our previous work [45]. In this study, which could be considered as a complement to the previous work, we were particularly interested by the fatigue damage evolution of these composite laminates as well as the influence of the loading type and ply orientation on their failure mechanisms. The experimental results of quasi-static bending tests are reported in Fig. 6. This figure shows the normalized stress–strain curves for two specimens that have different ply orientation. The stress and strain are calculated according to ASTM D790-03 using the load cell data and real
Fig. 4. Scheme of the applied displacement during fatigue testing. 4
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Fig. 5. Experimental set-up for fatigue testing.
details of the internal and external damage of specimens are observed. For the first specimen (L), fourteen layers were broken in tension and only five layers in compression. While for the second specimen T, seventeen layers were broken in compression and ten layers in tension. Thirteen of the fourteen layers broken in tension in the specimen L are oriented at 45°, 90° and 135°, and only one of the five layers broken in compression is oriented at 0°. All these layers were suddenly broken for a stress of 570 MPa and a strain of 2.3%. For the specimen T, the majority of the twenty-seven layers broken are oriented at 0°. The final failure of this specimen (T) is initiated by a sudden break of several compressed layers for a stress and strain of approximately 960 MPa and 1.9%, respectively. Moreover, the break of these layers was not produced exactly below the loading axis, but it has been produced in its sides. This is probably due to the fact that the large load applied perpendicularly to the middle of the specimen prevented its layers to break in this zone. It can also be seen in these figures that the predominant damage mechanism in the specimen L is the matrix cracking in off-axis plies, whereas for the specimen T it is the fiber breaking in on-axis plies. Since the ratio E11/E22 of the elementary layer is more than 15, the differences observed previously between the mechanical behavior and
layers as well as the number of 45° and 135° layers are the same in the two specimens, the difference observed in their mechanical behavior is mainly attributed to the large difference that exists in the number of 90° and 0° layers. The mechanical properties of L and T specimens are summarized in Table 2. The results presented in this table revealed a low variability in the mechanical properties of specimens that have the same ply orientation, which reflects the reliability of specimens manufacturing and cutting processes. However, for the two different ply orientations, the elastic moduli measured by non-destructive technique (IET) are approximately 5–7% greater than that measured by bending tests. Several previous studies have reported the existence of a slight discrepancy between the elastic modulus measured by destructive and non-destructive methods, and they attributed this discrepancy to the fundamental differences that exist in the theoretical bases and practical conditions between the two methods [45,47–49]. Furthermore, the elastic moduli calculated by Classical Laminate Theory (CLT) are approximately 25% and 30% greater than those measured by IET and bending tests, respectively. The CT cross-sectional views of the L and T specimens, in the (YZ) and (XY) planes, are shown in Figs. 7 and 8 respectively, from which
Fig. 6. Three-point bending stress–strain curves of two specimens that have different ply orientation. 5
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Table 2 Mechanical properties of specimens. Specimens
L T
Elastic modulus (GPa) IET
Bending tests
CLT
30.03 ± 0.46 56.18 ± 0.54
27.87 ± 0.80 53.45 ± 0.73
36.26 70
Maximum stress (MPa)
Strain at failure (%)
546.93 ± 17.63 956.34 ± 23.28
2.25 ± 0.06 1.96 ± 0.05
100% at the first cycle of each displacement level of 10 000 cycles, and it was calculated from the residual force measured by the load-cell of the fatigue machine by the same equation used in the static tests, according to ASTM D790-03. After 10 000 loading cycles at 35–65% δf, no damage was observed on the outside surface of the specimen and the drop in the residual strength does not exceed 3% for the first four levels (35–50% δf) and 6% for the three following levels (55–65% δf). Nevertheless, from the first loading cycles at 70% δf, the rate of the residual strength decrease has significantly increased compared to the first seven levels, and after 3200 loading cycles, a macro damage generated in the compressive zone led to a suddenly drop of about 32.8% in the residual strength of specimen. Finally, at 75% δf, a significant
damage mechanisms of L and T specimens are mainly due to the difference in the number of on-axis plies in these specimens. Indeed, the behavior of the specimen L is dominated by the matrix and interface behavior, while that of the specimen T is dominated by the fiber behavior. 4.2. Fatigue bending tests 4.2.1. Residual strength degradation Fig. 9 shows the L-specimen residual strength degradation as function of cycle number for nine applied displacement levels (from 35% to 75% δf). The residual strength of specimen was considered equal to
Fig. 7. CT cross-sectional views in the YZ and XY planes of the specimen L. 6
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Fig. 8. CT cross-sectional views in the YZ and XY planes of the specimen T.
Fig. 9. Normalized residual strength degradation of a specimen broken after nine loading levels.
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Fig. 10. Distribution of AE events recorded during the fatigue test of the specimen broken after nine loading levels.
Fig. 11. CT cross-sectional view in the YZ plane of the specimen after the first three loading levels.
specimen lost about 50% and 70% of its initial strength, respectively. Almost the same curves were also obtained with other specimens.
decrease in the residual strength was observed after each cycle, up to about 8500 cycles where a drop of more than 27% in the residual strength, occurred after a sudden break of several layers in the tensile zone, leading to the final failure of the specimen. The blue and red arrows in Fig. 9 show that, after 10 000 cycles at 70% δf and 75% δf the
4.2.2. Acoustic emission monitoring and damage mechanisms evolution Fig. 10 shows the distribution of AE events recorded during the 8
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layers and the delamination of a single interface were observed in the compressive zone. These observations validate the results obtained through the analysis of the acoustic activity recorded during tests, where the events attributed to the interface delamination were recorded from 50% δf, and those attributed to the fiber breaking were recorded from 55% δf. Fig. 14 shows the CT cross-sectional views of the broken specimen after nine loading levels in the different planes, where the nature, location and size of the main damage mechanisms previously identified by acoustic emission technique, are shown. It can be seen that the final failure of the laminate results from the combination of all these different damage mechanisms which intervene at different scales and can interact with each other. Furthermore, it can also be observed that the number of layers broken in tension is much higher than that of layers broken in compression. This difference is mainly due to the non-symmetrical distribution of compressive and tensile strains with respect to the horizontal median plane of the specimen during loading.
