9 Joining of Composite Structures 9.1 Introduction Airframe structures consist essentially of an assembly of simple elements connected to form a load transmission path.The elements,which include skins, stiffeners,frames,and spars,form the major components such as wings,fuselage, and empennage.The connections or joints are potentially the weakest points in the airframe so can determine its structural efficiency. In general,it is desirable to reduce the number and complexity of joints to minimize weight and cost.A very important advantage of composite construction is the ability to form unitized components,thus minimizing the number of joints required.However,the design and manufacture of the remaining joints is still a major challenge to produce safe,cost-effective,and efficient structures. This chapter is concerned with joints used to connect structural elements made of advanced fiber composite laminates,mainly carbon/epoxy (carbon/epoxy),to other composite parts or to metals.Sections 9.3 and 9.4 deal,respectively,with bonded and mechanical joints typical of those used in the manufacture of airframe components.Joints are also required to repair structural damage;this topic is dealt with in Chapter 10.Both design and materials aspects are considered.The aim of this chapter,when discussing design,is to outline simple analytical procedures that provide a physical insight into the behavior of joints involving composites.The materials aspects covered will be those essential to the manufacture of sound joints. Joint types used in airframe construction can be broadly divided into joints that are mechanically fastened using bolts or rivets,adhesively bonded using a polymeric adhesive,or that feature a combination of mechanical fastening and adhesive bonding. In mechanical joints,loads are transferred between the joint elements by compression on the internal faces of the fastener holes with a smaller component of shear on the outer faces of the elements due to friction.In bonded joints,the loads are transferred mainly by shear on the surfaces of the elements.In both cases,the load transmission elements (fastener or adhesive) are stressed primarily in shear along the joint line;however,the actual stress distribution will be complex. Joints can be classified as single or multiple load path.Single load path joints are joints in which failure would result in catastrophic loss of structural 289
9 Joining of Composite Structures 9.1 Introduction Airframe structures consist essentially of an assembly of simple elements connected to form a load transmission path. The elements, which include skins, stiffeners, frames, and spars, form the major components such as wings, fuselage, and empennage. The connections or joints are potentially the weakest points in the airframe so can determine its structural efficiency. In general, it is desirable to reduce the number and complexity of joints to minimize weight and cost. A very important advantage of composite construction is the ability to form unitized components, thus minimizing the number of joints required. However, the design and manufacture of the remaining joints is still a major challenge to produce safe, cost-effective, and efficient structures. This chapter is concerned with joints used to connect structural elements made of advanced fiber composite laminates, mainly carbon/epoxy (carbon/epoxy), to other composite parts or to metals. Sections 9.3 and 9.4 deal, respectively, with bonded and mechanical joints typical of those used in the manufacture of airframe components. Joints are also required to repair structural damage; this topic is dealt with in Chapter 10. Both design and materials aspects are considered. The aim of this chapter, when discussing design, is to outline simple analytical procedures that provide a physical insight into the behavior of joints involving composites. The materials aspects covered will be those essential to the manufacture of sound joints. Joint types used in airframe construction can be broadly divided into joints that are mechanically fastened using bolts or rivets, adhesively bonded using a polymeric adhesive, or that feature a combination of mechanical fastening and adhesive bonding. In mechanical joints, loads are transferred between the joint elements by compression on the internal faces of the fastener holes with a smaller component of shear on the outer faces of the elements due to friction. In bonded joints, the loads are transferred mainly by shear on the surfaces of the elements. In both cases, the load transmission elements (fastener or adhesive) are stressed primarily in shear along the joint line; however, the actual stress distribution will be complex. Joints can be classified as single or multiple load path. Single load path joints are joints in which failure would result in catastrophic loss of structural 289
