Impact of process induced residual stresses on interlaminar fracture toughness in carbon epoxy composites

Journal Pre-proofs Impact of process induced residual stresses on interlaminar fracture toughness in carbon epoxy composites M.A. Umarfarooq, P.S. Shivakumar Gouda, G.B. Veeresh kumar, N.R. Banapurmath, Abhilash Edacherian PII: DOI: Reference:

S1359-835X(19)30401-4 https://doi.org/10.1016/j.compositesa.2019.105652 JCOMA 105652

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Composites: Part A

Received Date: Revised Date: Accepted Date:

12 June 2019 10 August 2019 1 October 2019

Please cite this article as: Umarfarooq, M.A., Gouda, P.S.S., kumar, G.B.V., Banapurmath, N.R., Edacherian, A., Impact of process induced residual stresses on interlaminar fracture toughness in carbon epoxy composites, Composites: Part A (2019), doi: https://doi.org/10.1016/j.compositesa.2019.105652

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IMPACT OF PROCESS INDUCED RESIDUAL STRESSES ON INTERLAMINAR FRACTURE TOUGHNESS IN CARBON EPOXY COMPOSITES Umarfarooq M A1, P S Shivakumar Gouda1* , Veeresh kumar G B2, N R Banapurmath3 Abhilash Edacherian4 1Department

of Mechanical Engineering, SDM College of Engineering & Technology, Dharwad, India. of Mechanical Engineering, National Institute of Technology Andhra Pradesh, Tadepalligudem, Andhra Pradesh, India 3Department of Mechanical Engineering, B.V.B. College of Engineering and Technology, KLE Technological University, Hubballi, India 4Department of Mechanical Engineering, College of Engineering, King Khalid University, Kingdom of Saudi Arabia 2Department

*Corresponding author: [email protected]

Abstract Residual stresses are induced in composite laminates during the process of manufacturing due to thermal mismatch between composite constituents during curing. Processes induced residual stresses effects the safety of structural components and can play a main role in matrix cracking, crack initiation and propagation. The objective of this research is to study the influence induced residual stresses on interlaminar fracture toughness (GI) of Carbon/Epoxy laminates. Initially laminates were cured under room temperature followed by post curing at various temperatures using a chosen cure cycle. Slitting method was employed to determine the residual stress distribution in post cured laminates and its effects on GI was estimated using Double Cantilever beam (DCB) test. The results show a gradual increase in GI with increase in compressive residual stresses in composite laminate. Further, the fracture surfaces of laminates were carefully studied using scanning electron microscope to know the interfacial adhesion of matrix and fiber. 1. Introduction Composite materials are multiphase materials, one of the phase being matrix which is continuous and surrounds the other phase, reinforcements. High strength to weight ratio, high stiffness, low density, ease of manufacturing and tailor made properties have propelled the application of the composites ranging from aerospace, marine, automotive industries to the domestic appliances. Composite are heterogeneous and anisotropic by nature. One of the issue associated with the manufacturing of the polymer composites is the inheritance of the residual stresses induced during curing process. Residual stresses are induced in the multiphase composite material due to the difference in the thermal coefficient of expansion of its phases and also due to the chemical shrinkage of the matrix phase Residual stresses may combine with service stresses and can play a 1