fatigue test of a specimen broken after nine loading levels according to their main features (amplitude, number of counts and duration). The choice of the applied filters and the attribution of AE events to the damage mechanisms are based on the results of a previous study on the same material [45] and some other works [50–53]. From the first displacement level (35% δf), which is well below the static damage threshold (50% δf), several events with amplitude of less than 60 dB were recorded. According to the results of the static study [36], the source mechanism of these events is matrix cracking. Up to 45% δf, the amplitude of almost all recorded events has not exceeded 60 dB. For all previous displacement levels, the number of events attributed to the matrix cracking increases with displacement level increasing. At 50% δf, some events with amplitude of 60–70 dB were recorded. These events are attributed to the interfaces delamination within the specimen tensile zone. From 55% δf the amplitude of some recorded events has exceeded 70 dB, which indicates the beginning of fiber failure. Finally, during the last four displacement levels (60, 65, 70 and 75% δf), the number of events attributed to the three damage mechanisms previously detected (matrix cracking, delamination, fibers breaking) increases rapidly with raising the displacement level until the final failure of the specimen.
4.2.4. Through-thickness strain field evolution To characterize the evolution of the longitudinal strain through the thickness of the laminate during different fatigue loading levels, the DIC technique was used and five virtual strain gauges were created across the thickness of the specimen. Fig. 15 shows the positioning of virtual strain gauges and the evolution of the longitudinal strain through the thickness of specimen at various loading levels. The curves intersection point is inside the compressive zone of the specimen and the neutral plane moves towards this zone with increasing loading level. For the first four loading level, the longitudinal strain increases almost linearly with increasing load and from the neutral plane toward the surface. However, from the fifth level (55% δf), the increase of the longitudinal strain becomes nonlinear. The appearance of some macro-damage on the surface of specimens are at the origin of this nonlinearity and the reason for the inability to monitor the longitudinal strain evolution of the last four levels.
4.2.3. X–ray computed tomography and damage mechanisms evolution Fig. 11 shows the CT cross-sectional view of the L-specimen in YZ plane after the first three loading levels (35, 40 and 45% δf). Despite the fact that the three loading levels applied are below the static damage threshold (50% δf) and no damage is observed on the outside surface of the specimen, some transverse matrix cracks in the off-axis plies of the laminate can be clearly observed from this figure. Therefore, this damage mechanism is the main responsible for the residual strength drop during the first three loading levels as it is also the main source of the AE events recorded. The CT cross-sectional views in the (YZ) and (XY) planes of the fatigued specimen after the first seven loading levels are shown in Fig. 12 and Fig. 13, respectively. Where matrix cracks in the off-axis plies and some delamination at the 135°/90°, 0°/90° and 135°/45° interfaces as well as some fibers break in the 90° layers are observed. One can also noticed that the damage in the tensile zone of the specimen is greater than that in the compressive zone. In the tensile zone, matrix cracking of thirteen layers, delamination of two interfaces and some fiber breaking were observed, whereas the matrix cracking of only three
4.3. Loading type influence Fig. 16 correspond to CT cross-sectional views in the YZ plane of broken specimens in static and cyclic loading. The figure shows the most damaged area of the specimens, which is in the vicinity of the loading axis. In the static loading, the majority of damage occurred in a
Fig. 12. CT cross-sectional view in the YZ plane of the specimen after the first seven loading levels. 9
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Fig. 13. CT cross-sectional view in the XY plane of the specimen after the first seven loading levels.
Fig. 14. CT cross-sectional views in the YZ and XY planes of the broken specimen after nine loading levels.
specimen thickness, which is reflected by the linearity of the longitudinal strain variation.
plane perpendicular to the y-axis and at 5 mm to the left of the loading axis, whereas in the cyclic loading, the damages occurred in the compressive zone and in the tensile zone are not on the same plane. The number of broken layers is the same for both types of loading, but the predominant damage mechanism in static loading is matrix cracking in off-axis plies, while in cyclic loading delamination at the 0°/90°, 90°/ 135° and 45°/135° interfaces is the most predominant one. Fig. 17 shows the positioning of virtual strain gauges and the variation of the longitudinal strain through the specimen thickness at 55%δf in static and cyclic loading. For cyclic loading, the strain is represented after 10 000 cycles. The figure shows that in cyclic loading, the neutral plane of the specimen is not exactly in the middle of the specimen thickness as in static loading. Furthermore, in the static loading, the longitudinal strain increases non-linearly from the neutral plane toward the surface, while in cyclic loading, the longitudinal strain increases almost linearly. The preceding observations can be explained by the following hypothesis: After several loading cycles at the same displacement amplitude, the evolution of the different damage mechanisms in the laminate leads to an attenuation of the stress concentration and a more uniform distribution of stresses through the
5. Discussion and conclusion The main objective of this study is to provide a thorough experimental investigation of fatigue damage mechanisms and evolution in thick carbon/epoxy laminate subjected to bending load. To achieve this goal, three non-destructive evaluation techniques were used. Furthermore, the evaluation of the advantages and limitations of each used technique, and the influence of the loading type and ply orientation on the laminate failure process were the second objective of the present study. - After 10 000 loading cycles at 35–45% δf, that are below the static damage threshold (50% δf), no damage was observed on the outside surface of the specimen, although a drop in the residual strength that does not exceed 3% was observed and several events with amplitude less than 60 dB were recorded. The CT cross-sectional views of the specimen showed that some transverse matrix cracks in 10
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Fig. 15. Evolution of the longitudinal strain through the thickness of specimen at various loading levels.
