8 Properties of Composite Systems 8.1 Introduction The mechanical properties of simple unidirectional continuous fiber composites depend on the volume fraction and properties of the fibers (including flaw and strength distribution),the fiber/matrix bond strength,and the mechanical properties of the matrix.The alignment (waviness)of the fibers also has a significant effect on some properties-notably,compression strength. Elevated temperature and moist environments also significantly affect properties dependent on matrix properties or interfacial strength. Because of these and several other factors,the efficiency of translation of fiber properties into those of the composite is not always as high as may be expected. Stiffness can be predicted more reliably than strength,although static tensile strength is easier to predict than other strength properties.Chapter 2 provides some elementary equations for estimating the mechanical properties of unidirectional composites,which are reasonably accurate in estimating elastic properties,providing fiber alignment is good.The equations can also provide ball-park figures for the matrix-dominated shear and transverse elastic properties and the fiber-dominated tensile strength properties.However,estimation of matrix-dominated or fiber/matrix bond strength-dominated strength properties, including shear and compression,is complex.Prediction problems also arise when the fibers are sensitive to compression loading,as is the case for aramid fibers,as discussed later. Chapter 7 describes the experimental procedures for measuring the mechanical properties,including those for assessing tolerance to damage and fatigue.These tests are used to develop a database for design of aerospace components and as part of the information required for airworthiness certification,as described in Chapter 12. Table 8.1 lists relevant mechanical and physical properties of the composites discussed in this chapter.Details of aerospace structural alloys aluminum 2024 T3 and titanium 6Al4V are also provided for comparison.The nomenclature used for the properties is similar to that used in Chapter 2.The data provided for the composites can be used as an estimate of ply properties for making an approximate prediction of laminate properties. The first four sections of this chapter provide an overview of the mechanical properties of composite systems based on glass,boron,aramid,or carbon fibers. 239
8 Properties of Composite Systems 8.1 Introduction The mechanical properties of simple unidirectional continuous fiber composites depend on the volume fraction and properties of the fibers (including flaw and strength distribution), the fiber/matrix bond strength, and the mechanical properties of the matrix. The alignment (waviness) of the fibers also has a significant effect on some properties--notably, compression strength. Elevated temperature and moist environments also significantly affect properties dependent on matrix properties or interfacial strength. Because of these and several other factors, the efficiency of translation of fiber properties into those of the composite is not always as high as may be expected. Stiffness can be predicted more reliably than strength, although static tensile strength is easier to predict than other strength properties. Chapter 2 provides some elementary equations for estimating the mechanical properties of unidirectional composites, which are reasonably accurate in estimating elastic properties, providing fiber alignment is good. The equations can also provide ball-park figures for the matrix-dominated shear and transverse elastic properties and the fiber-dominated tensile strength properties. However, estimation of matrix-dominated or fiber/matrix bond strength-dominated strength properties, including shear and compression, is complex. Prediction problems also arise when the fibers are sensitive to compression loading, as is the case for aramid fibers, as discussed later. Chapter 7 describes the experimental procedures for measuring the mechanical properties, including those for assessing tolerance to damage and fatigue. These tests are used to develop a database for design of aerospace components and as part of the information required for airworthiness certification, as described in Chapter 12. Table 8.1 lists relevant mechanical and physical properties of the composites discussed in this chapter. Details of aerospace structural alloys aluminum 2024 T3 and titanium 6A14V are also provided for comparison. The nomenclature used for the properties is similar to that used in Chapter 2. The data provided for the composites can be used as an estimate of ply properties for making an approximate prediction of laminate properties. The first four sections of this chapter provide an overview of the mechanical properties of composite systems based on glass, boron, aramid, or carbon fibers. 239
