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Manufacturing defects in composites and their effects on performance
Manufacturing defects in composites and their effects on performance
5
R. Talreja Texas A&M University, College Station, TX, USA
5.1
Introduction
Lightweight structures in aerospace applications are mostly made of polymer matrix composites. The size and shape of the structure, and its usage, e.g. as primary or secondary structure in aircraft, dictate the choice of the manufacturing process. While the early aerospace applications in defence were mainly performance driven, the cost of manufacturing today is of increasing concern. Therefore, the traditional design approach of ‘defect-free’ structures must be revisited. In fact, no structure is without defects; a low threshold of measurable defects is essentially what is taken to define the defect-free condition. If the cost of manufacturing is to be managed, i.e. reduced in a controlled way, then the effects of defects must be assessed. This requires a mechanics-based knowledge base on characterization of defects and quantification of their effects on specific performance characteristics. Reducing cost of manufacturing further requires that the manufacturing process be quantified with parameters that can be varied to minimize cost. This chapter reviews a few common manufacturing defect types for illustration. A cost-effective manufacturing strategy is then described, followed by a discussion of the role of manufacturing defects in affecting the performance of the composite part that has been produced. Three examples are given to illustrate how the manufacturing defects can be characterized and their effect on the relevant performance characteristics can be evaluated. This mechanics-based analysis methodology can be utilized in other cases and can become what may be called a defect-engineering framework for costeffective manufacturing of composite structures.
5.2
Defects in composite materials
Composite materials can be manufactured by a variety of methods. Polymer matrix composites (PMCs), for instance, can be processed by compression moulding, liquid moulding, injection moulding, resin infusion, etc. and can be joined by adhesive bonding, mechanical fastening, fusion bonding, etc. For a good overview of processing and manufacturing of PMCs, see Månson et al. [1]. Each processing and manufacturing route will induce defects. A systematic way to categorize defects is Polymer Composites in the Aerospace Industry. http://dx.doi.org/10.1016/B978-0-85709-523-7.00005-0 Copyright © 2015 Elsevier Ltd. All rights reserved.
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to group them into matrix, fibre and interface defects. Thus, matrix defects can be incomplete curing and voids; fibre defects can be misalignments, waviness and broken fibres and irregularities of fibre distribution in the matrix; and interfaces can carry defects as unbonded regions on fibre surfaces and between layers (delamination). These defects will be discussed next. Voids are the most common type of matrix defects and are found in virtually all PMC parts, whether manufactured by autoclave, liquid compression moulding or resin transfer moulding (RTM). Void formation can to some extent be controlled by manufacturing process parameters such as vacuum pressure, resin viscosity, cure temperature and consolidation pressure. In one study [2], for example, void morphology and spatial distribution in a circular disc of glass/epoxy was studied in an RTM process in which the resin was injected under pressure into a mould containing fibre bundle preform. The voids were found to vary in size, shape and spatial location. These characteristics were found to depend on the resin flow kinematics in the mould. Figure 5.1 taken from that work shows representative images of voids. Figure 5.2 shows voids in a unidirectional carbon/epoxy composite made by the autoclave technique [3]. The two sections, cut parallel and perpendicular to the fibre
Figure 5.1 Representative microscopic images of voids resulting in an E-glass/epoxy composite from an RTM process. From Ref. [2].
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Figure 5.2 Voids in a unidirectional carbon/epoxy composite made by autoclave process. From Ref. [3].
direction, display the voids that appear as elongated cylindrical-shaped entities lying mostly between the plies. These voids are typical of the autoclave process where they form as a consequence of trapped air in ply interfaces during the composite layup or by evaporation of water and volatiles inside the prepregs during curing. The vacuum and applied pressure during the curing process cause vapour and volatiles to diffuse into existing air pockets. Reducing voids in the autoclave process generally increases the cost of manufacturing and can become prohibitive for large structures. For this reason, the autoclave process is not used for structures such as wind turbine blades. One of the ways considered to reduce the cost of manufacturing by the autoclave technique is to eliminate external pressure and use vacuum pressure only. One study looked at the effect of moisture formation resulting from dissolved moisture in the prepregs [4]. Significant voids resulted in this case when the pressure was not applied. Figure 5.3 shows an example of the voids formed. On the other hand, eliminating vacuum in the autoclave technique also produces voids. An example of this is shown in Figure 5.4 [5]. Fibre defects, as noted above, are fibre misalignment and waviness, and broken fibres. In composites, where fibres are assumed to be straight, parallel and oriented in intended directions, deviations due to misalignment and waviness can reduce initial properties, particularly compression strength and stiffness and lead to reductions in aircraft design limit load and design ultimate load capability in service. Figure 5.5 illustrates the level of misalignment found in a typical as-delivered unidirectional prepreg [6]. A study of the effect of fibre waviness in unidirectional composites under axial compression [7] showed that the stiffness and strength reduced severely due to this type of defect. Stress analysis and experimental observations indicated that the interlaminar shear stress developed due to fibre waviness was responsible for delamination and subsequent failure. Figure 5.6 taken from that work illustrates occurrence of delamination (separation of individual layers that were intentionally made with wavy fibres) and final failure in axial compression. In an experimental and analytical
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Figure 5.3 Voids resulting in a composite laminate from dissolved moisture in the prepregs in an autoclaving process where vacuum bagging was used but no external pressure was applied. From Ref. [4].