laminate is matrix cracking in off-axis plies, whereas for the second laminate fiber breaking in on-axis plies is the most predominant one. For the loading type influence, a few damages were detected from the first cyclic loading level (35% δf) which was well below the static damage threshold (50% δf). The number of broken layers is the same for both types of loading, but the nature and location of damage differ substantially between the static and cyclic loading. So that, the predominant damage mechanism in static loading is matrix cracking in off-axis plies, while in cyclic loading delamination at the 0°/90°, 90°/135° and 45°/135° interfaces is the most predominant one. - The use of X-ray tomography in this study allowed visualization of all damage present in the studied laminates, which made it possible to identify, quantify and locate them precisely and therefore, to identify the physical origin of residual strength decrease and acoustic emissions. In general, this powerful technique allowed us to provide a complete and precise characterization of damage of the studied laminates. - The analysis of acoustic emission activity recorded during the tests provided many information about the evolution, identification and localization of the damage. However, the determination of the depth, size and distribution of internal damage was not possible with this technique. - The DIC technique used in this study made it possible the measurement of the superficial displacement field and the calculation of the strain field. These data allowed us to make a good characterization of the mechanical behavior of material and to monitor the
the off-axis plies of the laminate are the main responsible for the residual strength drop and the main source of the AE events recorded. For the four following levels (45–65% δf), some delamination and a drop of less than 6% in the residual strength were observed. Moreover, the amplitude of some recorded events has exceeded 70 dB. The CT cross-sectional views of the specimen showed that several matrix cracks in the off-axis plies and some delamination at the 135°/90°, 0°/90° and 135°/45° interfaces as well as some fibers break in the 90° layers are the physical origin of the previous observations. Finally, for the last two loading levels (70 and 75% δf), a significant decrease in the residual strength was observed after each cycle, and a sudden rupture of several layers in the compressive zone and then in the tensile zone of the specimen, which leads to a drop of more than 27% in the residual strength, were seen with the naked eye. In parallel, a large number of events attributed to the main three damage mechanisms (matrix cracking, delamination, fibers breaking) was recorded. The CT cross-sectional views of the specimen showed the nature, location and size of these different damage mechanisms previously identified by AE technique. - The ply orientation and loading type had a great influence on the mechanical and damage behavior of composite laminates. Indeed, the behavior of the laminate consisting of a large number of off-axis plies was governed by the matrix and interface behavior, while that of the laminate consisting of a large number of on-axis plies it was governed by the fiber behavior. Moreover, the CT cross-sectional views showed that the predominant damage mechanism in the first
Fig. 16. CT cross-sectional views in the YZ plane of broken specimens in static and cyclic loading. 11
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Fig. 17. Variation of the longitudinal strain through the specimen thickness at 55% δf in static and cyclic loading.
strain field evolution. [13]
Finally, we believe that this work contributes to a better understanding of the fatigue damage evolution in thick carbon/epoxy laminates subjected to bending load and a better evaluation of the advantages and limitations of each non-destructive technique used. The main perspectives of this work is to provide an accurate characterization and description of the mechanical and damage behavior of these laminates under different loading conditions, and to evaluate the effectiveness of other non-destructive techniques to characterize and control these materials during their lifetime. These perspectives are aimed at constructing models that predict, as reliably as possible, the response of thick laminated composite structures to different external loadings and the establishment of specific standards and criteria for NDE of thick composite structures.
[14] [15] [16] [17] [18]
[19] [20] [21]
Appendix A. Supplementary data [22]
Supplementary data to this article can be found online at https:// doi.org/10.1016/j.compscitech.2019.107815
[23]
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Contents lists available at ScienceDirect
Composites Science and Technology journal homepage: www.elsevier.com/locate/compscitech
Fatigue damage evolution in thick composite laminates: Combination of Xray tomography, acoustic emission and digital image correlation
T
Abderrahmane Djabalia,b,c, Lotfi Toubalb,∗, Redouane Zitounec, Saïd Rechaka a
Mechanical Engineering and Development Laboratory, Ecole Nationale Polytechnique, 10 Av. Hassen Badi, El-Harrach, 16200, Algiers, Algeria Mechanical Engineering Department, Université Du Québec à Trois-Rivières, 3351 Boul. Des Forges, Trois-Rivières, G9A 5H7, Québec, Canada c Institut Clément Ader (ICA), CNRS UMR 5312, IUT-A of Toulouse Université, 133c, Av. de Rangueil, 31077, Toulouse Cedex, France b
A R T I C LE I N FO
A B S T R A C T
Keywords: Thick composite laminates Fatigue damage evolution X-ray computed tomography Acoustic emission Digital image correlation
The main purpose of this study is to provide a thorough experimental investigation of fatigue damage mechanisms and evolution in thick carbon/epoxy laminate subjected to bending load. The use of X-ray computed tomography (CT) in this study has allowed the visualization of all damage present in the studied laminates, which made it possible to identify, quantify and locate them precisely and therefore, to identify the physical origin of residual strength decrease and acoustic emissions (AE). Furthermore, the results of the AE analysis have provided very valuable information about the nature and evolution of damage. However, the determination of the depth and size of internal damage was not possible with this technique. The displacement field measured by digital image correlation (DIC) made it possible to determine and monitor the strain field evolution during the experiments. The combination of the results of the three non-destructive techniques used in this work has allowed better characterization of fatigue damage evolution in the studied laminates, and provide a complete and accurate description of the different mechanisms involved during their damage process.
