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Quasi-static and Fatigue Behaviour of Composite Bolted Joints Roman Starikov Department of Aeronautics Royal Institute of Technology SE-100 44 Stockholm, Sweden

Report 2001-9 ISSN 0280-4646

Quasi-static and Fatigue Behaviour of Composite Bolted Joints

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Preface This doctoral thesis is a result of my research activities over the last four years. The work was carried out between 1997 and 2001 at the Department of Aeronautics at KTH (the Royal Institute of Technology) and the Structures Department at FFA (the Swedish Defence Research Agency). My work on the project was supported by Saab Military Aircraft (SMA) and the Material Defence Agency (FMV). This support is very much appreciated. First of all, I especially acknowledge my supervisor, Dr Joakim Schön, to whom I am indebted for his help, great support, encouragement, and guidance over these years. Thank you, Joakim! I would like to thank all my colleagues who have greatly contributed to my research by providing an inspiring, creative and friendly environment. Especially I wish to thank the laboratory staff for their assistance with the testing equipment. I also take this opportunity to express my gratitude to Professor Jan Bäcklund for accepting me as a graduate student and to Professor Anders Blom who allowed me to work within the project. I am thankful to a range of my friends at the Department of Aeronautics, Andrey Shipsha, Lance McGarva, Gordon Kelly, Steven Ribeiro-Ayeh, and Per Wennhage, as well as to numerous friends throughout the world for being an inexhaustible source of support, inspiration, fun, and enjoyable atmosphere over the last years. A special note of gratitude is extended to my parents for their countless sacrifices for the sake of my education. Finally, I express my profound gratitude to my wife Katya for her support, devotion, and patience throughout these years.

Stockholm, April 2001 Roman Starikov


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Abstract Since modern aircraft composite structures often are fastened by means of bolted joints, the structural integrity of such connections has to be validated according to the certification requirements. The behaviour of full-scale composite structures can be represented on sub-structural elements including mechanically fastened joints. The objective of the dissertation is to improve the understanding of the quasi-static and fatigue behaviour of bolted joints with different configurations. In the experimental procedure composite joints with single and multi-row, single and double overlap configurations were tested. The influences of protruding and countersunk-bolt heads, fastener system, quasi-isotropic and highly orthotropic lay-ups on the joint strength and fatigue performance were studied. Strain gauge measurements were carried out in order to measure strain distribution and to calculate load transfer between bolt rows on specimens tested in quasi-static and fatigue. An experimental method of bolt-movement measurement was used to analyse the local behaviour of the bolt and the surrounding composite during loading. Fractographic SEM and optical microscopy were employed to investigate fracture and damage developed in the joint system during static and fatigue testing. As the obtained test results show, increasing the number of bolt rows to the joint configuration, improves its ability to support static and fatigue loading. A linear dependence was found between the number of bolt rows and the fatigue life of those joints, which failure mode was due to fatigue fracture of bolts. The fatigue resistance of single and double lap joints was similar. Among different fastener systems, specimens joined by protruding-head bolts displayed better static and fatigue performance than those bolted by countersunk fasteners. When comparing applied strain, joints with a quasi-isotropic lay-up yielded higher static and fatigue strength than bolted laminates with a 0◦ -dominated lay-up. Load-transfer measurements of bolt rows showed that the distribution of applied load on the bolt rows was not the same. The outer bolt rows appeared to be transferring a slightly larger amount of the applied load than the inner bolt row. Strain distribution between bolt rows was found to be greatly affected by the bolt presence and its behaviour during static and fatigue testing. The bolt behaviour was affected by reduction in pre-stress in the bolt, friction forces between the composite parts, and the level of damage sustained by the fastener and composite. Fractographic observation showed that bolted joints subjected to fatigue loading experienced severe damage around the bolt holes. However, the degree of damage developed at different bolt holes varied since the individual bolts were exposed to different load levels during testing. The joint configuration, the level of pre-stress in the fasteners, load transfer of the applied load between the joint parts, and the extent of damage developed during loading are very important issues which should be considered during the design procedure of composite mechanically fastened joints.