study of the failure of unidirectional composites loaded in combined compression and shear, Vogler et al. [8] also found significant effect of fibre waviness on strength. The mechanism of failure in this case was microbuckling-induced kink band formation. A severe example of fibre defect is shown in Figure 5.7 where tangled fibres
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Figure 5.4 An example of voids formed between prepregs when vacuum is eliminated from the standard autoclave technique for manufacturing composite laminates. From Huang et al. [5].
Figure 5.5 Level of fibre misalignment in a typical as-delivered unidirectional prepreg. From Ref. [6].
are seen in a carbon/epoxy prepreg [9]. It is difficult to predict quantitatively the effects of such defects on design stresses except by testing to measure the degradation. If the manufacturing process control makes occurrence of such defects possible, then designers must ensure that design ultimate strength can be achieved in the presence of such defects. Interface defects are largely caused by inadequate conditions for generating bonds at interfaces between fibres and matrix and between layers of composites during manufacturing. If resin is infused into a fibre preform, it may not wet the entire fibre
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Figure 5.6 Illustration of the fibre waviness effect on failure under axial compression of a unidirectional carbon/epoxy composite. Top: graded waviness of fibres before loading; Middle: occurrence of delamination under compression; Bottom: final failure. From Ref. [7].
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Figure 5.7 A region of entangled fibres in a carbon/epoxy prepreg. From Ref. [8].
surface, resulting in unbonded regions of the fibre surface. If pre-impregnated layers of composites (prepregs) are stacked in a manufacturing process, then air can be trapped between the layers. The pockets of air in the interlaminar region can flatten during consolidation of layers resulting in planes of no contact, and therefore no bond, between layers. These defects are often called delaminations, although this term more appropriately describes separation (debonding) of bonded layers at interfaces.
5.3
Modelling with defects
A given manufacturing process results in a composite part that is specific to that process and can be described by its material state. The conventional representation of the material state is by homogenization that produces direction-dependent overall properties (e.g. anisotropic elastic constants). Explicit knowledge of the size and distribution of fibres is not retained in this representation. This is adequate and efficient if the purpose is to determine initial deformation characteristics, such as deflections and rotations of the produced composite part. When service loading is applied, the composite part suffers damage that in many cases initiates from the manufacturinginduced defects. Stress analysis using homogenized material state is then inadequate in analysing the formation and progression of failure. A new approach to representation of the material state is needed (discussed next). Instead of the conventional homogenization, the material state description should retain knowledge of the heterogeneities in a statistical sense to the degree that their effects on the local stress states generated at the micro level are properly represented. Additionally, the defects described in section 5.2 should be incorporated in the material state description along with the other heterogeneities (e.g. fibres and differently oriented layers) in a manner that the perturbations in the local stress field are explicitly attributable to the defects. This can allow evaluation of the defects from the viewpoint of failure initiation and progression.
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Implications on cost-effective manufacturing
Figure 5.8 depicts the overall methodology of a cost-effective design. Manufacturing plays a central role in this methodology since a substantial part of the total cost of a structure lies in steps such as processing, machining and assembly. On completion of these steps, the manufactured part is characterized by its geometry and the material state described in section 5.3. This characterization may require various efforts such as polymer process modelling, microscopy and nondestructive inspection. The quantified material state enters into the next group of items placed in the box labelled as properties/performance evaluation. Of particular note is the explicit separation of information concerning the material state (‘microstructure’) into the ‘homogenized’ and ‘defects’ parts. This is a departure from the traditional design approach where defects are not considered as a variable in the material state but only as a threshold for ‘accept/reject’ criteria. For instance, in some aerospace applications, a 2% void content is used as a threshold to accept or reject a manufactured part. In the strategy proposed here, the defects are viewed as a result of a welldefined manufacturing process, and it is assumed that they can be varied by controlled variation of the manufacturing parameters. The role of defects in determining the properties and performance (i.e. degradation of properties in service) is a key to assessing the effectiveness, and thereby the cost-effectiveness, of the manufacturing process. However, to ensure that each aircraft can achieve design ultimate and design limit strength values, the manufacturing process must be well in statistical process control producing a level of defects that will guarantee that design ultimate and limit strengths are consistently achieved. The outcome of the properties/performance evaluation enters in the last piece of the iterative process for cost reduction where cost/performance trade-offs are done. Conducting these trade-offs
Manufacturing • Process modelling and simulation • Tooling, machining, assembly
Materials characterization • Material state ('microstructure') Idealized (homogenized)
Cost/performance trade-offs
Defects
Properties/performance evaluation • Stiffness • Integrity, durability
Figure 5.8 Cost-effective design process for composite structures.
