Reinforcement of CFRP joints with fibre metal laminates and additional adhesive layers

Accepted Manuscript Reinforcement of CFRP joints with fibre metal laminates and additional adhesive layers D.G. dos Santos, R.J.C. Carbas, E.A.S. Marques, L.F.M. da Silva PII:

S1359-8368(18)32224-8

DOI:

https://doi.org/10.1016/j.compositesb.2019.01.096

Reference:

JCOMB 6578

To appear in:

Composites Part B

Received Date: 17 July 2018 Revised Date:

28 November 2018

Accepted Date: 25 January 2019

Please cite this article as: Santos DGd, Carbas RJC, Marques EAS, da Silva LFM, Reinforcement of CFRP joints with fibre metal laminates and additional adhesive layers, Composites Part B (2019), doi: https://doi.org/10.1016/j.compositesb.2019.01.096. This is a PDF file of an unedited manuscript that has been accepted for publication. As a service to our customers we are providing this early version of the manuscript. The manuscript will undergo copyediting, typesetting, and review of the resulting proof before it is published in its final form. Please note that during the production process errors may be discovered which could affect the content, and all legal disclaimers that apply to the journal pertain.

ACCEPTED MANUSCRIPT Reinforcement of CFRP joints with fibre metal laminates and additional adhesive layers D.G. dos Santos1, R.J.C. Carbas1,2, E.A.S. Marques2, L.F.M. da Silva1 Department of Mechanical Engineering, Faculty of Engineering, University of Porto, Portugal 2

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Institute of Science and Innovation in Mechanical and Industrial Engineering (INEGI),

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Faculty of Engineering, University of Porto, Portugal

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Abstract

The use of fibre metal laminates (FMLs) offers significant improvements over more traditional materials applied in aircraft structures, such as metallic alloys. The use of this type of materials offers weight reduction, improved damage tolerance

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characteristics and enhanced safety due to a synergetic combination of the advantageous properties of both composites and metallic alloys. The main objective of this work was to study the effect of reinforcing a basic carbon

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fibre reinforced polymer (CFRP) joint with titanium laminates and/or with additional

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adhesive layers in the interfaces between titanium and composite, following different lay-up configurations. The lay-up configuration that led to the best results in terms of failure mode and failure load, was found to be the configuration using titanium laminates on a basic CFRP substrate and additional layers of film adhesive in between the metal laminate and the CFRP. A numerical model using finite element analysis with cohesive zone elements was also developed, with the aim of studying the performance of the different proposed configurations, correlate the models with the experimental

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ACCEPTED MANUSCRIPT results, and to aid the identification of the optimal material to use for reinforcement of the CFRP adherend.

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Keywords: Fibre metal laminates (FML); mechanical; numerical; adhesive bonding; hybrid; CFRP laminates

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1. Introduction

Composite materials and, more specifically, fibre reinforced plastics (FRP) have gained

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importance for use in structural applications during the last few decades. These materials were first applied in military applications in the aircraft industry after World War II and their usage has now expanded to encompass the commercial aircraft and aerospace industries. The use of innovative materials in these industries is mainly driven

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by the need to achieve lighter aerostructures, reducing both the fuel consumption and the environmental impact of this sector [1, 2]. Composite materials are therefore replacing more traditional structural materials, such as steel or aluminium, because of

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their strength/stiffness weight ratio and the flexibility to manufacture complex

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aerodynamic structures. Additionally, they provide excellent fatigue properties and corrosion resistance [3, 4]. However, aircraft manufacturers must balance weight reduction against structural integrity, price, and durability and, despite the increase in composite materials usage, metallic materials still continue to play a key role. Examples of the growing importance of composite materials in commercial aircraft manufacture are the Airbus A350 XWB, with 53% of composite material in weight and the Boeing 787 Dreamliner, where composite materials make up around 50% of the aircraft structural weight, enhancing fuel economy, reducing maintenance times, and operating costs [5, 6]. 2

ACCEPTED MANUSCRIPT Adhesive bonding technology has been developed side by side with composite materials, as it is key factor for enabling large scale use of composites. Composite materials exhibit a major decrease in their mechanical properties when holes are needed for joining using rivets or bolts due to the low bearing and shear strengths and the

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higher notch sensitivity when compared to metals [7]. In contrast, adhesive bonding is an ideal technology to avoid these inherently damaging joining techniques, allowing higher stiffness, more uniform stress distribution while still presenting an extremely low

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added weight. In the case of composite bonded substrates, there is an additional concern - the peel stresses, which reach peak values at the ends of the overlap of a single lap

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joint and can cause interlaminar failure by delamination near the points of the singularities due to the low transverse tensile strength in the through-thickness direction [8 - 15].