1. Introduction Faced with an annual growth rate of more than 4% in the air traffic according to the International Civil Aviation Organization (ICAO), and a global community increasingly aware of the amount of greenhouse gases emitted by aircraft (carbon dioxide, nitrogen oxides) and their detrimental effects on our planet as well as the depletion of fossil fuels, the reduction of fuel consumption in future aircraft has become one of the main challenges of the aviation industry. These new environmental requirements are in line with the economic interest of the actors, since the cost of fuel represents more than 25% of aircraft operating cost and its price is highly unstable. Consequently, aircraft manufacturers have targeted significant reductions in fuel consumption and pollutant emissions for future aircraft, which can only be achieved with simultaneous work on increasing the engines energy efficiency, improving the aircraft aerodynamics and lightning their structure. As the lifetime of new aircraft is between 35 and 40 years, a reduction of 1 kg of their weight allows us, during this period, to save several thousand kilograms of fuel and avoid the emission of tons of CO2 into the atmosphere. Due to their lightness, high strength and stiffness, and their good resistance to corrosion and fatigue, composite materials have been the great alternative to metallic materials in the aeronautical and
∗
aerospace industry for a few decades. In recent years, the proportion of composite materials used in aircraft structures has exceeded 50% and the substitution of metal materials by composites in aeronautics is no longer reserved for secondary structures, but it was extended to some primary structures, such as the fuselage and wings of the Airbus A350 and Boeing 787, and the fan blades of the LEAP engine. This progress is considered as a major step towards the aim of building a whole aircraft structure from composite materials. However, during this development, the thickness of the composite laminates used in this area has greatly increased and led to several questions about the mechanical and failure behavior of these materials which are little studied in the literature. In the literature, there is no agreement on the definitions or thickness of what can be called a thick composite laminate. Nevertheless, most previous work in this area considers a composite laminate of more than 6 mm thick as a thick composite [1–4]. The study of the mechanical and failure behavior of thick composites has been the subject of several research papers [5–15]. However, most of these papers have been focused on the effect of specimen thickness on the mechanical properties of the composites, and more particularly on their compressive strength. In this context, an extensive literature review of size and scale effects in composites is given by Wisnom [16]. In most of these works, a significant decrease in the strength of the thick specimens
Corresponding author. E-mail address: lotfi[email protected] (L. Toubal).
https://doi.org/10.1016/j.compscitech.2019.107815 Received 30 May 2019; Received in revised form 30 August 2019; Accepted 3 September 2019 Available online 04 September 2019 0266-3538/ © 2019 Elsevier Ltd. All rights reserved.
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compared to the thin specimens was reported. According to these works, the observed decrease is due to the influence of test configuration that are not designed for thick composites [5–7], the thermal gradients that develop during the curing process [9–12], the interlaminar shear, the manufacturing defects and matrix dominated failures [13–15]. Furthermore, the effect of impact damage on the mechanical properties and delamination behavior of thick composite laminates, can be found in Refs. [4,17,18]. It was found that a near edge impact affects the compressive strength of specimens and gives a large area of delamination, whereas a central impact affects the tensile strength of specimens and gives a great number of fiber breaks. Furthermore, it has been found that when a preload is applied, the cumulative surface area of delamination increases, and the projected damaged area decreases. On the other hand, some authors have investigated the damage behavior of thick composites under compressive loading [19–21], they reported that the first failure mechanism appears is shear failure in the matrix precipitated by pre-existing fiber misalignment, and that the kink bands and delamination are predominant failure mechanisms. Nevertheless, only very few papers have reported on the behavior of thick composites under fatigue loading. These papers have investigated the effects of thickness [22–24], impact damage [23], curing cycles [11], and heat generation [25] on the fatigue-behavior of these materials. It was mentioned that the fatigue life of laminates under compression loading decrease with increasing thickness [22]. However, under tension-compression loading, the thicker composites perform more reliably in fatigue, and that even when they are subjected to a low-energy impact [23]. In the last few decades, the non-destructive techniques (NDT) have almost become indispensable to the study and in-service control of composite materials and structures. Among the various NDT used in the study of composite materials, the acoustic emission (AE) is the most common and effective technique for damage detection, identification, location and real-time monitoring [26–33]. Furthermore, with good penetrating ability in most materials and high spatial resolution, the Xray computed tomography (CT) has recently become a powerful and potential tool for materials characterization, it was used in several studies on composite materials as a non-destructive technique and efficient tool for damage, voids and fiber orientation characterization [34–43]. However, a literature review on non-destructive evaluation (NDE) of thick-section composites and sandwich structures [44] shows that the NDE of thick composite materials appears to be relatively immature compared to that of thin composites, and that there are no references in the literature on inspection reliability or acceptance criteria for thick composite structures. The increased use of the thick composite laminates in so-called vital structures of the aircraft requires a full characterization of their static and especially dynamic mechanical and failure behavior. Nonetheless, to date, the modeling, sizing and non-destructive evaluation and control of thick composite materials and structures are mainly based on the results of studies conducted on thin composites. In this regard, the aim of the present study is to provide a thorough experimental investigation of fatigue damage evolution in thick carbon/epoxy laminate subjected to bending load using three non-destructive evaluation and monitoring techniques, as well as to evaluate the potentialities and the limits of each of these techniques. The choice of bending load was motivated by the fact that most primary aircraft structures, such as the fuselage and wings are mainly subjected to bending loads. Furthermore, these tests allow to have a complex stress field, which enrich the study and make it closer to reality.
