Materials And Design

Materials and Design 32 (2011) 3671–3685
Contents lists available at ScienceDirect
Materials and Design
journal homepage: www.elsevier.com/locate/matdes
Review
A review: Fibre metal laminates, background, bonding types and applied
test methods
Tamer Sinmazçelik a,b,⇑, Egemen Avcu a, Mustafa Özgür Bora a, Onur Çoban a
a
b
Kocaeli University, Mech. Eng. Dept., Umuttepe Campus, 41380 Izmit, Kocaeli, Turkey
TUBITAK-MRC, Materials Institute, P.K. 21, 41470 Gebze, Kocaeli, Turkey
a r t i c l e
i n f o
Article history:
Received 28 January 2011
Accepted 4 March 2011
Available online 12 March 2011
Keywords:
B. Sandwich structures
C. Surface treatments
E. Mechanical
a b s t r a c t
During the past decades, increasing demand in aircraft industry for high-performance, lightweight
structures have stimulated a strong trend towards the development of refined models for fibre-metal
laminates (FMLs). Fibre metal laminates are hybrid composite materials built up from interlacing layers of thin metals and fibre reinforced adhesives. The most commercially available fibre metal laminates (FMLs) are ARALL (Aramid Reinforced Aluminium Laminate), based on aramid fibres, GLARE
(Glass Reinforced Aluminium Laminate), based on high strength glass fibres and CARALL (Carbon Reinforced Aluminium Laminate), based on carbon fibres. Taking advantage of the hybrid nature from
their two key constituents: metals (mostly aluminium) and fibre-reinforced laminate, these composites offer several advantages such as better damage tolerance to fatigue crack growth and impact
damage especially for aircraft applications. Metallic layers and fibre reinforced laminate can be
bonded by classical techniques, i.e. mechanically and adhesively. Adhesively bonded fibre metal laminates have been shown to be far more fatigue resistant than equivalent mechanically bonded
structures.
Ó 2011 Elsevier Ltd. All rights reserved.
1. Introduction
Composite materials have been subject of permanent interest of
various specialists during the last decades. Firstly, military applications in the aircraft industry triggered off the commercial use of
composites after the Second World War. The innovations in the
composite area have allowed significant weight reduction in structural design. Composites offer many advantages when compared to
metallic alloys, especially where high strength and stiffness to
weigh ratio is concerned. Additionally, they provide excellent fatigue properties and corrosion resistance in applications [1]. With all
these advantages, composite structures have gained widespread
use in the aerospace industry during the last decades [2–6].
In the fifties, an important goal of aircraft materials development was to improve the crack growth properties of structural
materials [7]. Competing materials like advanced aluminium alloys
and fibre reinforced composites have potential to increase the cost
⇑ Corresponding author at: Kocaeli University, Mech. Eng. Dept., Umuttepe
Campus, 41380 Izmit, Kocaeli, Turkey. Tel.: +90 262 303 3006; fax: +90 262 303
3003.
E-mail address: [email protected] (T. Sinmazçelik).
0261-3069/$ - see front matter Ó 2011 Elsevier Ltd. All rights reserved.
doi:10.1016/j.matdes.2011.03.011
effectiveness of the structure. These materials still have their
advantages and disadvantages, like the poor fatigue strength of
the aluminium alloys and the poor impact and residual strength
properties of carbon fibre reinforced composites. At the end of
the seventies, the idea of using the two materials to form a hybrid
composite structural material to overcome most of the disadvantages of both materials was born [8]. At the Delft University of
Technology, it was found that the fatigue crack growth rates in
adhesive bonded sheet materials can be reduced, if they are built
up by laminating and adhesively bonding thin sheets of the material, instead of using on one thick monolithic sheet [7]. The advantage becomes highly evident if cracks start in one of the sheets of
the laminate only, the adhesive layers behaving as crack dividers.
Under these circumstances, the sheets that are still uncracked
reduce the crack growth rate in the cracked sheet. The reduction
in the crack growth rates persists until a crack is initiated in the
neighbouring sheet also. Based upon all the research, initial Fibre
Metal Laminates material, ARALL (Aramid Fibre Reinforced
Aluminium Laminate) was introduced at 1978 in the Faculty of
Aerospace Engineering at the Delft University of Technology
(DUT) [3]. ARALL consists of alternating thin aluminium alloy
layers (0.2 ± 0.4 mm) and uniaxial or biaxial aramid fibre prepreg,
as shown in Fig. 1 [8].
3672
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
Fig. 1. Schematic presentation of Fibre Metal Laminate (ARALL 2).
Fibre metal laminates (FMLs) are hybrid composite structures
based on thin sheets of metal alloys and plies of fibre reinforced
polymeric materials [6]. The fibre/metal composite technology
combines the advantages of metallic materials and fibre reinforced
matrix systems. Metals are for instance isotropic, have a high bearing strength and impact resistance and are easy to repair, while full
composites have excellent fatigue characteristics and high strength
and stiffness. The fatigue and corrosion characteristics of metals
and the low bearing strength, impact resistance and reparability
of composites can be overcome by the combination [5,9].
These material systems are created by bonding composite laminate plies to metal plies [8]. The concept usually applied to alu-
minium with aramid and glass fibres, but also can be applied to
other constituents [10]. Fig. 2 gives a classification of FML based
on metal plies. The most commercially available FMLs are ARALL1,
based on aramid fibres and GLARE1 based on high strength glass
fibres [8]. These two grades of FMLs are going to be explained in
detail Chapter 3.
1.1. Historical development of fibre metal laminates
During the last three decades, there has been a search for lightweight materials that can replace the traditional aluminium alloys
in aerospace structures [11]. For an optimal structural design, a
Fig. 2. Classification of FMLs based on metal plies.
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
new material is needed which combines high strength, low density
and high elasticity modulus with improved toughness, corrosion
resistance and fatigue properties. Fibre reinforced composites
materials almost cover all these demands, except for fracture
toughness.
In 1978, researches were carried out to increase the fatigue performance of aluminium alloys at the National Aerospace Laboratory and at the Delft University of Technology in Netherland. An
improvement of the fatigue behaviour in laminate sheet materials
was obtained by introducing a high strength aramid fibre into the
adhesive layers. As a result of all these studies, they introduced
ARALL, first fibre metal laminate at the Faculty of Aerospace Engineering at the Delft University of Technology in Netherland [3]. In
1984, two international patents were accepted and a pilot production of four different types of standardized ARALL was started by
the Alcoa Company after sufficient confidence in this material
had been gained [7]. The trades ARALL 1 and ARALL 2 were standardized. Arall 1 is a variant with aluminium 7075 layers and Arall
2 uses aluminium 2024 layers and it was in the as-cured condition
[1].
Later, a much stiffer ARALL which consists carbon fibres instead
of aramid fibres, the CARALL Laminates, had been investigated in
DUT [12]. Recent research has shown that CARALL laminates also
have fibre failure occurred during flight-simulation fatigue tests
at elevated stress levels, which resulted in poor fatigue performance. The limited failure strain of the carbon fibres (0.5–2.0%)
was thought to be a disadvantage. Thus, it is sensitive to notch
behaviour comparing to monolithic aluminium alloy. Due to the
problem of galvanic corrosion between the carbon fibres and the
aluminium sheet in moisture environment, more research has to
be done.
In 1990, another attempt to improve ARALL laminates, adopting
high strength glass fibre instead of aramid fibres, called GLARE
(Glass reinforced, or ARALL with glass fibres) was developed successfully [7]. Glass reinforced aluminium (GLARE) is the successor
of Aramid aluminium laminate (ARALL) [12]. A partnership between AKZO and ALCOA started to operate in 1991 to produce
and commercialize GLARE [1].
Finally, a new concept for fibre metal laminates (ARALL &
GLARE), the Spliced Laminates, is launched by the Structural Laminates Company (SLC) in 1992. Spliced laminates are defined as
FML in which the aluminium layers are interrupted such that the
dimensions of the spliced sheets depend on the autoclave dimensions only [7].
1.2. Advantages and disadvantages of fibre metal laminates
Fibre metal laminates take advantages of metal and fibre-reinforced composites, providing superior mechanical properties to
the conventional lamina consisting only of fibre-reinforced lamina
or monolithic aluminium alloys [9]. Table 1 summarizes all the
advantages of fibre metal laminates depending on previous
studies.
The major disadvantage associated with epoxy based fibre
metal laminates is the long processing cycle to cure the polymer
matrix in the composite plies [6]. This problem increases the cycle
time of whole production and decreases productivity. This
increases labour costs and overall cost of FMLs.
1.3. Production of FMLs
To produce FMLs, as for polymeric composite materials, the
most common process involves autoclave processing. The overall
generic scenario for production of FMLs involves about five major
activities.
3673
1. Preparation of tools and materials. During this step, the aluminium layer surfaces are pre-treated by chromic acid or phosphoric acid, in order to improve the adhesive bonding between
metallic layer and fibre reinforced laminate (metallic surface
treatment is explained in detail in Chapter 2).
2. Material deposition, including cutting, lay-up and debunking.
3. Cure preparation, including the tool cleaning and the part transferring in some cases, and the vacuum bag preparation in all
cases.
4. Cure, including the flow-consolidation process, the chemical
curing reactions, as well as the bond between fibre/metal
layers.
5. Post stretching, after the hot-curing (200 °C) cycle in the autoclave, the fibre metal laminates carry a residual stress system
over the thickness of the material with a small tensile stress
in the aluminium sheets and compression in the fibres, this situation adversely effects fatigue resistance of FMLs. Post stretching operation after curing cycle can reverses the residual stress
system and solves this problem.
6. Inspection, usually by ultrasound, X-ray, visual techniques and
mechanical tests [1,7]. Mechanical tests of FMLs are explained
widely in Chapter 4.
1.4. Applications
Due to their advantages aforementioned above, FMLs are finding great use in most commonly in aerospace applications. A number of companies have interest in substitute the traditional
aluminium components by FML composites [5]. Both ARALL and
GLARE laminates are now being used as structural materials in aircrafts. Fibre Metal Laminates have been successfully introduced
into the Airbus A380 [4]. The Fig. 3 shows FML composite applications in the Airbus A380 airplane [1,18].
ARALL has been developed for the lower wing skin panels of the
former Fokker 27 aircraft and the cargo door of the Boeing C-17
[8,10]. ARALL 3 material is currently in production and flight test
on the C-17 cargo doors and GLARE is selected for the Boeing
777 impact resistant bulk cargo floor.