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Dissertation The present thesis includes an introduction to the area of research and five appended research papers: Paper A Starikov R., and Schön J. “Quasi-static behaviour of composite joints with protruding-head bolts” Composite Structures 2001;51(4):411–425. Paper B Starikov R., and Schön J. “Quasi-static behaviour of composite joints with countersunk composite and metal fasteners” Accepted for publication in Composites Part B: Engineering. Paper C Starikov R., and Schön J. “Fatigue resistance of composite joints with countersunk composite and metal fasteners” Submitted to International Journal of Fatigue. Paper D Starikov R., and Schön J. “Experimental study on fatigue resistance of composite joints with protruding-head bolts” Submitted to Composite Structures. Paper E Starikov R., and Schön J. “Local fatigue behaviour of CFRP bolted joints” Submitted to Composite Science and Technology.


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Division of Work Between Authors Paper A Starikov performed the quasi-static tests, experimental measurements, and fractographic observations. Starikov and Schön wrote the paper. Paper B Starikov carried out the quasi-static tests, experimental measurements, and fractographic observations. Starikov and Schön wrote the paper. Paper C The fatigue tests and experimental measurements were done by Starikov. Starikov and Schön wrote the paper. Paper D Starikov performed the fatigue tests and experimental measurements. Starikov and Schön wrote the paper. Paper E The measurements and fractographic work were carried out by Starikov. Starikov and Schön wrote the paper.

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Contents Preface

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Abstract

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Dissertation

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Division of Work Between Authors

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Introduction

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Background

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Fatigue tolerance of aircraft composite structures

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Mechanical joints in aircraft composite structures Macroscopic failure modes in bolted joints . . . . . . Microscopic damages in bolted joints . . . . . . . . . Design parameters . . . . . . . . . . . . . . . . . . . Material parameters . . . . . . . . . . . . . . . . . . Fastening parameters . . . . . . . . . . . . . . . . .

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Future Outlook

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References

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Paper A

A1–A21

Paper B

B1–B21

Paper C

C1–C16

Paper D

D1–D18

Paper E

E1–E20


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Introduction Background As the last decades have shown, the use of advanced composites in primary and secondary aircraft structures is steadily growing. The increased application of composites in the aircraft industry is motivated as these materials offer a number of advantages compared to conventional engineering materials such as steel and aluminium. Firstly, a good combination of low density with high stiffness and strength of advanced composites is to be mentioned. Through this the structural weight can be reduced, and as a result this enables aircraft to fly farther and faster. That leads to more economical commercial flying by introducing larger aircraft, payload, and by fuel savings. Advanced composites are recognized as uniquely flexible in design capabilities and ease in manufacturing. Using advanced manufacturing techniques it is possible to produce large identical composite structures with complex shapes, thus providing a high degree of parts integration and lowering manufacturing and assembling cost (see Fig. 1).

Figure 1: The upper composite wing skin for the Joint Strike Fighter X32A concept demonstrator in the progress of manufacturing. Photograph courtesy of the Boeing Company, USA.

Large composite parts can further be fastened by means of bolted joints to aid replacement in the event of damage. Therefore, this eliminates necessity for large number of small components to be manually adjusted to fit with each other if metal parts would be used. In addition, the ability to manufacture complexly shaped structures makes it possible, on one hand, to improve the aerodynamic and aircraft performance through achieving very smooth and accurate profiles. On the other hand, aircraft stealth properties can be enhanced, which is very important for military aircraft. Another significant advantage is that aircraft composites are known to possess better fatigue resistance than light metals. Since they are mostly used in the form of laminates, the propagation of initial damage, even that occurred very early in

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the fatigue life, is to be arrested by the internal composite structure. Whereas, in metals initiated damage (in most cases it is in the form of cracks) is generally to be considered unsafe since it leads to a rapid growth following by catastrophic failure. Today examples of aircraft composite components include complete vertical and horizontal stabilizer, wing skins and substructures, control surfaces, access doors, weaponry, and external fuel tanks [1, 2]. For instance, such application of composites resulted in 20 and 30 percent of the structural weight being advanced composites for the F-22 Advanced Tactical Fighter and the multi-role combat Saab Gripen aircraft, respectively [2]. Although advanced composites are used widely in the aircraft industry, more significant amounts of their application are limited by expensive raw materials and that automated lay-up techniques are still used to a limited degree. This eventually results in high production cost.

