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
INTRODUCTION AND OVERVIEW Table 1.4 Comparison of Carbon/Epoxy with Table 1.1 and Conventional Aluminum Alloys for Airframe Applications ·Weight Reduction ·Acquisition Cost saving 15-20%compared with -material cost increase aluminum alloys reduction due to high conversion rate cost of reduction $60-$100 per kg (low fly-to-buy ratio) reduction in number of joints reduction due to reduction in joints -fabrication cost generally increases ◆Performance ·Repair Costs smoother,more aerodynamic form -fatigue resistant,reduction -improved aeroelastic properties corrosion immune,reduction -more resistant to accoustic environment -fretting resistant,reduction -more resistant to service environment -impact sensitive,increase -improved fire containment -prone to delamination,increase improved crash resistance improved stealth properties Diffusion bonding can be employed to produce metal-matrix composites. Fibers are melt-coated,plasma is sprayed or interleaved with metal foil and then hot pressed.However,other than for the larger-diameter fibers such as boron and silicon carbide,excessive fiber breakage resulting from the high mechanical pressures used is a major problem with this approach.Additionally,if high temperatures are required to encourage metal flow,weakening of the fibers by solid-state chemical interactions is difficult to avoid. Fibers can be coated by electrodeposition or CVD to provide a continuous reinforced matrix without the need for subsequent consolidation (pressure). These approaches are much less severe than liquid metal or diffusion bonding and may be attractive for some applications.However,the range of alloys that can be produced by this route is limited,and the high-temperature properties of the matrix may be poor. The formation of a metal-matrix composite by hot pressing coated fibers is illustrated clearly in Figure 1.5,which shows an early metal-matrix composite silica fiber-reinforced aluminum,developed in the mid-1960s by Rolls Royce.The fibers are first individually coated with aluminum and then the coated fibers are hot pressed at a temperature of around 500C and a pressure of 60 MPa.In the example shown for illustrative purposes,only half of the sample has been consolidated. 1.5.3 Ceramics For much higher temperatures than can be achieved with polymer or metal matrices,the options are to employ a silica-based glass;a ceramic such as silicon
Table 1.4 INTRODUCTION AND OVERVIEW Comparison of Carbon/Epoxy with Table 1.1 and Conventional Aluminum Alloys for Airframe Applications 11 • Weight Reduction - saving 15-20% compared with aluminum alloys - cost of reduction $60-$100 per kg - reduction in number of joints • Performance - smoother, more aerodynamic form - improved aeroelastic properties - more resistant to accoustic environment - more resistant to service environment - improved fire containment - improved crash resistance - improved stealth properties • Acquisition Cost - material cost increase - reduction due to high conversion rate (low fly-to-buy ratio) - reduction due to reduction in joints - fabrication cost generally increases • Repair Costs - fatigue resistant, reduction - corrosion immune, reduction - fretting resistant, reduction - impact sensitive, increase - prone to delamination, increase Diffusion bonding can be employed to produce metal-matrix composites. Fibers are melt-coated, plasma is sprayed or interleaved with metal foil and then hot pressed. However, other than for the larger-diameter fibers such as boron and silicon carbide, excessive fiber breakage resulting from the high mechanical pressures used is a major problem with this approach. Additionally, if high temperatures are required to encourage metal flow, weakening of the fibers by solid-state chemical interactions is difficult to avoid. Fibers can be coated by electrodeposition or CVD to provide a continuous reinforced matrix without the need for subsequent consolidation (pressure). These approaches are much less severe than liquid metal or diffusion bonding and may be attractive for some applications. However, the range of alloys that can be produced by this route is limited, and the high-temperature properties of the matrix may be poor. The formation of a metal-matrix composite by hot pressing coated fibers is illustrated clearly in Figure 1.5, which shows an early metal-matrix composite silica fiber-reinforced aluminum, developed in the mid-1960s by Rolls Royce. 5 The fibers are first individually coated with aluminum and then the coated fibers are hot pressed at a temperature of around 500 °C and a pressure of 60 MPa. In the example shown for illustrative purposes, only half of the sample has been consolidated. 1.5.3 Ceramics For much higher temperatures than can be achieved with polymer or metal matrices, the options are to employ a silica-based glass; a ceramic such as silicon
