1 Introduction and Overview 1.1 General Since the first edition of this textbook'in 1986,the use of high-performance polymer-matrix fiber composites in aircraft structures has grown steadily, although not as dramatically as predicted at that time.This is despite the significant weight-saving and other advantages that these composites can provide. The main reason for the slower-than-anticipated take-up is the high cost of aircraft components made of composites compared with similar structures made from metal,mainly aluminum,alloys.Other factors include the high cost of certification of new components and their relatively low resistance to mechanical damage,low through-thickness strength,and (compared with titanium alloys) temperature limitations.Thus,metals will continue to be favored for many airframe applications. The most important polymer-matrix fiber material and the main subject of this and the previous book,Composite Materials for Aircraft Structures,is carbon fiber-reinforced epoxy(carbon/epoxy).Although the raw material costs of this and similar composites will continue to be relatively high,with continuing developments in materials,design,and manufacturing technology,their advantages over metals are increasing. However,competition will be fierce with continuing developments in structural metals.In aluminum alloys developments include improved toughness and corrosion resistance in conventional alloys;new lightweight alloys(such as aluminum lithium);low-cost aerospace-grade castings;mechanical alloying (high-temperature alloys);and super-plastic forming.For titanium,they include use of powder preforms,casting,and super-plastic-forming/diffusion bonding. Advanced joining techniques such as laser and friction welding,automated riveting techniques,and high-speed (numerically controlled)machining also make metallic structures more affordable. The growth in the use of composites in the airframes in selected aircraft is illustrated in Figure 1.1.However,despite this growth,the reality is,as illustrated in Figure 1.2 for the U.S.Navy F-18 fighter,that airframes (and engines)will continue to be a mix of materials.These will include composites of various types and a range of metal alloys,the balance depending on structural and economic factors
1 Introduction and Overview 1.1 General Since the first edition of this textbook 1 in 1986, the use of high-performance polymer-matrix fiber composites in aircraft structures has grown steadily, although not as dramatically as predicted at that time. This is despite the significant weight-saving and other advantages that these composites can provide. The main reason for the slower-than-anticipated take-up is the high cost of aircraft components made of composites compared with similar structures made from metal, mainly aluminum, alloys. Other factors include the high cost of certification of new components and their relatively low resistance to mechanical damage, low through-thickness strength, and (compared with titanium alloys) temperature limitations. Thus, metals will continue to be favored for many airframe applications. The most important polymer-matrix fiber material and the main subject of this and the previous book, Composite Materials for Aircraft Structures, is carbon fiber-reinforced epoxy (carbon/epoxy). Although the raw material costs of this and similar composites will continue to be relatively high, with continuing developments in materials, design, and manufacturing technology, their advantages over metals are increasing. However, competition will be fierce with continuing developments in structural metals. In aluminum alloys developments include improved toughness and corrosion resistance in conventional alloys; new lightweight alloys (such as aluminum lithium); low-cost aerospace-grade castings; mechanical alloying (high-temperature alloys); and super-plastic forming. For titanium, they include use of powder preforms, casting, and super-plastic-forming/diffusion bonding. Advanced joining techniques such as laser and friction welding, automated riveting techniques, and high-speed (numerically controlled) machining also make metallic structures more affordable. The growth in the use of composites in the airframes in selected aircraft is illustrated in Figure 1.1. However, despite this growth, the reality is, as illustrated in Figure 1.2 for the U.S. Navy F-18 fighter, that airframes (and engines) will continue to be a mix of materials. These will include composites of various types and a range of metal alloys, the balance depending on structural and economic factors
2 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES eungonns 4053025 JSF◆ B2 V-22,年-22 0 AV8B Rafale 15 F-18E/F Q A320 ◆ 6330 F-18A 5 ◆ A340 0 F-15AF-16A767737 MD11 C17A 1970 1980 1990 2000 2010 Approximate Year of Introduction Fig.1.1 Growth of use of advanced composites in airframe structures. In this introductory chapter,the incentives or drivers for developing improved materials for aircraft applications are discussed.This is followed by a brief overview of fiber composites,including polymer,metal,and ceramic-matrix composites as well as hybrid metal/composite laminates.Other than polymer- matrix composites,these composites are not considered elsewhere in this book and so are discussed in this chapter for completeness. PERCENT OF STRUCTURAL WEIGHT F/A-F/A- 18C/D 18E/F Aluminum 49 31 Steel 尔 14 Titanium 13 24 Carbon Eposy 10 19 Other 3 的 100 100 Fig.1.2 Schematic diagram of fighter aircraft F-18 E/F.For comparison details of the structure of the earlier C/D model are also provided in the inset table