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requires balancing the manufacturing cost with the properties and the level of the achieved performance.
5.5 5.5.1
Mechanics-based analysis of defects Effect of voids on effective elastic properties
Traditionally the analysis of the effects of voids on the effective elastic properties is conducted by using inclusion theories where the inclusions are assigned zero properties [10]. Typically, the voids are distributed as embedded entities in the homogenized composite and the reduction in the elastic properties caused by their presence is estimated. Simulating the geometry of actual voids in a unidirectional composite, Huang and Talreja [11] showed that such estimates were generally not accurate. Basically, instead of cutting out a volume in a homogenized composite and replacing it with a void, as is commonly done, this work displaced the fibres locally in the composite to accommodate the void. This approach produced the effect of voids more accurately. A parametric study of the effects of voids in that work also showed that the void shape generally has significant effects on the matrix-dominated elastic properties, such as the modulus transverse to fibres (both in-plane and out-of-plane) and the shear modulus. An example of the through-thickness elastic modulus reduction [12] and its comparison with model prediction [11] is shown in Figure 5.9.
Figure 5.9 Percentage reduction in the through-thickness elastic modulus (Ezz) induced by voids. Modelling prediction [11] compared with test data from [12].
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5.5.2
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Effect of voids on delamination growth
If a delamination exists in a composite part either as an initial interlaminar defect or induced by service loading, such as lateral impact, then it is common to assess the effect on the failure of the part by fracture mechanics methods. These methods typically analyse the propensity of the delamination, viewed as a crack, to grow unstably to failure. The driving force for crack growth, or the strain energy release rate (SERR) of the delamination crack front, is compared with the fracture toughness, i.e. the resistance to crack growth, of the material in the interface in which the delamination lies. Thus if G is the SERR of a given delamination crack, then G ¼ Gc is the criterion for incipient unstable crack growth (fracture). Gc, the interface fracture toughness, is evaluated by certain standard experimental techniques. Generally, the loading conditions on the structural part could result in a mixed-mode delamination growth, in which case the SERR and the fracture toughness will have different values corresponding to each mode. For most PMCs, the mode I (opening mode) fracture toughness is the lowest of the three possible crack growth modes. To evaluate the failure potential of an existing delamination, therefore, one commonly uses the opening mode fracture criterion GI ¼ GIc. Depending on the manufacturing process, voids can be produced in the interfacial region between the composite layers, or in the layers themselves, or both. For delamination growth the overall void content is of little consequence. What matters is the presence of voids in the delaminated (cracked) interface. Ricotta et al. [13] demonstrated this in a study of delamination growth by analysing the so-called double cantilever beam (DCB) specimen containing voids ahead of the crack tip (Figure 5.10). Figure 5.11 shows the computed effect of the presence of circular voids in the interface as a function of the distance D of the nearest void from the crack tip. As seen, the effect of the nearest void has the most influence on the SERR, while adding more voids has increasingly less effect. In Figure 5.12, the effect of multiple voids as
Figure 5.10 A schematic of a DCB specimen with voids ahead of the crack tip.
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1.3
GI,v /GI
D
2 mm
1.2 3 voids R = 0.1 mm 2 voids R = 0.1 mm 1 void R = 0.1 mm 1.1
1 0
1 2 3 Distance from the crack tip D (mm)
4
Figure 5.11 Variation of SERR (GI,v) with the distance D from the crack tip of the nearest circular void of 0.1 mm radius. Effect of up to three voids of mutual distance 2.0 mm is also shown.
1.11
l
c
c
GI,v /GI
1.10
1.09 3 voids R = 0.1 mm 2 voids R = 0.1 mm 1.08 0
8 2 4 6 Distance from the first void c (mm)
10
Figure 5.12 Effect on the SERR (GI,v) of the mutual spacing c of the voids with the distance of the nearest void being fixed.
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a function of the distance of the additional voids from the nearest void to the crack tip is illustrated. The results shown for a fixed distance of the first void from the crack tip indicate that the additional voids enhance the SERR, but as they move away from the first void, that enhancement decreases. More results of the parametric study of the effects of the voids on SERR are given in [13].