12 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Fig.1.5 Photograph of a (half)hot-pressed silica fiber/aluminum matrix composite, and (right)microstructure of consolidated side showing fibers,aluminum matrix, and boundary between original fiber coatings.Taken from Ref.5. carbide,silicon nitride,or alumina;or a carbon matrix.These are called ceramic- matrix composites (CMCs). In the case of the high-modulus ceramic matrices,the fibers provide little stiffening;their purpose is to increase toughness.This is achieved mainly by blunting and deflecting cracks in the matrix and contributing to increased fracture energy through the various energy-absorbing mechanisms,such as crack bridging and fiber pull-out. Several techniques are used to form composites with ceramic matrices.These include infiltration of aligned fibers by 1)CVD,2)impregnation of fibers with a fine powder and consolidating,and 3)impregnation of fibers with a liquid ceramic precursor,generally a polymer,and converting to ceramic at elevated temperature.The powders may be added to the aligned fibers or fiber preforms by injection molding or by sol-gel techniques.Densification of powder coatings may be achieved by hot-pressing,sintering,hot isostatic pressing,or superplastic forging.In most respects,the precursor route is the most promising for ceramics because dense matrices can be produced at low temperatures without causing fiber damage,and complex components can be formed directly. Glass and glass-ceramic matrices are readily formed by consolidation of fiber preforms impregnated with fine powders applied from a dispersion or gel.The glass melts easily and flows between the fibers to form a continuous pore-free matrix.The procedure is similar to that adopted for thermoplastic matrix composites.In glass- ceramic matrices,the matrix may subsequently be crystallized by heat treatment, greatly enhancing performance at elevated temperatures. Carbon matrices may also be formed by CVD of carbon from high-carbon content gases,such as methane,propane,and benzene into a fiber preform.They can also be formed by liquid phase impregnation of fibers followed by pyrolytic
12 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Fig. 1.5 Photograph of a (half) hot-pressed silica fiber/aluminum matrix composite, and (right) microstructure of consolidated side showing fibers, aluminum matrix, and boundary between original fiber coatings. Taken from Ref. 5. carbide, silicon nitride, or alumina; 6 or a carbon matrix. These are called ceramicmatrix composites (CMCs). In the case of the high-modulus ceramic matrices, the fibers provide little stiffening; their purpose is to increase toughness. This is achieved mainly by blunting and deflecting cracks in the matrix and contributing to increased fracture energy through the various energy-absorbing mechanisms, such as crack bridging and fiber pull-out. Several techniques are used to form composites with ceramic matrices. These include infiltration of aligned fibers by 1) CVD, 2) impregnation of fibers with a fine powder and consolidating, and 3) impregnation of fibers with a liquid ceramic precursor, generally a polymer, and converting to ceramic at elevated temperature. The powders may be added to the aligned fibers or fiber preforms by injection molding or by sol-gel techniques. Densification of powder coatings may be achieved by hot-pressing, sintering, hot isostatic pressing, or superplastic forging. In most respects, the precursor route is the most promising for ceramics because dense matrices can be produced at low temperatures without causing fiber damage, and complex components can be formed directly. Glass and glass-ceramic matrices are readily formed by consolidation of fiber preforms impregnated with fine powders applied from a dispersion or gel. The glass melts easily and flows between the fibers to form a continuous pore-free matrix. The procedure is similar to that adopted for thermoplastic matrix composites. In glassceramic matrices, the matrix may subsequently be crystallized by heat treatment, greatly enhancing performance at elevated temperatures. Carbon matrices may also be formed by CVD of carbon from high-carbon content gases, such as methane, propane, and benzene into a fiber preform. They can also be formed by liquid phase impregnation of fibers followed by pyrolytic