2 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES 40 E X e < 35 30 25 20 15 10 5 0 1970 AV8B F-18A • •• F-15~ F-16A 767 A320 B2 t 737 V-22 4~.22 Rafale F-18E/F A~0 ~77 MD11 C17A JSF-O 1980 1990 2000 2010 Approximate Year of Introduction Fig. 1.1 Growth of use of advanced composites in airframe structures. In this introductory chapter, the incentives or drivers for developing improved materials for aircraft applications are discussed. This is followed by a brief overview of fiber composites, including polymer, metal, and ceramic-matrix composites as well as hybrid metal/composite laminates. Other than polymermatrix composites, these composites are not considered elsewhere in this book and so are discussed in this chapter for completeness. PERCENT OF STRUCTURAL WEIGHT ~ Aluminum I Steel IB Titanium I Carbon Ep Other F/A- F/A- 18C/D 18E/F An 11 Fig. 1.2 Schematic diagram of fighter aircraft F-18 E/F. For comparison details of the structure of the earlier C/D model are also provided in the inset table
INTRODUCTION AND OVERVIEW 3 1.2 Drivers for Improved Airframe Materlals Weight saving through increased specific strength or stiffness is a major driver for the development of materials for airframes.2 However,as listed in Table 1.1,there are many other incentives for the introduction of a new material. A crucial issue in changing to a new material,even when there are clear performance benefits such as weight saving to be gained,is affordability.This includes procurement (up front)cost(currently the main criterion)and through- life support cost(i.e.,cost of ownership,including maintenance and repair).Thus the benefits of weight savings must be balanced against the cost.Approximate values that may be placed on saving 1 kilogram of weight on a range of aircraft types are listed in Table 1.2. In choosing new materials for airframe applications,it is essential to ensure that there are no compromises in the levels of safety achievable with con- ventional alloys.Retention of high levels of residual strength in the presence of typical damage for the particular material (damage tolerance)is a critical issue. Durability,the resistance to cyclic stress or environmental degradation and damage,through the service life is also a major factor in determining through-life support costs.The rate of damage growth and tolerance to damage determine the frequency and cost of inspections and the need for repairs throughout the life of the structure. 1.3 High-Performance Fiber Composite Concepts The fiber composite approach can provide significant improvements in specific(property/density)strength and stiffness over conventional metal alloys. As summarized in Table 1.3,the approach is to use strong,stiff fibers to reinforce a relatively weaker,less stiff matrix.Both the fiber and matrix can be a polymer,a metal,or a ceramic. Table 1.1 Drivers for Improved Material for Aerospace Applications ·Weight Reduction Improved Performance -increased range -smoother,more aerodynamic form -reduced fuel cost -special aeroelastic properties -higher pay load -increased temperature capability -increased maneuverability -improved damage tolerance .Reduced Acquisition Cost -reduced detectability -reduced fabrication cost Reduced Through-Life Support Cost -improved“"Hy-to-buy”ratio resistance to fatigue and corrosion -reduced assembly costs -resistance to mechanical damage
INTRODUCTION AND OVERVIEW 3 1.2 Drivers for Improved Airframe Materials Weight saving through increased specific strength or stiffness is a major driver for the development of materials for airframes. 2 However, as listed in Table 1.1, there are many other incentives for the introduction of a new material. A crucial issue in changing to a new material, even when there are clear performance benefits such as weight saving to be gained, is affordability. This includes procurement (up front) cost (currently the main criterion) and throughlife support cost (i.e., cost of ownership, including maintenance and repair). Thus the benefits of weight savings must be balanced against the cost. Approximate values that may be placed on saving 1 kilogram of weight on a range of aircraft types are listed in Table 1.2. In choosing new materials for airframe applications, it is essential to ensure that there are no compromises in the levels of safety achievable with conventional alloys. Retention of high levels of residual strength in the presence of typical damage for the particular material (damage tolerance) is a critical issue. Durability, the resistance to cyclic stress or environmental degradation and damage, through the service life is also a major factor in determining through-life support costs. The rate of damage growth and tolerance to damage determine the frequency and cost of inspections and the need for repairs throughout the life of the structure. 1.3 High-Performance Fiber Composite Concepts The fiber composite approach can provide significant improvements in specific (property/density) strength and stiffness over conventional metal alloys. As summarized in Table 1.3, the approach is to use strong, stiff fibers to reinforce a relatively weaker, less stiff matrix. Both the fiber and matrix can be a polymer, a metal, or a ceramic. Table 1.1 Drivers for Improved Material for Aerospace Applications • Weight Reduction • Improved Performance - increased range - smoother, more aerodynamic form - reduced fuel cost - special aeroelastic properties - higher pay load - increased temperature capability - increased maneuverability - improved damage tolerance • Reduced Acquisition Cost - reduced detectability - reduced fabrication cost • Reduced Through-Life Support Cost - improved "fly-to-buy" ratio - resistance to fatigue and corrosion - reduced assembly costs - resistance to mechanical damage
COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Table 1.2 Approximate Actual(US$/kg)Values of Saving One Unit of Weight: Costing Based on Some Late 1980s Estimates ●small civil$80 advanced fighter $500 .civil helicopter $80-$200 ·VTOL$800 military helicopter $400 ·SST$1500 ·large transport$300 Space Shuttle $45,000 large commercial $500 Chapter 2 describes the basic principles(micromechanics)of fiber composite materials.As an example,to a good first approximation,the stiffness under loading in the fiber direction(unidirectional fibers)may be determined by the simple law of mixtures.This is simply a sum of the volume (or area)fraction of the fibers and the matrix multiplied by the elastic modulus.The strength estimation is similar(for a reasonably high fiber-volume fraction)but with each elastic modulus multiplied by the breaking strain of the first-failing component. In the case of carbon fiber/epoxy composites,this is generally the fiber-breaking strain.If,however,the lowest failure strain is that of the matrix,the first failure event may be the development of extensive matrix cracking,rather than total fracture.This damage may or may not be defined as failure of the composite. However,toughness is usually much more than the sum of the toughness of each of the components because it depends also on the properties of the fiber/ matrix interface.Therefore,brittle materials such as glass fibers and polyester resin,when combined,produce a tough,strong composite,most familiarly known as fiberglass,used in a wide range of structural applications. Control of the strength of the fiber/matrix interface is of paramount importance for toughness,particularly when both the fiber and the matrix are brittle.If the interface is too strong,a crack in the matrix can propagate directly through fibers in its path.Thus it is important that the interface is able to disbond Table 1.3 Summary of the Approach for Development of a High-Performance Fiber Composite ·Fibers ·Polymer Matrix 。Composite stiff/strong/brittle/low -low stiffness and strength toughness through density ductile or brittle synergistic action -high temperature -can be polymer,metal, (woodlike) capability or ceramic -high strength and able to carry major load -transmits load to and stiffness in fiber as reinforcement from fiber direction,weak at -usually continuous -forms shape and protects angles to fiber axis -oriented for principal fiber tailor fiber directions to stresses optimize properties
COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Table 1.2 Approximate Actual (US$/kg) Values of Saving One Unit of Weight: Costing Based on Some Late 1980s Estimates • small civil $80 • civil helicopter $80-$200 • military helicopter $400 • large transport $300 • large commercial $500 • advanced fighter $500 • VTOL $800 * SST $1500 • Space Shuttle $45,000 Chapter 2 describes the basic principles (micromechanics) of fiber composite materials. As an example, to a good first approximation, the stiffness under loading in the fiber direction (unidirectional fibers) may be determined by the simple law of mixtures. This is simply a sum of the volume (or area) fraction of the fibers and the matrix multiplied by the elastic modulus. The strength estimation is similar (for a reasonably high fiber-volume fraction) but with each elastic modulus multiplied by the breaking strain of the first-failing component. In the case of carbon fiber/epoxy composites, this is generally the fiber-breaking strain. If, however, the lowest failure strain is that of the matrix, the first failure event may be the development of extensive matrix cracking, rather than total fracture. This damage may or may not be defined as failure of the composite. However, toughness is usually much more than the sum of the toughness of each of the components because it depends also on the properties of the fiber/ matrix interface. Therefore, brittle materials such as glass fibers and polyester resin, when combined, produce a tough, strong composite, most familiarly known as fiberglass, used in a wide range of structural applications. Control of the strength of the fiber/matrix interface is of paramount importance for toughness, particularly when both the fiber and the matrix are brittle. If the interface is too strong, a crack in the matrix can propagate directly through fibers in its path. Thus it is important that the interface is able to disbond Table 1.3 Summary of the Approach for Development of a High-Performance Fiber Composite • Fibers t Polymer Matrix • Composite - stiff/strong/brittle/low - low stiffness and strength - toughness through density ductile or brittle synergistic action - high temperature - can be polymer, metal, (woodlike) capability or ceramic - high strength and - able to carry major load - transmits load to and stiffness in fiber as reinforcement from fiber direction, weak at - usually continuous - forms shape and protects angles to fiber axis - oriented for principal fiber - tailor fiber directions to stresses optimize properties
INTRODUCTION AND OVERVIEW 5 at a modest stress level,deflecting the crack and thereby avoiding fiber failure. However,if the interface is too weak,the composite will have unacceptably low transverse properties.As discussed in more detail in Chapter 2,several other mechanisms contribute to energy absorbed in fracture and thus to toughness, including fiber disbonding and pullout,matrix deformation,and bridging of the cracked region by unbroken fibers. The composite structure is arranged (tailored)during manufacture of the component with the fibers orientated in various directions in sufficient concentrations to provide the required strength and stiffness(Chapter 12).For in-plane loading,this is usually achieved using a laminated or plywood type of construction consisting of layers or plies of unidirectional or bi-directional orientated fibers.This concept is illustrated in Figure 1.3 for an aircraft wing. Alternatively,the fibers may be arranged by a variety of advanced textile techniques,such as weaving,braiding,or filament winding. Thus to obtain the desired mechanical properties,the fiber layers or plies in a laminate are arranged at angles from 0 to 90 relative to the 0 primary loading direction.However,certain sequence and symmetry rules must be obeyed to avoid distortion of the component after cure or under service loading (as described in Chapters 6 and 12).For simplicity the plies are most often based on combinations of0°，±45°，and90°orientations.The laminate is stiffest and strongest(in-plane)in the direction with the highest concentratio of 0 fibers, Ref.Axis (spanwise) Torque Spanwise bending moment Shear Fig.1.3 Tailoring of fiber directions for the applied loads in a composite wing skin Taken from Ref.1