Fatigue tolerance of aircraft composite structures The aircraft industry has to cope with the highest safety in operation as well as economic efficiency today. The aim of modern design philosophy is to combine these contrasts in the best perfection giving the highest priority to safety of people. In the aircraft industry two design approaches have been used in order to build safe aircraft structures, that is safe-life and damage-tolerance concepts [3]. According to the safe-life concept, a structure is designed to have a minimum life during which it is known that no catastrophic failure will occur. The damage-tolerance design means that a damage tolerant structure must withstand all expected stresses in the case of damage until the flaw is detected at a regular scheduled inspection. Generally, it is more economical to design some parts of the structure to be damage tolerant than to have a long safe life since such components can be lighter. Although composites are widely used in the aircraft industry, they are recognized as a comparatively recent technology having been introduced in the 1970s. Current design approach for fatigue tolerance of composite structures is based on the safe-life concept [4]. In comparison with aluminium, which has been used since the beginning of the last century, the well-established certification procedures for light-metals have to be modified for composites taking into account their anisotropy and sensitivity to defects. Therefore, prior to the application to real aircraft structures, damage of composites, their general applicability, and validity are studied at different levels involving coupons, sub-elements, sub-structures, and finally full-scale structures [5– 9]. This design method is known as “building-block approach” (see Fig. 2).

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full-scale structures

preproduction and sertification tests

1 10 100

sub-structures

cupons

structural design development and verification material selection, manufacturing processes, and material properties

1000

Number of tests

components

sertification tests

4

10

Figure 2: Structural building block test approach. The given numbers based on data from Ref. [10].

The building block approach involves a step-by-step validation from a large number of specimens tested at coupon levels to a limited number of sub-structures and full-scale structures tested at structural levels (see Fig. 3).

Figure 3: The Boeing X-32A concept demonstrator is being in structural tests to confirm the integrity of the aircraft’s structural design by simulating flight and landing conditions. Photograph courtesy of the Boeing Company, USA.

Validation tests on full-scale structures are performed to simulate more accurately and to study secondary effects due to damage [6, 9]. The concept of the building-block approach is based on coupons and sub-elements can accurately represent the damage resistance and the behaviour of full-scale structure [6]. Through this the total number of expensive full-scale structures can be reduced. Therefore, a good understanding of the behaviour of composite sub-components and sub-structures, which include a large number of mechanical joints [10], shall be achieved.

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Mechanical joints in aircraft composite structures In aircraft composite components are usually relatively thick. Such structural parts are design to be able to withstand a high level of structural loads. Therefore, in order to provide an efficient aircraft design the thick structural parts are joined by means of mechanically fastened joints. Although there are many different joint configurations in the aircraft industry, their application is driven by service requirements applied to particular structures to be joined. As has already been mentioned, the use of bolted joints lets the connected structural components to be further disassembled, thus increasing the degree of components integration. In addition, component disassembly makes it possible to gain access to the interior of the structure for inspection and repair purposes. The functioning principle of bolted joints is based on micro- and macroscopic mechanical interference, such as friction between the joined parts, shear or tensile transfer forces in fasteners, and contact forces between the joined components. Through this dissimilar materials can be fastened by means of mechanical joints. This feature is used intensively in the aircraft industry to join titanium or aluminium components with composite structures. Figure 4 shows the first wing structure for the first F-22 Advanced Tactical Fighter.

Figure 4: The wing component for the F-22 Advanced Tactical Fighter. Photograph courtesy of the Boeing Company, USA.

The upper composite wing skin is attached by mechanical fastening to the internal wing sub-structure in the form of composite and titanium spars. Despite the above mentioned advantages, mechanical fastened joints possess several disadvantages. The major joints imperfection is based on a concentration of high stresses at the bolt hole induced by so-called “notch effect” due the hole presence. Secondly, aluminium and stain-less steel fasteners result in potential for galvanic corrosion being installed in CFRP-laminates. Hole generation requires specific drilling techniques taking into account the possibility of mechanically and thermally induced defects [2, 11]. Finally, thousands of fasteners installed to join aircraft structural components result in a large weight penalty.