2. Adhesive bonding
Adhesive bonding process has been used in the manufacture of
aircraft structures and components for 30–40 years [19]. Bonding
techniques are used in an ever-growing number of applications:
electronics (multilayer boards, bounded components), aerospace
and aeronautical industries (F-18 bonded wings), automotive
industries and many more organizations. Bonding structural components with adhesives offers many advantages over conventional
mechanical fasteners: lower structural weight, lower fabrication
cost, and improved damage tolerance [20]. Light weight sandwich
construction and structural bonded joints form a major proportion
of modern aircraft [21]. Modern metallic bonded structures using
longitudinal lap joints in commercial aircrafts were introduced
with the advent of the Airbus A300 [19]. As stated in Ref. [19], this
technique is currently migrating from secondary to primary structural applications as one of the most interesting ways to fasten
structural parts with a high level of confidence. For example, structural adhesive bonding is mainly used for attaching stringers and/
or tear straps to the fuselage and wing skins, to stiffen the structures against buckling. It is also applied to skin-to-core bonding
in metallic honeycomb structures, such as elevators, ailerons,
spoilers, and so on [22]. Bonded structures have been shown to
be far more fatigue resistant than equivalent mechanically fastened structures and when designed correctly, can sustain higher
load levels than equivalent mechanically fastened joints [21].
3674
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
Table 1
Advantages of fibre metal laminates.
Material
behaviour
1. High fatigue resistance: It is achieved by the intact bridging fibres in the wake of the crack, which restrain crack opening. FMLs have excellent
fatigue characteristics [5,13]
2. High strength: As mentioned before, FMLs are hybrid structures based on thin sheets of metal alloy and plies of fibre-reinforced polymeric
materials. Combination of metal alloys which have high bearing strength and fibre reinforced composites which have high strength and stiffness.
FMLs are structural materials with high strength [4,5]
3. High fracture toughness: Depending on the studies, FMLs have better fracture toughness than their constituent alloys. Based on this behaviour and
accounting for the lower fatigue crack growth rates presented by these materials, they can be a very attractive option for a wide set of structural
applications [14]
4. High impact resistance: Unlike composites, the damage tolerance behaviour of FML is comparable to conventional aluminium alloys. The same
types of damage and plastic deformation are observed, only at higher impact energy levels. Impact deformation is actually a significant advantage of
FML, especially when compared to composites, because this visible damage significantly increases inspect ability and detect ability. Therefore, the
use of FML allows for repair criteria and techniques [8,15]
5. High energy absorbing capacity: Based upon researches, FMLs are capable of absorbing significant energy through localized fibre fracture and shear
failure in the metal plies [6,16]
Physical
properties
1. Low density: Owing to epoxy based polymer matrix and low density aluminium sheets, FMLs are a weight saving structural material compare to
others [7]
Durability
1. Excellent moisture resistance: The moisture absorption in FML composites is slower when compared with polymer composites, even under the
relatively harsh conditions, due to the barrier of the aluminium outer layers [1]. Additionally pregreg layers are able to act as moisture barriers
between the various aluminium layers inside of the FMLs [13]
2. Excellent corrosion resistance: As mentioned above excellent moisture resistance of FMLs and high corrosion resistance of polymer based fibre
laminates ensures FMLs excellent corrosion resistance [7,8,10,13]
3. Lower material degradation: FMLs have excellent moisture and corrosion resistance, FMLs degradation as result of environmental aspects is
significant lower as compared to either metallic structures, or composite structures [5]
Safety
1. Fire resistance: The high melting point of the fibres in FML (for instance, glass fibres in GLARE laminates can withstand 1100 °C) prevents
penetration of the fire to the inner layers. Hence, the fire resistance of FMLs are much better than monolithic aluminium alloys depending on their
fibre melting point. FMLs are being used as fuselage materials in aircrafts. With their good fire resistance, FMLs ensures enough time to passengers
for evacuating the aircraft safely [8,17]
Cost saving
1. FML offer substantial weight savings relative to current metallic structures. Further, the number of parts required to build a component may be
dramatically less than the number of parts needed to construct the same component of metal alloy. This can lead to labour savings. Sometimes
offsetting the higher price of the present materials
2. Owing to fatigue insensitivity of FMLs, less repair and longer maintenance periods are sufficient for FMLs. These advantages reduce maintenance
costs of FMLs [7]
As reported in Ref. [19], early experiences in bonding techniques demonstrated that surface treatment prior to bonding is
the single most critical step which can not be disregarded, even
for tertiary-loaded structures, since it is essential to achieve
long-term service capability [21,23]. A particular surface treatment
tends to modify the substrate surface by delivering the following
features: free from contamination; wettable with either primer
or adhesive; highly roughened; and mechanically and hydrolytically stable [23]. Surface treatment technologies need to be explained from the viewpoint of adhesively bonded fibre-metal
laminate research.
3. Surface treatments for adhesive bonding
Fig. 3. Metal/fibre applications in A380 airplane from Airbus [1].
As shown in Table 2, all the treatments for modification of metal
surfaces can be grouped as:
One of the most important aspect of designing aircraft structures is to consider skin cracks resulting from fatigue that the
cracks will be discovered within a particular inspection interval
prior to their reaching a critical crack length that can lead to catastrophic failure. As noted in Ref. [7], at the Delft University of Technology, it was found that the fatigue crack growth rates in adhesive
bonded sheet materials can be reduced. If they are built up by laminating and adhesively bonding thin sheets of the material, instead
of using on one thick monolithic sheet. The advantage becomes
highly evident if cracks start in one of the sheets of the laminate
only, the adhesive layers behaving as crack dividers. Under these
circumstances the sheets that are still uncracked reduce the crack
growth rate in the cracked sheet. The reduction in the crack growth
rates persists until a crack is initiated in the neighboring sheet also.
The research on adhesive-bonded sheet metal laminates is then became the predecessor of a new class of materials; the fibre-metal
laminates [7].
1.
2.
3.
4.
5.
Mechanical.
Chemical.
Electrochemical.
Coupling agent.
Dry surface treatments.
Solvent degreasing is important, because it removes contaminant materials which inhibit the formation of the chemical bonds.
However, solvent degreasing, while providing a clean surface, does
not promote the formation of acceptable surface conditions for
longer term bond durability [21]. The degreasing stage usually
makes use of chlorinated solvents such as trichloroethylene,
1,1,1-trichloroethane, perchloroethylene, or dichloromethane, or
alternatively, non-chlorinated solvents including methyl ethyl
ketone, methanol, isobutanol, toluene or acetone [24]. All aluminium alloy sheets were initially degreased prior to further surface
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
Table 2
All treatments for modification of metal surfaces.
Treatments
Nature of
treatments
Ref.
Grit-blasting
Chromic–sulphuric acid (CAE)
Sulfo-ferric acid (P2)
Forest Product Laboratory (FPL)
Alkaline
Chromic acid anodizing (CAA)
Phosphoric acid anodizing (PAA)
Sulphiric acid anodizing (SAA)
Boric-sulphiric acid anodizing (BSAA)
Phosphoric acid anodizing (AC-PAA)
Sulphiric acid anodizing (AC-SAA)
Silane
Sol–gel
Excimer laser texturing
Plasma sprayed coating
Ion beam enchanced deposition
(IBED)
Mechanical
Acid etching
Acid etching
Acid etching
Etching
DC-anodizing
DC-anodizing
DC-anodizing
DC-anodizing
AC-anodizing
AC-anodizing
Coupling/oxidation
Coupling/oxidation
Mechanical
Ablation/oxidation
Ablation/oxidation
[26–29]
[30,33]
[31]
[32]
[25,30]
[46]
[23,34,36]
[36,45]
[24]
[36]
[36]
[28,47,49–52]
[53–57]
[54,58–62]
[35,63–67,69]
[62,68]
pre-treatment steps. The first step in the fabrication of baseline
test specimens was methyl ethyl ketone (MEK)-wiping of aluminium substrates with lint-free tissues to degrease the surface [25].
Each surface treatment technologies which have been used to
modify metal surfaces will be explained detail in subtitles of Section 3.
3.1. Mechanical treatment
As a preliminary preparation step in the multi-stage schedules,
mechanical abrasion has been used to produce a macro-roughened
surface, different roughness levels of the surface textures and to remove an undesirable oxide layer, respectively [19]. This method
typically involves abrasive scrubbing of the substrate surface with
sand paper. This mechanical treatment would introduce physicochemical changes which yield a wettable surface and modify the
surface topography, i.e., a macro-roughened surface [27].
Mechanical treatments include abrasion methods usually combined with degreasing. Pretreatment by blasting using alumina or
silica grit or glass beads change the topography and chemical state
of an aluminium adherend by the introduction of a ‘‘peak-and-valley’’ type morphology [24]. In general terms, degreasing is the minimum pretreatment that is usually carried out prior to bonding.
Grit-blasting or other mechanical abrasion methods are recognized
as providing a useful increase in initial adhesion levels [24]. Several
researches have highlighted different aspects of grit blasting or
other mechanical abrasion methods for aluminium alloys. Gritblasting, which uses alumina grit with clean, dry, propellant air,
is employed as a logical surface activation step prior to modify
the metallic surface for satisfactory bonding results [26–29].
3.2. Chemical treatment
The most commonly applied chemical treatments are based on
a chromic–sulphuric acid etch [23,37]. This treatment consists of
immersion of the substrate in a solution of sulphuric acid and
potassium dichromate. Typically, chemical treatment, i.e., acidetching, is an intermediate production step between degreasing,
alkaline cleaning, and electrochemical treatment [19]. Three classical acid-etching solutions were introduced to modify the metallic
surfaces: chromic–sulphuric acid (CAE) [38], Forest Product Laboratory (FPL) [25], and sulfo-ferric acid (P2) etches. The most effective etches incorporate mixed chromic and hydrofluoric acids.
However, non-chromated acid etchants have been demonstrated
to provide good adhesion results [23,39].The surfaces treated with
3675
acid-etch were consistently reported to be out-performed by anodized surfaces so that the inconsistent results made them unsuitable
for stand-alone treatment in primary bonded structures [23].
Chemical acid etches are given in Table 3 with concentration of
compounds, used temperature and immersion time.
3.3. Electrochemical treatment
Good adhesion between the metal and the adhesive could be
obtained when the aluminium surfaces were degreased in alkaline solutions or organic solvents and subsequently etched in
aqueous chromic–sulphuric acid solutions. However, this surface
pre-treatment was found to be insufficient in certain nonbonded areas where corrosion occurred. The corrosion susceptibility was reduced if the aluminium surface, after chromic–sulphuric acid etching, was anodized before bonding [34]. In
order to obtain optimum durability of adhesive bonded aluminium joints, anodizing processes are extensively used in the aerospace industry. Anodizing produces a thin oxide film, again with
a high degree of micro-roughness and an oxide form which is
highly resistant to hydration [21]. Anodic oxidation in solutions
of chromic acid anodizing (CAA) or phosphoric acid anodizing
(PAA) is the preferred stabilizing treatment for the structural
adhesive bonding of high-strength aluminium alloys in critical
applications such as aircraft components [35]. Chromic acid
anodizing (CAA) was found to give a relatively thin and ductile
oxide layer and was established as an effective pretreatment
for adhesive bonding with good durability properties in service.
The European aviation industry is still using this method [36].