Table 1 Mechanical properties of HexPly T700-M21-GC. Mechanical properties of material (T700-M21) Young's modulus in the L direction (GPa) Young's modulus in the T direction (GPa) Shear modulus (GPa) Poisson's ratio Ply thickness (mm) Fiber content (%)
E1 = 142 E2 = 8.4 G12 = 3.8 ν12 = ν13 = 0.33 ν21 = 0.023 h = 0.26 Vf = 59
T700GC) supplied by Hexcel Composites. This prepreg is already used in the aeronautical industry to manufacture primary and secondary structures of some aircraft, such as the A380 and A400M. The mechanical properties of the used prepreg are listed in Table 1 [45]. The laminate was prepared in a controlled atmosphere (white room) and compaction was carried out using vacuum pump. A mould for the laminate was prepared and placed in a vacuum bag and evacuated to 0.7 bars. Curing was then conducted at 180 °C (raising at the rate of 2 °C/min) for 120 min during which the pressure was maintained at 7 bar in an autoclave and 0.7 bar of vacuum inside the mould. The composite plate of 300 × 250 × 9.4 mm3 is made up of 36 layers. For the stacking sequence, a typical aeronautical sequence was adopted [45°/135°/90°/0°/45°/90°/90°/135°/90°/90°/135°/90°/90°/45°/0°/ 90°/135°/45°]s. Using the abrasive water jet process, 17 longitudinal (L) and 3 transverse (T) specimens were trimmed from this plate into a rectangular shape with dimensions of 182 × 15.2 × 9.4 mm3 according to ASTM D790-03 (Fig. 1). According to X-ray diffraction measurements and acid digestion tests, the void and fiber contents of the specimens are about 1.6% and 58%, respectively. However, as shown in Fig. 2, the void content is higher in the middle plies than near-surface plies, and the ply thickness is not constant and depends on several factors, such as fiber orientation and position in the specimen thickness. 2.2. Instrumentation 2.2.1. Acoustic emission A two-channel MISTRAS data acquisition system from Physical Acoustics Corporation with a sampling rate of 4 MHz and a 40 dB preamplification was used to record AE data. AE measurements were achieved using two resonant MICRO-80 sensors with an operating frequency range of 100–1000 kHz. The sensors are coupled on the faces of the specimens with a thin layer of silicon grease and grasped with two small clamps. A threshold of 35 dB was used to suppress background noise. The peak definition time (PDT), hit definition time (HDT) and hit lockout time (HLT) were set at 50, 150 and 300 μs, respectively. A three pencil lead break tests were performed before each test out to calibrate the source location of AE events and check the coupling state between the specimen and the sensors. 2.2.2. X-ray tomography The X-ray CT measurements were carried out using micro-computed tomography scanner EasyTom 130 (Fig. 3). The specimens tested were exposed to radiation produced by X-ray tube and rotated through 360° to obtain a sequence of 2D tomograms for third-dimension reconstruction. The X-ray voltage and current were set to 130 kV and 300 mA, respectively. The scanner has sealed micro-focus X-ray tube that has a spot size of 3 μm and high-resolution flat panel detector with 1920 × 1536 pixels of 127 μm. The maximum possible resolution, under the measurement conditions of this study, was around 18 μm.
2. Experimentation
2.2.3. Digital image correlation A two-dimensional (2D) digital image correlation system (StrainMaster, LaVision Inc.) was used for in situ measurement of displacement and surface strain fields in the thickness of the specimens
2.1. Materials and specimens All specimens of the present study were made from a unidirectional T700 carbon-fiber/M21 epoxy-resin system prepreg (M21/35%/268/ 2
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Fig. 1. Dimensions and stacking sequence of the composite plate and specimens.
the ASTM D790-03 standard. A three longitudinal and three transverse specimens were tested. The testing machine was an Instron model LMU150, equipped with a 150 kN load cell. The fatigue bending tests were conducted in an MTS-370.10 servohydraulic testing machine equipped with a 100 kN load cell and hydraulic clamping jaws. All fatigue tests were performed in sinusoidal displacement control at room temperature and constant displacement ratio (R) of 0. The displacement ratio is defined as δmin/δmax, where δmin and δmax are the minimum and maximum applied displacement, respectively. The specimens were subjected to three-point bending tests with displacement ranging from 35% δf to 75% δf (δf is the static failure displacement), with a step size of 5% and each for blocks of 10 000 cycles. To prevent thermal effects, the tests frequency was set at 5 Hz [46] and the load blocks were interspersed with rest period of 10 min to let the specimens returned to its initial temperature (Fig. 4). The quasistatic and dynamic tests were instrumented with two acoustic emission sensors and CCD camera (Fig. 5). The elastic modulus of specimens
during testing. The CCD cameras of this system have a sensor of 7.2 × 5.4 mm2, with a resolution of 1628 × 1236 pixels and rate of 14 frames per second. A speckle pattern provided by the system manufacturer was applied on the side surface of the test specimens. The post processing of the results was performed by means of the commercial software Davis 8.2.3 (LaVision). The correlation parameters used for the measurement of the displacement field are 9 pixels for the scanning step and 16 pixels for the digital gauge. The calculation of the strain field is carried out after the measurement of the displacement field. 3. Experimental setup and procedure A quasi-static three-point bending tests were initially performed to determine the elastic properties of laminate, the fatigue tests parameters and the loading type and ply orientation influence on the laminate failure. The tests were performed at room temperature with a loading rate of 1 mm/min and a span length of 150.4 mm, according to
Fig. 2. (a) 3D reconstructed CT scan showing the distribution of the segmented voids in the specimen. (b) Optical micrograph showing the variation of ply thickness through the specimen thickness. 3
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Fig. 3. Experimental set-up for X-ray CT scanning.