INTRODUCTION AND OVERVIEW 5 at a modest stress level, deflecting the crack and thereby avoiding fiber failure. However, if the interface is too weak, the composite will have unacceptably low transverse properties. As discussed in more detail in Chapter 2, several other mechanisms contribute to energy absorbed in fracture and thus to toughness, including fiber disbonding and pullout, matrix deformation, and bridging of the cracked region by unbroken fibers. The composite structure is arranged (tailored) during manufacture of the component with the fibers orientated in various directions in sufficient concentrations to provide the required strength and stiffness (Chapter 12). For in-plane loading, this is usually achieved using a laminated or plywood type of construction consisting of layers or plies of unidirectional or bi-directional orientated fibers. This concept is illustrated in Figure 1.3 for an aircraft wing. Alternatively, the fibers may be arranged by a variety of advanced textile techniques, such as weaving, braiding, or filament winding. Thus to obtain the desired mechanical properties, the fiber layers or plies in a laminate are arranged at angles from 0 ° to 90 ° relative to the 0 ° primary loading direction. However, certain sequence and symmetry rules must be obeyed to avoid distortion of the component after cure or under service loading (as described in Chapters 6 and 12). For simplicity the plies are most often based on combinations of 0 °, + 45 °, and 90 ° orientations. The laminate is ~tiffest and strongest (in-plane) in the direction with the highest concentratio'~ of 0 ° fibers, Ref. Axis (spanwtse) ~-~ Torque Vertk Sheer Fig. 1.3 Tailoring of fiber directions for the applied loads in a composite wing skin. Taken from Ref. 1
6 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES but it will have much reduced strength and stiffness in other directions-the laminate is then said to be orthotropic. When the ply configuration is made of equal numbers of plies at0°，±45°，and 90the in-plane mechanical properties do not vary with loading direction and the composite is then said to be quasi-isotropic.A similar situation arises with a 060 ply configuration.The quasi-isotropic ply configuration is used when in-plane loading is bi-directional.Because the quasi-isotropic configuration has a stress concentration factor(similar to that of an isotropic material),it is also used where local stresses are high,such as in a mechanical joint.However,for most cases,the quasi-isotropic configuration is an inefficient use of the composite material. 1.4 Fiber Reinforcements As described in Chapter 3,continuous strong,stiff fibers can be made from the light elements;carbon and boron,and the compounds silicone oxide(silica and silica-based glasses),silicon carbide,and silicon nitride.Fibers can also be made from organic materials based on long-chain molecules of carbon,hydrogen,and nitrogen.Such fibers include aramid (Kevlar)fibers.Fibers may be available in the form of single large-diameter filaments or as tows (or rovings)consisting of many thousands of filaments.For example,boron fibers formed by chemical vapor deposition (CVD)are produced as single filaments with a diameter of over 100 um.Carbon fibers,formed by pyrolysis of a polymer precursor (polyacrylonitrile;PAN),are produced as a filament diameter of about 8 um and supplied in a tow (bundle of filaments)with up to 2.5 x 10 filaments. Chemical Vapor deposition and other techniques can make short ultra-strong and stiff fibers called whiskers.These are filamentary single crystals having diameters in the range 1-10 um and length-to-diameter ratios up to 10,000.With the correct deposition techniques,whiskers can have strengths approaching the theoretical maximum of one tenth of the Young's modulus.This high level of strength results from the perfection of the crystal structure and freedom from cracklike flaws.Whiskers can be made from various materials,including SiC, Al2O3,C,and B4C. In the early 1990s,a new form of carbon called carbon nanotubes was discovered.These are essentially sheets of hexagonal graphite basal plane rolled up into a tube,with a morphology determined by the way in which the sheet is rolled up.The tube walls may be made of single or double layers;typically, length is in the range 0.6-8 nm.They can be produced by a variety of processes, including arc-discharge and CVD.As may be expected,carbon in this form has exceptionally high strength and stiffness.Elastic moduli of over 1000 GPa (1 TPa)and strengths over 100 GPa are quoted,although the minute dimensions and wall geometry of the tubes makes measurement extremely difficult