5.5.3
Effect of defects on progressive intralaminar cracking
The first failure mechanism in multidirectional composite laminae is matrix cracking along fibres within plies. This mechanism described as transverse cracking or intralaminar cracking is often the basis of design, i.e. the applied load level is kept below the threshold for initiation of these cracks. These cracks do not cause failure of laminate but lead to other mechanisms that eventually cause failure. The transverse cracking process in laminates progresses in stages commonly described as noninteractive and interactive. In the noninteractive stage, the cracks initiate from defects in the plies, the largest defect (with most favourable orientation and size) initiating the first crack. As the load is increased, more cracks initiate from smaller defects at random locations. The noninteractive stage shows increasing rate of crack formation with loading until the cracks begin interacting, i.e. the perturbation in the local stress field induced by one crack is affected by the presence of another crack. The subsequent cracks form between the preexisting cracks at locations where the stress exceeds the material strength (with or without defects). Thus in this interactive stage the role of manufacturing defects diminishes as the stress field becomes increasingly affected by the cracks of reducing mutual spacing. In Figure 5.13 the transverse cracks as seen from the edge of a specimen are shown. Figure 5.13(a) shows an array of parallel cracks in the 90 plies of a cross-ply laminate, while in Figure 5.13(b) one sees the same type of parallel cracks within transverse bundles of an 8-harness woven fabric composite. To investigate the effect of manufacturing defects on the transverse cracking process, a study was recently conducted to test how different irregular manufacturing processes influence the progressive transverse cracking in cross-ply laminates [5]. Composite laminates with a thermosetting polymer matrix are manufactured in the aerospace industry commonly by the autoclaving process. In this process a prespecified temperature and pressure variation is followed along with application of vacuum to draw out air from the composite part. The process cost is elevated due to the requirement of placing the mould in a closed space in which the part is subjected to controlled temperature, pressure and vacuum. The cost could be reduced if the application of vacuum or pressure could be omitted without adversely affecting the part quality. It is expected that omitting vacuum or pressure will result in voids and possibly other defects such as inadequate wetting of the fibres by resin prior to curing. Three plates of a (0/90)s laminate were fabricated in an autoclave using prepregs of HexPlyÒ M10/38%/UD300/CHS. Plate 1 was produced by the standard
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Figure 5.13 (a) Transverse cracks in 90 plies of a cross-ply laminate; (b) cracks within transverse bundles of an 8-harness woven fabric composite.
process with the exception that the air was intentionally left entrapped between prepregs by not applying consolidation pressure during the lamination (stacking of prepregs) process. For plate 2, standard process was followed, except no vacuum was applied during the curing process. For plate 3, standard process was strictly followed. The coupons cut from the laminated plates were subjected to monotonic axial tension. The study [5] focused on the noninteractive stage of the transverse cracking process and accounted for the random defects by a statistical description of strength. Specifically, Weibull distribution was used to represent the random strength, and the two parameters of the probability distribution function were estimated from the test data. The result of the analysis is presented in Figure 5.14, where the crack density (average number of cracks per unit specimen length) is plotted against the applied axial stress. Three different curves illustrate the differences due to the defects induced by the three manufacturing processes. At the same applied stress, plate 1 made with air-entrapped layups consistently has highest crack density, followed by plate 2 cured without applying vacuum, while plate 3 made by standard process has the lowest crack density. Also, the onset stress for the initiation of transverse crack is much lower for plate 1.
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Figure 5.14 Variation of the density of transverse cracks in cross-ply laminates produced by three manufacturing processes. The manufacturing process for the lowest curve (Plate 3) is the standard autoclave process (specified by the manufacturer), while in Plate 2 vacuum was not applied, and Plate 1 was not consolidated by external pressure, trapping air between layers.
5.6
Summary
This chapter has briefly reviewed some of the common forms of defect found in composites structures and their effects on some mechanical properties. A good engineering design should not only meet the performance specifications required of the composite structure but it should do so at the lowest cost feasible for that structure. Traditional design focuses on meeting performance requirements and uses quality control measures to accept or reject the manufactured part. This is not going to be feasible for acceptance of large wings and fuselage structures e the cost will be prohibitive and instead service performance of composite aircraft structures must rely on manufacture with consistent levels of defect severity produced by a process that is strictly within manufacturing process control. This can be accomplished if the effects of defects on mechanical performance are evaluated. Three examples given here illustrate that average measures of defects, e.g. void volume fraction, are inadequate for this purpose. Instead, the size, shape and location of defects need to be accounted for to do proper assessment of performance. The manufacture process must be understood sufficiently well to allow process modification to control the level of defects produced.
