INTRODUCTION AND OVERVIEW 19 A more sophisticated approach to self-healing is to use glass-forming silicon- and boron-based particulate materials in the carbon matrix,which reacts with oxygen to form a glass.The glass flows into cracks,sealing them off from oxygen penetration.More sophisticated approaches involve the inclusion of organic compounds in the matrix precursor materials to inhibit oxidation.If successful barrier layers could be developed,carbon/carbon could be used extensively in gas turbines and in the airframes of future hypersonic aircraft. Finally,carbon/carbon is widely used for aircraft brake disk pads,where its combination of low weight,high-temperature capability,thermal conductivity, and excellent wear resistance results in considerable weight savings. 1.8 Hybrid Metal/PMC Composites Structural metals,such as aluminum alloys and composites,including carbon/ epoxy,have a variety of advantages and disadvantages for airframe applications.For example,metals are prone to fatigue cracking but PMCs are not;PMCs are easily damaged by low-energy mechanical impacts but metals are not.Thus,the potential exists to combine these materials in such a way as to get the best of both materials. One such approach is the aluminum/fiber composite hybrid laminate,3 which consists of thin sheets of aluminum alloy bonded with a fiber-reinforced adhesive.When a crack grows through the metal,the fibers,which are highly resistant to fatigue damage,are left spanning or bridging the crack in its wake (Fig.1.7).The result is a reduction in crack growth rate by approximately one order of magnitude and an insensitivity to crack length.However,the fibers have little infuence on crack initiation and,indeed,because the hybrid composite has relatively low modulus,the increased strain in the aluminum alloy can encourage earlier crack initiation.The fibers also significantly increase the post- yield strength compared with unreinforced aluminum alloy,and the composite has a much higher damping capacity. Disadvantages of these materials include sensitivity to blunt notches due to the inability of the fibers to withstand very high strain levels.Thus,the notch insensitivity of metals is not retained in the hybrid.Also,depending on the reinforcement used,the elastic modulus of the hybrid is generally lower than aluminum alloys,however,this is compensated for by a reduction of specific gravity of between 10-15%.Another problem is cost,which is typically 7-10 times that of standard aerospace-grade aluminum alloys. The aluminum alloy is generally either 2024 T3 or 7475 T761,0.2-0.4 mm thick.The composite is aramid (Kevlar)or glass fibers in an epoxy nitrile adhesive,around 0.2 mm thick for unidirectional reinforcement,or 0.25-0.35 mm thick for(glass reinforcement only)cross-ply.With aramid reinforcement,the laminate is called ARALL (aramid reinforced aluminum laminate),and with glass fiber,GLARE.Because of the sensitivity of aramid fibers to compressive
INTRODUCTION AND OVERVIEW 19 A more sophisticated approach to self-healing is to use glass-forming siliconand boron-based particulate materials in the carbon matrix, which reacts with oxygen to form a glass. The glass flows into cracks, sealing them off from oxygen penetration. More sophisticated approaches involve the inclusion of organic compounds in the matrix precursor materials to inhibit oxidation. If successful barrier layers could be developed, carbon/carbon could be used extensively in gas turbines and in the airframes of future hypersonic aircraft. Finally, carbon/carbon is widely used for aircraft brake disk pads, where its combination of low weight, high-temperature capability, thermal conductivity, and excellent wear resistance results in considerable weight savings. 1.8 Hybrid Metal/PMC Composites Structural metals, such as aluminum alloys and composites, including carbon/ epoxy, have a variety of advantages and disadvantages for airframe applications. For example, metals are prone to fatigue cracking but PMCs are not; PMCs are easily damaged by low-energy mechanical impacts but metals are not. Thus, the potential exists to combine these materials in such a way as to get the best of both materials. One such approach is the aluminum/fiber composite hybrid laminate, 13 which consists of thin sheets of aluminum alloy bonded with a fiber-reinforced adhesive. When a crack grows through the metal, the fibers, which are highly resistant to fatigue damage, are left spanning or bridging the crack in its wake (Fig. 1.7). The result is a reduction in crack growth rate by approximately one order of magnitude and an insensitivity to crack length. However, the fibers have little influence on crack initiation and, indeed, because the hybrid composite has relatively low modulus, the increased strain in the aluminum alloy can encourage earlier crack initiation. The fibers also significantly increase the