6 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES but it will have much reduced strength and stiffness in other directions--the laminate is then said to be orthotropic. When the ply configuration is made of equal numbers of plies at 0 °, ___ 45 °, and 90 ° the in-plane mechanical properties do not vary with loading direction and the composite is then said to be quasi-isotropic. A similar situation arises with a 0 ° +__ 60 ° ply configuration. The quasi-isotropic ply configuration is used when in-plane loading is bi-directional. Because the quasi-isotropic configuration has a stress concentration factor (similar to that of an isotropic material), it is also used where local stresses are high, such as in a mechanical joint. However, for most cases, the quasi-isotropic configuration is an inefficient use of the composite material. 1.4 Fiber Reinforcements As described in Chapter 3, continuous strong, stiff fibers can be made from the light elements; carbon and boron, and the compounds silicone oxide (silica and silica-based glasses), silicon carbide, and silicon nitride. Fibers can also be made from organic materials based on long-chain molecules of carbon, hydrogen, and nitrogen. Such fibers include aramid (Kevlar) fibers. Fibers may be available in the form of single large-diameter filaments or as tows (or rovings) consisting of many thousands of filaments. For example, boron fibers formed by chemical vapor deposition (CVD) are produced as single filaments with a diameter of over 100 p~m. Carbon fibers, formed by pyrolysis of a polymer precursor (polyacrylonitrile; PAN), are produced as a filament diameter of about 8 Ixm and supplied in a tow (bundle of filaments) with up to 2.5 x 10 4 filaments. Chemical Vapor deposition and other techniques can make short ultra-strong and stiff fibers called whiskers. These are filamentary single crystals having diameters in the range 1 - 10 Izm and length-to-diameter ratios up to 10,000. With the correct deposition techniques, whiskers can have strengths approaching the theoretical maximum of one tenth of the Young's modulus. This high level of strength results from the perfection of the crystal structure and freedom from cracklike flaws. Whiskers can be made from various materials, including SiC, A1203, C, and B4C. In the early 1990s, a new form of carbon called carbon nanotubes was discovered. 3 These are essentially sheets of hexagonal graphite basal plane rolled up into a tube, with a morphology determined by the way in which the sheet is rolled up. The tube walls may be made of single or double layers; typically, length is in the range 0.6-8 nm. They can be produced by a variety of processes, including arc-discharge and CVD. As may be expected, carbon in this form has exceptionally high strength and stiffness. Elastic moduli of over 1000 GPa (1 TPa) and strengths over 100 GPa are quoted, although the minute dimensions and wall geometry of the tubes makes measurement extremely difficult
INTRODUCTION AND OVERVIEW 7 Whiskers(with some exceptions)are expensive and difficult to incorporate into composites with high degrees of orientation and alignment.So,despite their early discovery,they have not been exploited in any practical composites. Although nanotubes are also expensive and similarly difficult to process into composites,they have such attractive mechanical properties and potential for relatively cheap manufacture that many R&D programs are focused on their exploitation.However,significant technological developments will be required to make composites based on these materials practically and economically feasible. Textile technology has been developed to produce special reinforcing fabrics from continuous fibers,mainly glass,carbon,or aramid.Small-diameter fiber tows may be woven to produce a wide range of fabrics;simple examples are plain weave or satin weave cloths.Fabrics can also be woven from two or more types of fiber,for example,with carbon fibers in the 0 or warp direction (the roll direction)and glass or aramid in the 90 weft direction. To avoid fiber crimping (waviness)associated with weaving,a textile approach can be used in which the fibers are held in place by a knitting yarn.The resulting materials are called non-crimp fabrics,and these can contain fibers orientated at0°，90°，and±45°in any specified proportions..Because of the elimination of fiber waviness,composites based on non-crimp fabric show a significant improvement in compression strength compared with those based on woven materials.Stiffness in both tension and compression is also improved by around 10%. Fiber preforms ready for matrix impregnation to form the component can be produced by several techniques including weaving,braiding,and knitting. Advanced weaving and braiding techniques are used to produce preforms with 3-D reinforcement,as described in Chapter 14.Three-dimensional weaving is extensively employed for the manufacture of carbon/carbon composites, described later. 1.5 Matrices The matrix,which may be a polymer,metal,or ceramic,forms the shape of the component and serves the following additional functions:1)transfers load into and out of the fibers,2)separates the fibers to prevent failure of adjacent fibers when one fails,and 3)protects the fiber from the environment.The strength of the fiber/matrix interfacial bond is crucial in determining toughness of the composite.The interface,known as the interphase,is regarded as the third phase in the composite because the matrix structure is modified close to the fiber surface.The interface is even more complex in some fibers,notably glass fibers, which are pre-coated with a sizing agent to improve bond strength,to improve environmental durability,or simply to reduce handling damage