INTRODUCTION AND OVERVIEW 13 decomposition of a precursor with a high carbon content.Suitable precursors include phenolic resin,pitch,and tar-based materials,all of which can have over 40%yield of carbon on pyrolysis.The fibers are generally carbon and the composite called carbon/carbon.Silicon carbide fibers are also used in some applications as an alternative to carbon,particularly where improved resistance to oxidation is required. With the resin-based route,standard polymer-matrix composite manufacturing processes,such as filament winding or braiding,can be used before pyrolysis. The precursor route is the most efficient for making carbon matrix composites; however,multiple impregnations and pyrolysis steps are required to produce a matrix with an acceptably low porosity level.This is a slow process resulting in high component costs.The CVD process is even slower,therefore it is mainly used to fill-in fine interconnected near-surface voids in composites produced by pyrolysis.The CVD is,however,suited to manufacture of thin-wall components. PMCs are extensively used in aerospace structures;however,carbon/epoxy is by far the most exploited so is the main focus of this book.Some current airframe applications are described in Chapter 12.Based on the drivers set out in Table 1.1,a comparison of carbon/epoxy with conventional aluminium alloys is provided in Table 1.4. 1.6 Polymer Matrix Composites The nomenclature used in the U.S.identifies the composite in the format fiber/ matrix.For example,the main composites discussed in this book are carbon fibers in an epoxy resin matrix and are referred to as carbon/epoxy or graphite/epoxy (also c/ep and gr/ep).Other common composite systems are carbon/BMI, carbon/polyimide,glass/epoxy,aramid/epoxy,and boron/epoxy.This notation can readily be expanded to specific composite systems;for example,a well-known commercial composite system,Hercules AS fibers in a 3501-6 epoxy resin matrix,is AS/3501-6.In the U.K.the terminology for carbon/epoxy is carbon fiber reinforced epoxy,or more usually,carbon fiber reinforced plastic(CFRP). 1.7 Non-polymeric Composite Systems In this section,some of the important non-polymeric composite systems are briefly discussed. 1.7.1 Metal-Matrix Composites Metal-matrix composites(MMCs),4.7.8 with continuous or discontinuous fiber reinforcement have been under development for well over 30 years,but have yet to be widely exploited. The main MMCs based on continuous fibers,and their advantages and disadvantages compared with PMCs,are listed in Table 1.5.Potential aircraft
INTRODUCTION AND OVERVIEW 13 decomposition of a precursor with a high carbon content. Suitable precursors include phenolic resin, pitch, and tar-based materials, all of which can have over 40% yield of carbon on pyrolysis. The fibers are generally carbon and the composite called carbon/carbon. Silicon carbide fibers are also used in some applications as an alternative to carbon, particularly where improved resistance to oxidation is required. With the resin-based route, standard polymer-matrix composite manufacturing processes, such as filament winding or braiding, can be used before pyrolysis. The precursor route is the most efficient for making carbon matrix composites; however, multiple impregnations and pyrolysis steps are required to produce a matrix with an acceptably low porosity level. This is a slow process resulting in high component costs. The CVD process is even slower, therefore it is mainly used to fill-in fine interconnected near-surface voids in composites produced by pyrolysis. The CVD is, however, suited to manufacture of thin-wall components. PMCs are extensively used in aerospace structures; however, carbon/epoxy is by far the most exploited so is the main focus of this book. Some current airframe applications are described in Chapter 12. Based on the drivers set out in Table 1.1, a comparison of carbon/epoxy with conventional aluminium alloys is provided in Table 1.4. 1,6 Polymer Matrix Composites The nomenclature used in the U.S. identifies the composite in the format fiber/ matrix. For example, the main composites discussed in this book are carbon fibers in an epoxy resin matrix and are referred to as carbon/epoxy or graphite/epoxy (also c/ep and gr/ep). Other common composite systems are carbon/BMI, carbon/polyimide, glass/epoxy, aramid/epoxy, and boron/epoxy. This notation can readily be expanded to specific composite systems; for example, a well-known commercial composite system, Hercules AS fibers in a 3501-6 epoxy resin matrix, is AS/3501-6. In the U.K. the terminology for carbon/epoxy is carbon fiber reinforced epoxy, or more usually, carbon fiber reinforced plastic (CFRP). 1.7 Non-polymeric Composite Systems In this section, some of the important non-polymeric composite systems are briefly discussed. 1.7.1 Metal-Matrix Composites Metal-matrix composites (MMCs), 4'7'8 with continuous or discontinuous fiber reinforcement have been under development for well over 30 years, but have yet to be widely exploited. The main MMCs based on continuous fibers, and their advantages and disadvantages compared with PMCs, are listed in Table 1.5. Potential aircraft