postyield strength compared with unreinforced aluminum alloy, and the composite has a much higher damping capacity. Disadvantages of these materials include sensitivity to blunt notches due to the inability of the fibers to withstand very high strain levels. Thus, the notch insensitivity of metals is not retained in the hybrid. Also, depending on the reinforcement used, the elastic modulus of the hybrid is generally lower than aluminum alloys, however, this is compensated for by a reduction of specific gravity of between 10-15%. Another problem is cost, which is typically 7-10 times that of standard aerospace-grade aluminum alloys. The aluminum alloy is generally either 2024 T3 or 7475 T761, 0.2-0.4 mm thick. The composite is aramid (Kevlar) or glass fibers in an epoxy nitrile adhesive, around 0.2 mm thick for unidirectional reinforcement, or 0.25-0.35 mm thick for (glass reinforcement only) cross-ply. With aramid reinforcement, the laminate is called ARALL (aramid reinforced aluminum laminate), and with glass fiber, GLARE. Because of the sensitivity of aramid fibers to compressive
20 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Fig.1.7 Schematic diagram of hybrid consisting of thin (~0.4 mm)aluminum alloy sheets bonded with an epoxy film adhesive reinforced with glass or aramid fibers.The fibers are left spanning or bridging fatigue cracks if they develop in the aluminum sheets, vastly reducing the rate of crack growth.Taken from Ref.13. stresses and the favorable residual strength that is produced,ARALL may be pre-stretched.This also overcomes,at a cost,the adverse residual stresses arising from the differences in thermal expansion coefficient between aramid,or glass, and aluminum.GLARE does not require pre-stretching as the high-strain glass fiber used is less susceptible to compressive stresses.Consequently,the glass fibers can be cross-plied to give crack growth resistance in two orthogonal directions as may be required for a fuselage structure.Although GLARE has a lower modulus than conventional aluminum alloys,with a reduction of around 20%(particularly with cross-plied fibers),it has the best resistance to fatigue crack growth. Significant weight savings-20%or so-can be achieved in fatigue-prone regions such as pressurized fuselage skins and stiffeners and lower wing skins by the use of these materials.The hybrid composites are also suited to high-impact regions such as leading edges and inboard flaps and to components subject to mishandling,such as doors. For applications requiring higher stiffness and strength,as well as a higher temperature,capability studies have been conducted3 on hybrid laminates made of thin sheets of titanium alloy (Ti-6Al-4V)and a low-modulus carbon fiber composite.The matrix for the composite and adhesive is a thermoplastic (PEEK). This laminate is reported to have excellent resistance to fatigue crack growth as well as good blunt-notch strength
20 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES r Fig. 1.7 Schematic diagram of hybrid consisting of thin (~0.4 mm) aluminum alloy sheets bonded with an epoxy film adhesive reinforced with glass or aramid fibers. The fibers are left spanning or bridging fatigue cracks if they develop in the aluminum sheets, vastly reducing the rate of crack growth. Taken from Ref. 13. stresses and the favorable residual strength that is produced, ARALL may be pre-stretched. This also overcomes, at a cost, the adverse residual stresses arising from the differences in thermal expansion coefficient between aramid, or glass, and aluminum. GLARE does not require pre-stretching as the high-strain glass fiber used is less susceptible to compressive stresses. Consequently, the glass fibers can be cross-plied to give crack growth resistance in two orthogonal directions as may be required for a fuselage structure. Although GLARE has a lower modulus than conventional aluminum alloys, with a reduction of around 20% (particularly with cross-plied fibers), it has the best resistance to fatigue crack growth. Significant weight savings--20% or so---can be achieved in fatigue-prone regions such as pressurized fuselage skins and stiffeners and lower wing skins by the use of these materials. The hybrid composites are also suited to high-impact regions such as leading edges and inboard flaps and to components subject to mishandling, such as doors. For applications requiring higher stiffness and strength, as well as a higher temperature, capability studies have been conducted 13 on hybrid laminates made of thin sheets of titanium alloy (Ti-6A1-4V) and a low-modulus carbon fiber composite. The matrix for the composite and adhesive is a thermoplastic (PEEK). This laminate is reported to have excellent resistance to fatigue crack growth as well as good blunt-notch strength