However, as a result of delamination in the boundary zone and
corrosion after only a few years of operation, an improvement
of the usual FPL process was established. Boeing developed the
phosphoric acid anodizing (PAA), which led to a clearly better
joint durability [23,34].
As stated in Ref. [24], in almost all studies, however, the DC
electrochemical pretreatments give the best levels of initial adhesion and durability [23,36,40,41]. The three most widely used
Table 3
Chemical acid etches and their application process.
Chemical
etch type
Application process
Ref.
CAE
Immersion in a water solution with 330 ml/l chromic–
sulphuric acid (97% v/v) and 50 g/l potassium
dichromate, at 60 oC for 15 min, and rinsed in tap water
Immersion in a water solution with 185 ml/l chromic–
sulphuric acid (97% v/v) and 127 g/l ferric sulphate, at
65 °C for 8 min, and rinsed in tap water (percentages by
weight: 48% H2O, 37% H2SO4 and 15% FeSO4)
Immersion at 15 min in a 65 °C, followed by tap and
de-ionised water rinsing
[30]
[30]
[33]
P2
Involves treatment of the adherend surface with an
aqueous acidic solution containing Fe(III). (1)
immersing the adherend in the P2 solution at 65 °C for
8 min, (2) rinsing the specimen in DI water for 2–3 min,
and (3) drying the aluminium bars in an oven at 60 °C
for 30 min. The P2 solution was prepared by dissolving
122.5 g Fe2(SO4)3 4H2O and 0.185 l of concentrated
sulphuric acid in enough water to make a liter of
solution
[31]
FPL
Treat with optimized FPL for 30 min at 62 + 2 °C. After
treatment in the acid, the aluminium was washed for
20 min in cold running tap water and finally dried in an
oven at 40 °C for 30 min
[32]
Alkaline
etch
Immersion in 100 g/l NaOH solution, at 60 °C for 1 min,
and rinsed in tap water
Immersion in %10 wt. NaOH solution for 2 min at 60 °C
[30]
[25]
3676
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
electrochemical processes used in prebond applications are phosphoric acid anodizing (PAA), chromic acid anodizing (CAA) and sulphuric acid anodizing (SAA). Other electrolytes have been studied
but to a much lesser extent. CAA and PAA are both extensively used
in aerospace applications [42,43]. All DC anodizing procedures are
complex multi-stage operations incorporating degreasing, desmutting and deoxidizing stages [34]. The boric–sulphuric acid anodizing (BSAA) process has advantages associated with both the CAA
and PAA. A BSAA process has been patented by Boeing as a direct
replacement to CAA to meet environmental regulations [44]. AC
anodizing is a highly feasible, robust, and environmentally friendly
process, which is free from the hexavalent chromium found in the
conventional CAA. A distinct difference from DC anodizing is
hydrogen gas generation on the treated surface during the hot
AC anodizing process [19]. Electrochemical treatments which are
used for improving the metallic surfaces as an adhesive layer are
shown in Table 4.
3.4. Coupling agent treatment
There is an ongoing need to develop environmentally friendly
coupling agent techniques, i.e., silane or sol–gel, which can be used
with the current generation of materials, i.e., primer and film adhesive [19]. Silane based products are becoming an interesting material for pre-treatment deposition, because, for the environmental
compatibility, they can be used as substitutes of traditional pretreatments like chromates [47]. Silane coatings are environmentally sound, multi-metal surface pre-treatments, considered for
use on various metals such as aluminium and its alloys, copper,
iron and steel, zinc and magnesium containing alloys [48]. It is
sufficient to mention that a large number of studies have demonstrated the usefulness of silanes such as c-glycidoxypropyltrimethoxysilane when used as a primer to improve the durability of
structural aluminium bonds. This has been attributed to the formation of stable, covalent bonds between the metal(oxide) and the
silane and the possibility for interphase formation giving a region
of intermediate modulus between the metal and polymer which
facilitates stress transfer [49,50]. Silanes may be used as standalone processes or part of a more complex pretreatment [28,51].
Bis-1,2-(triethoxysilyl)ethane (BTSE) has received much attention
as it is, after hydrolysis, highly reactive towards (covalent) metal/
film bonding and cross-link formation for the creation of barrier
properties [52]. Fedel et al. [47] examined the effect of water-based
silane pre-treatments on galvanized steel. Three different silanes
were used for the pre-treatment deposition (glycidoxypropiltrimethoxysilane, tetraethoxysilane and methyltriethoxysilane). The
electrochemical tests showed that the silane layer acts not only
as a coupling agent between the inorganic substrate and the organic coating, but it also ensures a good barrier effect against water
and oxygen. Rider and Arnott [28] described that the silane layer
Table 4
AC and DC Electrochemical Treatment.
Pretreatment
Electrolyte
(wt.%)
Voltage (V)
Time
(min)
Temperature
(°C)
Ref.
PAA
SAA
10 (H3PO4)
16 (H2SO4)
20
20
25
20
[36]
[36]
BSAA
5.0–10.0
H3BO3/30.0–
50.0 H2SO4
10 (DC)
1.5 (DC)
(current
density A/dm2)
15 ± 1 (DC)
18–
22
26.7 ± 2.2
[24]
22
35–
45
30 s
12 s
15
40.0 ± 2.0
[45]
[46]
50
80
[36]
[36]
SAA
CAA
2.5–3.0 CrO3
PAA-AC
SAA-AC
10 (H3PO4)
15 (H2SO4)
15 (DC)
40.0 ± 1.0/
50.0 ± 1.0 (DC)
4 (AC)
10 (AC)
promotes the hydrolytic stability of the adhesively bonded joint,
leading to increased durability in wedge tests.
Sol–gels are organic–inorganic polymers formed by hydrolysis/
condensation reactions of alkoxide precursors, primarily silanes,
which have found applications as electronic, optical and protective
coatings. These coatings possess important characteristics such as
chemical stability, physical strength and scratch resistance. The
sol–gel process can be used to form nanostructured inorganic films
(typically 200 nm to 10 lm in overall thickness) that are more
resistant than metals to oxidation, corrosion, erosion and wear
while also possessing good thermal and electrical properties [53].
As noted in Ref. [19], sol–gel which involves the growth of metal–oxo polymers through both hydrolysis and a condensation
reaction to form inorganic polymer networks in thicknesses ranging from 50 to 200 nm [54,55]. The Boegel EPII sol–gel agent consists of a dilute aqueous epoxy-silane and zirconium alkoxide. This
mixture provides a thin inorganic zirconium oxide film incorporated with the epoxy-silane that is applied to the metallic surface
[55,56]. In this process, zirconium reacts with the metallic surface
to produce a covalent chemical bond, while the epoxy-silane offers
a reactive organic group for bonding with the adhesive [55,57].
3.5. Dry surface treatments
Several dry treatments for aluminium alloy surfaces have been
developed to replace chemical wet treatment process: excimer laser texturing [54,58–62], plasma-sprayed coating [35,63–67], and
ion beam enhanced deposition (IBED) [62,68].
As reported in Ref. [19], laser texturing was utilized to modify
an aluminium substrate’s morphology and micro-structure, resulting in an increased bond strength and durability [58,59,62]. IBED is
a process that cleans and modifies the surface by sputtering with
high energy argon ions under vacuum. This process requires a surface activation step, particularly grit-blasting, prior to IBED. Good
initial bond strengths were obtained and the improvements in
wedge durability compared with PAA were found [62]. In the aeronautic industry, a complex and critical process is used in order to
enhance both wettability and adhesive properties of aluminium alloy surfaces.
Cold plasma treatment represents an alternative to conventional processes in terms of environmental impact. Plasma is an
ionized gas containing both charged and neutral particles, such
as electrons, ions, atoms, molecules and radicals. There are two
kinds of plasma: cold and hot. The working pressure of a cold plasma is generally lower than atmospheric pressure which allows the
production of chemical reactive species at low temperatures
(<773 K). Since hot plasma pressure is higher than atmospheric,
the temperature is about 104–105 K [35]. Cold plasma treatment
represents an efficient, clean and economic alternative to activate
aluminium surfaces [67]. Plasma spray coatings were also used
as pretreatments prior to adhesive bonding. Davis et al. [63] reported that a 60Al–Si/40 polyester (by wt.) plasma spray coating
could be engineered for specific applications and was better suited
for the adhesive bonding process of aluminium substrates, by
yielding a performance equivalent to that of the PAA treatment
with some epoxy adhesives in the wedge test. De Iorio et al. [69]
examined cold plasma treatments to modify the surface properties
of polymeric materials (polyoxyphenylene–polyammide, polycarbonate–ABS) and A1 6061 alloy, with the aim of increasing the surface adhesion in structural joints.
4. Mechanical bonding
As reported at Ref.[73], the increasing requirements for weight
reduction demand more and more use of composite materials in
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
aerospace applications [70]. There has been a progressive increase
in the number of metal parts and structures replaced by composite
materials, not only in military, but also in civil aircraft design. For
instance, composite materials comprise 22% of the total weight of
the Airbus A380 [71]. The centre wing box, the tail cone, the pressure bulkheads and the vertical and horizontal tails are only some
of numerous parts of this modern aircraft [72] being made of composite materials. Structural coupling leads to an increase in structural complexity and to additional local stress factors, but also to a
considerable reduction in global optimizing possibilities, which, in
turn, decreases the lightweight potential of composite design.
Thus, the development of advanced joining technologies designed
to minimize the weight penalties produced by structural coupling
gains a new meaning in the present situation. This is especially
true for concentrated load transmissions, as it is the case for
mechanically fastened joints [73]. Mechanical fastening is still
one of the main methods currently used for joining composite
components [70], with the advantage of no special surface preparations, easy disassembly and inspection.
Hybrid bolted bonded (HBB) joining technology in lap joints
combines both classical techniques, i.e. bolting and bonding. The
advantages of HBB joints, compared to bolted joints, could be summerized in two main points: (i) continuous instead of discrete load
transfer distribution along the overlap and (ii) decrease of load
transferred by fasteners. Furthermore, in terms of security, the
existence of fasteners in HBB joints could ensure the functioning
of the structure, even if the adhesive layer failure occurs. In manufacturing industries in which numerous bolted and riveted joints
are used, the use of HBB technology could be considered as a potential solution, in order to reduce the mass as well as the manufacturing cost [74].
As stated at Ref. [73], the mechanical behaviour of composite
bolted joints has been extensively studied in the past by means of
experimental, analytical and numerical approaches [75–81], mainly
focusing on the determination of the load capability, the load and
stress distributions and failure criteria of single- and multi-row
bolted joints under the influence of varying laminate configurations
and joint geometries. One of the most effective ways to improve the
load capacity of composite bolted joints entails the local reinforcement of the composite laminate with high-strength metal layers
[82–84], thus clearly improving its bearing and shear capabilities.