displacement of specimens. The stress is calculated by the following equation, which yield the apparent strength based on homogeneous beam theory: σf = 3PL/2bd 2 . Where P, L, b and d are load, support span, width and depth of the specimen, respectively. The strain is calculated by the following equation: εf = 6Dd/ L2 . Where εf and D are strain in the outer surface and maximum deflection of the center of the specimen measured by DIC, respectively. The elastic moduli were determined by the Chord-modulus Method. The first part of the curves shows a linear elastic behavior. At a strain of approximately 1% and 1.35% for L and T specimens, respectively, the stress–strain curves start to deviate from the elastic line and the composite behaves as a quasi-elastic material until failure where the composite has a quasi-brittle behavior. The final failure of the first specimen (L) occurred when several layers in the tensile and compressive zones of the specimen were suddenly and simultaneously broken, while for the second specimen (T) the final failure is initiated by the sudden breaking of several layers in the compressive zone followed by the breaking of almost the same number of layers in the tensile zone. Although damage initiation in the specimen L is earlier than that in the specimen T, the strain at failure of the specimen L is much greater than that of the specimen T. Since the total number of
were non-destructively measured before testing using the impulse excitation technique (IET) according to ASTM E1876-09, and all specimens were scanned after testing with a micro-computed tomography scanner (EasyTom 130). Fig. 1 illustrates the orientation of specimens during the X‐ray scan. 4. Results and discussions 4.1. Quasi-static bending tests The mechanical behavior and damage of thick carbon/epoxy laminates subjected to static and dynamic bending loads were already investigated in our previous work [45]. In this study, which could be considered as a complement to the previous work, we were particularly interested by the fatigue damage evolution of these composite laminates as well as the influence of the loading type and ply orientation on their failure mechanisms. The experimental results of quasi-static bending tests are reported in Fig. 6. This figure shows the normalized stress–strain curves for two specimens that have different ply orientation. The stress and strain are calculated according to ASTM D790-03 using the load cell data and real
Fig. 4. Scheme of the applied displacement during fatigue testing. 4
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Fig. 5. Experimental set-up for fatigue testing.
details of the internal and external damage of specimens are observed. For the first specimen (L), fourteen layers were broken in tension and only five layers in compression. While for the second specimen T, seventeen layers were broken in compression and ten layers in tension. Thirteen of the fourteen layers broken in tension in the specimen L are oriented at 45°, 90° and 135°, and only one of the five layers broken in compression is oriented at 0°. All these layers were suddenly broken for a stress of 570 MPa and a strain of 2.3%. For the specimen T, the majority of the twenty-seven layers broken are oriented at 0°. The final failure of this specimen (T) is initiated by a sudden break of several compressed layers for a stress and strain of approximately 960 MPa and 1.9%, respectively. Moreover, the break of these layers was not produced exactly below the loading axis, but it has been produced in its sides. This is probably due to the fact that the large load applied perpendicularly to the middle of the specimen prevented its layers to break in this zone. It can also be seen in these figures that the predominant damage mechanism in the specimen L is the matrix cracking in off-axis plies, whereas for the specimen T it is the fiber breaking in on-axis plies. Since the ratio E11/E22 of the elementary layer is more than 15, the differences observed previously between the mechanical behavior and
layers as well as the number of 45° and 135° layers are the same in the two specimens, the difference observed in their mechanical behavior is mainly attributed to the large difference that exists in the number of 90° and 0° layers. The mechanical properties of L and T specimens are summarized in Table 2. The results presented in this table revealed a low variability in the mechanical properties of specimens that have the same ply orientation, which reflects the reliability of specimens manufacturing and cutting processes. However, for the two different ply orientations, the elastic moduli measured by non-destructive technique (IET) are approximately 5–7% greater than that measured by bending tests. Several previous studies have reported the existence of a slight discrepancy between the elastic modulus measured by destructive and non-destructive methods, and they attributed this discrepancy to the fundamental differences that exist in the theoretical bases and practical conditions between the two methods [45,47–49]. Furthermore, the elastic moduli calculated by Classical Laminate Theory (CLT) are approximately 25% and 30% greater than those measured by IET and bending tests, respectively. The CT cross-sectional views of the L and T specimens, in the (YZ) and (XY) planes, are shown in Figs. 7 and 8 respectively, from which
Fig. 6. Three-point bending stress–strain curves of two specimens that have different ply orientation. 5
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Table 2 Mechanical properties of specimens. Specimens
L T
Elastic modulus (GPa) IET
Bending tests
CLT
30.03 ± 0.46 56.18 ± 0.54
27.87 ± 0.80 53.45 ± 0.73
36.26 70
Maximum stress (MPa)
Strain at failure (%)
546.93 ± 17.63 956.34 ± 23.28
2.25 ± 0.06 1.96 ± 0.05
100% at the first cycle of each displacement level of 10 000 cycles, and it was calculated from the residual force measured by the load-cell of the fatigue machine by the same equation used in the static tests, according to ASTM D790-03. After 10 000 loading cycles at 35–65% δf, no damage was observed on the outside surface of the specimen and the drop in the residual strength does not exceed 3% for the first four levels (35–50% δf) and 6% for the three following levels (55–65% δf). Nevertheless, from the first loading cycles at 70% δf, the rate of the residual strength decrease has significantly increased compared to the first seven levels, and after 3200 loading cycles, a macro damage generated in the compressive zone led to a suddenly drop of about 32.8% in the residual strength of specimen. Finally, at 75% δf, a significant
damage mechanisms of L and T specimens are mainly due to the difference in the number of on-axis plies in these specimens. Indeed, the behavior of the specimen L is dominated by the matrix and interface behavior, while that of the specimen T is dominated by the fiber behavior. 4.2. Fatigue bending tests 4.2.1. Residual strength degradation Fig. 9 shows the L-specimen residual strength degradation as function of cycle number for nine applied displacement levels (from 35% to 75% δf). The residual strength of specimen was considered equal to
Fig. 7. CT cross-sectional views in the YZ and XY planes of the specimen L. 6
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Fig. 8. CT cross-sectional views in the YZ and XY planes of the specimen T.