器 Table 8.1 Unidirectional Properties(Mostly Approximate)of Various Composites Considered in This Chapter and,for Comparison, Airframe Titanium and Aluminum Alloys.Based Largely on Ref.2 Glass fiber Carbon fiber composites composites COMPOSITE Units E s Boron Aramid K 49 HT HM UHM Al Ti SG 2.1 2.0 2.0 1.38 1.58 1.64 1.7 2.76 4.4 ai Lm/C 7.1 6.3 4.5 -1 -0.16 MATERIALS 02 um/C 20 70 24 二 二 23 9 23 Gitu MPa 1020 1620 1520 1240 1240 760 620 454 1102 FOR Glcu MPa 620 690 2930 275 1100 690 620 280* 1030 02u MPa 40 40 70 30 41 28 21 441 1102 Tu MPa 70 80 90 60 80 70 60 275 640 ILS MPa 70 80 90 60 80 70 60 AIRCRAFT E GPa 45 55 210 76 145 220 290 110 GPa 12 16 19 5.5 10 6.9 6.2 7 110 G12 GPa 5.5 7.6 4.8 2.1 4.8 4.8 4.8 4 V12 0.28 0.28 0.25 0.34 0.25 0.25 0.25 0.33 .31 Etu 0.022 0.029 0.006 0.016 0.01 0.03 0.02 0.12 0.06 E2u 0.4 0.4 0.4 0.5 0.4 0.4 0.3 STRUCTURES Notes:Ti Ti 6Al4V;Al 2024 T3 *yield value V60%in the composites;SG,specific gravity:ILS,interlaminar shear.See Chapters 2 and 12 for definition of other terms
240 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES ~3 "a f~ [.. J j r,j E .| ! r~ ,,,., o m- t'q ,.-~ O r., o ..= Cq t'4 .8 q= ~;,.., q= III ~
PROPERTIES OF COMPOSITE SYSTEMS 241 Further information on the properties of carbon-fiber composites is also provided throughout this book.In the last three sections,more generic discussion is provided on important impact,fatigue,and environmental properties,while a focus on carbon-fiber systems is maintained. 8.2 Glass-Fiber Composite Systems As described in Chapter 3,several types of glass reinforcements are suitable for the manufacture of aircraft and helicopter composite components.E-glass composites are used extensively in gliders and in non-structural components that do not require high stiffness,such as radomes.S-glass composites have better mechanical properties and therefore are used in more demanding applications.A third type of reinforcement known as D-glass has good dielectric properties and is occasionally used in aircraft to minimize the impact of lighting strikes.E-and S-glass are used in the form of epoxy-based pre-preg or as fabrics containing unidirectional,woven,or chopped strand filaments. A major advantage of E-glass fibers over the other types of fibers used in aircraft is their low cost.Figure 8.1 compares typical material costs for E-glass composites against costs for carbon,aramid (trade name,Kevlar),and boron/ epoxy composites;the relative cost of boron pre-preg shown is divided by a factor of 10 to make the chart readable.Costs are given for composites made of pre-preg or fabric(woven roving,chopped strand mat).The costs are approximate and do not include the expense of fabricating the composite into an aircraft component, which is usually much higher than the raw material cost.E-glass composites are 6 3 2 0 E-Glass Aramid HS Carbon IM Carbon Boron Fig 8.1 Relative costs of some fiber composite systems used in aerospace applications.Boron is shown at about 1/10 of its actual relative cost
PROPERTIES OF COMPOSITE SYSTEMS 241 Further information on the properties of carbon-fiber composites is also provided throughout this book. In the last three sections, more generic discussion is provided on important impact, fatigue, and environmental properties, while a focus on carbon-fiber systems is maintained. 8.2 Glass-Fiber Composite Systems As described in Chapter 3, several types of glass reinforcements are suitable for the manufacture of aircraft and helicopter composite components. E-glass composites are used extensively in gliders and in non-structural components that do not require high stiffness, such as radomes. S-glass composites have better mechanical properties and therefore are used in more demanding applications. A third type of reinforcement known as D-glass has good dielectric properties and is occasionally used in aircraft to minimize the impact of lighting strikes. E- and S-glass are used in the form of epoxy-based pre-preg or as fabrics containing unidirectional, woven, or chopped strand filaments. A major advantage of E-glass fibers over the other types of fibers used in aircraft is their low cost. Figure 8.1 compares typical material costs for E-glass composites against costs for carbon, aramid (trade name, Kevlar), and boron/ epoxy composites; the relative cost of boron pre-preg shown is divided by a factor of 10 to make the chart readable. Costs are given for composites made of pre-preg or fabric (woven roving, chopped strand mat). The costs are approximate and do not include the expense of fabricating the composite into an aircraft component, which is usually much higher than the raw material cost. E-glass composites are 6 0} .