The special feature of this reinforcement technique consists of only
embedding the metal layers into the bolted joining area locally,
which is accomplished either by ply-addition (metallic layers are inserted between the composite plies [82]) or ply-substitution techniques (composite plies are replaced by metallic layers [84]).
Hence, the total load capability of this reinforcement approach depends not only directly on the load capability of the bolted joint,
but also on the strength of the transition zone between the pure fibre composite material and the hybridized laminate region.
multiple layers of thin aluminium alloy sheets. A schematic presentation of ARALL is shown in Fig. 1. Combination of high strength
metals (aluminium layers) and strong fibres (aramid layers) generates a new composite material with a unique set of properties.
ARALL laminates offer many advantages such as high strength
and excellent fatigue properties. Moreover, they retain the advantages of aluminium alloys, namely lower cost, easy machining,
forming, and mechanical fastening abilities, as well as substantial
ductility [1,3,85,86].
In the ARALL concept, unidirectional fibre prepregs are used.
This concept is very tolerant of material inconsistencies. The fibre
orientation is chosen into the direction of the main load. The laminates are designed such that the fibres do not fail, when fatigue
cracks develop. That means they remain intact behind the tip of
the propagating crack in the metallic layers. They hinder the opening of the crack, and consequently, the crack tip stress intensity in
the aluminium sheet is reduced. Fig. 4 shows the crack bridging
mechanism of the ARALL laminates.
The fibres are insensitive to the fatigue loading, which key benefit of this material regarding crack propagation. The pregreg
transfers loads over the crack, decreasing the stress intensity at
the crack tip and reducing the crack growth rate. As a result of
crack bridging mechanism, the crack growth life of FMLs is significantly extended compared to monolithic aluminium. Crack bridging is related to the crack opening, the loads are initially mainly
transferred through the aluminium sheets until a certain crack
opening is reached [7,12,87,88]. This behaviour can lead to
improvement in the crack growth rates by a factor of a hundred
and even more, as compared to monolithic aluminium alloy sheet.
This improvement also allows for weight savings up to 30% in fatigue-critical aircraft components. Comparison of ARALL laminates
on a structural level with other aircraft materials show that, ARALL
laminates are very attractive materials, especially for fatigue
dominated structural parts, such as the lower wing skin and the
5. Fibre metal laminates
5.1. Arall
The idea of putting fibre reinforcement in the adhesive bond
lines between aluminium alloys has been researched at many laboratories over the past three decades. These researches made it
possible to develop first fibre metal laminates. Finally in 1978,
ARALL has been developed at the Faculty of Aerospace Engineering
at the Delft University of Technology in Netherland. In 1982 the
first commercial product under the trade name ARALL was
launched by ALCOA.
ARALL laminates are made of high strength aramid fibres
embedded in a structural epoxy adhesive sandwiched between
3677
Fig. 4. Crack bridging mechanism of ARALL laminates.
3678
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
pressurized fuselage cabin of an aircraft. In this way a new hybrid
material ARALL has been obtained. In 1984 two international patents were accepted and a pilot production of four different types of
standardized ARALL was started by the Alcoa Company after sufficient confidence in this material had been gained [7]. Commercialized ARALL laminates are given in Table 5 [89].
All four standard ARALL products employ a thermoset adhesive
system impregnated with unidirectional aramid fibres in a fibre to
resin weight ratio of 50:50. The fibres are oriented parallel to the
aluminium sheet rolling direction. ARALL laminates are produced
under very rigid quality control standards. The aluminium sheets
are anodized and primed according to existing aerospace industry
specifications. Both chromic acid anodizing and phosphoric acid
anodizing treatments are used. The outer aluminium surfaces can
be left bare for later processing after an ARALL part has been
formed. Table 6 summarizes production parameters of ARALL
laminates.
ARALL panels are laid up and cured under the appropriate times
and temperatures either in mechanical press or an autoclave. After
curing, the aluminium layers are in a slight tensile residual stress
state and the pregreg layers in a compensating compressive residual stress state. The ARALL 1 and 3 laminates are given a postcure
0.4% nominal stretch to reverse these residual stress signs. The
ARALL 2 and 4 laminates are not postcured stretched as a standard
practice, although they can be supplied that way [86].
Table 5
Commercially available ARALL laminates [89].
Metal type
Fibre
Metal
thickness layer
(mm)
(mm)
Characteristics
Fibre
direction
(°)
ARALL 1
7075-T6
0.3
0.22
0/0
ARALL 2
2024-T3
0.3
0.22
0/0
ARALL 3
7475-T76
0.3
0.22
0/0
ARALL 4
2024-T8
0.3
0.22
0/0
Fatigue,
strength
Fatigue,
formability
Fatigue,
strength,
exfoliation
Fatigue,
elevated
temperature
Table 6
Production parameters of ARALL laminates [86].
Metal
type
Cure
resin
Cure
temp.
Stretching
Characteristics
ARALL
1
7075T6
AF163-2
120 °C
0.4% permanent
stretch
Superior fatigue
resistance
High strength
ARALL
2
2024T3
AF163-2
120 °C
With or without
0.4% stretch
Excellent fatigue
resistance
Increased
formability
Damage tolerant
ARALL
3
7475T76
AF163-2
120 °C
0.4% permanent
stretch
Superior fatigue
resistance
Controlled
toughness
ARALL
4
2024T8
AF191
175 °C
With or without
0.4% stretch
Excellent fatigue
resistance
Increased
elevated
Temperature
properties
As mentioned before; ARALL has been developed for the lower
wing skin panels of the former Fokker 27 aircraft and the cargo
door of the Boeing C-17. Later with new developments, ARALL
has been used in wide range of fields and applications. They were
used in fuselage and wing structures in aerospace applications and
used as a ballistic material in military applications. However, they
are mostly used as a structural material in cargo doors of aircrafts
[8,10,90].
5.2. Glare
GLARE laminates belong to fibre metal laminates family, they
consist of alternating layers of unidirectional glass fibre reinforced
pregregs and high strength aluminium alloy sheets. At first, they
were developed for aeronautical applications as an improvement
of ARALL with advanced glass fibre and introduced at the Technical
University of Delft in Netherlands in 1990. Later, a partnership between AKZO and ALCOA started to operate in 1991 to produce and
commercialize GLARE [1,7,91,92]. Fig. 5 schematically illustrates a
cross-ply GLARE laminate.
The biggest differentiation of GLARE compared to ARALL is that
GLARE consists of glass fibres instead of aramid fibres. This diversity gives superior properties to GLARE laminates. Table 7 highlights the advantages and disadvantages of various fibres for FMLs.
The specific stiffness and strength in the fibre direction of
GLARE are enhanced over the high strength aluminium alloy used
for the metal layers, which significantly contributes to weight savings in the designs of tension-dominated structural components.
Aforementioned fibre bridging mechanism impedes the growth
and propagation of cracks in the aluminium-alloy layers under tensile fatigue loading conditions [91].
GLARE has a better adhesion between the glass fibres compared
to ARALL. Moreover, glass fibres are much more resistant to compression loading. As a consequence, fibre failure in the glass fibres
has rarely been observed during in fatigue load. Other advantages
of GLARE over ARALL are its higher tensile and compressive
strength, better impact behaviour and better residual strength.
Better adhesion between glass fibre and resin makes GLARE laminates with fibres build up in two directions is possible. This being
is much more suitable for some constructions where biaxial stresses occur. These properties seem to make GLARE has a wider range
of potential application [7].
Nowadays, GLARE materials are commercialized in six different
standard grades (Table 8). They are all based on unidirectional
glass fibres embedded with epoxy adhesive resulting in prepregs
with a nominal fibre volume fraction of 60%. During fabrication
of composites the prepregs are laid-up in different fibre orientations between aluminium alloy sheets, resulting in different standard GLARE grades. For the GLARE 1, GLARE 2, GLARE 4 and
GLARE 5 the composite laminate, i.e. the fibre/resin layer, are
stacked symmetrically. In the case of GLARE 3 composite, the composite laminate have a cross-ply fibre layer stacked to the nearest
outer aluminium layer of the laminate, in relation to the rolling
direction of the aluminium. For the GLARE 6 composite, the composite layers are stacked at +45° and 45°. Table 8 shows these
grades, including the most important material advantages [1].
GLARE is produced with a standard bonding technology. After the
hot-curing (120 °C) cycle in the autoclave, the fibre-metal laminates carry a residual stress system over the thickness of the material, with a small tensile stress in the Al-sheets and compression in
the fibres: this is not favourable for fatigue behaviour. A poststretching operation after the curing cycle can reverses the residual
stress system and solve the problem, thus a tension stress occurs in
fibres and compression in the Al-sheets. But post-stretching causes
extra production cost and it is practically impossible for wide panels [7]. Due to unique combination of properties, GLARE can be ap-
3679
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
Fig. 5. Schematic illustration of a cross-ply GLARE laminates.
Table 7
Identified advantages and disadvantages of various fibres for FMLs [93].
Fibre
Advantage
Disadvantage
Aramid
Outstanding toughness
Weak in bending,
buckling, compression
loading and transverse
tension
Absorb moisture
Excellent fatigue
resistance in bothtensile and flexural
fatigue loading
High Young’s modulus
Conventional
FML
ARALL
Do not form strong
bonds with other
materials such as
composite matrices
Low weight
Glass
High tensile strength
High failure strain
Do not absorb moisture
High weight
GLARE
Low stiffness
plied to wide range of applications. Manufacturers are giving lots
of attention on GLARE. However, it has been mostly used in aerospace applications so far. Recently, GLARE is used in the main fuselage skin and the leading edges of the horizontal and vertical tail
planes of new, high capacity Airbus A380. The latter application
is a consequence of the excellent impact resistance of the FML
concept utilizing the high strength glass fibres with strain rate effect [9,10,86,87]. With excellent impact characteristics, GLARE is
being evaluated for use as cockpit crown, forward bulkheads, the
leading edge and the flame-resistant capability of GLARE makes
it suitable for flame sensitive areas such as; fire walls, cargo-liners,
etc. in aerospace applications [7].
5.3. Carall
Fibre metal laminates consist of alternating layers of aluminium
alloy and adhesive impregnated fibres pregregs. Their superior
properties depend strongly on the type of reinforcing fibres used
in pregregs. Each type has its advantages and limitations. As mentioned before ARALL laminates are based on aramid fibres. Aramid
fibres provide good specific strength and modulus, high impact
resistance and outstanding toughness to ARALL laminates. However, the poor compressive strength is a major limitation for these
hybrid composites. In this point of view, CARALL laminates has
developed as an improvement of ARALL laminates. They contain
different amount of carbon/epoxy pregregs instead of aramid/
epoxy pregregs [93–95]. A schematic illustration of CARALL laminates is shown in Fig. 6.
Compared with aramid/epoxy, carbon/epoxy composites possess higher specific modulus, but relatively low values of specific
strength, strain to failure and impact resistance. In terms of fatigue,
it was recognized that aramid fibre composites have better low cycle fatigue performance but worse high cycle fatigue performance
Table 8
Commercially available GLARE grades [1,89].