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Macroscopic failure modes in bolted joints The term “macroscopic failure” usually implies a damage state of a specific structure, when it is not able to carry any additional increase in the applied load or to continue realizing the intended tasks. For some applications, this means catastrophic breakdown of the component, whereas in other instances failure may involve, for example, the unacceptable loss of stiffness. Figure 5 shows basic failure modes which may occur in composite bolted joints.

(a)

(b)

(c)

(d)

Figure 5: Basic failure modes in composite bolted joints: (a) net-section failure; (b) bearing failure; (c) shear-out failure; (d) bolt failure.

Net-section mode shown in Fig. 5(a), is caused by tangential or compressive stresses at the hole edge. This mode primarily occurs when the ratio of bolt hole diameter, D (see Fig. 6), to plate width, w, is high, or when the ratio of by-pass load (i.e., the load applied to the plate) to bearing load (i.e., the load which passes through the bolt) is high. Bearing failure mode shown in Fig. 5(b) is governed by compressive stresses acting on the hole surface. This failure occurs when the ratio D/w is low or when the ratio of bearing load to by-pass one is high. Shear-out failure mode illustrated in Fig. 5(c) is caused by shear stresses acting in shear-out planes on the hole boundary in the principal load direction. This mode primarily occurs when the end distance, e (see Fig. 6), is short. Shear-out mode is quite common for highly orthotropic laminates, where this failure mode is independent of the end distance.

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t

2λ

P

D d

w e Figure 6: Geometry of a filled hole.

Bolt failure mode shown in Fig. 5(d) is a result of high shear stresses acting in the bolt shank. This failure is governed by the properties of the bolt to support shear loading and depends on the plate and bolt geometry. Since the bearing failure mode usually develops slowly, giving plenty of warning before the ultimate failure occurs, it is recognized as a desirable failure mode in bolted laminates. The other failure modes, i.e., net section and shear out, usually occur suddenly and are therefore being catastrophic. The interaction between the composite joint’s behaviour during quasi-static loading and a specific failure mode is investigated on specimens with different configuration and fastener system. The achieved results are presented in Papers A and B included in this thesis. A single-bolt configuration of bolted laminates appears to have been most widely used in both analytical and experimental approaches. However, a multi-row solution is most frequently used in reality. Hart-Smith [12–14] has studied a relationship between the geometrical ratio, D/w, and which specific failure mode that occurred. This relationship was further developed in the form of design charts for bolted joints having fasteners with single and multi-row configuration. Although these charts already exist and are available for the designers, they are based on static test data. Therefore, a similar design approach for fatigue design of single and multi-row bolted joints would be of interest. In order to try to fill this niche, an experimental study on the fatigue resistance of composite bolted joints with single and multi-row configuration has been carried out. The obtained results and achieved conclusions are included in Paper D of the present thesis. Microscopic damages in bolted joints For safety reasons macroscopic failures are never allowed to take place in aircraft during their service life. A macroscopic failure is the final stage in the process of damage development, involving damage initiation and its propagation to the final rupture. The damage initiation occurs on a microscopic level. This level means those damages, which are not or barely seen with a naked eye. In order to inspect their presence various non-destructive techniques (NDT) are usually used [15]. However,