= ~5 .= a. 4 o o3 .m ~2 1 E-Glass Aramid HS Carbon IM Carbon Fig 8.1 Relative costs of some fiber composite systems used applications. Boron is shown at about 1/10 of its actual relative cost. Boron in aerospace
242 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES by far the cheapest,particularly when chopped strand mat or woven fabric is used.S-glass composites are much more expensive than the E-glass composites and only marginally less expensive than carbon/epoxy. Figures 8.2 and 8.3 provide comparisons'of the strength and stiffness of some of the available forms of E-glass fiber materials.The forms are chopped-stand mat,woven rovings,and unidirectional pre-preg material.The comparisons in these figures are based on relativities that will also be relevant to the other fiber types if made from similar geometrical forms. Table 8.1 provides relevant physical,thermal,and mechanical property data for unidirectional E-and S-glass/epoxy composites.Glass fibers have a specific gravity of about 2.5 g cm,which is slightly lower than the density of boron fibers (2.6g cm)but is appreciably higher than carbon (~1.8 g cm)and Kevlar(1.45 g cm)fibers.The specific gravity of thermoset resins is around 1.3 gcm,and as a result,glass/epoxy composites have a specific gravity that is higher than for other types of aerospace composites(except boron/epoxy)with the same fiber volume content.However,depending on the fiber volume fraction, it is still somewhat lower than that of aircraft-grade aluminum alloys (2.8 g cm).The Young's moduli and strengths of both E-and S-glass composites are lower than those of other aerospace structural composites and metals.The combined effects of low stiffness and high specific gravity makes glass/epoxy or 60 UD unidirectional 50 WR woven rovings CSM=chopped strand mat UD/epoxy 40 sninpow 30 20 WR/polyester 10 CSM/polyester 0 0 20 40 60 80 100 Glass Content by weight Fig 8.2 Typical Young's modulus for various types of glass-fiber composites. Adapted from Ref.1
242 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES by far the cheapest, particularly when chopped strand mat or woven fabric is used. S-glass composites are much more expensive than the E-glass composites and only marginally less expensive than carbon/epoxy. Figures 8.2 and 8.3 provide comparisons 1 of the strength and stiffness of some of the available forms of E-glass fiber materials. The forms are chopped-stand mat, woven rovings, and unidirectional pre-preg material. The comparisons in these figures are based on relativities that will also be relevant to the other fiber types if made from similar geometrical forms. Table 8.1 provides relevant physical, thermal, and mechanical property data for unidirectional E- and S-glass/epoxy composites. Glass fibers have a specific gravity of about 2.5 g cm -3, which is slightly lower than the density of boron fibers (2.6 g cm -3) but is appreciably higher than carbon (~ 1.8 g cm -3) and Kevlar (1.45 g cm -3) fibers. The specific gravity of thermoset resins is around 1.3 g cm -3, and as a result, glass/epoxy composites have a specific gravity that is higher than for other types of aerospace composites (except boron/epoxy) with the same fiber volume content. However, depending on the fiber volume fraction, it is still somewhat lower than that of aircraft-grade aluminum alloys (2.8 g cm-3). The Young's moduli and strengths of both E- and S-glass composites are lower than those of other aerospace structural composites and metals. The combined effects of low stiffness and high specific gravity makes glass/epoxy or 60 50 e~ t~ 40 "g 3o -g ~ 2o 0 10 UD = unidirectional WR = woven rovings CSM = chopped strand mat 0 20 40 60 80 100 Glass Content % by weight Fig 8.2 Typical Young's modulus for various types of glass-fiber composites. Adapted from Ref. I