Grade
Sub
Metal Type
Metal Thickness (mm)
Fibre Layer (mm)
Pregreg orientation in each fibre layer (°)
Characteristics
GLARE 1
–
7475-T761
0.3–0.4
0.266
0/0
Fatigue, strength, yield stress
GLARE 2
GLARE 2A
GLARE 2B
2024-T3
2024-T3
0.2–0.5
0.2–0.5
0.266
0.266
0/0
90/90
Fatigue, strength
Fatigue, strength
GLARE 3
–
2024-T3
0.2–0.5
0.266
0/90
Fatigue, impact
GLARE 4
GLARE 4A
GLARE 4B
2024-T3
2024-T3
0.2–0.5
0.2–0.5
0.266
0.266
0/90/0
90/0/90
Fatigue, strength, in 0° direction
Fatigue, strength, in 90° direction
GLARE 5
–
2024-T3
0.2–0.5
0.266
0/90/90/0
Impact, Shear, off-axis properties
GLARE 6
GLARE 6A
GLARE 6B
2024-T3
2024-T3
0.2–0.5
0.2–0.5
0.266
0.266
+45/-45
-45/+45
Shear, off-axis properties
Shear, off-axis properties
3680
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
Fig. 6. A schematic illustration of CARALL laminates.
than carbon fibre composites. Moreover, the high stiffness of carbon fibres allows for extremely efficient crack bridging and therefore very low crack growth rates [90,96]. CARALL is produced
similar to ARALL and GLARE laminates. Before the curing process,
aluminium surfaces are treated for an optimal adhesion between
aluminium alloy and epoxy resin. Then, they are cured in hot press.
The combination of high stiffness and strength with good impact
properties gives CARALL laminates a great advantage for space
applications. Other applications for this laminate are impact
absorbers for helicopter struts and aircraft seats [96].
6. Test methods of fibre metal laminate composites
6.1. Bending tests
Mechanical properties of composite materials are governed by
the adhesion between fibre and matrix. Beside this, same properties of FMLs are governed by the interface bond between composite ply and metal ply. Determination of this adhesion is not an easy
task therefore various test methods have been proposed in the literature for this purpose: interlaminar shear and interfacial fracture
tests. These methods give quality control information and are not
suitable for design specifications.
Interlaminar shear tests have been used to determine the interlaminar shear strength (ILSS) of FMLs for a long time [97–105].
According to the literature survey, there are three kinds of test
methods according to the interlaminar shear loading types: I. Compression loading [97], II. Three and five-point bending [98–
101,103,105] and III. Short beam shear loading [102,104]. These
tests will be explained individually.
Hinz et al. [97] investigated interlaminar shear properties of
FMLs by means of a double-notch shear test (DNS) according to
ASTM D3846 at room temperature. The interlaminar shear load between the notches was primarily applied by compression on both
ends of the DNS specimen as illustrated schematically in Fig. 7.
According to author’s advice; for controlled deflection, a linear
guided sliding carriage should be used. In this work a step motor
applied the compression load with a constant carriage speed
0.06 mm/min. It is reported that the load was measured using a
load cell with 500 N load capacities and recorded by means of a
Fig. 7. Schematic view of compression loading on FML DNS specimen.
Kyowa CDV-700A measuring amplifier. The compressive stress is
calculated using the cross section of the specimen. To calculate
the applied interlaminar shear stress the effective area of the shear
plane was used.
The other test methods applied for ILSS investigation of FMLs
are three and five-point bending tests (Fig. 8).
Lawcock et al. determined the ILSS of FMLs by three and fivepoint bending tests with a span of 10 mm at a crosshead speed
of 1.3 mm/min [99]. The authors calculated the ILSS by Eq. (1) for
three-point bending and five-point bending:
ILSS ¼ ð33Pc Þ=ð64WtÞ
ð1Þ
where ‘‘Pc’’ is the initial load for interlaminar failure and ‘‘W’’ and ‘‘t’’
are the width and thickness of the specimens, respectively. In another work; Lawcock et al. calculated ILSS of FMLs from three-point
and five-point bending tests as follows [101]:
—Three-point bending :
—Five-point bending :
ILSS ¼ ð3Pmax Þ=4Wt
ð2Þ
ILSS ¼ ð33Pmax Þ=ð64WtÞ
ð3Þ
where ‘‘Pmax’’ is the failure load and ‘‘W’’ and ‘‘t’’ are the width and
thickness of the specimens, respectively.
Reyes and Cantwell also reported their investigation about ILSS
of FMLs by means of three-point bending test [103]. These tests
were undertaken at crosshead displacement rates between 0.1
and 3 m/s and were stopped once a visible crack had propagated
from one of the starter defects. In all these studies; the authors
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
3681
Fig. 8. Schematic view of three and five-point bending.
were not mentioned about test standards. However; Khalili et al.
[105] reported that three-point bending tests were carried out on
Zwick 1484 by using specifications of ASTM D790 M-93.
The last test method for investigating ILSS of FMLs is short beam
shear test. Park et al. performed ILSS tests to evaluate the extent of
hygrothermal degradations as a function of water absorption and
thermal cycles [102]. ILSS of FMLs was evaluated by the ASTM
D2344 standard specimen through a short beam shear test. They
reported that tests were carried out with a crosshead speed of
1.3 mm/min. Botelho et al. [104] also used the same test standard
(ASTM D2344) for investigating the ILSS of FMLs. Authors reported
that the tests were performed with Instron mechanical testing machine using a test speed of 1.3 mm/min.
The next test method group for investigating the interface
bonding between composite and metal plies is as mentioned above
the interfacial fracture tests. In these tests, the degree of adhesion
between the composite and metal plies was investigated using the
single cantilever beam (SCB) geometry [103,106,107]. This geometry yields mixed-mode I/II loading (tension/shear) conditions at the
crack tip in the sample. It is reported in these papers that SCB specimens with dimensions of 20 mm 200 mm were removed from
the moulded plates. The SCB specimens were clamped in the test
fixture shown in Fig. 9 and tested at crosshead displacement rate
of 2 mm/min.
According to these studies the interfacial fracture energy was
determined using an experimental compliance method where the
interfacial fracture energy, ‘‘Gc’’ is given by:
2
Gc ¼ ð3P 2 ka Þ=2B
ð4Þ
where ‘‘P’’ is the applied force, ‘‘B’’ the specimen width, ‘‘C’’ the
specimen compliance and ‘‘a’’ the crack length. Also it is mentioned
that the specimen compliance was determined as a function of
crack length and a curve fit of the following form applied:
C ¼ C 0 þ ka
3
ð5Þ
where ‘‘C0’’ and ‘‘k’’ are constants for a given specimen. Here, it is
told that the parameter ‘‘k’’ was determined by measuring the gradient of the plot of the compliance ‘‘C’’ versus the cube of the crack
length ‘‘a3’’ whereas it was not necessary to determine the values of
‘‘C0’’.
6.2. Fatigue tests
The fatigue properties of fibre metal laminates have been evaluated in numerous test programs. Generally fatigue tests were
conducted according to the ASTM standards. The shape and dimensions of specimens for fatigue testing were based on ASTM D3479;
the specimen length L = 200 mm, the gauge length LG = 100 mm,
the specimen width W = 25 mm [9,91,108]. Beside this Reyes and
Kanga were conducted ASTM E466 test standard for analyzing fatigue behaviour of FML [109]. Except these papers there are many
studies that fatigue performance of FMLs was not investigated
according to these standards [8,87,89,96,106,110–119]. In these
papers specimens were prepared as notched or cracked for the initiation of crack. For fatigue tests; servo hydraulic test machines
with approximately a load capacity of 5, 10 and 25 metric tons
were used. All constant amplitude fatigue tests were conducted
in tension–tension or tension–compression loading at a frequency
of 10 Hz in lab-air at room temperature [9,87,89,91,108,106,
117,118]. Beside this in some papers fatigue tests were conducted
with sinusoidal waves of frequency 5 Hz [113,114]. During tests;
the fatigue cracks were observed by various ways; charge coupled
device (CCD) camera [110,115], visual observation [111], digital
camera [112], clip gage [114], travelling optical microscope
[106,116,119].
6.3. Tensile tests
The tensile behaviour of FMLs was generally investigated
according to ASTM D3039 [93,99,105,107,120–126]. Beside this
Kawai and Arai [123] were benefit from JIS K7073 standard for conducting tensile testing of FMLs. In all papers dealt with tensile
properties of FMLs, the test specimen shapes and dimensions are
different. Some of them used rectangular tensile specimens
[6,106,121] and some of them used dog bone shape tensile specimens [101,109,118,124]. Beside this; there are a few papers that
tensile test specimen dimensions were prepared according to the
standards. In these papers; the authors were not referenced their
tensile specimen dimensions. For tensile tests; universal and servo
hydraulic testing machines were used: Instron tensile testing machines (4469, 1195, 8502, 4204, 1125, 1251) [96,105,107,109,120–
122,127], MTS tensile testing machines (810, RT/50) [99,121,123–
126] and Shimadzu tensile testing machine [121]. Tensile tests
were performed at room temperature at various crosshead speeds
of 0,1 mm/min [124], 0,5 mm/min [96], 1 mm/min [103,93,109,
120,121,123,127], 1,27 mm/min [126], 2 mm/min [6,106,125,36],
5 mm/min [105]. However; in Carrillo and Cantwell’s studies tensile behaviour of FMLs were investigated at a constant strain rate
of 0.04 min1 on an Instron 4204 universal test machine according
to the ASTM D3039 test standard [107,122]. It is worthy to say that
during tensile tests the Young’s modulus of each specimen was
recorded using a clip-on extensometer which was incapable of
measuring the complete stress–strain curve.
6.4. Low and High velocity impact and blast loading tests
Fig. 9. Schematic illustration of SCB geometry.
Due to the importance of understanding the dynamic loading
effect on the FMLs, low and high velocity and also blast loading
tests should be performed. According to the literature survey, it
is well understood that there are many studies dealt with these
test methods applied to the FMLs. Among these papers some of
3682
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
them explain low velocity impact behaviour of FMLs
[103,106,13,128,129], some of them dealt with high velocity impact performance of FMLs [13,128–131] and the others explain
the blast loading effect on the FMLs [132–137].
a blast; the swing of the pendulum could be used to calculate
the impulse applied to the panel by detonating the explosive.
7. Conclusions
6.4.1. Low velocity impact tests
First of all; the low velocity impact studies shall be mentioned.
In these studies it is seen that all the low velocity impact tests were
performed by drop weight impact tester. Tests were carried out
using drop hammers that have various masses: 2000 g [103,106],
1160 g [128] and 575 g [13,129]. In addition to this, indenter types
were chosen as similar (hemispherical) but the diameters are differing from each other as following: 5 mm [107], 10 mm [106],
12.7 mm [128], 15 mm [13,129]. For the investigation of low velocity impact performance; square FML samples were generally prepared as 75 mm 75 mm, 100 mm 100 mm, 125 mm 125 mm and clamped between two square or circular frames.