INTRODUCTION AND OVERVIEW 7 Whiskers (with some exceptions) are expensive and difficult to incorporate into composites with high degrees of orientation and alignment. So, despite their early discovery, they have not been exploited in any practical composites. Although nanotubes are also expensive and similarly difficult to process into composites, they have such attractive mechanical properties and potential for relatively cheap manufacture that many R&D programs are focused on their exploitation. However, significant technological developments will be required to make composites based on these materials practically and economically feasible. Textile technology has been developed to produce special reinforcing fabrics from continuous fibers, mainly glass, carbon, or aramid. Small-diameter fiber tows may be woven to produce a wide range of fabrics; simple examples are plain weave or satin weave cloths. Fabrics can also be woven from two or more types of fiber, for example, with carbon fibers in the 0 ° or warp direction (the roll direction) and glass or aramid in the 90 ° weft direction. To avoid fiber crimping (waviness) associated with weaving, a textile approach can be used in which the fibers are held in place by a knitting yam. The resulting materials are called non-crimp fabrics, and these can contain fibers orientated at 0 °, 90 °, and _ 45 ° in any specified proportions. Because of the elimination of fiber waviness, composites based on non-crimp fabric show a significant improvement in compression strength compared with those based on woven materials. Stiffness in both tension and compression is also improved by around 10%. Fiber preforms ready for matrix impregnation to form the component can be produced by several techniques including weaving, braiding, and knitting. Advanced weaving and braiding techniques are used to produce preforms with 3-D reinforcement, as described in Chapter 14. Three-dimensional weaving is extensively employed for the manufacture of carbon/carbon composites, described later. 1.5 Matrices The matrix, which may be a polymer, metal, or ceramic, forms the shape of the component and serves the following additional functions: 1) transfers load into and out of the fibers, 2) separates the fibers to prevent failure of adjacent fibers when one fails, and 3) protects the fiber from the environment. The strength of the fiber/matrix interfacial bond is crucial in determining toughness of the composite. The interface, known as the interphase, is regarded as the third phase in the composite because the matrix structure is modified close to the fiber surface. The interface is even more complex in some fibers, notably glass fibers, which are pre-coated with a sizing agent to improve bond strength, to improve environmental durability, or simply to reduce handling damage
8 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Properties of the composite that are significantly affected by the properties of the matrix (matrix-dominated properties)include:1)temperature and environmental resistance,2)longitudinal compression strength,3)transverse tensile strength,and 4)shear strength. The matrix may be brittle or tough.Figure 1.4 shows the inherent toughness of some candidate materials. Economic production requires that the techniques used for matrix introduction allow simple low-cost formation of the composite without damaging or misaligning the fibers.The simplest method is to infiltrate an aligned fiber bed with a low-viscosity liquid that is then converted by chemical reaction or by cooling to form a continuous solid matrix with the desired properties. Alternatively,single fibers,tows of fibers,or sheets of aligned fibers may be coated or intermingled with solid matrix or matrix precursor and the continuous matrix formed by flowing the coatings together (and curing if required)under heat and pressure. 1.5.1 Polymers Chapter 4 discusses the thermosetting or thermoplastic polymers that are used for the matrix of polymer composites.Thermosetting polymers are long-chain molecules that cure by cross-linking to form a fully three-dimensional network and cannot be melted and reformed.They have the great advantage that they allow fabrication of composites at relatively low temperatures and pressures since they pass through a low-viscosity stage before polymerization and cross- 3 Aluminium Thermoplastics Toughened Polymethyl Unmodified Glass Alloys Epoxies methacrylate Epoxies Fig.1.4 Toughness of some materials used as matrices in advanced fiber composites
8 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Properties of the composite that are significantly affected by the properties of the matrix (matrix-dominated properties) include: 1) temperature and environmental resistance, 2) longitudinal compression strength, 3) transverse tensile strength, and 4) shear strength. The matrix may be brittle or tough. Figure 1.4 shows the inherent toughness of some candidate materials. Economic production requires that the techniques used for matrix introduction allow simple low-cost formation of the composite without damaging or misaligning the fibers. The simplest method is to infiltrate an aligned fiber bed with a low-viscosity liquid that is then converted by chemical reaction or by cooling to form a continuous solid matrix with the desired properties. Alternatively, single fibers, tows of fibers, or sheets of aligned fibers may be coated or intermingled with solid matrix or matrix precursor and the continuous matrix formed by flowing the coatings together (and curing if required) under heat and pressure. 1.5.1 Polymers Chapter 4 discusses the thermosetting or thermoplastic polymers that are used for the matrix of polymer composites. Thermosetting polymers are long-chain molecules that cure by cross-linking to form a fully three-dimensional network and cannot be melted and reformed. They have the great advantage that they allow fabrication of composites at relatively low temperatures and pressures since they pass through a low-viscosity stage before polymerization and crossLu Alumlnlum Thermoplastics Alloys Fig. 1.4 Toughness of some composites. Toughened Polymethyl Unmodified Glass Epoxies methacrylate Epoxies materials used as matrices in advanced fiber