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References [1]
[2] [3] [4] [5]
[6]
[7]
[8] [9] [10] [11] [12] [13]
Månson JAE, Wakeman MD, Bernet N. Composite processing and manufacturing e an overview. In: Talreja R, Månson JAE, editors. Kelly A, Zweben C, series editors, Comprehensive composite materials, vol. 2. Elsevier; 2000. pp. 577e607. Hamidi YK, Aktas L, Cengiz Altan M. Formation of microscopic voids in resin transfer moulded composites. J Eng Mater Technol Trans ASME 2004;126:420e6. Bowles KJ, Frimpong S. Void effects on the interlaminar shear strength of unidirectional graphite-fibre-reinforced composites. J Compos Mater 1992;26:1487e509. Grunenfelder LK, Nutt SR. Void formation in composite prepregs e effect of dissolved moisture. Compos Sci Technol 2010;70:2304e9. Huang Y, Varna J, Talreja R. The effect of manufacturing quality on transverse cracking in cross ply laminates. In: Bhattacharrya D, Lin RJT, Srivatsan TS, editors. Proceedings of processing and fabrication of advanced materials XIX, PFAM-19. New Zealand: Auckland; January 14e17, 2011. pp. 552e9. Potter K, Khan B, Wisnom M, Bell T, Stevens J. Variability, fibre waviness and misalignment in the determination of the properties of composite materials and structures. Compos Part A: Appl Sci Manuf 2008;39:1343e54. Hsiao HM, Daniel IM. Effect of fibre waviness on stiffness and strength reduction of unidirectional composites under compressive loading. Compos Sci Technol 1996;56: 581e93. Vogler TJ, Hsu S-Y, Kyriakides S. Composite failure under combined compression and shear. Int J Solids Struct 2000;37:1765e91. Adams RD, Cawley P. A review of defect types and non-destructive testing techniques for composites and bonded joints. NDT Int 1988;21:208e22. Berryman JG. Role of porosity in estimates of composite elastic constants. Trans ASME J Energy Resour Tech 1994;116:87e96. Huang H, Talreja R. Effect of void geometry on elastic properties of unidirectional fibre reinforced composites. Compos Sci Technol 2005;65:1964e81. Guerdal Z, Tomasino AP, Biggers SB. Effects of processing induced defects on laminate response: interlaminar tensile strength. SAMPE J 1991;27:3e49. Ricotta M, Quaresimin M, Talreja R. Mode 1 strain energy release rate in composite laminates in the presence of voids. Compos Sci Technol 2008;68:2616e23.
5
R. Talreja Texas A&M University, College Station, TX, USA
5.1
Introduction
Lightweight structures in aerospace applications are mostly made of polymer matrix composites. The size and shape of the structure, and its usage, e.g. as primary or secondary structure in aircraft, dictate the choice of the manufacturing process. While the early aerospace applications in defence were mainly performance driven, the cost of manufacturing today is of increasing concern. Therefore, the traditional design approach of ‘defect-free’ structures must be revisited. In fact, no structure is without defects; a low threshold of measurable defects is essentially what is taken to define the defect-free condition. If the cost of manufacturing is to be managed, i.e. reduced in a controlled way, then the effects of defects must be assessed. This requires a mechanics-based knowledge base on characterization of defects and quantification of their effects on specific performance characteristics. Reducing cost of manufacturing further requires that the manufacturing process be quantified with parameters that can be varied to minimize cost. This chapter reviews a few common manufacturing defect types for illustration. A cost-effective manufacturing strategy is then described, followed by a discussion of the role of manufacturing defects in affecting the performance of the composite part that has been produced. Three examples are given to illustrate how the manufacturing defects can be characterized and their effect on the relevant performance characteristics can be evaluated. This mechanics-based analysis methodology can be utilized in other cases and can become what may be called a defect-engineering framework for costeffective manufacturing of composite structures.
5.2
Defects in composite materials
Composite materials can be manufactured by a variety of methods. Polymer matrix composites (PMCs), for instance, can be processed by compression moulding, liquid moulding, injection moulding, resin infusion, etc. and can be joined by adhesive bonding, mechanical fastening, fusion bonding, etc. For a good overview of processing and manufacturing of PMCs, see Månson et al. [1]. Each processing and manufacturing route will induce defects. A systematic way to categorize defects is Polymer Composites in the Aerospace Industry. http://dx.doi.org/10.1016/B978-0-85709-523-7.00005-0 Copyright © 2015 Elsevier Ltd. All rights reserved.
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to group them into matrix, fibre and interface defects. Thus, matrix defects can be incomplete curing and voids; fibre defects can be misalignments, waviness and broken fibres and irregularities of fibre distribution in the matrix; and interfaces can carry defects as unbonded regions on fibre surfaces and between layers (delamination). These defects will be discussed next. Voids are the most common type of matrix defects and are found in virtually all PMC parts, whether manufactured by autoclave, liquid compression moulding or resin transfer moulding (RTM). Void formation can to some extent be controlled by manufacturing process parameters such as vacuum pressure, resin viscosity, cure temperature and consolidation pressure. In one study [2], for example, void morphology and spatial distribution in a circular disc of glass/epoxy was studied in an RTM process in which the resin was injected under pressure into a mould containing fibre bundle preform. The voids were found to vary in size, shape and spatial location. These characteristics were found to depend on the resin flow kinematics in the mould. Figure 5.1 taken from that work shows representative images of voids. Figure 5.2 shows voids in a unidirectional carbon/epoxy composite made by the autoclave technique [3]. The two sections, cut parallel and perpendicular to the fibre
Figure 5.1 Representative microscopic images of voids resulting in an E-glass/epoxy composite from an RTM process. From Ref. [2].