In the search for an optimal design applicable to aviation structures, studies have been

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made during the last decades in order to replace the aluminium alloys widely used in the past, replacing it with a material that combines high strength, low density, high elasticity modulus, improved toughness, corrosion resistance and fatigue properties.

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Fibre reinforced plastics therefore appear as an ideal material for this purpose with

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exception on their fracture toughness, which is quite low. This fact has stirred the development of innovative hybrid composite materials that could overcame most of the disadvantages of using aluminium alloys or fibre reinforced plastics alone [16 - 19]. In 1978, at the Faculty of Aerospace Engineering at Delft University of Technology, while studying methods to increase fatigue performance in aluminium alloys, an improvement in fatigue properties was observed by introducing a high strength aramid fibre into the adhesive layers of a laminate sheet. These studies led to development of ARALL (Aramid Reinforced Aluminium Laminates), the first fibre metal laminate 3

ACCEPTED MANUSCRIPT (FML), that consists of alternating thin aluminium alloy layers with uniaxial or biaxial aramid fibre prepreg. It was patented in 1984 and later industrially manufactured by the Alcoa Company [20]. Nowadays, there are several standard configurations of ARALL materials, each one employing a different aluminium alloy in the laminate. In FMLs, the

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resistance to crack growth is enhanced by the occurrence of fatigue cracks bridging. When a crack occurs in the aluminium layers, some limited delamination is observed at the metal-fibres interface and the stresses are redistributed from the metal to the

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unbroken fibres. This phenomenon of crack bridging, supported by the strength of the fibres present between the aluminium layers, limits crack opening and reduces crack

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growth in the metal layers [21, 22]. With later developments of this technology, other types of laminate materials were adopted, such as CARALL, that differs from ARALL in the utilization of carbon fibres instead of aramid fibres, resulting in a much stiffer material with low crack growth rates, and BARALL, that uses basalt fibres as

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reinforcement. Regarding CARALL, its performance is still limited by some unsolved problems associated with the direct contact between carbon fibres and aluminium. This pair of materials is highly susceptible to galvanic corrosion due to highly dissimilar

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electrochemical potentials. A related, but distinct issue is caused by the thermal expansion coefficients of aluminium and carbon, as these differences induce high

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thermal residual stresses on the final component [23, 24]. Due to these problems and additional limitations regarding operation at temperature and damage tolerance of aluminium based FML, efforts have been made to develop titanium based FML, also known as Hybrid Titanium Composite Laminates (HTCL), with existing studies indicating that such laminates can provide a stronger, stiffer and more damage-tolerant alternative for high temperature use [25, 26].

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ACCEPTED MANUSCRIPT In summary, FMLs combine the good properties of either metal and fibre-reinforced composites, presenting better mechanical properties when compared to the conventional lamina, using only fibre-reinforced lamina, or to the monolithic aluminium alloys. The long processing cycle, necessary to cure the matrix of the composite plies, appears as a

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major disadvantage of this material, increasing manufacturing and overall costs of FMLs [27].

The aim of this study is to increase the peel strength of composite materials and increase

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the joint strength of composite adhesive joints by using different lay-ups of the

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adherends of an adhesively bonded joint based on reinforcements made using either concept similar to the FML concept or a combination of FML and additional adhesive layers. Several material configurations were suggested for a CFRP adherend reinforced with titanium plies and adhesive layers and comparisons are made regarding the joint performance, allowing the identification of the best performing configurations with

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regards to failure load and failure mode. Titanium was chosen as the metallic material for reinforcing the CFRP adherend because of its lower value of thermal expansion

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coefficient, when compared to the value for the aluminium present in the original FML technology, avoiding the aforementioned problems of galvanic corrosion and thermally

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induced stresses. Tensile and shear properties of the adhesive, as well as the fracture energies in mode I and II were determined, through tensile, TAST (Thick Adherend Shear Test) and DCB (Double Cantilever Beam) and ENF (End Notched Flexure) tests, respectively. A numerical study was carried out using a Finite Element Analysis (FEA) with cohesive zone models (CZMs) to simulate the fracture behaviour of CFRP and hybrid laminates titanium carbon-fibre joints.

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ACCEPTED MANUSCRIPT 2. Characterization of the adhesive The adhesive used in this study is designated commercially as AF 163-2K, with 25.28 kg/m2 of weight per area (nominal weight of 0.06 lb/ft2 according to the manufacturer), supplied by 3M Scotch-Weld (Maplewood, MN, USA). This is a modified epoxy

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structural adhesive with a knit supporting carrier, whose main function is to ensure the thickness of the adhesive layer. The cure of this adhesive was performed in a hot-press machine at 120 °C for 1.5 hour. In order to fully characterize the adhesive, bulk tensile

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test (tensile strength, Young´s modulus), TAST (shear strength), DCB (fracture energy

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in pure mode I) and ENF (fracture energy in pure mode II) tests were performed.