INTRODUCTION AND OVERVIEW 9 linking.The processes used to manufacture components from thermosetting polymer composites are described in detail in Chapter 5 and include: Impregnating a fiber preform with liquid resin,which is then cured(resin- transfer molding;RTM).This process requires the resin to transition through a period of low viscosity (similar to light oil). Infusing a melted resin film into a fiber preform under pressure and then curing (resin-film infusion;RFI). Pre-impregnating fiber sheet bundles or tows with a"staged"liquid resin(pre- preg)for subsequent arrangement(stacking)followed by consolidation and cure under temperature and pressure. Epoxies have excellent mechanical properties,low shrinkage and form adequate bonds to the fibers.Importantly,they pass through a low-viscosity stage during the cure,so allow the use of liquid resin-forming techniques such as RTM.Epoxy systems curing at 120C and 180Chave respectively upper service temperatures of100°Cand130-150C. Bismaleimide resins (BMIs)have excellent formability and mechanical properties similar to epoxies and can operate at higher temperatures;however, they are more costly.BMI systems curing at about 200C have upper service temperatures above 180C. High-temperature thermosetting polymers such as polyimides,curing at around 270C,allow increases up to 300C.However,they are even more expensive and much more difficult to process. Thermosetting materials generally have relatively low failure strains.This results in poor resistance to through-thickness stresses and mechanical impact damage that can cause delaminations(ply separations)in laminated composites. They also absorb atmospheric moisture,resulting in reduced matrix-dominated properties in the composite,such as elevated temperature shear and compressive strength.Recent developments have resulted in much tougher thermoset systems, some with improved moisture resistance,through modifications in resin chemistry or alloying with tougher polymeric systems,including rubbers and thermoplastics. Thermoplastic polymers,linear (none-cross-linked)polymers that can be melted and reformed,are also suitable for use as matrices.High-performance thermoplastics suitable for aircraft applications include polymers such as polyetheretherketone (PEEK),application approximately to 120C;polyether- ketone (PEK),to 145C;and polyimide (thermoplastic type),to 270C. Thermoplastic polymers have much higher strains to failure because they can undergo extensive plastic deformations resulting in significantly improved impact resistance. Because these polymers are already polymerized,they form very high viscosity liquids when melted.Thus fabrication techniques are based on processes such as resin-film(or resin-fiber)infusion and pre-preg techniques.The main approach is to coat the fibers with the resin (from a solvent solution)and
INTRODUCTION AND OVERVIEW 9 linking. The processes used to manufacture components from thermosetting polymer composites are described in detail in Chapter 5 and include: • Impregnating a fiber preform with liquid resin, which is then cured (resintransfer molding; RTM). This process requires the resin to transition through a period of low viscosity (similar to light oil). • Infusing a melted resin film into a fiber preform under pressure and then curing (resin-film infusion; RFI). • Pre-impregnating fiber sheet bundles or tows with a "staged" liquid resin (prepreg) for subsequent arrangement (stacking) followed by consolidation and cure under temperature and pressure. Epoxies have excellent mechanical properties, low shrinkage and form adequate bonds to the fibers. Importantly, they pass through a low-viscosity stage during the cure, so allow the use of liquid resin-forming techniques such as RTM. Epoxy systems curing at 120 °C and 180 °C have respectively upper service temperatures of 100°C and 130-150°C. Bismaleimide resins (BMIs) have excellent formability and mechanical properties similar to epoxies and can operate at higher temperatures; however, they are more costly. BMI systems curing at about 200°C have upper service temperatures above 180 °C. High-temperature thermosetting polymers such as polyimides, curing at around 270°C, allow increases up to 300°C. However, they are even more expensive and much more difficult to process. Thermosetting materials generally have relatively low failure strains. This results in poor resistance to through-thickness stresses and mechanical impact damage that can cause delaminations (ply separations) in laminated composites. They also absorb atmospheric moisture, resulting in reduced matrix-dominated properties in the composite, such as elevated temperature shear and compressive strength. Recent developments have resulted in much tougher thermoset systems, some with improved moisture resistance, through modifications in resin chemistry or alloying with tougher polymeric systems, including rubbers and thermoplastics. Thermoplastic polymers, linear (none-cross-linked) polymers that can be melted and reformed, are also suitable for use as matrices. High-performance thermoplastics suitable for aircraft applications include polymers such as polyetheretherketone (PEEK), application approximately to 120°C; polyetherketone (PEK), to 145°C; and polyimide (thermoplastic type), to 270°C. Thermoplastic polymers have much higher strains to failure because they can undergo extensive plastic deformations resulting in significantly improved impact resistance. Because these polymers are already polymerized, they form very high viscosity liquids when melted. Thus fabrication techniques are based on processes such as resin-film (or resin-fiber) infusion and pre-preg techniques. The main approach is to coat the fibers with the resin (from a solvent solution) and