290 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES capability.Multiple load path joints are joints in which failure of a single element results in the load being carried by other load-carrying members.An apparently multiple load path joint would be classified as single load path if failure of one of the paths leads to an unacceptable reduction in the load capacity of the joint. The alignment of the load path and the geometry of the structural elements are important considerations in the design of joints.Airframe structural elements are usually intended to be loaded in either tension/compression or shear.Primary bending is avoided by keeping the loading as close as possible to collinear. However,secondary bending induced by minor eccentricity of the loads occurs in many types of joint(and structure)and can cause serious problems. Compared with metals,laminated fiber composites have relatively low through-thickness strength and bearing strength under concentrated loads.Thus metals,usually titanium alloys,are sometimes required to transmit loads in and out of highly loaded composite structure,particularly where stress fields are complex. Typical design parameters for carbon/epoxy airframe components(for a high- performance military aircraft)are: .Ultimate design strain:+3000 to 4000 microstrain for mechanically fastened structure,up to+5000 microstrain for bonded honeycomb structure Operating temperature -55 C to +105C Service fluids;presence of moisture,hydraulic oil,fuel,and (limited exposure to)paint stripper Strain,rather than strength,is generally used as the basis for comparison of the structural capacity of composite structure because composites of differing stiffness tend to fail at a similar strain level-particularly when damaged. Microstrain is strain x 10-6. 9.2 Comparison Between Mechanically Fastened and Adhesively Bonded Joints The advantages and disadvantages of forming joints by adhesive bonding and bolting or riveting are summarized in Table 9.1 Although there are many advantages for bonding composites from the performance view point,there are also many limitations or disadvantages that must be considered with each potential application.For a relatively thin-skinned structure,particularly where fatigue may be a problem,bonding is very attractive indeed.However,the use of suitable pre-bonding surface treatments and adhesives is essential to develop the required strength level and maintain it during a service life,which could be more than 30 years. A high level of quality control is very important to obtain reliable adhesive bonding.This is because current non-destructive inspection(NDI)procedures are able to detect only gross defects such as severe voids and disbonds in bonded
290 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES capability. Multiple load path joints are joints in which failure of a single element results in the load being carded by other load-carrying members. An apparently multiple load path joint would be classified as single load path if failure of one of the paths leads to an unacceptable reduction in the load capacity of the joint. The alignment of the load path and the geometry of the structural elements are important considerations in the design of joints. Airframe structural elements are usually intended to be loaded in either tension/compression or shear. Primary bending is avoided by keeping the loading as close as possible to collinear. However, secondary bending induced by minor eccentricity of the loads occurs in many types of joint (and structure) and can cause serious problems. Compared with metals, laminated fiber composites have relatively low through-thickness strength and bearing strength under concentrated loads. Thus metals, usually titanium alloys, are sometimes required to transmit loads in and out of highly loaded composite structure, particularly where stress fields are complex. Typical design parameters for carbon/epoxy airframe components (for a highperformance military aircraft) are: • Ultimate design strain: _ 3000 to 4000 microstrain for mechanically fastened structure, up to + 5000 microstrain for bonded honeycomb structure • Operating temperature - 55 °C to + 105 °C • Service fluids; presence of moisture, hydraulic oil, fuel, and (limited exposure to) paint stripper Strain, rather than strength, is generally used as the basis for comparison of the structural capacity of composite structure because composites of differing stiffness tend to fail at a similar strain level--particularly when damaged. Microstrain is strain × 10 -6. 9.2 Comparison Between Mechanically Fastened and Adhesively Bonded Joints The advantages and disadvantages of forming joints by adhesive bonding and bolting or riveting are summarized in Table 9.1 Although there are many advantages for bonding composites from the performance view point, there are also many limitations or disadvantages that must be considered with each potential application. For a relatively thin-skinned structure, particularly where fatigue may be a problem, bonding is very attractive indeed. However, the use of suitable pre-bonding surface treatments and adhesives is essential to develop the required strength level and maintain it during a service life, which could be more than 30 years. A high level of quality control is very important to obtain reliable adhesive bonding. This is because current non-destructive inspection (NDI) procedures are able to detect only gross defects such as severe voids and disbonds in bonded
JUINING UF CUMPUSIIE SIHUCIUHES 291 Table 9.1 A Comparison of the Advantages and Disadvantages of Adhesively Bonded and Bolted Composite Joints Advantages Disadvantages Bonded Joints Small stress concentration in Limits to thickness that can be joined with adherends simple joint configuration Stiff connection Inspection other than for gross flaws difficult Excellent fatigue properties Prone to environmental degradation No fretting problems Sensitive to peel and through-thickness stresses Sealed against corrosion Residual stress problems when joining to metals Smooth surface contour Cannot be disassembled Relatively lightweight May require costly tooling and facilities Damage tolerant Requires high degree of quality control May be of environmental concern Bolted Joints Positive connection,low initial risk Considerable stress concentration Can be disassembled Prone to fatigue cracking in metallic component No thickness limitations Hole formation can damage composite Simple joint configuration Composites's relatively poor bearing properties