2.1. Specimens fabrication i) Bulk test specimens

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Adhesive plates were produced according to the French NF T 76-142 standard [28]. In order to produce 2 mm thick adhesive plates, a total of nine 0.24 mm thick film layers were cut and sequentially stacked. To avoid any defects, such as voids in the adhesive

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plate, the cure was performed using an appropriate mould with a uniform pressure of 2 MPa in a hot plate press. A silicone rubber frame was used to ensure the final thickness

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of the adhesive plate (Figure 1).

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Mould lid Adhesive plate

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Silicone rubber frame

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Mould base

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Figure 1 – Schematic view of the mould and the silicone rubber frame The adhesive plate specimens were then machined into the bulk tensile specimens with

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standardized dimensions (Figure 2), in accordance to the BS 2782 standard [29].

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Figure 2 – Bulk tensile test specimen geometry (dimensions in mm).

ii) TAST specimens

TAST specimens were manufactured according to the ISO 11003-2 standard [30]. Steel adherends are used since the adherend deformation and rotation are reduced when in comparison to aluminium ones. Figure 3 shows the geometry used. Steel adherends with 12 mm of thickness are used in order to avoid plastic deformation of adherends.

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iii) DCB and ENF specimens

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Figure 3 – TAST specimen geometry (dimensions in millimetres)

The DCB and ENF tests use the same type of specimens, schematically represented in Figure 4. Steel adherends were used in order to ensure that the crack grows stably along

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the adhesive layer without plastic deformation of the adherends. The surface of the specimens was grit blasted and then degreased with acetone. In order to ensure the

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thickness of the adhesive, and despite the fact that the adhesive in study was knit supported, calibrated steel spacers were placed at both ends of the bonded length. These

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spacers include a blade at mid-thickness, with the purpose of inducing a pre-crack in the adhesive layer. This pre-crack is produced prior to the to DCB and ENF tests, by preloading the specimens for a short period of time until crack propagation is detected. This prevents blunt crack effect and ensures stable propagation of the crack through the adhesive. This initial crack length (a) was always measured and registered before the DCB and ENF tests were performed, its values varying between 55 to 60 mm.

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Figure 4 – DCB and ENF specimen geometry.

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The adhesive thickness has a major importance during this test, since it influences the mechanical properties of the adhesive. This is due to the fact that different thicknesses

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cause different constrains on deformations around the crack tip, which will naturally affect the way the fracture process zone (FPZ) is propagated [31].

Tensile test

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2.2.

The tensile tests for the bulk specimens were performed in an INSTRON® (Norwood, MA, USA) 3367 electro-mechanical testing machine, equipped with a 25 mm strain

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gauge and a 30 kN load cell, with a displacement rate of 1mm/min and under laboratory ambient conditions (room temperature of 23ºC, relative humidity of 55%). Five tests

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were performed per sample, and the three most relevant test were considered. The stress-strain curves obtained from tensile testing are shown in Figure 5.

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50 45 40

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30 25 20 15

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Tensile Stress [MPa]

35

10

0 0

0.02

0.04

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5

0.06

0.08

0.1

Specimen 1 Specimen 2 Specimen 3 0.12

0.14

Strain [mm/mm]

Figure 5 - Tensile stress-strain curves of AF 163-2K.

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The values for Young´s modulus were calculated from the tangent at the origin to the tensile stress-strain curve by making use of a polynomial approximation of the curve. The Young’s modulus was about 1.5 GPa and tensile strength of 46.9 MPa, which is

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consistent with the manufacturer supplied data.

2.3.

TAST

The TAST specimens were tested on the same electro-mechanical testing machine, according to the ISO 11003-2 standard [30]. The tests were performed at room temperature and a displacement rate of 1 mm/min was used. A shear strength of 46.9 MPa was determined through these shear tests.

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ACCEPTED MANUSCRIPT 2.4.

Pure mode I fracture toughness

The DCB test was selected and performed for determining the fracture energy in mode I (GIC). This test consists on loading the adhesive under mode I. During crack propagation, the values of force (P) and displacement (δ) are recorded. In order to

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calculate the fracture toughness, the compliance based beam method (CBBM) was used, a technique developed by de Moura et al. [32]. In this method, an equivalent crack length is computed based only on the P-δ curve during the crack propagation phase,

Data analysis

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i)

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which means the crack length measurement during the test is not required.