Fig. 9. Normalized residual strength degradation of a specimen broken after nine loading levels.
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Fig. 10. Distribution of AE events recorded during the fatigue test of the specimen broken after nine loading levels.
Fig. 11. CT cross-sectional view in the YZ plane of the specimen after the first three loading levels.
specimen lost about 50% and 70% of its initial strength, respectively. Almost the same curves were also obtained with other specimens.
decrease in the residual strength was observed after each cycle, up to about 8500 cycles where a drop of more than 27% in the residual strength, occurred after a sudden break of several layers in the tensile zone, leading to the final failure of the specimen. The blue and red arrows in Fig. 9 show that, after 10 000 cycles at 70% δf and 75% δf the
4.2.2. Acoustic emission monitoring and damage mechanisms evolution Fig. 10 shows the distribution of AE events recorded during the 8
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layers and the delamination of a single interface were observed in the compressive zone. These observations validate the results obtained through the analysis of the acoustic activity recorded during tests, where the events attributed to the interface delamination were recorded from 50% δf, and those attributed to the fiber breaking were recorded from 55% δf. Fig. 14 shows the CT cross-sectional views of the broken specimen after nine loading levels in the different planes, where the nature, location and size of the main damage mechanisms previously identified by acoustic emission technique, are shown. It can be seen that the final failure of the laminate results from the combination of all these different damage mechanisms which intervene at different scales and can interact with each other. Furthermore, it can also be observed that the number of layers broken in tension is much higher than that of layers broken in compression. This difference is mainly due to the non-symmetrical distribution of compressive and tensile strains with respect to the horizontal median plane of the specimen during loading.
fatigue test of a specimen broken after nine loading levels according to their main features (amplitude, number of counts and duration). The choice of the applied filters and the attribution of AE events to the damage mechanisms are based on the results of a previous study on the same material [45] and some other works [50–53]. From the first displacement level (35% δf), which is well below the static damage threshold (50% δf), several events with amplitude of less than 60 dB were recorded. According to the results of the static study [36], the source mechanism of these events is matrix cracking. Up to 45% δf, the amplitude of almost all recorded events has not exceeded 60 dB. For all previous displacement levels, the number of events attributed to the matrix cracking increases with displacement level increasing. At 50% δf, some events with amplitude of 60–70 dB were recorded. These events are attributed to the interfaces delamination within the specimen tensile zone. From 55% δf the amplitude of some recorded events has exceeded 70 dB, which indicates the beginning of fiber failure. Finally, during the last four displacement levels (60, 65, 70 and 75% δf), the number of events attributed to the three damage mechanisms previously detected (matrix cracking, delamination, fibers breaking) increases rapidly with raising the displacement level until the final failure of the specimen.
4.2.4. Through-thickness strain field evolution To characterize the evolution of the longitudinal strain through the thickness of the laminate during different fatigue loading levels, the DIC technique was used and five virtual strain gauges were created across the thickness of the specimen. Fig. 15 shows the positioning of virtual strain gauges and the evolution of the longitudinal strain through the thickness of specimen at various loading levels. The curves intersection point is inside the compressive zone of the specimen and the neutral plane moves towards this zone with increasing loading level. For the first four loading level, the longitudinal strain increases almost linearly with increasing load and from the neutral plane toward the surface. However, from the fifth level (55% δf), the increase of the longitudinal strain becomes nonlinear. The appearance of some macro-damage on the surface of specimens are at the origin of this nonlinearity and the reason for the inability to monitor the longitudinal strain evolution of the last four levels.
4.2.3. X–ray computed tomography and damage mechanisms evolution Fig. 11 shows the CT cross-sectional view of the L-specimen in YZ plane after the first three loading levels (35, 40 and 45% δf). Despite the fact that the three loading levels applied are below the static damage threshold (50% δf) and no damage is observed on the outside surface of the specimen, some transverse matrix cracks in the off-axis plies of the laminate can be clearly observed from this figure. Therefore, this damage mechanism is the main responsible for the residual strength drop during the first three loading levels as it is also the main source of the AE events recorded. The CT cross-sectional views in the (YZ) and (XY) planes of the fatigued specimen after the first seven loading levels are shown in Fig. 12 and Fig. 13, respectively. Where matrix cracks in the off-axis plies and some delamination at the 135°/90°, 0°/90° and 135°/45° interfaces as well as some fibers break in the 90° layers are observed. One can also noticed that the damage in the tensile zone of the specimen is greater than that in the compressive zone. In the tensile zone, matrix cracking of thirteen layers, delamination of two interfaces and some fiber breaking were observed, whereas the matrix cracking of only three
4.3. Loading type influence Fig. 16 correspond to CT cross-sectional views in the YZ plane of broken specimens in static and cyclic loading. The figure shows the most damaged area of the specimens, which is in the vicinity of the loading axis. In the static loading, the majority of damage occurred in a
Fig. 12. CT cross-sectional view in the YZ plane of the specimen after the first seven loading levels. 9
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Fig. 13. CT cross-sectional view in the XY plane of the specimen after the first seven loading levels.