PROPERTIES OF COMPOSITE SYSTEMS 243 900 800 700 UD=unidirectional WR woven rovings 600 CSM chopped strand mat 500 400 300 WR/polyester 200 100 CSM/polyester 0 0 20 40 60 80 100 Glass Content by Weight Fig 8.3 Typical strengths of various types of glass-fiber composites.Adapted from Ref.1. other glass fiber composites unattractive for use in weight-critical load-bearing primary structures on larger aircraft. 8.2.1 Fatigue Performance of Glass-Fiber Systems Another drawback of using glass/epoxy composites in aircraft structures is their relatively poor fatigue performance compared with the other composites discussed in this chapter.Glass/epoxy composites are more prone to fatigue- induced damage (e.g.,microscopic cracks,delaminations)and failure than other aerospace composite materials.Figure 8.4 shows a typical fatigue-life curve for a unidirectional glass/epoxy composite that was tested under cyclic tension- tension loading.Fatigue-life curves for unidirectional carbon/epoxy and Kevlar/ epoxy laminates that were also tested under tension-tension loading are shown for comparison.In the figure,the normalized fatigue strain (er/e)is the maximum applied cyclic tensile strain (er)divided by the static tensile failure strain of the composite().Of the three materials,the fatigue-life curve for the glass/epoxy
PROPERTIES OF COMPOSITE SYSTEMS 243 900 800 700 600 t~ ~" =E 500 400 ~ 300 " 200 o 1- 100 0 UD = unidirectional WR = woven rovings CSM = chopped strand mat II O 0 20 40 60 80 100 Glass Content % by Weight Fig 8.3 Typical strengths of various types of glass-fiber composites. Adapted from Ref. 1. other glass fiber composites unattractive for use in weight-critical load-bearing primary structures on larger aircraft. 8.2.1 Fatigue Performance of Glass-Fiber Systems Another drawback of using glass/epoxy composites in aircraft structures is their relatively poor fatigue performance compared with the other composites discussed in this chapter. Glass/epoxy composites are more prone to fatigueinduced damage (e.g., microscopic cracks, delaminations) and failure than other aerospace composite materials. Figure 8.4 shows a typical fatigue-life curve for a unidirectional glass/epoxy composite that was tested under cyclic tensiontension loading. Fatigue-life curves for unidirectional carbon/epoxy and Kevlar/ epoxy laminates that were also tested under tension-tension loading are shown for comparison. In the figure, the normalized fatigue strain (ef/eo) is the maximum applied cyclic tensile strain (ef) divided by the static tensile failure strain of the composite (eo). Of the three materials, the fatigue-life curve for the glass/epoxy
244 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES 1.0 °0.8 0.6 0.4 Glass/epoxy Kevlar/epoxy 0.2 49441010”40 Carbon/epoxy 0.0 10° 10 102 103 10 105 10 10 Number of Cycles Fig 8.4 Fatigue-life curves for unidirectional composites subject to tension-tension loading. composite drops the most rapidly,with increasing number of cycles.This indicates that glass/epoxy is the most susceptible to fatigue-induced failure under tension-tension loading,and this is due to the low stiffness of the glass reinforcement,resulting in damaging strains in the matrix,as discussed in Section 8.8. The fatigue performance of glass/epoxy composites is degraded further when cyclic loading occurs in a hostile environment,such as in hot and wet conditions. The microscopic cracks and delaminations caused by fatigue loading create pathways for the rapid ingress of moisture into the composite.Moisture can then cause stress-corrosion damage to the glass fibers,which may dramatically reduce the fatigue life.Cracks and delaminations caused by fatigue also cause large reductions to the stiffness and strength of glass/epoxy.Figure 8.5 shows that the static tensile modulus and strength of [0/90]s glass/epoxy composites decrease rapidly with increasing number of load cycles before reaching a constant level.The residual modulus and strength remain relatively constant until near the end of the fatigue life.For some glass/epoxy materials,a reduction in stiffness and strength can occur within the early stage of the fatigue process,when the damage is not visible. Finally,glass fibers composites and other composites having fibers with low thermal conductivity and low stiffness are prone to heat damage under cyclic