The tests were performed on the same previously mentioned electro-mechanical universal testing machine, at room temperature and with a displacement rate of 0.2 mm/min. The load-displacement curves obtained from the test are represented in Figure 6, while the R-curves resultant from the CBBM data reduction are shown in Figure 7.

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With the crack propagation the critical strain energy converges to a plateau, corresponding to the fracture toughness of the adhesive in mode I. The mode I fracture

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energy was approximately 4.1 N/mm. The failure was found to be cohesive through the

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entire bonded length.

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3000 Specimen 1 Specimen 2

2500

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2000

1500

1000

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0 2

4

6

8

10

12

14

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0

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Load [N]

Specimen 3

Displacement [mm]

Figure 6 - Load-displacement curves for the DCB test

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6

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4

3

2

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GIC [N/mm]

5

Specimen 1 Specimen 2

1

Specimen 3

0

40

60

80

100

120

140

aeq [mm]

Figure 7 - Experimental R-curves obtained for the DCB test using the CBBM

2.5.

Pure mode II toughness 12

ACCEPTED MANUSCRIPT The ENF test was conducted to determine the fracture energy in mode II. This consists of of a simple three-point bending test (a beam simply supported on the extremes and loaded at mid length) of a bonded specimen, with a pre-crack on one edge of the bonding length, placing the adhesive under a pure mode II loading. In order to calculate

i)

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the critical energy release rate in mode II, GIIC, the CBBM method was used [33]. Data analysis

The specimens were tested in the same electro-mechanical universal testing machine.

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All the specimens were tested at room temperature and with a displacement rate of 0.2 mm/min. The load-displacement curves obtained from the test are represented in Figure

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8 and the R-curves resultant from the CBBM data reduction are shown in Figure 9. It is possible to identify a plateau in the R-curves for one single curve (specimen 2), corresponding to stable crack propagation during the test. For the other specimens the plateau is not well pronounced and, in these cases, the GIIC was determined by

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considering the average point of the R-curve that corresponds to a reduction of the load in the load-displacement curve. By identifying this plateau, it is then possible to determine the fracture toughness of the adhesive in mode II. The mode II fracture

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toughness was approximately 9.8 N/mm. The failure was cohesive in the adhesive, but

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close to the interface.

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20

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Load [kN]

15

10

Specimen 1

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0 0

0.5

1

1.5

2

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Specimen 2

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Specimen 3

3

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Displacement [mm]

Figure 8 – Load-displacement curves for the ENF test 14

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GIIC [N/mm]

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4

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2

Specimen 1 Specimen 2 Specimen 3

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75

85

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105

115

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Figure 9 – experimental R-curves obtained for the ENF test using the CBBM A summary of the mechanical properties determined for the adhesive are shown in Table 1. These values were then used for the numerical analysis of the adhesive, characterizing both elastic and cohesive elements.

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ACCEPTED MANUSCRIPT Table 1 - Summary of the mechanical properties of AF 163-2K 46.93 ± 0.63

Young’s Modulus [GPa]

1.52 ± 0.12

Shear strength [MPa]

46.86 ± 2.57

GIc [N/mm]

4.05 ± 0.07

GIIc [N/mm]

9.77 ± 0.21

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Tensile strength [MPa]

3. Experimental details CFRP

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3.1.

The CFRP used was a unidirectional 0° carbon-epoxy composite, commercial reference HS 160 T700, supplied in a prepreg roll by Composite Materials (Legnano, Italy). For the adherends preparation, prepreg sheets of 300 by 300 mm were cut and manually

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stacked. Each CFRP layer of HS1600 T700 has 0.15 mm of thickness, so the final thickness of CFRP depends on the number of layers that are stacked and on the configuration that is being studied. After the introduction of metal laminates, the

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specimens were cured in a hot plate press, to obtain the final hybrid laminate.

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The mechanical properties of the CFRP are presented in Table 2, as determined by Campilho et al [34].

Table 2: Orthotropic components for a unidirectional CFRP ply [34]

Ex

Ey

Ez νzy

[MPa]

[MPa]

[MPa]

109000

8819

8819

0.342

νyz

0.342

Gxy

Gyz

Gxz

[MPa]

[MPa]

[MPa]

4315

4315

3200

νxz

0.380

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3.2.

Titanium alloy

The titanium alloy used for the fibre metal laminates was the Ti-6Al-4V alpha-beta

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(Grade 5) 16 alloy, annealed, and was provided by Smiths Metal Centres Ltd (Biggleswade, UK) [35]. This specific titanium alloy is extensively used in aerospace applications such as bolts, seat rails (in airframes) and fan blades (in engines) [36]. The alloy was supplied in 300 by 300 mm sheets with 0.8 mm of thickness. The mechanical

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properties of this titanium alloy are presented in Table 3.

