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