244 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES 1.0 °0.8 Cr "~ 0.6 0.4 O Z 0.2 0.0 ........ ' 10 0 101 Fig 8.4 loading. ......................................... ........ Kevlar/epoxy ............. Carbon/epoxy , , ,,,,,,I , , ,.,i,,I , ,,,,,,J , , ,ll,.I i II'''''! , IlllJ 10 2 10 3 10 4 10 s 10 6 I O r Number of Cycles Fatigue-life curves for unidirectional composites subject to tension-tension composite drops the most rapidly, with increasing number of cycles. This indicates that glass/epoxy is the most susceptible to fatigue-induced failure under tension-tension loading, and this is due to the low stiffness of the glass reinforcement, resulting in damaging strains in the matrix, as discussed in Section 8.8. The fatigue performance of glass/epoxy composites is degraded further when cyclic loading occurs in a hostile environment, such as in hot and wet conditions. The microscopic cracks and delaminations caused by fatigue loading create pathways for the rapid ingress of moisture into the composite. Moisture can then cause stress-corrosion damage to the glass fibers, which may dramatically reduce the fatigue life. Cracks and delaminations caused by fatigue also cause large reductions to the stiffness and strength of glass/epoxy. Figure 8.5 shows that the static tensile modulus and strength of [0/90]s glass/epoxy composites decrease rapidly with increasing number of load cycles before reaching a constant level. The residual modulus and strength remain relatively constant until near the end of the fatigue life. For some glass/epoxy materials, a reduction in stiffness and strength can occur within the early stage of the fatigue process, when the damage is not visible. Finally, glass fibers composites and other composites having fibers with low thermal conductivity and low stiffness are prone to heat damage under cyclic
PROPERTIES OF COMPOSITE SYSTEMS 245 1.0 0.8 Nomalized Modulus 0.6 pazlleuoN 0.4 Nomalized Strength 0.2 0.0 0 2 4 6 8 10 Number of Cycles (x10) Fig 8.5 Effect of number of tensile load cycles on the Young's modulus and strength of a [0/901s glass/epoxy composite. loading at high frequencies2 above around 5 Hz.This is because the heat generated by stress/strain hysteresis in the polymer matrix cannot be easily dissipated.The problem increases in thick composites,in which heat dissipation is even more difficult and with off-angle fibers where matrix strains are higher. The performance of composites under cyclic loading is discussed further in Section 8. 8.2.2 Impact Strength of Glass-Fiber Systems Although many mechanical and fatigue properties of glass/epoxy composites are lower than those of other carbon/epoxy and aramid/epoxy materials,they generally have a superior ability to absorb energy during impact.Figure 8.6 illustrates the relative energies for failure under impact of glass fiber and other composites considered in this chapter and some aluminum alloys,as measured with the Charpy test method.The exceptionally high impact toughness of S-glass fibers has led to their use in ballistic protective materials. As shown,glass/epoxy composites have the highest impact energies,with S-glass/epoxy composites being 4-7 times more impact-resistant than high- strength carbon/epoxy laminates and about 35 times more resistant than high-modulus carbon/epoxy materials.Glass/epoxy composites are even 9-11 times more impact-resistant on this basis than aircraft-grade aluminum alloy
PROPERTIES OF COMPOSITE SYSTEMS 245 1.01 0.8 = 0.6 ~0.4 o 0.2 Nomalized Strength 0,0 u I ~ I = I ~ I u I 0 2 4 6 8 10 Number of Cycles (xlO 4) Fig 8.5 Effect of number of tensile load cycles on the Young's modulus and strength of a [0/90]s glass/epoxy composite. loading at high frequencies 2 above around 5 Hz. This is because the heat generated by stress/strain hysteresis in the polymer matrix cannot be easily dissipated. The problem increases in thick composites, in which heat dissipation is even more difficult and with off-angle fibers where matrix strains are higher. The performance of composites under cyclic loading is discussed further in Section 8. 8.2.2 Impact Strength of Glass-Fiber Systems Although many mechanical and fatigue properties of glass/epoxy composites are lower than those of other carbon/epoxy and aramid/epoxy materials, they generally have a superior ability to absorb energy during impact. Figure 8.6 illustrates the relative energies for failure under impact of glass fiber and other composites considered in this chapter and some aluminum alloys, as measured with the Charpy test method. The exceptionally high impact toughness of S-glass fibers has led to their use in ballistic protective materials. As shown, glass/epoxy composites have the highest impact energies, with S-glass/epoxy composites being 4-7 times more impact-resistant than highstrength carbon/epoxy laminates and about 35 times more resistant than high-modulus carbon/epoxy materials. Glass/epoxy composites are even 9-11 times more impact-resistant on this basis than aircraft-grade aluminum alloy