Young’s Yield Strength Modulus

Coefficient of

Elongation

Poisson’s ratio

[Mpa]

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[Gpa]

900

0.342

[%]

thermal expansion [µm/m.K-1]

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8.6

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113.8

Lay-up configurations

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3.3.

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Table 3: Mechanical properties of Ti-6Al-4V alpha-beta, annealed [35]

To study of the effect of different lay-up configurations on the failure mode and failure load of a fibre metal laminate SLJ, three different lay-up configurations were suggested. Serving as the reference specimens, SLJs using CFRP only adherends were manufactured, to assess the performance gains provided by the novel specimen configurations; In addition to the basic configuration, other two configurations were tested employing the FML concept, using the Ti-6Al-4V alpha-beta (Grade 5), annealed, as the laminate 16

ACCEPTED MANUSCRIPT metal. In one configuration this was combined with the introduction of additional adhesive layers in the Ti-CFRP interfaces: One, designated as Ti-CFRP-Ti, where the CFRP adherend is reinforced on both extremities by a 0.8 mm thick ply of titanium. This configuration was suggested based

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on the search for a configuration that used the titanium plies in direct contact with the main adhesive layer of the joint, in order to avoid the occurrence of delamination, using

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a balanced ratio of titanium and CFRP for the adherend (50% Ti-50% CFRP);

An additional configuration was suggested where two additional layers of AF 163-2K

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adhesive, with an approximate thickness of 0.2 mm, were used between the interfaces CFRP and titanium, to avoid any adhesion problems that could occur; Figure 10 shows a schematic representation of lay-up configurations used in this study – including the basic CFRP only configuration and the lay-ups using titanium laminates

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a)

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and additional adhesive layers as reinforcements for the CFRP basic adherend.

Ti-6Al-4V Titanium alloy

c)

AF 163-2K – Adhesive

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CFRP

b)

Figure 10: Lay-up configurations a) CFRP only b) Ti-CFRP-Ti c) Ti-Adh-CFRP-AdhTi.

3.4.

SLJ Specimen Manufacture

All the manufactured SLJs are approximately 3.2 mm thick, 25 mm wide and have an overlap length of 50 mm. The dimensions of the SLJs are shown in Figure 11. 17

Figure 11:

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SLJs geometry schematization.

The manufacturing process, directly related with that of the CFRP, is described below: The prepreg roll is removed from the freezer and left to warm until it reaches

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I.

II.

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room temperature (about 26°C);

While the prepreg is defrosting, the mould components used to manufacture the

laminates are cleaned and degreased. The cleaning and degreasing process is performed first using fine sandpaper to remove the solid impurities, followed by cleaning with an

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organic solvent (acetone). The mould preparation is completed with the application of a release agent to the mould components so that the plate’s removal may be easier at the end of the cure cycle. The product used for this purpose was Loctite® Frekote 770-NC,

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provided by Henkel (Dusseldorf, Germany); When the prepreg reaches the room temperature, several 300 mm by 300 mm

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sheets are cut to use in different configurations and the roll is re-stored in the freezer. For the stacking process of the final adherends a hand lay-up method was used, by stacking the metal laminates, the CFRP and the adhesive, in such a way, that the proposed configurations were obtained. To obtain a unidirectional final plate, every single composite layer must be stacked with the fibres oriented in the same direction. The number of layers stacked in each configuration varied according with the lay-up, in such a way that the final FML could be approximately 3.2 ± 0.1 mm thick for all the specimens manufactured. 18

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3.5 Titanium surface preparation The surface of the titanium laminates was grit blasted alternately on both sides, to

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prevent bending due the unbalanced residual stresses. Grit-blasting is a mechanical surface treatment that is used to produce a clean macroscopically rough surface and to remove surface contaminants. This mechanical process uses a machine that projects an

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abrasive material against the surface, such as alumina (particle size of 45 µm), with several passes at a distance of about 15-20 cm to the surface being treated, under high

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pressure [37].

The surface was at first cleaned with acetone to remove some ink marks present on the sheet surface, followed by the grit-blasting process. Afterwards, the laminates were degreased and cleaned again with acetone to remove the last impurities. The machine

Italy).

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used in this work was a grit blaster model 705 GM, produced by de Laurentiis (Pescara,

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The manufacture of the final FMLs was performed immediately after the grit blasting

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process was completed, to avoid any kind of contamination.

3.6.