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Figure 5.2 Voids in a unidirectional carbon/epoxy composite made by autoclave process. From Ref. [3].
direction, display the voids that appear as elongated cylindrical-shaped entities lying mostly between the plies. These voids are typical of the autoclave process where they form as a consequence of trapped air in ply interfaces during the composite layup or by evaporation of water and volatiles inside the prepregs during curing. The vacuum and applied pressure during the curing process cause vapour and volatiles to diffuse into existing air pockets. Reducing voids in the autoclave process generally increases the cost of manufacturing and can become prohibitive for large structures. For this reason, the autoclave process is not used for structures such as wind turbine blades. One of the ways considered to reduce the cost of manufacturing by the autoclave technique is to eliminate external pressure and use vacuum pressure only. One study looked at the effect of moisture formation resulting from dissolved moisture in the prepregs [4]. Significant voids resulted in this case when the pressure was not applied. Figure 5.3 shows an example of the voids formed. On the other hand, eliminating vacuum in the autoclave technique also produces voids. An example of this is shown in Figure 5.4 [5]. Fibre defects, as noted above, are fibre misalignment and waviness, and broken fibres. In composites, where fibres are assumed to be straight, parallel and oriented in intended directions, deviations due to misalignment and waviness can reduce initial properties, particularly compression strength and stiffness and lead to reductions in aircraft design limit load and design ultimate load capability in service. Figure 5.5 illustrates the level of misalignment found in a typical as-delivered unidirectional prepreg [6]. A study of the effect of fibre waviness in unidirectional composites under axial compression [7] showed that the stiffness and strength reduced severely due to this type of defect. Stress analysis and experimental observations indicated that the interlaminar shear stress developed due to fibre waviness was responsible for delamination and subsequent failure. Figure 5.6 taken from that work illustrates occurrence of delamination (separation of individual layers that were intentionally made with wavy fibres) and final failure in axial compression. In an experimental and analytical
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Figure 5.3 Voids resulting in a composite laminate from dissolved moisture in the prepregs in an autoclaving process where vacuum bagging was used but no external pressure was applied. From Ref. [4].
study of the failure of unidirectional composites loaded in combined compression and shear, Vogler et al. [8] also found significant effect of fibre waviness on strength. The mechanism of failure in this case was microbuckling-induced kink band formation. A severe example of fibre defect is shown in Figure 5.7 where tangled fibres
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Figure 5.4 An example of voids formed between prepregs when vacuum is eliminated from the standard autoclave technique for manufacturing composite laminates. From Huang et al. [5].
Figure 5.5 Level of fibre misalignment in a typical as-delivered unidirectional prepreg. From Ref. [6].
are seen in a carbon/epoxy prepreg [9]. It is difficult to predict quantitatively the effects of such defects on design stresses except by testing to measure the degradation. If the manufacturing process control makes occurrence of such defects possible, then designers must ensure that design ultimate strength can be achieved in the presence of such defects. Interface defects are largely caused by inadequate conditions for generating bonds at interfaces between fibres and matrix and between layers of composites during manufacturing. If resin is infused into a fibre preform, it may not wet the entire fibre
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Figure 5.6 Illustration of the fibre waviness effect on failure under axial compression of a unidirectional carbon/epoxy composite. Top: graded waviness of fibres before loading; Middle: occurrence of delamination under compression; Bottom: final failure. From Ref. [7].
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Figure 5.7 A region of entangled fibres in a carbon/epoxy prepreg. From Ref. [8].
surface, resulting in unbonded regions of the fibre surface. If pre-impregnated layers of composites (prepregs) are stacked in a manufacturing process, then air can be trapped between the layers. The pockets of air in the interlaminar region can flatten during consolidation of layers resulting in planes of no contact, and therefore no bond, between layers. These defects are often called delaminations, although this term more appropriately describes separation (debonding) of bonded layers at interfaces.
5.3
Modelling with defects
A given manufacturing process results in a composite part that is specific to that process and can be described by its material state. The conventional representation of the material state is by homogenization that produces direction-dependent overall properties (e.g. anisotropic elastic constants). Explicit knowledge of the size and distribution of fibres is not retained in this representation. This is adequate and efficient if the purpose is to determine initial deformation characteristics, such as deflections and rotations of the produced composite part. When service loading is applied, the composite part suffers damage that in many cases initiates from the manufacturinginduced defects. Stress analysis using homogenized material state is then inadequate in analysing the formation and progression of failure. A new approach to representation of the material state is needed (discussed next). Instead of the conventional homogenization, the material state description should retain knowledge of the heterogeneities in a statistical sense to the degree that their effects on the local stress states generated at the micro level are properly represented. Additionally, the defects described in section 5.2 should be incorporated in the material state description along with the other heterogeneities (e.g. fibres and differently oriented layers) in a manner that the perturbations in the local stress field are explicitly attributable to the defects. This can allow evaluation of the defects from the viewpoint of failure initiation and progression.