10 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES then consolidate the part under high temperature and pressure.Alternatively, sheets of thermoplastic film can be layered between sheets of dry fiber or fibers of thermoplastic can be woven through the fibers and the composite consolidated by hot pressing. Because thermoplastics absorb little moisture,they have better hot/wet property retention than thermosetting composites.However,they are generally more expensive and are more costly to fabricate because they require elevated- temperature processing.In addition,with improvements in thermosets,even the toughness advantage is being eroded.There is little doubt that thermoplastics will be used extensively in the future for aircraft structures,particularly in areas subject to mechanical damage. 1.5.2 Metals The light metals,magnesium,aluminum,and titanium alloys (including titanium aluminides),are used to form high-performance metal-matrix composites.4 These materials offer the possibility of higher temperature service capabilities-approximately 150C,300C,500C,and >700C, respectively-and have several other advantages,as discussed later,over polymer-matrix composites.However,these advantages are offset by more costly,complex,and limited fabrication techniques. Metals often react chemically with and weaken fibers during manufacture or in service at elevated temperatures,so translation of fiber properties is often poor.The tendency for a metal to react with the fiber is termed fiber/matrix compatibility.Generally,because of compatibility problems,ceramic fibers such SiC,Al2O3,and Borsic(boron fibers coated with silicon carbide)are most suited for reinforcing metals.However,carbon fibers may be used with aluminum or magnesium matrices,provided that exposure to high temperature is minimized. Methods based on infiltration liquid metal have many advantages for aluminum,provided damaging chemical interaction between the metal and fibers does not occur and the metal is able (or is forced under pressure)to wet the fibers.The process of squeeze casting is attractive because time in contact with liquid metal is limited,minimizing chemical interaction,and the high pressure overcomes wetting difficulties.Another major advantage of this process is that alloys other than casting alloys can be employed.If the fiber does not react readily with molten metal but is easily wetted,for example, silicon carbide fibers in aluminum,more conventional casting techniques such as investment casting may be used.Conventional casting has the major advantage that the size of the component that can be formed is much less limited and requires only simple equipment.Even carbon fibers can be used if the casting process is very rapid,particularly if the fibers are coated with a barrier layer such as silicon carbide,thus minimizing reaction with the molten metal
10 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES then consolidate the part under high temperature and pressure. Alternatively, sheets of thermoplastic film can be layered between sheets of dry fiber or fibers of thermoplastic can be woven through the fibers and the composite consolidated by hot pressing. Because thermoplastics absorb little moisture, they have better hot/wet property retention than thermosetting composites. However, they are generally more expensive and are more costly to fabricate because they require elevatedtemperature processing. In addition, with improvements in thermosets, even the toughness advantage is being eroded. There is little doubt that thermoplastics will be used extensively in the future for aircraft structures, particularly in areas subject to mechanical damage. 1.5.2 Metals The light metals, magnesium, aluminum, and titanium alloys (including titanium aluminides), are used to form high-performance metal-matrix composites. 4 These materials offer the possibility of higher temperature service capabilities--approximately 150°C, 300°C, 500°C, and >700°C, respectively--and have several other advantages, as discussed later, over polymer-matrix composites. However, these advantages are offset by more costly, complex, and limited fabrication techniques. Metals often react chemically with and weaken fibers during manufacture or in service at elevated temperatures, so translation of fiber properties is often poor. The tendency for a metal to react with the fiber is termed fiber/matrix compatibility. Generally, because of compatibility problems, ceramic fibers such SiC, A1203, and Borsic (boron fibers coated with silicon carbide) are most suited for reinforcing metals. However, carbon fibers may be used with aluminum or magnesium matrices, provided that exposure to high temperature is minimized. Methods based on infiltration liquid metal have many advantages for aluminum, provided damaging chemical interaction between the metal and fibers does not occur and the metal is able (or is forced under pressure) to wet the fibers. The process of squeeze casting is attractive because time in contact with liquid metal is limited, minimizing chemical interaction, and the high pressure overcomes wetting difficulties. Another major advantage of this process is that alloys other than casting alloys can be employed. If the fiber does not react readily with molten metal but is easily wetted, for example, silicon carbide fibers in aluminum, more conventional casting techniques such as investment casting may be used. Conventional casting has the major advantage that the size of the component that can be formed is much less limited and requires only simple equipment. Even carbon fibers can be used if the casting process is very rapid, particularly if the fibers are coated with a barrier layer such as silicon carbide, thus minimizing reaction with the molten metal