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the exact location and the extent of damages are better to be observed with fractographic methods [16]. There are a few basic types of damage which may occur in bolted composite laminates. They include fracture of fibres, fibre pull-out, matrix microcracking, and debonding [17]. Delaminations between plies and cracks within plies may also occur in laminated composites. It has been observed that the failure process in bolted laminates begins far before the final fracture will occur [18]. The progress of failure began with matrix cracking occurring at 25% of the quasi-static failure load. As the applied load was increased up to 35% of the ultimate failure load, fibre fracture in the form of bands induced by matrix cracking, which destabilized the fibres, was noticed in the outer layers. At 70% of the failure load, delaminations being extended from the hole edge between the outer layers and between layers in the middle of the laminate’s thickness were detected. In addition, the damage process started to develop far below the load level at which non-linearities in the loading curves were seen, which would indicate a loss in the stiffness [18]. Hole generation in laminated composites can be a source of damage as well. Delaminations, chip-outs of fibres/matrix, and matrix overheating are damages to be expected if the drilling would not have been properly done. It has been reported that among composite laminates with holes generated using three different drilling methods, those ones yielded better strength and fatigue resistance, which were less affected by the presence of hole machining defects [11]. Composite laminates with filled holes subjected to fatigue loading have a different pattern of damage development. In addition to the above mentioned, several other types of fatigue damage around bolt holes may develop during fatigue testing: hole wear, damage of the bolted laminates in the contact surfaces with the bolt, the initiation and growth of delamination around holes, and shear damage of several plies [19, 20]. The hole wear is a result of the erosion and re-deposition processes due to the presence of friction between the composites parts (see “worn areas” in Fig. 7).

worn area

hole edge hole edge

worn area Figure 7: Fatigue damage of the hole edges and the adjacent composite.

In bolted joints applied loads are transferred via the fasteners and some part of the applied load is carried by the friction forces between the joint elements. When the

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fasteners are loaded, the load distribution along their length is not uniform [18, 21]. This will result in high contact forces localized at the top and bottom hole edges (see Fig. 8). F/2 F

F/2 friction forces for the outer plates

contact forces between bolts and composite

Figure 8: Bolt bending induces high localized contact forces in the adjacent composite.

The composite laminate at these locations suffers from severe degradation induced by bolt bending under loading (see Fig. 7). Such fatigue degradation results in hole elongation during fatigue testing, that eventually increases the clearance between the hole surface and the bolt shank (see 2λ in Fig. 6). Delaminations in composite laminates can be initiated during the drilling process and/or during assembly of joint components [11, 19]. However, other types of damage, such as cross-ply cracks, debonding, and cracks in the longitudinal plies (i.e., those plies, which are parallel to the direction of applied load) create a stress concentration [22]. Because of the presence of interlaminar stresses, a delamination crack will initiate from the stress concentration. It will grow along the ply interface as the number of cycles increases. The presence of delaminations prevent the distribution of load between composite plies, thus lowering the stiffness of the laminated composite. Figure 9 shows an example of long interlaminar delaminations developed during fatigue loading of bolted laminates. delaminations

delaminations

shear damage

crushing of hole edge

shear damage

Figure 9: Delaminations and shear damage occurred during fatigue testing of a bolted joint.

The same figure also shows extensive shear damage. It has been observed that this


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severe damage leads to an appreciable reduction in bearing stiffness of bolted laminates [20]. However, as the literature overview has shown, there is no available information about the development of fatigue damage in composite bolted joints with multi-row configuration. Paper E of the present thesis reports some results of a fractographic investigation of multi-row bolted composite joints tested in fatigue. Design parameters The influence of design parameters on the structural integrity of composite bolted joints has been extensively investigated for many years. The design parameters include joint geometry (see Fig. 6), joint configuration (single overlap, double lap, single bolt row, or multi-bolt row), and loading condition. The basic influence of the joint geometry on a specific failure mode has already been mentioned above. Depending on the amount of load being transferred, composite joints may be divided into lightly and heavily loaded. The first group is usually realized in the from of single-row joints, whereas the latter type of mechanical joints is implemented as multi-row joints. Although non-catastrophic bearing failures are quite common for single-row joints, net-section failure is recognized as unavoidable for multi-row joints [12–14]. For multi-row joints spacing between the fasteners would be an additional geometrical parameter to be taken into account. Another significant difference between single and multi-row bolted joints is distribution of applied load between the fasteners. In single-row joints all the fasteners are carrying equal loads, whereas the outer bolt rows in multi-row joints transfer more load than the inner bolt rows [13, 23–25]. Despite that this issue has been studied by numerical and analytical approaches, very little experimental data have been found [25]. Moreover, there have not been published any information dealing with the load-transfer distribution between different bolt rows in joints subjected to fatigue loading so far. Since repeated loads are unavoidable in reality, the subject of load transfer in multi-row joints under cyclic loading shall be highlighted. The present thesis includes some achievements in solving the above problems. Papers A and B present the obtained experimental results on the load-transfer distribution in multi-row joints subjected to quasi-static loading; whereas Papers C and D highlight the same subject for multi-row joints tested in fatigue. Mechanically fastened joints may be implemented either with single overlap or double lap configuration. The first configuration is recognized as prone to secondary bending during loading (see Fig. 10).