major role in structural failure of composites. Residual stresses are found to affect the dimensional stability, performance of the composite and may result in warpage, matrix cracking, interface debonding, stiffness and strength reduction of composite laminas [1-6]. As the residual stresses impacts the performance of the composite, the magnitude of these stresses and its effects on the mechanical and fracture properties of the composite need to be investigated. Earlier works have revealed that the residual stresses have a substantial effect on mechanical as well as on fracture properties of composites. The state of residual stresses in the composite can be varied either by applying different post curing curves [7-9] or by using different cooling rates [10-12]. Nairn [13] suggested that the toughness measured in DCB specimen ignoring the residual stresses is the apparent rather than actual toughness. Corrected Beam theory was used to predict the toughness by considering residual stresses. The error in the critical energy release rate of DCB specimen on a various graphite/epoxy laminates by neglecting residual stresses varied in the range of -54.85 % to + 76.36 %. Guo et al [14] studied the effects of compressive and tensile residual stresses on interfacial strain energy release rate in a triple layered specimen using analytical method and finite element analysis where in residual stress induced energy release rate is always higher for adhesive with compressive residual stress than one with tensile residual stress. Sicot et al [10] determined the induced residual stresses in post cured Carbon/Epoxy composite using Hole drilling method and studied its influence on mechanical behavior, damage initiation and its subsequent development using acoustic emission technique. Different cure cycles were employed and found to have negligible effect on the longitudinal, transverse modulus and the in-plane Poisson coefficient but higher residual stresses were found, which are responsible to accelerate damage initiation and growth. Laik et al [15] studied the influence of residual stresses on the inplane shear strength, damage evolution and the mode I delamination fracture toughness of Carbon/epoxy composite prepared by vacuum-assisted resin transfer molding where in residual stresses induced in composites for different cure cycles were obtained using finite element methods. Cure induced residual stresses were also found to have a positive effect on strain energy release rate whereas a detrimental effect on in-plane shear strength. Gillespie et al [16] developed a numerical model to study the effect of various parameters such as thermal conductivity, plate thickness, thermal conductivity, cooling rate, and orientation of plies on development of residual stresses for a thermoplastic composite and attributed a 35 % decline in critical strain energy release rate with higher residual stress state. 2

Few researchers have shown influence of residual stresses on the interlaminar fracture toughness based on analytical and finite element methods [13-16]. Hence, the present study aims at finding the influence of residual stresses on interlaminar fracture toughness of the Carbon fiber reinforced polymer composite through experimental methods. 2. Experimental Details 2.1 Materials Unidirectional (UD) Carbon fibers of 0.35 mm thickness was used as reinforcement. Epoxy (Araldite LY-556) and curing catalyst (Aradur 917) applied in volumetric ratio of 9:1 has been used as matrix. 2.2 Preparation of Carbon/Epoxy Composite laminates UD Carbon/Epoxy composite laminate [0]8 were fabricated by hand lamination technique which were initially cured under room temperature. DCB Laminates for Mode I interlaminar test were manufactured by inserting a Teflon sheet of 14 µm to generate pre crack in-between the 4th and 5th plies. Specimens for tensile and fracture toughness testing have been cut as per ASTM D3039 [17] and ASTM D 5528[18] standards respectively and laminates were post cured using different temperature cycles. 2.3 Post Curing Cycles: For the present investigation the state of residual stress were changed using different post curing regime. Post curing were carried using two different designs. Three sets of laminates were post cured at constants temperatures of 90 °C, 135 °C and 180 °C for 6 hours followed by direct cooling under room temperature with cooling rate of 20°C /min and two set of composites were post cured as per cure cycle [9] shown in Fig. 1. The end cooling conditions of post curing cycle are Condition A: The cooling of sample is carried out at room temperature with cooling rate of 20°C /min resulting in normal cooling. Condition B: Cooling of sample is achieved by keeping the sample in oven resulting in slow cooling with cooling rate of 0.5°C/min..

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Fig. 1. Cure cycle.

The composite laminates has been coded with respect to post curing cycles and corresponding post curing condition are listed in Table 1. Table 1: Laminate Code and post curing condition. Laminate Code

Post curing condition

CE-90-6

Post cured at 90 °C for 6 hours and cooled under room temperature with cooling rate of 20°C /min

CE-135-6

Post cured at 135 °C for 6 hours and cooled under room temperature with cooling rate of 20°C /min

CE-180-6

Post cured at 180 °C for 6 hours and cooled under room temperature with cooling rate of 20°C /min

CE-AC

Post cured as per Condition A of Fig. 1 (Normal Cooling)

CE-OC

Post cured as per Condition B of Fig. 2. (Slow Cooling)