Cure cycle

After the stacking process was completed, the stacked specimens were assembled in the manufacturing mould and placed on the hot plate press, in order to initiate the cure cycle. All the supplier recommendations were followed, including a heating rate of 4°C/min until a temperature of 130°C was achieved. Once the cure cycle was completed and the 19

ACCEPTED MANUSCRIPT plate cooled to room temperature, it was removed from the press and then cut to the desired dimensions.

Testing conditions

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3.7.

For each suggested configuration, three SLJs specimens were manufactured and tensile tested in a servo-hydraulic machine, MTS® (Eden Prairie, MN, USA) model 810, with a

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load cell of 100 kN. Tests were performed at a constant crosshead speed of 1mm/min, at laboratory ambient conditions (room temperature of 23ºC, relative humidity of 55%).

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The SLJs were fixed using clamps that hold the free extremities of each specimen. Dowel pins were also used to align the sample and the clamps. The 4 bolts that hold the two parts of each clamp were tightened using a torque wrench. The torque was applied

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progressively, up to 40 N.m, ensuring an even clamping force.

4. Experimental results

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All the SLJ specimens manufactured were tested in tension, with the load and displacement being registered. In this section the typical load-displacement curve and

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the value for average failure load obtained for each configuration will be presented and compared against the reference configuration. The failure mode was also analysed, and it will be shown for each configuration. Regarding the reinforcement of the basic CFRP only adherend with titanium laminates and with additional adhesive layers, the typical load-displacement curves obtained for each configuration studied are shown in Figure 12.

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Load [kN]

35 30 25

CFRP only

20

Ti-CFRP-Ti

15

Ti-Adh-CFRP-Adh-Ti

5 0 0

0.2

0.4

0.6 0.8 Displacement [mm]

1

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10

1.2

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Figure 12: Typical load-displacement curves for the configurations under study.

All the titanium configurations tested presented improvements both in terms of failure load and failure mode when compared with the original lay-up, with the Ti-Adh-CFRPAdh-Ti configuration achieving the best results. With this configuration, it was possible

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to obtain an improvement in terms of average failure load approximately 30 % and no adhesion problems were detected. This improvement in failure load found by testing with this configuration is thought to be related with the fact that, with the introduction

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of additional adhesive in the adherend, layers with lower rigidity are created, allowing the stresses to be redistributed and absorbed.

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Analysing the failure surface of the tested CFRP only specimens, shown in Figure 13 a), it is possible to observe delamination of the CFRP fibres, an expected phenomenon, that is recurrent in this type of adhesively jointed materials, as referred in the introductory section of this document. The value for the average failure load for CFRP only SLJs was 29.44 ± 0.82 kN. These results for the CFRP only specimens, will serve as the baseline for the assessing the performance of the suggested novel lay-up configurations.

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ACCEPTED MANUSCRIPT For all the configurations under study, and when using the suggested concept of reinforcing the CFRP adherend with titanium laminates and with additional adhesive layers, the failure mode observed was cohesive failure in the adhesive layer, as can be

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seen in Figure 13.

Figure 13: Failure mode of a SLJ using the configuration a) CFRP only b) Ti-CFRP-Ti c) Ti-Adh-CFRP-Adh-Ti.

It can be therefore concluded that the usage of metal layers for manufacturing an FML,

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in all the suggested configurations, can change the failure mode from delamination of the composite to cohesive failure in the adhesive which makes the control of the failure process much simpler from a mechanical design approach. Moreover, an increase of the

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joint strength was obtained when hybrid joints were used.

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5. Numerical study 5.1.

Analysis of the effect of the adherends stiffness under peel stresses

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Before a detailed numerical analysis was performed to assess the joint performance of the novel FML concept-based configurations, it was considered more relevant to start by evaluating the distribution of the peel stresses in a more basic FML configuration.

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For this purpose, a numerical model with elastic elements was created in Abaqus®. The model is based on a 2D planar deformable shell part, using the material properties listed

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above in the material section.

In terms of boundary conditions, at the right end of the SLJ a constant load of 30 kN was applied for every stiffness level studied, while the left end was fixed in every direction. Also, for both edges, the movement was limited in the vertical direction,

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representing the forces imposed by the griping system of the test machine, Figure 14.

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Figure 14: Schematic view of the physical boundary conditions in Abaqus®, quasistatic conditions, for the study of the effect of the adherend’s rigidity on the peel stresses.

Starting from the referred lay-up configuration shown in Figure 15, the stiffness of the material used in the extremities was modified and the peel stresses were evaluated using two “paths” running along the overlap length, one at the middle of adhesive layer and another at a depth corresponding to the bottom of the first layer of CFRP pre-prepreg. 23

ACCEPTED MANUSCRIPT This procedure allows to understand the effect of the material used in an FML and to

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assess if using titanium as the reinforcing material was in fact the correct approach.