Figure 10: Secondary bending in a joint with single overlap due to the load eccentricity.

The secondary bending causes bending deflection of bolt and joint plates which may result in non-uniform stress distributions through the thickness of the composite

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laminates [18]. However, in real aircraft structures composite components, such as wing skins, for instance, are usually supported by stiffners, spars, or wing boxes. Through this the secondary bending may be avoided. In testing procedures, single overlap joints are tested being mounted with special lateral supports to simulate this situation. The resistance of composite joints with single and double lap configuration to quasi-static and fatigue loading has been studied and compared. The achieved results are presented in Papers A and D appended to the thesis. Bolted joints can be subjected either to shear, see Fig. 5, or to tensile loading, as shown in Fig. 11.

Figure 11: An example of bolted joint loaded in tension.

In the latter case, pull-through strength of composite in bolted joints needs to be taken into account when designing such connections [26]. In the case of shear loading, bolted joints can be subjected either to tension or compression, or to a combination of both tension and compression during cyclic loading. It has been reported a difference in load transfer through a filled hole loaded in tension or compression. During compressive loading an additional load path develops via the fastener shank, thus reducing the local stress concentration at the hole edge [14, 27]. As a result peak net-section stresses grow more slowly during loading. In the case of tensile loading, tension loads go around the filled hole, being confined to the net section. This effect may have influence on the behaviour and the resistance of bolted joints during static or fatigue testing [28]. To highlight this issue, the behaviour of bolts and their influence on local strain in the adjacent composite have been studied using bolt movement and strain gauge measurements during quasi-static and fatigue testing. The obtained results are concluded in Papers A, B, and E. Material parameters In the aircraft industry composites are generally manufactured by stacking layers (or plies) of prepreg consisting of unidirectional reinforcing fibres embedded in a resin matrix. A combination of epoxy matrices, which offer improved mechanical and thermal properties, with carbon fibres, which possess very high strength and stiffness together with improved tolerance to high temperatures and corrosive environments, makes carbon fibre/epoxy composites in the form of prepregs very attractive to be


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used in the aircraft industry. Although carbon fibre/epoxy prepregs are dominated, aramid and glass as matrix reinforcement are used as well [2]. Stacking sequence of a laminate is recognized to have a significant influence on the resistance of bolted joints. It is known that interlaminar stresses are very high at the laminate edges, cut-outs, and holes [22]. These stresses may initiate delamination and thus accelerate the failure process. However, the stacking sequence in laminates has been shown to affect the nature as well as the magnitude of the interlaminar stresses. Thus, by selecting a right sequence of stacked plies it is possible to reduce the effect of these stresses on the joint resistance to static and fatigue loading [18, 29–32]. The influence of lay-up on a specific failure mode has already been mentioned above. This subject is highlighted in Papers A and D including the quasi-static and fatigue testing results on composite joints with quasi-isotropic and 0◦ -dominated lay-ups. Fastening parameters Since the main objective of mechanically fastened joints is to transfer applied loads from one part to another via fasteners, the right configuration of fastening system is therefore quite important. It may be characterized by protruding or countersunk-bolt head, the presence of countersink, washers, nuts, clamping force, and the bolt/hole clearance (see 2λ in Fig. 6). The right fastening system means the best combination of these contributors, which would enhance the performance of joined structures. Today there are many different types of fasteners available in the aircraft industry. However, a combination of some metals with carbon fibre/epoxy composite, which is very common in manufacturing of aircraft composite structures, can lead to galvanic corrosion. It was observed that stainless-steel and aluminium fasteners used to connect carbon fibre/epoxy skins to the inner substructure were seriously damaged due to the fastener material was galvanically corroded [33]. This type of structural damage is very expensive to repair since it involves redrilling and installation of oversized fasteners from an appropriate material. The mechanism of the galvanic corrosion is based on the difference in electrode potential between the aforementioned metals and the composite material. A solution to avoid this problem is to use metals with less difference in electrode potential to carbon fibre/epoxy composites. Titanium alloys have already been found as the most suitable material to manufacture fastening systems for composites in the aircraft industry. Since thousands of fasteners are usually used in an aircraft to join different structural components, this incurs a large weight and cost penalty. Therefore, an optimum design of the fastener system must be developed and approved prior to the application to real aircraft structures. Another solution to avoid the galvanic corrosion is to use composite fasteners. These fasteners are a new type of mechanical joining for the aircraft industry being developed very recently. They can be made of carbon, glass fibers embedded in polyimide, PEEK, or epoxy matrices [33–35]. An example of fasteners made of PEEK/long carbon fibers reinforced composite material is shown in Fig. 12(a).