Square frames have 75 mm 75mm or 100 mm 100 mm internal openings [106,128]. Beside this circular frames have internal
diameters of 70 mm [107], 75 mm [103] and 80 mm [13]. The impact tests are instrumented as the contact force is measured with a
load cell in the nose of the impactor and the strain in the specimen
is measured with a strain gauge. After an impact, the carriage was
caught in order to avoid secondary impacts.
6.4.2. High velocity impact tests
After an explanation about low velocity impact tests, in this section high velocity impact tests applied to FMLs shall be explained
in detail. According to literature survey, it is very clear that generally a gas gun consisted of a pressure vessel was used for high
velocity impact tests [13,128–131]. Generally square plates were
clamped in a steel support square aperture so that the impactor
can hit the centre of the test specimen. In this test method; after
burning through a membrane of the gas gun, the expanding gas
or air accelerates a projectile or a steel ball bearing to the velocities
ranging between 25 and 100 m/s. The velocity of impactor prior to
impact can be measured as it exited the barrel by the interruption
of two laser beams a known distance apart by using light emitting
diode photovoltaic cell pairs.
The ballistic limit is defined as the lowest velocity for complete
penetration or perforation. Impact testing should be conducted
over a range of impact energies until complete perforation of the
FML target was achieved. The resulting perforation energy was
then could be used to calculate the specific perforation energy of
the target by normalizing the measured perforation energy by
the areal density of the target [131]. After testing, the specimens
could be sectioned, polished and then viewed under an optical
microscope in order to elucidate the failure mechanisms during
impact.
6.4.3. Blast loading tests
Numerous studies [132,134–138] have investigated the localized blast loading of FMLs. Blast loading was created using a plastic
explosive (PE4) shaped into a circular disc and positioned at the
centre of the plate, on top of a polystyrene foam (12–14 mm)
pad to attenuate the blast. PE4 consists of 88% RDX and 12% lithium grease, has a detonation velocity of 8200 m/s and a TNT equivalency of 1.3 [132]. The detonator should be attached to the centre
of the disc using 1 g of explosive, referred to as a ‘leader’. The impulse could be varied by varying the diameter of the explosive
strips. The FML panels 300 mm 300 mm [136], 400 mm 400 mm [134,138] were clamped between two steel frames leaving an exposed area of 200mm 200 mm [132,136], 300 mm 300mm [134,135,138]. The frames were mounted onto a ballistic
pendulum. Each FML panel should be weighted and the thickness
should be measured at four different points before testing. After
In this paper, historical development, advantages, disadvantages, production and applications of FMLs were presented by
examining recent studies. In 1978 initial FMLs, ARALL laminates
which consist alternating thin aluminium layers and aramid fibre
pregreg were introduced. Later GLARE laminates based on high
strength glass fibres and CARALL laminates based on carbon fibres
were developed. GLARE laminates are the successor of ARALL laminates and commercially production of GLARE laminates were
started in 1991. ARALL, GLARE and CARALL laminates are now
being used as structural materials in aircrafts. Yet, it can be said
that GLARE laminates found more use than ARALL and CARALL
laminates in aircraft applications. Both laminates production involves five major activities; (i) surface treatment of metallic surface in order to improve bonding between metallic layer and
fibre reinforced pregreg, (ii) material deposition, (iii) cure preparation, (iv) cure process, (v) post stretching for reserving residual
stress of FML which caused by curing process. FMLs provides superior mechanical properties compared to fibre reinforced composites and aluminium alloys. High strength, high elasticity modulus
with improved toughness and excellent fatigue properties can be
said as major advantages of FMLs. ARALL, GLARE and CARALL laminates explained and compared with each other in following
chapters.
In addition to the importance of reinforcement and matrix in
polymer composites, the bonding between the composite laminae
and the metallic layer is a key issue for the overall metal–fibre laminate performance. An adequate surface treatment of the metallic
layer is required to assure a good mechanical and adhesive bond
between the composite laminates and metal surfaces. In this paper
the effects of surface treatments which are improved the metallic
surface morphology for a better bonding with composite laminates
are investigated. Surface treatment methods, such as mechanical,
chemical, electrochemical, coupling agents and dry surface treatments are introduced and showed comparable performance to improve fibre metal laminates.
Mechanical properties of FMLs are being enhanced by the interface bond between composite and metal plies. This enhancement
could be controlled by various test methods. This control reports
give quality information and are suitable for design specifications.
In this review; test methods of bending, fatigue, tensile, low and
high velocity impact and blast loading tests for determining the
mechanical properties of FMLs and research studies utilized from
these test methods were explained in detail.
8. Developments for the future
FMLs consist of metallic alloy and fibre reinforced prepreg.
Mostly commercially avaible GLARE, ARALL and CARALL consist
various aluminium alloys. Many researches have been trying to
use possible metallic alloys such as magnesium, titanium, etc. instead of aluminium alloys. It is expected that this diversity gives
optimum mechanical properties. Same efforts have been examined
for engineering polymeric materials to replace fibre reinforced
prepreg.
Long processing cycle to cure the polymer matrix in the composite plies increases the cycle time of whole production and decreases productivity. New processing methods are investigated
for improving the productivity of curing process and decreasing
the labour costs of FMLs. These improvements will make FMLs very
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
attractive to various industrial applications (military, automotive
and aircraft).
Last decades many researchers have been investigating the possibility of using thermoplastic materials instead of thermoset
materials as matrix materials in FMLs. By using thermoplastic matrix, FMLs will find new application areas. However low compatibility of thermoplastic matrix with metal surfaces needs to be
improved by surface modification methods explained in this
review.
References
[1] Botelho EC, Silva RA, Pardini LC, Rezende MC. A review on the development
and properties of continuous fiber/epoxy/aluminum hybrid composites for
aircraft structures. Mater Res 2006;9(3):247–56.
[2] Bernhardt S, Ramulu M, Kobayashi AS. Low-velocity impact response
characterization of a hybrid titanium composite laminate. J Eng Mater
Technol 2007;129:220–6.
[3] Villanueva GR, Cantwell WJ. The high velocity impact response of composite
and
FML-reinforced
sandwich
structures.
Compos
Sci
Technol
2004;64:35–54.
[4] Beumler T, Pellenkoft F, Tillich A, Wohlers W, Smart C. Airbus costumer
benefit from fiber metal laminates. Airbus Deutschland GmbH; 2006 May,
Ref. no: L53pr0605135-Issue 1. p. 1–18.
[5] Alderliesten RC, Benedictus R. Fiber/metal composite technology for future
primary aircraft structures. In: 48th Aiaa/Asme/Asce/Ahs/Asc structures,
structural dynamics, and materials conference 15th; April 23–26, 2007;
Honolulu, Hawaii; 2007. p. 1–12.
[6] Cortes P, Cantwell WJ. The prediction of tensile failure in titanium-based
thermoplastic fibre–metal laminates. Compos Sci Technol 2006;66:2306–16.
[7] Asundi A, Choi Alta YN. Fiber metal laminates: an advanced material for
future aircraft. J Mater Process Technol 1997;63:384–94.
[8] Vogelesang LB, Vlot A. Development of fibre metal laminates for advanced. J
Mater Process Technol 2000;103:1–5.
[9] Chang PY, Yeh PC, Yang JM. Fatigue crack initiation in hybrid boron/glass/
aluminum fiber metal laminates. Mater Sci Eng 2008;A 496:273–80.
[10] Alderliesten R. On the development of hybrid material concepts for aircraft
structures. Recent Patents Eng 2009;3:25–38.
[11] Remmers JJC, Borst RD. Delamination buckling of fibre–metal laminates.
Compos Sci Technol 2001;61:2207–13.
[12] Schut J, Alderliesten RC. Delamination growth rate at low and elevated
temperatures in glare. In: 25th International congress of the aeronautical
sciences, 2006 September 3–8; Hamburg, Germany: 2006. p. 1–7.
[13] Vlot A. Impact loading on fibre metal laminates. Int J Impact Eng
1996;18(3):291–307.
[14] Castrodeza EM, Bastian FL, Perez JEI. Critical fracture toughness, Jc and D5c, of
unidirectional fibre–metal laminates. Thin-Walled Struct 2003;4:1089–101.
[15] Vogelesang B, Gunnink JW, Roebroeks GHJJ, Muè Ller RPG. Toward the
supportable and durable aircraft fuselage structure. In: Grandage JM, Jost GS,
editors. Proceedings of the 18th symposium of the international committee
on aeronautical fatigue, 1995 May 3–5, Melbourne, Australia; 1995. p. 257–
272.
[16] Cortes P, Cantwell WJ. Fracture properties of a fiber–metal laminates based
on magnesium alloy. J Mater Sci 2004;39:1081–3.
[17] Roebroeks GHJJ. Fibre–metal laminates recent developments and
applications. Fatigue 1994;16(1):33–42.
[18] Yehia A, Bahei-El-Din A, Amany G, Botrous B. Analysis of progressive fiber
debonding in elastic laminates. Int J Solids Struct 2003;40:7035–53.
[19] Park SY, Choi WJ, Choi HS, Kwon H, Kim SH. Recent trends in surface
treatment technologies for airframe adhesive bonding processing: a review
(1995–2008). J Adhes 2010;86:192–221.
[20] Guo YJ, Wu XR, Zhang ZL. Characterization of delamination growth behaviour
of hybrid bonded laminates. Fatigue Fract Eng Mater Struct
1997;20(12):1699–708.
[21] Davis M, Bond D. Principles and practices of adhesive bonded structural joints
and repairs. Int J Adhes Adhes 1999;19:91–105.
[22] di Scalea FL, Bonomo M, Tuzzeo D. Ultrasonic guided wave inspection of
bonded lap joints: noncontact method and photoelastic visualization. Res
Nondestr Eval 2001;13(3):153–71.
[23] Critchlow GW, Brewis DM. Review of surface pretreatments for aluminium
alloys. Int J Adhes Adhes 1996;16(4):255–75.
[24] Critchlow GW, Yendall KA, Bahrani D, Quinn A, Andrews F. Strategies for the
replacement of chromic acid anodising for the structural bonding of
aluminium alloys. Int J Adhes Adhes 2006;26:419–53.
[25] Sang Park SY, Choi WJ, Choi HS, Kwon H, Kim SH. Effects of surface pretreatment and void content on GLARE laminate process characteristics. J
Mater Process Technol 2010;210:1008–16.
[26] Liu J, Chaudhury MK, Berry DH, Seeberg JE, Osborne JH, Blohowiak KY. Effect
of surface morphology on crack growth at a sol–gel reinforced epoxy/
aluminium interface. J Adhes 2006;82:487–516.
[27] Harris AF, Beevers A. The effects of grit-blasting on surface properties for
adhesion. Int J Adhes Adhes 1999;19:445–52.