Simple manufacturing process Pone to fretting in metal Simple inspection procedure Prone to corrosion in metal Not environmentaly sensitive May require extensive shimming Provides through-thickness reinforcement;not sensitive to peel stresses No major residual stress problem components but are unable to detect weak or(due to environmental degradation) potentially weak bonds.The limitations of NDI are a major reason why adhesive bonding has rarely been used in critical primary joints in metallic airframe structure;bonded metal joints are particularly prone to environmental degradation if not adequately surface-treated. Mechanical fastening is usually the lower-cost option because of its simplicity and low-cost tooling and inspection requirements.However,hole-drilling can be highly labor intensive(unless automated)and,if not correctly done,can be highly damaging to the composite.Joints in aircraft usually require many thousands of expensive fasteners (usually titanium alloy),and extensive shimming may be required to avoid damage to the composite structure during bolt clamp-up.Thus adhesive bonding,despite the high tooling,process,and quality control costs,can in many cases offer significant cost savings
Table 9.1 J(.,)INING L)P L;L)MP(..,)~I I I:: ~51 HUG I UHI::S A Comparison of the Advantages and Disadvantages of Adhesively Bonded and Bolted Composite Joints 291 Advantages Disadvantages Small stress concentration in adherends Stiff connection Excellent fatigue properties No fretting problems Sealed against corrosion Smooth surface contour Relatively lightweight Damage tolerant Positive connection, low initial risk Can be disassembled No thickness limitations Simple joint configuration Simple manufacturing process Simple inspection procedure Not environmentaly sensitive Provides through-thickness reinforcement; not sensitive to peel stresses No major residual stress problem Bonded Joints Limits to thickness that can be joined with simple joint configuration Inspection other than for gross flaws difficult Prone to environmental degradation Sensitive to peel and through-thickness stresses Residual stress problems when joining to metals Cannot be disassembled May require costly tooling and facilities Requires high degree of quality control May be of environmental concern Bolted Joints Considerable stress concentration Prone to fatigue cracking in metallic component Hole formation can damage composite Composites' s relatively poor bearing properties Pone to fretting in metal Prone to corrosion in metal May require extensive shimming components but are unable to detect weak or (due to environmental degradation) potentially weak bonds. The limitations of NDI are a major reason why adhesive bonding has rarely been used in critical primary joints in metallic airframe structure; bonded metal joints are particularly prone to environmental degradation if not adequately surface-treated. Mechanical fastening is usually the lower-cost option because of its simplicity and low-cost tooling and inspection requirements. However, hole-drilling can be highly labor intensive (unless automated) and, if not correctly done, can be highly damaging to the composite. Joints in aircraft usually require many thousands of expensive fasteners (usually titanium alloy), and extensive shimming may be required to avoid damage to the composite structure during bolt clamp-up. Thus adhesive bonding, despite the high tooling, process, and quality control costs, can in many cases offer significant cost savings
292 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES 9.3 Adhesively Bonded Joints Symbols Shear modulus (also used for Young's modulus E strain energy release rate) Shear stress T Stress Shear strain Y Strain Thickness Displacement U Transfer length L Plastic zone size d Applied load P Step length N Transmitted load T Scarf angle Thermal expansion coefficient Temperature range △T AT=(service temperature-cure temperature) Subscripts/Superscripts Plastic condition Outer adherend (also for mode 1 opening) Elastic condition Inner adherend (also for mode 2 2 opening) Ultimate value Maximum value max Adhesive A Minimum value min Temperature T Balanced b Value at infinite length 0∞ Unbalanced un Critical value C 9.3.1 Introduction Bonded joints used in aerospace applications can be classified as single (primary)or multiple (secondary)load path joints,as indicated in Figure 9.1. This section describes simple design procedures and some materials' engineering aspects relevant to the application of these types of joint in airframe structures. In the design of bonded composite joints,consideration is given to each of the elements to be joined (adherends),including their geometry,size,materials of construction,actual or potential modes of failure,coefficients of thermal expansion,magnitude and nature of the loading involved,and operating environment. Potential modes of failure are: Tensile,compressive,or shear of the adherends Shear or peel in the adhesive layer .Shear or peel in the composite near-surface plies Shear or peel in the resin-rich layer on the surface of the composite Adhesive failure at the metal or composite/adhesive interface