3 Mechanical Tests 3. 1 Fracture test Mode I interlaminar fracture test was carried out using calibrated machine of capacity 10 kN with an crosshead speed of 2 mm/ min. DCB specimen used for fracture toughness test with initial pre crack of 50 mm from hinge position is shown in Fig. 2. Crack growth was recorded using a 50x magnification digital video camcorder. Modified beam theory was applied to determine the critical strain energy release rate in accordance with ASTM D 5528. Extension of crack (δ) and respective loads were recorded in the increments of 1 mm for initial 5 mm of crack growth and then readings (δ and P) were monitored for every 5 mm crack extension until the crack reached to 100 mm length. Equation (1) was used to determine the Mode I fracture toughness for initiation and propagation of crack. 3P GI 

2b(a   )

(1)

where P: the load, δ: the load point displacement, b: the specimen width, h: the specimen thickness, a: the delamination length, and Δ: correction factor to account for rotation of the DCB arms. Δ was obtained experimentally by plotting least square plot of the cube root of compliance (C1/3) with respect to crack length. 4

Fig. 2. Geometry of DCB Specimen

3.2 Tensile Test To determine the elastic constants; longitudinal elastic modulus (E11) and transverse elastic modulus (E22) a tensile test for the composite laminates was carried as per guidelines of ASTM D 3039. Testing was accomplished using a calibrated machines with 150 kN capacity with cross head speed of 3 mm/min. Dimensions of the specimen used for tensile testing are shown in Fig.3.

Fig. 3. Dimensions of specimens used for tensile testing to obtain elastic constants.

Elastic constants (E11 and E22) of all the post cured laminates have been determined as per ASTM 3039 and have been plotted in Fig.4 and Fig.5. Elastic constants for the room temperature (RT) cured composite laminates are given in Table 2. Table 2. Elastic constants of RT cured composite E11 (GPa) E22 (GPa) 55.16 2.60

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Fig.4. Longitudinal Elastic moduli E11

Fig.5. Transverse Elastic moduli E22

From the Fig.4 ,it can be observed that E11 has decined gradually with increase in post curing temperature. The composites post cured using the chosen cure cycle have also exhibited an decrease in E11 and slow cured laminates exhibit higer E11 compared to normal cooled laminate. Post curing has limited effect on E11 and the variations were within 15 % compared to RT cured composite.Transverse modulus (E22) also declined with post curing of the composite laminate as shown in Fig. 5. All post cured laminates depicted a decline in E22 and the CE-180-6 exhibited 26 % decline with respect to RT cured composite.

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5. Residual stress measurement using slitting method Residual stress measurement techniques have been grouped as destructive and non-destructive techniques. Hole drilling method and slitting method [19-30] are the most commonly used destructive methods of residual stresses measurement techniques. Basic principle of slitting method can be explained with typical slitting specimen shown in Fig. 6. Slitting method involves machining a thin slit on the top face of specimen in incremental cuts (X-Direction) and the strain gauge bonded on the other face exactly opposite to the slit measures the Y-direction relaxed strains from each cut. Residual stresses can be obtained using an appropriate computational method from the obtained residual strains.

Fig. 6. Typical Slitting method specimen with coordinate geometry

Relationship between the residual stresses and the measured deformations is of integral form and is given by Equation (2) ai

 yy ai   C x, ai  yy x 



(2)

0

Where, εyy (ai) is the strain measured in ‘y’ direction when the slit depth is ai. The kernel function C(x, ai) defines the strain sensitivity to a stress at depth ‘x’ within machined material of depth ai. In order to obtain the solution for Equation (2), it is required to assume an initial distribution of the residual stresses. The two most commonly used methods to obtain the stresses are “Pulse Method” and “Series Expansion method”. Pulse method have been applied in this work to approximate the residual stresses. In pulse method stresses are assumed to be uniform over each incremental cut and the unknown residual stress variation is approximated by a series of strip loads as shown in Fig.7. Therefore unknown residual stresses are measured using Equation (3) as  x j  

n

 U x j

j

(3)

j 1

Where σj is stress in the jth incremental cut and ‘n’ implies total number of incremental cuts.