3683
[28] Rider AN, Arnott DR. Boiling water and silane pre-treatment of aluminium
alloys for durable adhesive bonding. Int J Adhes Adhes 2000;20:209–20.
[29] Rider AN, Olsson-Jacques CL, Arnott DR. Influence of adherend surface
preparation on bond durability. Surf Interface Anal 1999;27:1055–63.
[30] Prolongo SG, Uren A. Effect of surface pre-treatment on the adhesive strength
of epoxy–aluminium joints. Int J Adhes Adhes 2009;29:23–31.
[31] Lefebvre DR, Ahn BK, Dillard DA, Dillard JG. The effect of surface treatments
on interfacial fatigue crack initiation in aluminum/epoxy bonds. Int J Fract
2002;114:191–202.
[32] Brewis DM, Critchlow GW. Locus of failure of T-peel joints formed between
aluminium and various adhesives. Int J Adhes Adhes 1997;17:33–8.
[33] Rider AN. Factors influencing the durabilityof epoxyadhesion to silane
pretreated aluminium. Int J Adhes Adhes 2006;26:67–78.
[34] Sheasby PG, Pinner R. The surface treatment and finishing of aluminium and
its alloys. 6th ed. Materials Park (OH): ASM International/Stevenage, Herts
(UK): Finishing Publications Ltd.; 2001.
[35] Carrino L, Napolitano G, Sorrentino L. Wettability improving of 2024
aluminium alloy by oxygen cold plasma treatment. Int J Adv Manufact
Technol 2006;31:465–73.
[36] Bjorgum A, Lapique F, Walmsley J, Redford K. Anodising as pre-treatment for
structural bonding. Int J Adhes Adhes 2003;2:401–12.
[37] Bishopp A. Handbook of adhesives and sealants. Amsterdam: Elsevier; 2005.
[38] Kinloch AJ, Little MSG, Watts JF. The role of the interphase in the
environmental failure of adhesive joints. Acta Mater 2000;48:4543–53.
[39] Digby RP, Packham DE. Pretreatment of aluminium: topography, surface
chemistry and adhesive bond durability. Int J Adhes Adhes 1995;15(2):61–71.
[40] Hadavinia H, Kinloch AJ, Little MSG, Taylor AC. The prediction of crack growth
in bonded joints under cyclic-fatigue loading I. Experimental studies. Int J
Adhes Adhes 2003;23:449–61.
[41] Vine K, Cawley P, Kinloch AJ. The correlation of non-destructive
measurements and toughness changes in adhesive joints during
environmental attack. J Adhes 2001;77(2):125–61.
[42] Minford JD. Handbook of aluminum bonding technology and data. New
York: Marcel Dekker; 1993.
[43] Lunder O, Olsen B, Nisancioglu K. Pre-treatment of AA6060 aluminium alloy
for adhesive bonding. Int J Adhes Adhes 2002;22:143–50.
[44] Domingues L, Fernandes JCS, Da Cunha Belo M, Ferreira MGS, Guerra-Rosa L.
Anodising of Al 2024-T3 in a modified sulphuric acid/boric acid bath for
aeronautical applications. Corros Sci 2003;45:149–60.
[45] Mertens T, Kollek H. On the stability and composition of oxide layers on pretreated titanium. Int J Adhes Adhes 2010;30:466–77.
[46] Oosting R. Toward a new durable and environmentally compliant adhesive
bonding process for aluminum alloys. PhD thesis, Delft University of
Technology, The Netherlands; 1995.
[47] Fedel M, Olivier M, Poelman M, Deflorian F, Rossi S, Druart ME. Corrosion
protection properties of silane pre-treated powder coated galvanized steel.
Prog Org Coat 2009;66:118–28.
[48] Zucchi F, Trabanelli G, Grassi V, Frignani A. Proceeding of the EUROCORR.
2001; 30 September–4 October; Riva del Garda, Italy; 2001.
[49] Abel ML, Digby RP, Fletcher IW, Watts JF. Evidence of specific interaction
between c-glycidoxypropyltrimethoxysilane and oxidized aluminium using
high-mass resolution to F-SIMS. Surf Interface Anal 2000;29(2):115–25.
[50] Hobbs PM, Kinloch AJ. The computational molecular modelling of
organosilane primers. J Adhes 1998;66:203–28.
[51] Crook RA, Sinclair JW, Poulter LW, Schulte KJ. An environmentally-friendly
process for bonding aluminum using aqueous metasilicate sol–gel and silane
adhesion promoters. J Adhes 1998;68(3 and 4):315–29.
[52] DeGraevea I, Tourwé EM, Biesemans M, Willem R, Terryn H. Silane solution
stability and film morphology of water-based bis-1,2-triethoxysilyl)ethane
for thin-film deposition on aluminium. Prog Org Coat 2008;63:38–42.
[53] Varma PCR, Colreavy J, Cassidy J, Oubaha M, Duffy B, McDonagh C. Effect of
organic chelates on the performance of hybrid sol–gel coated AA 2024-T3
aluminium alloys. Prog Org Coat 2009;66:406–11.
[54] Critchlow GW, Brewis DM, Emmony DC, Cottam CA. Initial investigation into
the effectiveness of CO2-laser treatment of aluminium for adhesive bonding.
Int J Adhes 1995;15(4):233–6.
[55] Mazza JJ, Gaskin GB, Piero WSD, Blohowiak KY. Sol-gel technology for lowVOC, nonchromated adhesive bonding applications. Report no.: AFRL-MLWP-TR-2004-4063; AFRL Wright-Patterson AFB; Ohio, 2004.
[56] Voast PV, Blohowiak K. Critical materials & processes bonded joint issues. In:
2004, FAA Workshop on Bonded Structures 2004 June 16–18; Seattle, WA;
2004.
[57] Osborne JH, Blohowiak KY, Sekits DF. Environmentally benign sol-gel surface
treatment for aluminum bonding applications. Report no.: WL-TR-95-4060;
AFRL Wright-Patterson AFB, Ohio; 1996.
[58] Critchlow GW, Cottam CA, Brewis DM, Emmony DC. Further studies into the
effectiveness of CO2-laser treatment of metals for adhesive bonding. Int J
Adhes Adhes 1997;17(2):143–50.
[59] Walters CT. Laser surface preparation for adhesive bonding II. Report no.
AFRL-ML-WP-TR-2006-4139; AFRL Wright-Patterson AFB, Ohio; 2004.
[60] Spadaro C, Sunseri C, Dispenza C. The influence of the nature of the surface
oxide on the adhesive fracture energy of aluminium bonded joints as
measured by T-peel tests. Int J Adhes Adhes 2007;76:1441–6.
[61] Wong RCP, Hoult AP, Kim JK, Yu TX. Improvement of adhesive bonding in
aluminium alloys using a laser surface texturing process. J Mater Process
Technol 1997;63:579–84.
3684
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
[62] Mazza JJ. Advanced surface preparation for metal alloys. In: 83rd Meeting of
the AGARD Structures and Materials Panel; 1996 September 2–6; Florence,
Italy; 1996. p. 10-1–10-12.
[63] Davis GD, Groff GB, Zatorski RA. Plasma spray coatings as treatments for
aluminum, titanium and steel adherends. Surf Interface Anal
1997;25(5):366–73.
[64] Fernandes JCS, Ferreira MGS, Haddow DB, Goruppa A, Short R, Dixon DG.
Plasma-polymerised coatings used as pre-treatments for aluminium alloys.
Surf Coat Technol 2002;154:8–13.
[65] Alexander MR, Zhou X, Thompson GE, Duc TM, McAlpine E, Tielsch BJ.
Functionalized plasma polymer coatings for improved durability of
aluminium–epoxy adhesive joints: fractography. Surf Interface Anal
2000;30:16–20.
[66] Anagreh N, Dorn L, Bilke-Krause C. Low-pressure plasma pretreatment of
polyphenylene sulfide (PPS) surfaces for adhesive bonding. Int J Adhes Adhes
2007;28(1–2):16–22.
[67] Polini W, Sorrentino L. Improving the wettability of aluminium alloy 2024 by
means of cold plasma. Appl Surf Sci 2003;214:232–42.
[68] Koch GH, Todd GL, Deutchman A, Partyka R. Nonchemical surface treatment
for aluminum alloys. Report no.: WL-TR-96-4084; AFRL Wright-Patterson
AFB, Ohio; 1996.
[69] De Iorio I, Leone C, Nele L, Tagliaferri V. Plasma treatments of polymeric
materials and Al alloy for adhesive bonding. J Mater Process Technol
1997;68:179–83.
[70] Niu MCY. Composite airframe structures, 2nd published. Hong
Kong: Conmilit Press Ltd.; 1996.
[71] Kupke M, Kolax M. CFRP-fuselage – ensuring future competitiveness. In:
Material & process technology – the driver for tomorrow’s improved
performance. Proceeding of the 25th jubilee international SAMPE Europe
conference 2004 of the society for the advancement of materials and process
engineering Paris EXPO; 2004 March 30–April 1, Porte de Versailles, Paris;
2004. p. 432–7.
[72] Jackson P. Airbus A380, fixed-wing civil. In: Jane’s all the world’s aircraft
2005–2006, 96th ed. Coulsdon, Surrey CRS 2YH, UK; 2005 p. 241–6.
[73] Kolesnikov B, Herbeck L, Fink A. CFRP/titanium hybrid material for improving
composite bolted joints. Compos Struct 2008;83:368–80.
[74] Hoang-Ngoc CT, Paroissien E. Simulation of single-lap bonded and hybrid
(bolted/bonded) joints with flexible adhesive. Int J Adhes Adhes
2010;30:117–29.
[75] Hollmann K. Failure analysis of bolted composite joints exhibiting in-plane
failure modes. J Compos Mater 1996;30(3):358–83.
[76] Hung CL, Chang FK. Strength envelope of bolted composite joints under
bypass loads. J Compos Mater 1996;30(13):1402–35.
[77] Oh JH, Kim YG, Lee DG. Optimum bolted joints for hybrid composite
materials. Compos Struct 1997;38(1–4):329–41.
[78] Camanho PP, Matthews FL. Delamination onset prediction in mechanically
fastened joints in composite laminates. J Compos Mater 1999;33(10):906–27.
[79] Dano ML, Gendron G, Picard A. Stress and failure analysis of mechanically
fastened joints in composite laminates. Compos Struct 2000;50:287–96.
[80] Starikov R, Schön J. Quasi-static behaviour of composite joints with
protruding-head bolts. Compos Struct 2001;51:411–25.
[81] Lawlor VP, McCarthy MA, Stanley WF. An experimental study of bolt-hole
clearance effects in double-lap, multi-bolt composite joints. Compos Struct
2005;71:176–90.
[82] Nadler MA, Yoshino SY, Darms FJ. Boron/epoxy support strut for non-integral
cryogenic tankage. In: Materials and processes, 15th SAMPE symposium, vol.