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5.4
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Implications on cost-effective manufacturing
Figure 5.8 depicts the overall methodology of a cost-effective design. Manufacturing plays a central role in this methodology since a substantial part of the total cost of a structure lies in steps such as processing, machining and assembly. On completion of these steps, the manufactured part is characterized by its geometry and the material state described in section 5.3. This characterization may require various efforts such as polymer process modelling, microscopy and nondestructive inspection. The quantified material state enters into the next group of items placed in the box labelled as properties/performance evaluation. Of particular note is the explicit separation of information concerning the material state (‘microstructure’) into the ‘homogenized’ and ‘defects’ parts. This is a departure from the traditional design approach where defects are not considered as a variable in the material state but only as a threshold for ‘accept/reject’ criteria. For instance, in some aerospace applications, a 2% void content is used as a threshold to accept or reject a manufactured part. In the strategy proposed here, the defects are viewed as a result of a welldefined manufacturing process, and it is assumed that they can be varied by controlled variation of the manufacturing parameters. The role of defects in determining the properties and performance (i.e. degradation of properties in service) is a key to assessing the effectiveness, and thereby the cost-effectiveness, of the manufacturing process. However, to ensure that each aircraft can achieve design ultimate and design limit strength values, the manufacturing process must be well in statistical process control producing a level of defects that will guarantee that design ultimate and limit strengths are consistently achieved. The outcome of the properties/performance evaluation enters in the last piece of the iterative process for cost reduction where cost/performance trade-offs are done. Conducting these trade-offs
Manufacturing • Process modelling and simulation • Tooling, machining, assembly
Materials characterization • Material state ('microstructure') Idealized (homogenized)
Cost/performance trade-offs
Defects
Properties/performance evaluation • Stiffness • Integrity, durability
Figure 5.8 Cost-effective design process for composite structures.
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requires balancing the manufacturing cost with the properties and the level of the achieved performance.
5.5 5.5.1
Mechanics-based analysis of defects Effect of voids on effective elastic properties
Traditionally the analysis of the effects of voids on the effective elastic properties is conducted by using inclusion theories where the inclusions are assigned zero properties [10]. Typically, the voids are distributed as embedded entities in the homogenized composite and the reduction in the elastic properties caused by their presence is estimated. Simulating the geometry of actual voids in a unidirectional composite, Huang and Talreja [11] showed that such estimates were generally not accurate. Basically, instead of cutting out a volume in a homogenized composite and replacing it with a void, as is commonly done, this work displaced the fibres locally in the composite to accommodate the void. This approach produced the effect of voids more accurately. A parametric study of the effects of voids in that work also showed that the void shape generally has significant effects on the matrix-dominated elastic properties, such as the modulus transverse to fibres (both in-plane and out-of-plane) and the shear modulus. An example of the through-thickness elastic modulus reduction [12] and its comparison with model prediction [11] is shown in Figure 5.9.
Figure 5.9 Percentage reduction in the through-thickness elastic modulus (Ezz) induced by voids. Modelling prediction [11] compared with test data from [12].
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5.5.2
Polymer Composites in the Aerospace Industry
Effect of voids on delamination growth
If a delamination exists in a composite part either as an initial interlaminar defect or induced by service loading, such as lateral impact, then it is common to assess the effect on the failure of the part by fracture mechanics methods. These methods typically analyse the propensity of the delamination, viewed as a crack, to grow unstably to failure. The driving force for crack growth, or the strain energy release rate (SERR) of the delamination crack front, is compared with the fracture toughness, i.e. the resistance to crack growth, of the material in the interface in which the delamination lies. Thus if G is the SERR of a given delamination crack, then G ¼ Gc is the criterion for incipient unstable crack growth (fracture). Gc, the interface fracture toughness, is evaluated by certain standard experimental techniques. Generally, the loading conditions on the structural part could result in a mixed-mode delamination growth, in which case the SERR and the fracture toughness will have different values corresponding to each mode. For most PMCs, the mode I (opening mode) fracture toughness is the lowest of the three possible crack growth modes. To evaluate the failure potential of an existing delamination, therefore, one commonly uses the opening mode fracture criterion GI ¼ GIc. Depending on the manufacturing process, voids can be produced in the interfacial region between the composite layers, or in the layers themselves, or both. For delamination growth the overall void content is of little consequence. What matters is the presence of voids in the delaminated (cracked) interface. Ricotta et al. [13] demonstrated this in a study of delamination growth by analysing the so-called double cantilever beam (DCB) specimen containing voids ahead of the crack tip (Figure 5.10). Figure 5.11 shows the computed effect of the presence of circular voids in the interface as a function of the distance D of the nearest void from the crack tip. As seen, the effect of the nearest void has the most influence on the SERR, while adding more voids has increasingly less effect. In Figure 5.12, the effect of multiple voids as
Figure 5.10 A schematic of a DCB specimen with voids ahead of the crack tip.
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1.3
GI,v /GI
D
2 mm
1.2 3 voids R = 0.1 mm 2 voids R = 0.1 mm 1 void R = 0.1 mm 1.1
1 0
1 2 3 Distance from the crack tip D (mm)
4
Figure 5.11 Variation of SERR (GI,v) with the distance D from the crack tip of the nearest circular void of 0.1 mm radius. Effect of up to three voids of mutual distance 2.0 mm is also shown.
1.11
l
c
c
GI,v /GI
1.10
1.09 3 voids R = 0.1 mm 2 voids R = 0.1 mm 1.08 0
8 2 4 6 Distance from the first void c (mm)
10
Figure 5.12 Effect on the SERR (GI,v) of the mutual spacing c of the voids with the distance of the nearest void being fixed.