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(a)

(b)

Fig. 7. (a) An unknown residual stress distribution on slit faces; (b) Unknown residual stress are approximated using series of uniform strip loads.

The pulse functions [Uj(x)] can be defined as follows:

Uj(x) =

{ 01

aj ― 1 ≤ x ≤ aj x < aj ― 1,x > aj

(4)

Considering a linear elastic deformation, the strain corresponding to the residual stress acting on the faces of the slit given by Equation (5)

C   

(5)

Each element of compliance matrix, Cij relates to the strains measured by the strain gauge for a slit of depth ai when normal stress of unit load is applied in the domain a j 1  x  a j . Cij   (a  ai , ( x)  U j ( x))

(6)

The method of measuring residual stresses using slitting method shown in flowchart Fig. 8. CNC milling used for slitting of composite laminate is shown in Fig.9 (a). Slitting technique setup along with strain recorder is shown in Fig. 9 (b). Fig. 9 (c-d) shows relative position of the strain gauge bonded on composite laminate and slitting cutter. The specifications of the specimen used for residual stress measurement are given in Table 3. Specimen face used for strain gauge bonding was thoroughly washed with acetone and the strain gauge TML UBFLA-03 was carefully bonded on specimen as shown in Fig.9 (c). Machining of slit was carried out on a specimen with one end clamped and the other end free using a CNC machine. Slit was made using a circular saw cutter of diameter 20 mm and thickness 0.2 mm with spindle speed of 5000 rpm and feed of 1 mm/s. The slits were performed in incremental cuts, depth of increment being 0.35 mm to depth of 2.45 mm. Strains released from each cut are recorded by the strain recorder and readings are noted once they are stabilized. The strains released after slitting of each laminate are listed in Table 4. Once the relaxed strains are obtained the next step is to determine the residual stresses using equation (7).

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Table 3. Specifications of the specimens used for slitting Length (mm) Width (mm) Thickness (mm) 20 2.8 100

Step 2 Step 1 Machine the slit on specimen in incremental cuts

Relaxed strains after slitting are measured using a strain gauge

Calculating compliance coefficients by simulating Slitting method using FEM Deriving relationship between {𝜎} and {𝜀} in terms of compliance coefficient matrix Obtaining residual stress distribution using computational method Fig. 8. Slitting method flowchart

Fig. 9. Images showing (a) & (b) Slitting method experimental setup; (c) & (d) Relative position of strain gauge and cutter Table 4: Recorded strains (µε) from slitting of all composite laminates 9

Lamina No.

Slit Depth (mm)

1 2 3 4 5 6 7

0.35 0.70 1.05 1.40 1.75 2.10 2.40

Laminate Code CE-135-6 CE-180-6 (µε) (µε) 10 -16 32 40 46 121 99 308 226 714 634 1001 1094 1107

CE-90-6 (µε) -21 -9 8 47 190 1043 1111

CE-AC (µε) 10 21 38 254 483 630 1260

CE-OC (µε) -3 -16 -28 -11 26 690 746

5.1 Determination of Compliance coefficients Compliance coefficients are required to calculate the residual stresses from the relaxed strains as per equation (7). Compliance coefficients for each laminate are obtained using ANSYS APDL by using the experimentally measured elastic constants. The specimen was meshed using a Solid185 element as shown in Fig. 11, the mesh is refined around the slit to capture the minute variation of the strains. FEM simulation is replicated similar to experimental slitting of specimen with one end constrained completely and an incremental cut is introduced at the slit location. Slitting method was simulated step by step by removing the elements of slit area and unit load was applied across the boundary of the slit. Then the obtained strains at the gauge location are averaged to obtain each element of compliance matrix. These Compliance coefficients will be used to obtain the residual stresses induced in each lamina due to post curing using Equation (7). The physical meaning of compliance matrix elements is shown in Fig. 10. Residual stresses listed in Table 5 are obtained by applying pulse method.