246 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES E 800 700 600 500 400 300 芯eaEs 200 100 0 S-glass/epoxy High strength carbon/epoxy E-glass/epoxy High modulus carbon/epoxy Kevlar 49/epoxy Boron/epoxy 7075-T6 aluminum Fig 8.6 Charpy impact energy absorption of some composite and,for comparison, non-composite materials,as indicated. 8.2.3 Stress and Environmental Effects As discussed in Chapter 3,glass fibers are prone to fracture when subjected to high stress for prolonged periods of time.This behavior,known as static fatigue or stress rupture,is exacerbated by exposure to moisture,as shown in Figure 8.7. 1 0.9 0.8 Dry 0.7 0.6 Wet 0.5 0.4 0.3 0.2 0.1 0 0 2 3 Time to Failure (Seconds) Fig 8.7 Stress rupture strength of E-glass fibers in air and water.Adapted from Ref.1
246 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES 800 E P ILl m e~ _E 700 600 500 400 300 200 100 0 ii~i17111111 Fig 8.6 Charpy impact energy absorption of some composite and, for comparison, non-composite materials, as indicated. 8.2.3 Stress and Environmental Effects As discussed in Chapter 3, glass fibers are prone to fracture when subjected to high stress for prolonged periods of time. This behavior, known as static fatigue or stress rupture, is exacerbated by exposure to moisture, as shown in Figure 8.7. 0.9 0.8 0.7 ¢n 0,6 -o 0.5 N "- 0.4 E 0,3 0 z 0.2 0.1 Wet I I I I I Fig 8.7 Ref. 1. 0 1 2 3 4 5 6 Time to Failure (Seconds) Stress rupture strength of E-glass fibers in air and water. Adapted from
PROPERTIES OF COMPOSITE SYSTEMS 247 The influence of moisture on fiber strength is much reduced if the fiber is embedded in a polymer matrix,but can still be of concern in highly loaded applications such as pressure vessels. Glass fiber composites,when exposed to moist environments or other aggressive environments,are also prone to degradation caused by weakening of the fiber/matrix interface.This weakening generally occurs by chemical attack at the fiber surface.The degree of weakening experienced depends on the matrix, the coating (called size or finish)used on the fiber,and the type of fiber. Weakening of the interface will result in significant loss in matrix-dominated mechanical properties such as shear,off-angle,and compression strength. Environmental degradation is thus of significant concern for structural applications in which the ability to carry high loads is required and particularly where the loading is sustained. 8.3 Boron Fiber Composite Systems Boron fibers(Chapter 3)were first discovered in 1959 and were subsequently developed during the 1960s into the first true high-performance fibers.Until that time,glass fiber was the only other high-strength fiber available in continuous lengths,and the low modulus of glass severely restricted its use in high- performance structures.The high-temperature capability of boron also provided the opportunity for producing metal-matrix composites,although it was a boron/ epoxy (b/ep)composite that produced much of the initial commercial success. These composites were used successfully in several important aircraft component programs during the 1970s including the skins of the horizontal stabilizers on the F-14 and the horizontal and vertical stabilizers and rudders on the F-15.Boron/ epoxy pre-preg materials are currently available in commercial quantities,and their unique properties make them suited to a range of specialized applications. Because of the presence of a dense tungsten boride core (Chapter 3),the diameter of boron fibers is significantly greater than that of carbon fibers,to minimize fiber density and to ensure the properties of the fiber are not greatly influenced by the properties of the core.Fibers are currently produced in 100-and 140-wm diameters and therefore boron fibers have a very high bending stiffness (proportional to the fourth power of the radius).This restricts the radius that the composite can be formed into.For the 100-m diameter fibers,a radius of around 30 mm is the practical limit.Although this is not of concern in the production of large,relatively flat aircraft components,it is sometimes a limiting factor in the selection of this composite system for the manufacture of a part with complex geometry. The large diameter of boron fibers means that it is virtually impossible to weave these fibers into a fabric in the same way that glass,kevlar,and carbon fibers can.It is,however,possible to hold parallel boron fibers together with a weft thread of polyester to form a dry unidirectional preform.Boron pre-pregs are