15, Los Angeles; 1969. p. 179–207.
[83] Gunjaev GM, Shelesina GF, Iltshenko SI. Metal laminates with aluminium and
titanium alloys. Aviation materials and technologies. Moscow: VIAM; 2002. p.
50–8 [in Russian].
[84] Kolesnikov B, Herbeck L, Fink A. Fortschrittliche Verbindungstechniken von
Faserverbunden. In: DGLR-Kongress, Dresden, Band II, 20–23 September,
2004. p. 1419–28.
[85] Sun CT, Dicken A, Wut HF. Characterization of impact damage in ARALL
laminates. Compos Sci Technol 1993;49:139–44.
[86] Lee SM, editor. Handbook of composite reinforcements. Wiley VCH Publishers
Inc.; 1993. pp. 1–6.
[87] Shima DJ, Alderliesten RC, Spearing SM, Burianek DA. Fatigue crack growth
prediction in GLARE hybrid laminates. Compos Sci Technol 2003;63:1759–67.
[88] Burianek DA, Giannakopoulos AE, Spearing SM. Modeling of facesheet crack
growth in titanium–graphite hybrid laminates, part I. Eng Fract Mech
2003;70:775–98.
[89] Khan SU, Alderliesten RC, Benedictus R. Post-stretching induced stress
redistribution in fibre metal laminates for increased fatigue crack growth
resistance. Compos Sci Technol 2009;69:396–405.
[90] Vlot A. Fibre metal laminates: an introduction. Kluwer Academic Publishers;
2001.
[91] Kawai M, Hachinohe A. Two-stress level fatigue of unidirectional fiber–metal
hybrid composite: glare 2. Int J Fatigue 2002;24:567–80.
[92] Castrodeza EM, Abdala MRWS, Bastian FL. Crack resistance curves of glare
laminates by elastic compliance. Eng Fract Mech 2006;73:2292–303.
[93] Moussavi-Torshizi SE, Dariushic S, Sadighic M, Safarpour P. A study on tensile
properties of a novel fiber/metal laminates. Mater Sci Eng A
2010;527:4920–5.
[94] Lin CT, Kao PW. Fatigue delamination growth in carbon fibre-reinforced
aluminium laminates. Composites A 1996;27A:9–15.
[95] Xia Y, Wang Y, Zhou Y, Jeelani S. Effect of strain rate on tensile behavior of
carbon fiber reinforced aluminum laminates. Mater Lett 2007;61:213–5.
[96] Lin CT, Kao PW, Yang FS. Fatigue behaviour of carbon fibre-reinforced
aluminium laminates. Composites 1991;22(2):135–41.
[97] Hinz S, Omoori T, Hojo M, Schulte K. Damage characterisation of fibre metal
laminates under interlaminar shear load. Composites A 2009;40:925–31.
[98] Cepeda-Jiménez CM, Alderliesten RC, Ruano OA, Carreño F. Damage tolerance
assessment by bend and shear tests of two multilayer composites: glass fibre
reinforced metal laminate and aluminium roll-bonded laminate. Compos Sci
Technol 2009;69:343–8.
[99] Lawcock GD, Ye L, Mai YW, Sun CT. Effects of fibre/matrix adhesion on
carbon-fibre-reinforced metal laminates I. Residual strength. Compos Sci
Technol 1997;57:1609–19.
[100] Afaghi-Khatibi A, Lawcock G, Ye L, Mai YW. On the fracture mechanical
behaviour of fibre reinforced metal laminates (FRMLs). Comput Methods
Appl Mech Eng 2000;185:173–90.
[101] Lawcock G, Ye L, Mai YW, Sun CT. The effect of adhesive bonding between
aluminum and composite prepreg on the mechanical properties of carbonfiber reinforced metal laminates. Compos Sci Technol 1997;1997:35–45.
[102] Park SY, Choi WJ, Choi HS. The effects of void contents on the long-term
hygrothermal behaviors of glass/epoxy and GLARE laminates. Compos Struct
2010;92:18–24.
[103] Reyes VG, Cantwell WJ. The mechanical properties of fibre-metal laminates
based on glass fibre reinforced polypropylene. Compos Sci Technol
2000;60:1085–94.
[104] Botelho EC, Rezende MC, Pardini LC. Hygrothermal effects evaluation using
the Iosipescu shear test for glare laminates. J Braz Soc Mech Sci Eng
2008;30(3):213.
[105] Khalili SMR, Mittal RK, Kalibar GS. A study of the mechanical properties of
steel/aluminium/GRP laminates. Mater Sci Eng A 2005;412:137–40.
[106] Cortés P, Cantwell WJ. The fracture properties of a fibre–metal laminate
based on magnesium alloy. Composites B 2006;37:163–70.
[107] Carrillo JG, Cantwell WJ. Mechanical properties of a novel fiber–metal
laminate based on a polypropylene composite. Mech Mater 2009;41:828–38.
[108] Kawai M, Kato K. Effects of R-ratio on the off-axis fatigue behavior of
unidirectional hybrid GFRP/Al laminates at room temperature. Int J Fatigue
2006:1226–38.
[109] Reyes G, Kang H. Mechanical behavior of lightweight thermoplastic fiber–
metal laminates. J Mater Process Technol 2007;186:284–90.
[110] Khan SU, Alderliesten RC, Benedictus R. Delamination growth in fibre metal
laminates under variable amplitude loading. Compos Sci Technol
2009;69:2604–15.
[111] Homan JJ. Fatigue initiation in fibre metal laminates. Int J Fatigue
2006;28:366–74.
[112] Rodi R, Alderliesten R, Benedictus R. Experimental characterization of the
crack-tip-opening angle in fibre metal laminates. Eng Fract Mech
2010;77:1012–24.
[113] Takamatsu T, Matsumura T, Ogura N, Shimokawa T, Kakuta Y. Fatigue crack
growth properties of a GLARE3-5/4 fiber/metal laminate. Eng Fract Mech
1999;63:253–72.
[114] Takamatsu T, Shimokawa T, Matsumura T, Miyoshi Y, Tanabe Y. Evaluation of
fatigue crack growth behavior of GLARE3 fiber/metal laminates using a
compliance method. Eng Fract Mech 2003;70:2603–16.
[115] Alderliesten R, Rans C. The meaning of threshold fatigue in fibre metal
laminates. Int J Fatigue 2009;31:213–22.
[116] Kim CW, Oh DJ. Progressive delamination with and without crack
propagation in aramid fiber reinforced metal laminates containing a
circular notch. Mater Sci Eng A 2008;483–484:251–3.
[117] Austin TSP, Singh MM, Gregson PJ, Powell PM. Characterisation of fatigue
crack growth and related damage mechanisms in FRP–metal hybrid
laminates. Compos Sci Technol 2008;68:1399–412.
[118] Johnson WS, Hammond MW. Crack growth behavior of internal titanium
plies of a fiber metal laminate. Composites Part A 2008;39:1705–15.
[119] Lin CT, Kao PW. Effect of fiber bridging on the fatigue crack propagation in
carbon fiber-reinforced aluminum laminates. Mater Sci Eng A
1995;190:65–73.
[120] Wu G, Tan Y, Yang JM. Evaluation of residual strength of notched fiber metal
laminates. Mater Sci Eng A 2007;457:338–49.
[121] Ye L, Afaghi-Khatibi A, Lawcock G, Mai YW. Effect of fibre/matrix adhesion on
residual strength of notched composite laminates. Composites Part A
1998;29A:1525–33.
[122] Carrillo JG, Cantwell WJ. Scaling effects in the tensile behavior of fiber-metal
laminates. Compos Sci Technol 2007;67:1684–93.
[123] Kawai M, Arai Y. Off-axis notched strength of fiber–metal laminates and a
formula for predicting anisotropic size effect. Composites Part A
2009;40:1900–10.
[124] de Vries TJ, Vlot A, Hashagen F. Delamination behavior of spliced fiber metal
laminates part 1. Experimental results. Compos Struct 1999;46:131–45.
[125] Iaccarino P, Langella A, Caprino G. A simplified model to predict the tensile
and shear stress–strain behaviour of fibreglass/aluminium laminates.
Compos Sci Technol 2007;67:1784–93.
[126] Botelho EC, Almeida RS, Pardini LC, Rezende MC. Elastic properties of
hygrothermally conditioned glare laminate. Int J Eng Sci 2007;45:163–72.
[127] Caprino G, Squillace A, Giorleo G, Nele L, Rossi L. Pin and bolt bearing
strength of fibreglass/aluminium laminates. Composites Part A
2005;36:1307–15.
T. Sinmazçelik et al. / Materials and Design 32 (2011) 3671–3685
[128] Lawcock GD, Ye L, Mai YW, Sun CT. Effects of fibre/matrix adhesion on
carbon-fibre-reinforced metal laminates – II. Impact behaviour. Compos Sci
Technol 1997;57:1621–8.
[129] Vlot A. Impact properties of fibre metal laminates. Compos Eng
1993;3(10):911–27.
[130] Michelle S, Fatt H, Lin C, Revilock Jr DM, Hopkins DA. Ballistic impact of
GLARE fiber–metal laminates. Compos Struct 2003;61:73–88.
[131] Abdullah MR, Cantwell WJ. The impact resistance of polypropylene-based
fibre–metal laminates. Compos Sci Technol 2006;66:1682–93.
[132] Langdon GS, Nurick GN, Cantwell WJ. The response of fibre metal laminate
panels subjected to uniformly distributed blast loading. Eur J Mech A/Solids
2008;27:107–15.
[133] Karagiozova D, Langdon GS, Nurick GN, Yuen KCS. Simulation of the response
of fibre–metal laminates to localised blast loading. Int J Impact Eng
2010;37:766–82.
3685
[134] Langdon GS, Lemanski SL, Nurick GN, Simmons MC, Cantwell WJ, Schleyer
GK. Behaviour of fibre–metal laminates subjected to localised blast loading:
part I—experimental observations. Int J Impact Eng 2007;34:1202–22.
[135] Lemanski SL, Nurick GN, Langdon GS, Simmons MS, Cantwell WJ, Schleyer
GK. Understanding the behaviour of fibre metal laminates subjected to
localised blast loading. Compos Struct 2006;76:82–7.
[136] Langdon GS, Cantwell WJ, Nurick GN. Localised blast loading of fibre–
metal laminates with a polyamide matrix. Composites Part B
2007;38:902–13.
[137] Langdon GS, Cantwell WJ, Nurick GN. The blast response of novel
thermoplastic-based fibre–metal laminates – some preliminary results and
observations. Compos Sci Technol 2005;65:861–72.
[138] Lemanski SL, Nurick GN, Langdon GS, Simmons MC, Cantwell WJ, Schleyer
GK. Behaviour of fibre metal laminates subjected to localised blast loading—
part II: quantitative analysis. Int J Impact Eng 2007;34:1223–45.