292 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES 9,3 Adhesively Bonded Joints Symbols Shear modulus (also used for G strain energy release rate) Shear stress ~- Shear strain y Thickness t Transfer length L Plastic zone size d Step length N Scarf angle 0 Young' s modulus E Stress o" Strain e Displacement U Applied load P Transmitted load T Thermal expansion coefficient a Temperature range AT AT = (service temperature----cure temperature) Subscripts/Superscripts Plastic condition p Outer adherend (also for mode 1 1 opening) Elastic condition e Inner adherend (also for mode 2 2 opening) Ultimate value u Maximum value max Adhesive A Minimum value min Temperature T Balanced b Value at infinite length co Unbalanced un Critical value C 9.3.1 Introduction Bonded joints used in aerospace applications can be classified as single (primary) or multiple (secondary) load path joints, as indicated in Figure 9.1. This section describes simple design procedures and some materials' engineering aspects relevant to the application of these types of joint in airframe structures. In the design of bonded composite joints, consideration is given to each of the elements to be joined (adherends), including their geometry, size, materials of construction, actual or potential modes of failure, coefficients of thermal expansion, magnitude and nature of the loading involved, and operating environment. Potential modes of failure are: • Tensile, compressive, or shear of the adherends • Shear or peel in the adhesive layer • Shear or peel in the composite near-surface plies • Shear or peel in the resin-rich layer on the surface of the composite • Adhesive failure at the metal or composite/adhesive interface
JOINING OF COMPOSITE STRUCTURES 293 Composite Aerospace Structural Joints Primary Joints Secondary Joints Reinforcements Laminates Single Double Step Lap Lap Lap Fuselage Splice Wing/Tail Root Stiffeners Doublers Patches Solid Laminates Honeycomb Laminates Stringers DoorFrame Repairs Skins D00r5 Longerons WindowFrame Helicopter Blades Control Surfaces Frames Skin Propellers Fairings Empennage Fig.9.1 Classification and applications of adhesively bonded joints used in airframe manufacture. The design aim is for the joint to fail by bulk failure of the adherends.A margin of safety is generally included in the design to provide tolerance to service damage and manufacturing defects in the bond line.Generally,the adhesive is not allowed to be or to become the weak link'because adhesive strength can be highly variable,and the growth of damage or defects in the adhesive layer can be very rapid under cyclic loading.For composite adherends,the very thin, relatively brittle resin bonding the near-surface plies is more prone to failure than the adhesive layer,so great care must also be taken to ensure that this does not become the weak link. The design input parameters include: Stiffness and strength(for metals usually the yield strength)of the adherends Shear modulus,yield strength,and strain-to-failure of the adhesive .Thermal expansion coefficient of the adherends .Magnitude and direction of the applied loads .Overlap length of the adherends Thickness of the adherends Thickness of the adhesive The properties used must be sufficient to handle the weakest state of the materials;for the adhesive and composite adherends,this is usually the hot/wet condition.It is most important to ensure that the strength of the adherend/ adhesive interface does not become significantly weakened as a result of environmental degradation.For a degraded interface,there is no way of
JOINING OF COMPOSITE STRUCTURES 293 Composite Aerospace Structural Joints J Reinfo I I Single Double Step Lap Lap Lap Fuselage Splice Wing/Tail Root I I I Primary Joints Secondary Joints I I ,cements Laminates i i I I Stiffeners Doublers Patches Solid Laminates Honeycomb Laminates Stringers DoorFrame Repai~ Skins Doors Longerons WindowFrame Helicopter Blades Control Surfaces Frames Skin Propellers Fairings Empennage Fig. 9.1 Classification and applications of adhesively bonded joints used in airframe manufacture. The design aim is for the joint to fail by bulk failure of the adherends. A margin of safety is generally included in the design to provide tolerance to service damage and manufacturing defects in the bond line. Generally, the adhesive is not allowed to be or to become the weak link I because adhesive strength can be highly variable, and the growth of damage or defects in the adhesive layer can be very rapid under cyclic loading. For composite adherends, the very thin, relatively brittle resin bonding the near-surface plies is more prone to failure than the adhesive layer, so great care must also be taken to ensure that this does not become the weak link. The design input parameters include: • Stiffness and strength (for metals usually the yield strength) of the adherends • Shear modulus, yield strength, and strain-to-failure of the adhesive • Thermal expansion coefficient of the adherends • Magnitude and direction of the applied loads • Overlap length of the adherends • Thickness of the adherends • Thickness of the adhesive The properties used must be sufficient to handle the weakest state of the materials; for the adhesive and composite adherends, this is usually the hot/wet condition. It is most important to ensure that the strength of the adherend/ adhesive interface does not become significantly weakened as a result of environmental degradation. For a degraded interface, there is no way of