Figure 15: Schematization of the configuration used to vary the rigidity of the material

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used to reinforce an FML SLJ.

The results obtained for the case in which the peel stresses were taken along the middle of the adhesive layer can be seen in the Figure 16. It can be noted that with the increase

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overlap tend to reduce.

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of the stiffness of the material used in the FML, the maximum peel stresses along the

Figure 16: Evolution of the maximum peel stresses along the overlap with the variation of the Young’s modulus of the material used in the FML.

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ACCEPTED MANUSCRIPT The results obtained for the case in which the peel stresses were measured on a location after the first layer of the CFRP prepreg can be analysed in the Figure 17, allowing to conclude that, with the increase of the stiffness of the material used in the FML, the maximum peel stresses tend to reduce. In aeronautical applications, as referred in the

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introductory section, one of the major aspects to consider is the weight of the structures. Thus, it is important to reach a point of balance between the density of the materials and, in this case, the peel stresses generated by the use of a given material. In Figure 17,

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a study is shown comparing the peel stresses taken in one of the cases above and the stiffness and density of the reinforcing material, where it can be seen that there is an

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optimal area, where it is possible to balance these factors.

Figure 17: Evolution of the maximum peel stresses in an FML with the variation of the Young’s modulus of the material used compared with the variation of material’s densities.

Considering the material under study in this project, Ti-6Al-4V (Young’s Modulus is 113.8 GPa), there is a suitable balance between the peel stresses given by the application of this material as the reinforcement for the CFRP. The plots also why more 25

ACCEPTED MANUSCRIPT rigid materials, such as steel would not be ideal for this application – although the use of steel would result in lower peel stresses on the adherend, its higher density makes it less compatible with the requirements for aeronautical uses.

Failure load and failure mode prediction of SLJ specimens

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5.2.

The numerical model used predict the performance of single lap joints was also developed using the Abaqus® software, with the main goal of numerically reproducing

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the results obtained with the experimental tests, both for failure load and failure mode.

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In this analysis, a traction-separation law was applied to the adhesive layers of the model in order to introduce damage evolution in cohesive elements during the analysis. A trapezoidal cohesive law was employed for this work, mainly due to the semi-ductile nature of the adhesive used [38-40].

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In the section of the model using CFRP, it was also necessary to include a cohesive zone with a traction separation law with the parameters present in the Table 4, with the aim of modelling composite delamination. This cohesive layer was modelled using a

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thickness of 0.02 mm and was placed 0.15 mm away from, depending on the configuration, the adhesive layer or the metal layer. This value of 0.15 mm was used

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because it corresponds to the thickness of a single CFRP prepreg layer. A schematic representation of the placement of the several cohesive layers used in the developed model is shown and described in Figure 18.

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Figure 18: Schematic representation of the location of several CZE layers throughout the SLJ in the numerical model.

Table 4: Cohesive parameters for CFRP interlaminar failure Mode II

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σR [MPa]

Mode I

0.66

1.13

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Gc [N/mm]

Two calculation steps were sequentially applied to the model, a first step modelling the

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curing process during the manufacture of the SLJ, followed by another related to the boundary conditions needed to simulate the tensile test. The thermal residual stresses due to the difference in the thermal expansion coefficient are a very important concern especially in the study regarding CFRP and aluminium. To take this factor into account, and to make the model as realistic as possible, a primary step was introduced before applying the displacement, in which a thermal field was imposed to the model, imposing at first, the cure temperature applied experimentally,

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ACCEPTED MANUSCRIPT followed by the imposition of room temperature conditions, therefore simulating the temperature changes that occur during the cure cycle process. The coefficients of thermal expansion used were 0 µm/m.K-1, for CFRP and 8.6 µm/m.K-1, for the titanium. The boundary conditions of the model were selected to closely match those that occur in

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a tensile test. A constant displacement was applied at the right end of the joint, while the left end was fixed in every direction. Also, for both edges, the movement was limited in the vertical direction, representing the forces imposed by the griping system of the test

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machine.

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The mesh used was refined until the element spacing was set at 0.2 mm, a value corresponding to the adhesive layer thickness. The elastic sections of the model were modelled using 8-node bilinear plane strain elements, CPE8R and the cohesive sections

5.3.

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employed a 4-node two-dimensional cohesive element, COH2D4.