INTRODUCTION AND OVERVIEW 21 References Hoskin,B.C.,and Baker,A.A.(eds.),Composite Materials for Aircraft Structure, AIAA Education Series,AIAA,New York,1986. 2Baker,A.A.,"Development and Potential of Advanced Fibre Composites for Aerospace Applications"Materials Forum,Vol.11,1988,pp.217-231. 3Thostenson,E.T.,Ren,Z.,and Chou,T.-W.,"Advances in the Science and Technology of Carbon Nanotubes and Their Composites:A Review,"Composites Science and Technology (UK),Vol.61,No.13,Oct.2001,pp.1899-1912. Cline,T.W.,and P.J.Withers,An Introduction to Metal-Matrix Composites, Cambridge,England,UK,1993. Cratchley,D.,Baker,A.A.,and Jackson,P.W.,"Mechanical Behaviour of a Fibre Reinforced Metal and Its Effect Upon Engineering Applications"Metal-Matrix Composites,American Society for Testing Materials STP,1967,p.438. Richerson,D.W.,"Ceramic Matrix Composites,"Composite Materials Handbook, edited by P.K.Mallick,Marcel Dekker,1997. Rawal,S.,"Metal-Matrix Composites for Space Application,"Journal of Metal,Vol. 53,2001,pp.14-17. 8Baker,A.A.,"Carbon Fibre Reinforced Metals-A Review of the Current Technology,"Materials Science and Engineering,Vol.17,1975,pp.177-208. Leyens,C.,Kocian.F.,Hausman,J.,and Kaysser,W.A.,"Materials and Design Concepts for High-Performance Compressor Components,"Aerospace Science and Technology,Vol.7,2003,pp.201-210. Lloyd,D.J."Particle Reinforced Aluminum and Magnesium Matrix Composites." International Materials Reviews,Vol.39,1994,p.1. Parlier,M..and Ritti,M.H.,"State of the Art and Perspectives for Oxide/Oxide Composites,"Aerospace Technology,Vol.7,2003,pp.211-221. 2Buckley J.D.,and Edie,D.D.(eds.),Carbon-Carbon Materials and Composites, Noyes,Park Ridge,NJ,1993. 13volt,A..and Willem,J.(eds.),Fibre Metal Laminates:An Introduction,Kluwer Academic Publishers,2001. Bibliography Chawla,K.K.,Composite Materials Science and Engineering,Springer-Verlag, New York. Kelly,A.,and Zweben,C.(eds.),Comprehensive Composite Materials,Elsevier,2000. Mallick,P.K.,(ed.),Composite Materials Handbook,Marcel Dekker,New York,1997. Middleton,D.H.(ed.),Composite Materials in Aircraft Structures,Longmans,UK,1990. Niu,M.C.Y.,Composite Airframe Structures,Comilit Press,Hong Kong,1992. Peel,C.J.."Advances in Materials for Aerospace,"Aeronautical Journal,Vol.100, 1996,pp.487-506
INTRODUCTION AND OVERVIEW 21 References I Hoskin, B. C., and Baker, A. A. (eds.), Composite Materials for Aircraft Structure, AIAA Education Series, AIAA, New York, 1986. 2Baker, A. A., "Development and Potential of Advanced Fibre Composites for Aerospace Applications" Materials Forum, Vol. 11, 1988, pp. 217-231. 3Thostenson, E. T., Ren, Z., and Chou, T.-W., "Advances in the Science and Technology of Carbon Nanotubes and Their Composites: A Review," Composites Science and Technology (UK), Vol. 61, No. 13, Oct. 2001, pp. 1899-1912. 4Cline, T. W., and P. J. Withers, An Introduction to Metal-Matrix Composites, Cambridge, England, UK, 1993. 5Cratchley, D., Baker, A. A., and Jackson, P. W., "Mechanical Behaviour of a Fibre Reinforced Metal and Its Effect Upon Engineering Applications" Metal-Matrix Composites, American Society for Testing Materials STP, 1967, p. 438. 6Richerson, D. W., "Ceramic Matrix Composites," Composite Materials Handbook, edited by P. K. MaUick, Marcel Dekker, 1997. 7Rawal, S., "Metal-Matrix Composites for Space Application," Journal of Metal, Vol. 53, 2001, pp. 14-17. 8Baker, A. A., "Carbon Fibre Reinforced Metals--A Review of the Current Technology," Materials Science and Engineering, Vol. 17, 1975, pp. 177-208. 9Leyens, C., Kocian. F., Hausman, J., and Kaysser, W. A., "Materials and Design Concepts for High-Performance Compressor Components," Aerospace Science and Technology, Vol. 7, 2003, pp. 201-210. 1°Lloyd, D. J., "Particle Reinforced Aluminum and Magnesium Matrix Composites." International Materials Reviews, Vol. 39, 1994, p. 1. 11Parlier, M., and Ritti, M. H., "State of the Art and Perspectives for Oxide/Oxide Composites," Aerospace Technology, Vol. 7, 2003, pp. 211-221. 12Bucldey J. D., and Edie, D. D. (eds.), Carbon-Carbon Materials and Composites, Noyes, Park Ridge, NJ, 1993. 13Volt, A., and WiUem, J. (eds.), Fibre Metal Laminates: An Introduction, Kluwer Academic Publishers, 2001. Bibliography Chawla, K. K., Composite Materials Science and Engineering, Springer-Verlag, New York. Kelly, A., and Zweben, C. (eds.), Comprehensive Composite Materials, Elsevier, 2000. Mallick, P. K., (ed.), Composite Materials Handbook, Marcel Dekker, New York, 1997. Middleton, D. H. (ed.), Composite Materials in Aircraft Structures, Longrnans, UK, 1990. Niu, M. C. Y., Composite Airframe Structures, Comilit Press, Hong Kong, 1992. Peel, C. J., "Advances in Materials for Aerospace," Aeronautical Journal, Vol. 100, 1996, pp. 487-506