PROPERTIES OF COMPOSITE SYSTEMS 247 The influence of moisture on fiber strength is much reduced if the fiber is embedded in a polymer matrix, but can still be of concern in highly loaded applications such as pressure vessels. Glass fiber composites, when exposed to moist environments or other aggressive environments, are also prone to degradation caused by weakening of the fiber/matrix interface. This weakening generally occurs by chemical attack at the fiber surface. The degree of weakening experienced depends on the matrix, the coating (called size or finish) used on the fiber, and the type of fiber. Weakening of the interface will result in significant loss in matrix-dominated mechanical properties such as shear, off-angle, and compression strength. Environmental degradation is thus of significant concern for structural applications in which the ability to carry high loads is required and particularly where the loading is sustained. 8.3 Boron Fiber Composite Systems Boron fibers (Chapter 3) were first discovered in 1959 and were subsequently developed during the 1960s into the first true high-performance fibers. Until that time, glass fiber was the only other high-strength fiber available in continuous lengths, and the low modulus of glass severely restricted its use in highperformance structures. The high-temperature capability of boron also provided the opportunity for producing metal-matrix composites, although it was a boron/ epoxy (b/ep) composite that produced much of the initial commercial success. These composites were used successfully in several important aircraft component programs during the 1970s including the skins of the horizontal stabilizers on the F-14 and the horizontal and vertical stabilizers and rudders on the F-15. Boron/ epoxy pre-preg materials are currently available in commercial quantities, and their unique properties make them suited to a range of specialized applications. Because of the presence of a dense tungsten boride core (Chapter 3), the diameter of boron fibers is significantly greater than that of carbon fibers, to minimize fiber density and to ensure the properties of the fiber are not greatly influenced by the properties of the core. Fibers are currently produced in 100- and 140-1xm diameters and therefore boron fibers have a very high bending stiffness (proportional to the fourth power of the radius). This restricts the radius that the composite can be formed into. For the 100-1xm diameter fibers, a radius of around 30 mm is the practical limit. Although this is not of concern in the production of large, relatively flat aircraft components, it is sometimes a limiting factor in the selection of this composite system for the manufacture of a part with complex geometry. The large diameter of boron fibers means that it is virtually impossible to weave these fibers into a fabric in the same way that glass, kevlar, and carbon fibers can. It is, however, possible to hold parallel boron fibers together with a weft thread of polyester to form a dry unidirectional preform. Boron pre-pregs are
248 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES unidirectional and have a fine polyester scrim material (similar to that in structural film adhesives)incorporated into the resin on one side of the fibers to provide some lateral strength to the pre-preg during handling. 8.3.1 Mechanical Properties of Boron-Fiber Systems Typical properties of unidirectional boron/epoxy composites are shown in Table 8.1.Boron composites typically have high compressive strength due to the large-fiber diameter,and this is one of their distinguishing features compared with carbon composites.Most of the advanced composite systems provide a significant improvement in specific stiffness over the conventional aircraft metallic materials,which have a common specific stiffness of around 25 GPa. Also apparent from Table 8.1 is the fact that although the density of cured boron composites is higher than carbon composites,it is appreciably lower than that of aluminum or titanium. There are several types of carbon fibers on the market,some of which have properties that the densites of exceed either the tensile modulus or strength of boron fibers.Boron fiber composites,however,still have a blend of tensile and compressive properties that no single carbon fiber type is able to match.A form of pre-preg is available in which boron and carbon fibers are mixed together in the same pre-preg and this is marketed by Textron Specialty Materials as Hy-Bor. The properties of this material exceed those of conventional boron/epoxy composite due to the higher volume fraction of fibers. 