Fig. 14. CT cross-sectional views in the YZ and XY planes of the broken specimen after nine loading levels.
specimen thickness, which is reflected by the linearity of the longitudinal strain variation.
plane perpendicular to the y-axis and at 5 mm to the left of the loading axis, whereas in the cyclic loading, the damages occurred in the compressive zone and in the tensile zone are not on the same plane. The number of broken layers is the same for both types of loading, but the predominant damage mechanism in static loading is matrix cracking in off-axis plies, while in cyclic loading delamination at the 0°/90°, 90°/ 135° and 45°/135° interfaces is the most predominant one. Fig. 17 shows the positioning of virtual strain gauges and the variation of the longitudinal strain through the specimen thickness at 55%δf in static and cyclic loading. For cyclic loading, the strain is represented after 10 000 cycles. The figure shows that in cyclic loading, the neutral plane of the specimen is not exactly in the middle of the specimen thickness as in static loading. Furthermore, in the static loading, the longitudinal strain increases non-linearly from the neutral plane toward the surface, while in cyclic loading, the longitudinal strain increases almost linearly. The preceding observations can be explained by the following hypothesis: After several loading cycles at the same displacement amplitude, the evolution of the different damage mechanisms in the laminate leads to an attenuation of the stress concentration and a more uniform distribution of stresses through the
5. Discussion and conclusion The main objective of this study is to provide a thorough experimental investigation of fatigue damage mechanisms and evolution in thick carbon/epoxy laminate subjected to bending load. To achieve this goal, three non-destructive evaluation techniques were used. Furthermore, the evaluation of the advantages and limitations of each used technique, and the influence of the loading type and ply orientation on the laminate failure process were the second objective of the present study. - After 10 000 loading cycles at 35–45% δf, that are below the static damage threshold (50% δf), no damage was observed on the outside surface of the specimen, although a drop in the residual strength that does not exceed 3% was observed and several events with amplitude less than 60 dB were recorded. The CT cross-sectional views of the specimen showed that some transverse matrix cracks in 10
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Fig. 15. Evolution of the longitudinal strain through the thickness of specimen at various loading levels.
laminate is matrix cracking in off-axis plies, whereas for the second laminate fiber breaking in on-axis plies is the most predominant one. For the loading type influence, a few damages were detected from the first cyclic loading level (35% δf) which was well below the static damage threshold (50% δf). The number of broken layers is the same for both types of loading, but the nature and location of damage differ substantially between the static and cyclic loading. So that, the predominant damage mechanism in static loading is matrix cracking in off-axis plies, while in cyclic loading delamination at the 0°/90°, 90°/135° and 45°/135° interfaces is the most predominant one. - The use of X-ray tomography in this study allowed visualization of all damage present in the studied laminates, which made it possible to identify, quantify and locate them precisely and therefore, to identify the physical origin of residual strength decrease and acoustic emissions. In general, this powerful technique allowed us to provide a complete and precise characterization of damage of the studied laminates. - The analysis of acoustic emission activity recorded during the tests provided many information about the evolution, identification and localization of the damage. However, the determination of the depth, size and distribution of internal damage was not possible with this technique. - The DIC technique used in this study made it possible the measurement of the superficial displacement field and the calculation of the strain field. These data allowed us to make a good characterization of the mechanical behavior of material and to monitor the
the off-axis plies of the laminate are the main responsible for the residual strength drop and the main source of the AE events recorded. For the four following levels (45–65% δf), some delamination and a drop of less than 6% in the residual strength were observed. Moreover, the amplitude of some recorded events has exceeded 70 dB. The CT cross-sectional views of the specimen showed that several matrix cracks in the off-axis plies and some delamination at the 135°/90°, 0°/90° and 135°/45° interfaces as well as some fibers break in the 90° layers are the physical origin of the previous observations. Finally, for the last two loading levels (70 and 75% δf), a significant decrease in the residual strength was observed after each cycle, and a sudden rupture of several layers in the compressive zone and then in the tensile zone of the specimen, which leads to a drop of more than 27% in the residual strength, were seen with the naked eye. In parallel, a large number of events attributed to the main three damage mechanisms (matrix cracking, delamination, fibers breaking) was recorded. The CT cross-sectional views of the specimen showed the nature, location and size of these different damage mechanisms previously identified by AE technique. - The ply orientation and loading type had a great influence on the mechanical and damage behavior of composite laminates. Indeed, the behavior of the laminate consisting of a large number of off-axis plies was governed by the matrix and interface behavior, while that of the laminate consisting of a large number of on-axis plies it was governed by the fiber behavior. Moreover, the CT cross-sectional views showed that the predominant damage mechanism in the first
Fig. 16. CT cross-sectional views in the YZ plane of broken specimens in static and cyclic loading. 11
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Fig. 17. Variation of the longitudinal strain through the specimen thickness at 55% δf in static and cyclic loading.
strain field evolution. [13]
Finally, we believe that this work contributes to a better understanding of the fatigue damage evolution in thick carbon/epoxy laminates subjected to bending load and a better evaluation of the advantages and limitations of each non-destructive technique used. The main perspectives of this work is to provide an accurate characterization and description of the mechanical and damage behavior of these laminates under different loading conditions, and to evaluate the effectiveness of other non-destructive techniques to characterize and control these materials during their lifetime. These perspectives are aimed at constructing models that predict, as reliably as possible, the response of thick laminated composite structures to different external loadings and the establishment of specific standards and criteria for NDE of thick composite structures.
[14] [15] [16] [17] [18]
[19] [20] [21]
Appendix A. Supplementary data [22]
Supplementary data to this article can be found online at https:// doi.org/10.1016/j.compscitech.2019.107815
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