2 Basic Principles of Fiber Composite Materials 2.1 Introduction to Fiber Composite Systems A fiber composite material consists of a filamentary phase embedded in a continuous matrix phase.The aspect ratio (i.e.,ratio of length to diameter)of the filaments may vary from about 10 to infinity (for continuous fibers).Their scale, in relation to the bulk material,may range from microscopic (e.g.,8-m diameter carbon fibers in an epoxy matrix)to gross macroscopic (e.g.,25-mm diameter steel bars in concrete). Composite constituents(fibers and matrices)can be conveniently classified according to their elastic moduli E and ductility.Within the composite,the fibers may,in general,be in the form of continuous fibers,discontinuous fibers,or whiskers (very fine single crystals with lengths of the order 100-1000 um and diameters of the order 1-10 um)and may be aligned to varying degrees or randomly orientated.This classification is depicted in Figure 2.1 for a number of common fibers and matrices;also listed are examples of composites formed from these materials. 2.2 Mlcromechanical Versus Macromechanical View of Composites Fiber composites can be studied from two points of view:micromechanics and macromechanics.Micromechanical analyses are aimed at providing an understanding of the behavior of composites,usually those with unidirectional fiber reinforcement,in terms of the properties of the fibers and matrices.Models of varying degrees of sophistication are used to simulate the microstructure of the composite and hence predict its properties(such as strength and stiffness)in terms of the properties and behavior of the constituents. Macromechanics is the approach used to predict'the strength and stiffness of composite structures,as well as other properties such as distortion,on the basis of the "average"properties of the unidirectional material;namely,the longitudinal modulus E1,transverse modulus E2,major Poisson's ratio v2i and the in-plane shear modulus G12,as well as the appropriate strength values.A full analysis also 23
2 Basic Principles of Fiber Composite Materials 2.1 Introduction to Fiber Composite Systems A fiber composite material consists of a filamentary phase embedded in a continuous matrix phase. The aspect ratio (i.e., ratio of length to diameter) of the filaments may vary from about 10 to infinity (for continuous fibers). Their scale, in relation to the bulk material, may range from microscopic (e.g., 8-1xm diameter carbon fibers in an epoxy matrix) to gross macroscopic (e.g., 25-mm diameter steel bars in concrete). Composite constituents (fibers and matrices) can be conveniently classified according to their elastic moduli E and ductility. Within the composite, the fibers may, in general, be in the form of continuous fibers, discontinuous fibers, or whiskers (very fine single crystals with lengths of the order 100-1000 ixm and diameters of the order 1-10 Ixm) and may be aligned to varying degrees or randomly orientated. This classification is depicted in Figure 2.1 for a number of common fibers and matrices; also listed are examples of composites formed from these materials. 2.2 Mlcromechanical Versus Macromechanical View of Composites Fiber composites can be studied from two points of view: micromechanics and macromechanics. Micromechanical analyses are aimed at providing an understanding of the behavior of composites, usually those with unidirectional fiber reinforcement, in terms of the properties of the fibers and matrices. Models of varying degrees of sophistication are used to simulate the microstructure of the composite and hence predict its properties (such as strength and stiffness) in terms of the properties and behavior of the constituents. Macromechanics is the approach used to predict 1 the strength and stiffness of composite structures, as well as other properties such as distortion, on the basis of the "average" properties of the unidirectional material; namely, the longitudinal modulus El, transverse modulus E2, major Poisson's ratio v21 and the in-plane shear modulus G12 , as well as the appropriate strength values. A full analysis also 23
g Continuous Discontinuous ceramic chopped fibres glass whiskers COMPOSITES metal (wires) nanotubes textile textile Carbon/Epoxy Matrices Carbon/BMI Fibres Aramid/Epoxy aligned Glass/Epoxy or random Carbon/Carbon High E LowE Carbon/Glass High E Low E SiC/SiC SiC/Ti Brittle Brittle SiC/Ti aluminide Brittle Brittle carbon epoxy W/superalloy carbon hair glass polyester SiC whiskers/SiC aramid flax phenolic ceramic boron cotton titanium-aluminide polyimide alumina HD polyethylene rubber silicon-carbide wood/bamboo plaster glass/polyester glass/nylon polyester/rubber Ductile steelrubber Ductile Ductile COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES aluminum Ductile steelconcrete tungsten wire nylon nickel polysulphone nylon/polyethylene steel wire polyethylene copper polypropolene hair/plaster titanium nylon Fig.2.1 Classification of composites according to fiber and matrix properties
24 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES ++°+~ I =i I I L °~, E ~3 E T I I l + +++ C I= O 0~