Fig. 10. Compliance matrix coefficients [Cij] physical meaning for the Slitting method. Table 5: Residual stress (MPa) distribution in different layers determined from relaxed strains of composite laminates Lamina No.

Slit Depth (mm)

1 2 3 4 5 6 7

0.35 0.70 1.05 1.40 1.75 2.10 2.45

CE-90-6 (MPa) 4.04 -6.34 0.36 -1.16 -4.93 -5.20 36.51

Laminate Code CE-135-6 CE-180-6 (MPa) (MPa) -3.35 3.99 0.34 -15.04 2.53 1.18 -3.11 -4.98 -6.64 -8.95 11.25 50.96 41.91 -2.24

CE-AC (MPa) -1.84 0.33 -0.16 -15.73 -18.64 78.41 -35.50

CE-OC (MPa) 0.75 1.55 -1.11 -3.76 -7.75 -1.73 31.64

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Fig. 11. Finite element model for compliance co-efficient determination.

In all the composite laminates the delamination is initiated between the 4th and 5th laminas, the average residual stresses induced in these layers is listed in Table 6. Table 6: Average Residual stress obtained from residual stresses induced in 4th and 5th lamina of composite laminate Laminate Code CE-90-6

Average Residual Stress (MPa) -3.05

CE-135-6

-4.88

CE-180-6

-6.97

CE-AC

-17.16

CE-OC

-5.78

6. Mode I Test results Experimental test setup for mode I test is shown in Fig. 12. Load versus displacement curves for all the composite laminates which are obtained from mode I test using DCB specimens are shown in Fig. 12. From the Fig. 13 it was observed that CE-AC exhibit higher load for crack initiation compared to other composite laminates. The load for crack initiation has increased with increase in post curing temperature. Interlaminar fracture toughness for crack initiation (GIC) and crack propagation (GIP) for all the composite laminates are summarized in Fig. 14. To study the effect of residual stresses induced during curing on interlaminar fracture toughness, the average residual stresses across the laminas constituting the mid plane where initial crack was introduced as delamination initiator for all the composites was considered and is listed in Table 6.

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Fig.12. Actual experimental setup for DCB specimen

Fig.13. Load versus displacement curves for all composite laminates

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Fig. 14. Mode I interlaminar fracture toughness GIC values (Left) and GIP values (Right) of all composite laminates

From Table 6 and Fig. 14, it can be observed that presence of the compressive residual stresses have positive effect on interlaminar fracture toughness. It can be seen that the energy release rate (GIC and GIP) increases with an increase in compressive residual stresses. GIC for composite laminate CE-90-6 was found to be 200.16 J/m2 with the induced residual stresses due to post curing has a magnitude of -3.05 MPa. An increase in post curing temperature to 135 °C and

180 °C an

increased stress value to -4.88 MPa and -6.97 MPa respectively which has resulted an enhancement in GIC by 9 % and 23.80 % respectively as compared to CE-90-6. Composites post cured using chosen cure cycle has also displayed an increase in GIC in comparison to CE-90-6. Residual stresses induced in CE-AC and CE-OC are of magnitude -17.16 MPa and -5.78 MPa respectively which has resulted an increase in GIC by 41.13 % and 16.07 % with respect to CE-90-6 respectively.

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From the Fig. 14, it can be observed that GIP of CE-90-6 is 245.48 J/m2 and the induced residual stresses due to post curing being -3.05 MPa as per Table 6. CE-135-6 and CE-180-6 exhibited an increase in GIP by 15.82 % and 32.29 % respectively as compared to CE-90-6 while the residual stresses induced have increased to -4.88 MPa and -6.97 MPa respectively. CE-AC and CE-OC has exhibited a rise in GIP by 83.32 % and 20.58 % respectively with respect to CE-90-6. Residual stresses can contribute to the energy release rate either by doing external work or by releasing strain energy as the crack propagates. Compressive residual stresses contribute to the energy release rate by doing external work [13]. As the level of compressive residual stresses increased across the laminas containing the initial crack, a steady rise in interlaminar fracture toughness during initiation and propagation was observed. 7. Scanning electron microscopic (SEM) analysis To get an information of morphologies on effect of post curing on failure mechanism of the composites under the Mode I fracture, surfaces of all post cured composites were examined using SEM and images are shown in Fig. 15. SEM microphotographs were taken from the regions closer to crack tip and away from it. JEOL machine operated at 20 keV using Microscope Control (SED) software was used to examine the delaminated fracture surfaces near and away from the crack initiation region of post cured composites laminates. The laminates were carefully cut with size of 10 mm x 10 mm and coated with gold using sputter coater machine before the SEM analysis to prevent built up surface charge.