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Figure 12: Composite bolt (a), protruding titanium bolt (b), and countersunk titanium bolt (c).

Composite fasteners can reduce the weight of fastening system by 80% compared with metal ones. In external aircraft applications, composite fasteners can lower radar cross-section and reduce lightning strike risk. For electronically sensitive applications (radar and instrument housings), the application of low dielectric constant of quartz fibres provide a minimum of electronic interference. Despite the mentioned advantages, the literature review has found only few publications showing the application of composite fasteners to join composite laminates [33, 35–37]. Moreover, among them only two deal with the fatigue durability of such joints [36, 37]. To reduce this gap an experimental investigation on the fatigue resistance of CFRP-joints with composite and titanium fasteners has been done. The achieved results are presented in this thesis in Paper C. There are two basic configurations of fastener heads: protruding and countersunk (see Fig. 12(b) and (c)). It has been observed a difference in stress distribution through the thickness for joints with countersunk and protruding-head bolts [13,18, 21]. It was shown by FEM-modeling that the stress distribution for both configurations is non-uniform along the hole surface due to bolt bending. However, in the case of countersunk fasteners, almost all load is carried by the cylindrical part of the fastener, thus resulting in more localized stresses there. For the protruding-head bolts, the bearing load is distributed over the thickness of the laminate [18,21]. This effect is even more pronounced for the fatigue resistance of composite joints bolted by countersunk and protruding-head bolts. From fatigue testing of composite laminates with the same geometry and bolted by both fastener configurations, joints with protruding-head bolts yielded better fatigue resistance than those bolted by countersunk fasteners [31,32,38]. A comparison between the quasi-static and fatigue behaviour of joints bolted by protruding and different countersunk fasteners is made in Papers B and C. The amount of clamping force provided from installed fasteners is quite important for the mechanical behaviour of bolted joints. As has been observed, mechanically fastened joints with high clamp-ups yielded better strength to applied static loading than joints with low clamp-ups [18,30,39–41]. The same influence of clampup on the fatigue durability of bolted joints has been reported [37,38,42]. The main reason for this is that in bolted joints with highly tightened fasteners a larger part of the applied load is transferred by the friction forces than in joints with low tight-


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ened fasteners. In addition, the clamping results in a lateral constrain in the washer area which prevent delamination [18]. The influence of clamp-up on the fatigue behaviour of bolted joints has been investigated using a method of bolt-movement measurement. The obtained results are presented in Paper E.


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Future Outlook The obtained results of experimental programs on the quasi-static and fatigue resistance of composite bolted joints point to a number of areas where additional work is required. Some of them are listed below. • The influence of bolt movement and friction forces on the joint behaviour has appeared to be very important during loading. Therefore, to study their influence on the load transfer of applied load between bolt rows and local strain using FEM approach would be of interest. • Local fatigue behaviour in composite joints with countersunk fasteners has to be analysed since the load transfer of applied load in bolted plates with countersunk and ordinary holes is different. • More detailed fractographic observations are needed in order to study the extent of fatigue damage in bolted laminates tested at different load levels. This will highlight the mechanisms of fatigue damage in the joint system and the influence of individual bolts on them during fatigue testing.


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