Numerical results

The numerical results obtained by running numerical analysis for each of the specimen

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lay-up configurations using the configurations regarding titanium and adhesive as the

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reinforcements for the CFRP basic joints under study will be presented. To exemplify the ability of the model to simulate the mechanical behaviour of the experimental joint in a tensile test, a comparison between the experimental and numerical loaddisplacement curve of one of the configurations is presented in Figure 19, in this case the Ti-CFRP-Ti configuration.

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Figure 19: Numerical P-δ curve vs experimental P-δ curve, for a 50 mm overlap SLJ, with the configuration Ti-CFRP-Ti.

In Figure 20 the numerical failure modes for the suggested configurations are shown, alongside those observed experimentally. The numerical failure mode presented by

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every titanium configuration was cohesive in the adhesive layer of the joint, which is coherent with the experimental results. For the CFRP only configuration the numerical result is also coherent with the obtained experimentally, being possible to check the

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occurrence of delamination.

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Figure 20: Comparison of the numerical and experimental failure modes obtained for

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the SLJs with the configuration a) CFRP only b) Ti-CFRP-Ti c) Ti-Adh-CFRP-Adh-Ti.

The numerically predicted failure loads are presented and compared with the experimental data in Figure 21.

Experimental

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45 40 35

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10 5 0

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Failure load (kN)

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15

Numerical

CFRP only

Ti-CFRP-Ti [0.8 mm]

Ti-Adh-CFRP-Adh-Ti [0.8 mm]

Figure 21: Comparison of the numerical and experimental failure load obtained.

For the models involving the proposed novel lay-ups, the developed model predicts the failure load relatively well, always on the safety side – with predicted numerical failure 30

ACCEPTED MANUSCRIPT load slightly lower than that obtained experimentally – which can be seen as a positive point when considering a joint design tool for structural applications. The configuration that achieved closer results between numerical and experimental data, was the TiCFRP-Ti lay-up, where the effect of the stacking of cohesive layers in the substrate is

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minimum. Comparing the density of the hybrid joints with that of the CFRP joints, there is an increase when Ti is used for reinforcement (see Figure 17). This could be optimized if Ti was only used to locally reinforce the area that shows a high level of

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peel strength.

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6. Conclusions The main objective of this project was to explore several lay-up configurations for reinforcing a basic CFRP substrate in order to increase the peel strength of a CFRP

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substrate, as well as the adhesively bonded joint strength itself when hybrid materials are used as adherends.

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The work employed an approach based on a concept similar to FML, using metal plies combined with the application of additional adhesive in the interfaces between the metal

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and the CFRP. To apply this concept, several novel lay-up configurations for SLJs with an overlap of 50 mm and a thickness of 3.2 mm were suggested, manufactured, tested and evaluated in terms of failure load and failure mode.

For the suggested novel configurations which use 0.8 mm thick titanium plies, the

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configuration where the best results were reached was the configuration defined as TiAdh-CFRP-Adh-Ti, where the average failure load presented by the specimens tested had a good improvement, when compared to the basic CFRP only configuration. For all

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the configurations tested, the failure modes obtained were also very satisfactory, exhibiting cohesive failure in the adhesive layer of the joint, being the occurrence of

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delamination eliminated.

Regarding the numerical analysis, an initial study was made to assess the effectivity of using titanium as the metal for the suggested FML, reinforcing the fact that, in the material selection process for aerospace applications, it is important to balance the stresses that will be present on the structures with the weight of the final structure. In terms of the simulation of the tensile testing of the joints, the numerical results were acceptably coherent with the experimental results as expected. Regarding the failure 32

ACCEPTED MANUSCRIPT mode obtained numerically for the configurations under study, the results were in every case coherent with the experimental data, being possible to obtain in each case a cohesive failure in the adhesive layer of the joint. To summarize, the experimental and numerical procedures undertaken, strongly indicate

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that it is possible to significantly reinforce a CFRP joint using titanium plies combined with the application of additional interfacial adhesive layers. The best results are obtained by the introduction of titanium laminates combined with adhesive in the Ti-

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CFRP interfaces, where a significant increase in the failure load and in the peel strength

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of a CFRP joint can be observed, with no occurrence of delamination.

Acknowledgments

R. J. C. Carbas would like to thank the Portuguese Foundation for Science and

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Technology (FCT) for supporting the work presented here, through the individual grant

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References

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SFRH/BPD/96992/2013.

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ACCEPTED MANUSCRIPT [5] The new-technology Boeing 787 dreamliner, which makes extensive use of composite materials, promises to revolutionize commercial air travel. Aviation Week & Space Technology Market Supplement, 2005;16201:S1-31. [6] Kolesnikov B, Herbeck L, Fink A. CFRP/titanium hybrid material for improving composite bolted joints. Composite Structures, 2008;83:368-80.

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