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a function of the distance of the additional voids from the nearest void to the crack tip is illustrated. The results shown for a fixed distance of the first void from the crack tip indicate that the additional voids enhance the SERR, but as they move away from the first void, that enhancement decreases. More results of the parametric study of the effects of the voids on SERR are given in [13].
5.5.3
Effect of defects on progressive intralaminar cracking
The first failure mechanism in multidirectional composite laminae is matrix cracking along fibres within plies. This mechanism described as transverse cracking or intralaminar cracking is often the basis of design, i.e. the applied load level is kept below the threshold for initiation of these cracks. These cracks do not cause failure of laminate but lead to other mechanisms that eventually cause failure. The transverse cracking process in laminates progresses in stages commonly described as noninteractive and interactive. In the noninteractive stage, the cracks initiate from defects in the plies, the largest defect (with most favourable orientation and size) initiating the first crack. As the load is increased, more cracks initiate from smaller defects at random locations. The noninteractive stage shows increasing rate of crack formation with loading until the cracks begin interacting, i.e. the perturbation in the local stress field induced by one crack is affected by the presence of another crack. The subsequent cracks form between the preexisting cracks at locations where the stress exceeds the material strength (with or without defects). Thus in this interactive stage the role of manufacturing defects diminishes as the stress field becomes increasingly affected by the cracks of reducing mutual spacing. In Figure 5.13 the transverse cracks as seen from the edge of a specimen are shown. Figure 5.13(a) shows an array of parallel cracks in the 90 plies of a cross-ply laminate, while in Figure 5.13(b) one sees the same type of parallel cracks within transverse bundles of an 8-harness woven fabric composite. To investigate the effect of manufacturing defects on the transverse cracking process, a study was recently conducted to test how different irregular manufacturing processes influence the progressive transverse cracking in cross-ply laminates [5]. Composite laminates with a thermosetting polymer matrix are manufactured in the aerospace industry commonly by the autoclaving process. In this process a prespecified temperature and pressure variation is followed along with application of vacuum to draw out air from the composite part. The process cost is elevated due to the requirement of placing the mould in a closed space in which the part is subjected to controlled temperature, pressure and vacuum. The cost could be reduced if the application of vacuum or pressure could be omitted without adversely affecting the part quality. It is expected that omitting vacuum or pressure will result in voids and possibly other defects such as inadequate wetting of the fibres by resin prior to curing. Three plates of a (0/90)s laminate were fabricated in an autoclave using prepregs of HexPlyÒ M10/38%/UD300/CHS. Plate 1 was produced by the standard
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Figure 5.13 (a) Transverse cracks in 90 plies of a cross-ply laminate; (b) cracks within transverse bundles of an 8-harness woven fabric composite.
process with the exception that the air was intentionally left entrapped between prepregs by not applying consolidation pressure during the lamination (stacking of prepregs) process. For plate 2, standard process was followed, except no vacuum was applied during the curing process. For plate 3, standard process was strictly followed. The coupons cut from the laminated plates were subjected to monotonic axial tension. The study [5] focused on the noninteractive stage of the transverse cracking process and accounted for the random defects by a statistical description of strength. Specifically, Weibull distribution was used to represent the random strength, and the two parameters of the probability distribution function were estimated from the test data. The result of the analysis is presented in Figure 5.14, where the crack density (average number of cracks per unit specimen length) is plotted against the applied axial stress. Three different curves illustrate the differences due to the defects induced by the three manufacturing processes. At the same applied stress, plate 1 made with air-entrapped layups consistently has highest crack density, followed by plate 2 cured without applying vacuum, while plate 3 made by standard process has the lowest crack density. Also, the onset stress for the initiation of transverse crack is much lower for plate 1.
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Figure 5.14 Variation of the density of transverse cracks in cross-ply laminates produced by three manufacturing processes. The manufacturing process for the lowest curve (Plate 3) is the standard autoclave process (specified by the manufacturer), while in Plate 2 vacuum was not applied, and Plate 1 was not consolidated by external pressure, trapping air between layers.
5.6
Summary
This chapter has briefly reviewed some of the common forms of defect found in composites structures and their effects on some mechanical properties. A good engineering design should not only meet the performance specifications required of the composite structure but it should do so at the lowest cost feasible for that structure. Traditional design focuses on meeting performance requirements and uses quality control measures to accept or reject the manufactured part. This is not going to be feasible for acceptance of large wings and fuselage structures e the cost will be prohibitive and instead service performance of composite aircraft structures must rely on manufacture with consistent levels of defect severity produced by a process that is strictly within manufacturing process control. This can be accomplished if the effects of defects on mechanical performance are evaluated. Three examples given here illustrate that average measures of defects, e.g. void volume fraction, are inadequate for this purpose. Instead, the size, shape and location of defects need to be accounted for to do proper assessment of performance. The manufacture process must be understood sufficiently well to allow process modification to control the level of defects produced.
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