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Fig. 15. SEM micrographs for mode-I fracture surface near crack initiation region in composites of

(a) CE-90-6

(b) CE-135-6 (c) CE-180 -6 (d) CE-AC (e) CE-OC. SEM micrographs for mode-I fracture surface away from the starter crack region in composite specimen of (f) CE-90-6 (g) CE-135-6 (h) CE-180 -6 (i) CE-AC (j) CE-OC

Based on Fig. 15, the differences between the fracture surfaces of post cured laminates at different temperature can be seen with respect to resin ductility and interfacial bonding strength are well agreement with the earlier findings on mode I delamination resistance by Bradley [31]. From the Fig. 15 (a, f), bare fibers are loosely embedded in matrix are clearly visible. Bare fibers with little or no matrix adhesion can be seen which indicates weak interfacial bonding between fibers and matrix. River line markings and fractured fiber can be seen from the Fig. 15 (b, g). Usually the inception of river line markings occurs across the fiber matrix interface indication of adhesive fracture [31, 32]. The broken fibers in Fig.15 (g) may be attributed to the fiber bridging effect during fracture. Fig. 15 (c, h) shows a starting crack tip and hackle marks can be seen across the matrix and the de-bonded fibers. The hackle marks implies an extensive matrix plastic deformation indicating an improved matrix/fiber bonding and an effective stress transfer between matrix and reinforced fibers. It can be observed from Fig. 15 (a-c & f-h) as post curing temperature increased the fracture failure became more ductile, a feature which can be further supported from the deformation witnessed in the matrix itself. The increase in the ductility resulted an increase in interfacial strength which leads to improved delamination resistance. Fig. 15 (d, i) and Fig. 15 (e, j) shows morphologies of CE-AC and CE-OC in which two different behaviors were observed for two end cooling condition. Composite cooled at faster rate shows regions of resin covering the fiber surface indicating a strong interfacial bond and the other shows almost a fiber surface with little/ no matrix adhesions to fiber. The strong interfacial bond in CE-AC composite is further appreciated by higher delamination resistance compared to all other post cured composites. 8. Conclusions Carbon epoxy composite laminates were prepared by hand lamination technique and were initially cured under room temperature. The state of residual stresses in composite was varied by post curing at different temperatures using specified curing cycle. Induced residual stresses were successfully measured experimentally by the aid of incremental slitting method using CNC milling machine integrated with strain gauge recorder. As the post curing temperature increased, the residual stresses are gradually increased in CE-AC composite and eventually exhibited higher residual stresses as compared to other post cured laminates. An increase in compressive residual 16

stresses from -3.25 MPa to -17.16 MPa leads to an improved fracture toughness GIC and GIP values by 41.13 % and 83.32 % respectively.. Based on these investigations it was concluded that the presence of residual stresses in composites have a positive effect on the fracture toughness under mode I loading.

ACKNOWLEDGMENTS Authors thankfully acknowledge the funding and support provided by Deanship of Scientific Research, King Khalid University, Abha-Asir, Kingdom of Saudi Arabia, with grant number R.G.P.2/6/38 under research group-Materials and Production, to complete the research work. The authors of this paper appraise their appreciation to the organization of SDM College of Engineering and Technology, Dharwad for their motivation and backing throughout the investigation and the author would like to acknowledge Directorate of Minorities, Government of Karnataka (GOKDOM), Bengaluru.

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