8.3.2 Handling and Processing Properties of Boron-Fiber Systems Boron is an extremely hard material with a Knoop value of 3200,which is harder than tungsten carbide and titanium nitride(1800-1880)and second only to diamond (7000).Cured boron composites can be cut,drilled,and machined with diamond-tipped tools,and the pre-pregs are readily cut with conventional steel knives.In practice,the knives cannot actually cut the hard fibers;however, gentle pressure fractures the fibers with one or two passes.Although it is possible to cut complex shapes with the use of templates,laser-cutting has been shown to be the most efficient way to cut a large amount of non-rectangular boron plies. Boron fibers are currently available in several forms.As well as the two fiber diameters,pre-pregs are available with either 120C or 175C curing epoxies. With the exception of the reduction in formability mentioned above,in most other aspects,boron pre-pregs handle and process in a similar fashion to the more common carbon pre-preg materials. 8.3.3 Aircraft Applications of Boron-Fiber Composites The fiber manufacturing process described in Chapter 3 shows that the fibers are produced as monofilaments on an expensive precursor filament,and this basic
248 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES unidirectional and have a fine polyester scrim material (similar to that in structural film adhesives) incorporated into the resin on one side of the fibers to provide some lateral strength to the pre-preg during handling. 8.3.1 Mechanical Properties of Boron-Fiber Systems Typical properties of unidirectional boron/epoxy composites are shown in Table 8.1. Boron composites typically have high compressive strength due to the large-fiber diameter, and this is one of their distinguishing features compared with carbon composites. Most of the advanced composite systems provide a significant improvement in specific stiffness over the conventional aircraft metallic materials, which have a common specific stiffness of around 25 GPa. Also apparent from Table 8.1 is the fact that although the density of cured boron composites is higher than carbon composites, it is appreciably lower than that of aluminum or titanium. There are several types of carbon fibers on the market, some of which have properties that the densites of exceed either the tensile modulus or strength of boron fibers. Boron fiber composites, however, still have a blend of tensile and compressive properties that no single carbon fiber type is able to match. A form of pre-preg is available in which boron and carbon fibers are mixed together in the same pre-preg and this is marketed by Textron Specialty Materials as Hy-Bor. The properties of this material exceed those of conventional boron/epoxy composite due to the higher volume fraction of fibers. 8.3.2 Handling and Processing Properties of Boron-Fiber Systems Boron is an extremely hard material with a Knoop value of 3200, which is harder than tungsten carbide and titanium nitride (1800-1880) and second only to diamond (7000). Cured boron composites can be cut, drilled, and machined with diamond-tipped tools, and the pre-pregs are readily cut with conventional steel knives. In practice, the knives cannot actually cut the hard fibers; however, gentle pressure fractures the fibers with one or two passes. Although it is possible to cut complex shapes with the use of templates, laser-cutting has been shown to be the most efficient way to cut a large amount of non-rectangular boron plies. Boron fibers are currently available in several forms. As well as the two fiber diameters, pre-pregs are available with either 120°C or 175°C curing epoxies. With the exception of the reduction in formability mentioned above, in most other aspects, boron pre-pregs handle and process in a similar fashion to the more common carbon pre-preg materials. 8.3.3 Aircraft Applications of Boron-Fiber Composites The fiber manufacturing process described in Chapter 3 shows that the fibers are produced as monofilaments on an expensive precursor filament, and this basic