294 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES quantifying minimum strength;even zero is a possibility.Environmental degradation of the interface in service is much more likely if the adherends are not given the correct surface treatment before bonding.Suitable methods will be discussed later. Fatigue damage or creep in the adhesive layer can be avoided,or at least minimized,by maintaining the adhesive in an elastic state for most of its service life.Ideally,significant plastic deformation of the adhesive should be permitted only when the joint is stressed to limit load.Limit load is the highest load expected during the service life of the aircraft.Even at ultimate load(1.5x limit), the strain in the adhesive should not approach the failure strain. The design aim is to maintain the adhesive in a state of shear or compression.Structural adhesive joints (and composites)have relatively poor resistance to through-thickness(peel)stresses and,where possible,this type of loading is avoided.The classical joint types,suitable for joining composites to either composites or metals2(Fig.9.2),are 1)the double lap,2)the single lap,3)the single scarf,4)double scarf,5)the single-step lap,and 6)the double-step lap. Figure 9.3,by Hart-Smith,3 illustrates schematically the load-carrying capacities of these joints and some simple design improvements. The single-lap joint is generally the cheapest of all joints to manufacture. However,because the loads are offset (eccentric),a large secondary bending moment develops that results in the adhesive being subjected to severe peel stresses.This type of joint is therefore only used for lightly loaded structure or is supported by underlying structure such as an internal frame or stiffener. The double-lap joint has no primary bending moment because the resultant load is collinear.However,peel stresses arise due to the moment produced by the unbalanced shear stresses acting at the ends of the outer adherends.The resulting stresses,although relatively much smaller in magnitude than in the single-lap joint,produce peel stresses limiting the thickness of material that can be joined. Peel(and shear)stresses in this region are reduced by tapering the ends of the joint.As shown in Figure 9.3,this markedly increases the load capacity of this joint. The scarf and step-lap joints,when correctly designed,develop negligible peel stresses and may be used (at least in principle)to join composite components of any thickness. To explore the feasibility of using primary lap joints that use only adhesive bonding,the USAF funded the Primary Adhesively Bonded Structure Technology (PABST)program which,although concerned with the bonding of aluminum alloy airframe components,must be mentioned as a landmark in the development of bonded joints for aeronautical applications;many of its conclusions are relevant to bonded composite construction.The Douglas Aircraft Company was the major contractor.The program (based on a full-scale section of fuselage for a military transport aircraft)demonstrated that significant improvements could be obtained in integrity,durability,weight,and cost in an
294 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES quantifying minimum strength; even zero is a possibility. Environmental degradation of the interface in service is much more likely if the adherends are not given the correct surface treatment before bonding. Suitable methods will be discussed later. Fatigue damage or creep in the adhesive layer can be avoided, or at least minimized, by maintaining the adhesive in an elastic state for most of its service life. Ideally, significant plastic deformation of the adhesive should be permitted only when the joint is stressed to limit load. Limit load is the highest load expected during the service life of the aircraft. Even at ultimate load (1.5x limit), the strain in the adhesive should not approach the failure strain. The design aim is to maintain the adhesive in a state of shear or compression. Structural adhesive joints (and composites) have relatively poor resistance to through-thickness (peel) stresses and, where possible, this type of loading is avoided. The classical joint types, suitable for joining composites to either composites or metals 2 (Fig. 9.2), are 1) the double lap, 2) the single lap, 3) the single scarf, 4) double scarf, 5) the single-step lap, and 6) the double-step lap. Figure 9.3, by Hart-Smith, 3 illustrates schematically the load-carrying capacities of these joints and some simple design improvements. The single-lap joint is generally the cheapest of all joints to manufacture. However, because the loads are offset (eccentric), a large secondary bending moment develops that results in the adhesive being subjected to severe peel stresses. This type of joint is therefore only used for lightly loaded structure or is supported by underlying structure such as an internal frame or stiffener. The double-lap joint has no primary bending moment because the resultant load is collinear. However, peel stresses arise due to the moment produced by the unbalanced shear stresses acting at the ends of the outer adherends. The resulting stresses, although relatively much smaller in magnitude than in the single-lap joint, produce peel stresses limiting the thickness of material that can be joined. Peel (and shear) stresses in this region are reduced by tapering the ends of the joint. As shown in Figure 9.3, this markedly increases the load capacity of this joint. The scarf and step-lap joints, when correctly designed, develop negligible peel stresses and may be used (at least in principle) to join composite components of any thickness. To explore the feasibility of using primary lap joints that use only adhesive bonding, the USAF funded the Primary Adhesively Bonded Structure Technology (PABST) program 4 which, although concerned with the bonding of aluminum alloy airframe components, must be mentioned as a landmark in the development of bonded joints for aeronautical applications; many of its conclusions are relevant to bonded composite construction. The Douglas Aircraft Company was the major contractor. The program (based on a full-scale section of fuselage for a military transport aircraft) demonstrated that significant improvements could be obtained in integrity, durability, weight, and cost in an
JOINING OF COMPOSITE STRUCTURES 295 Double overlap Single overlap Scarf Double scarf Stepped lap Double-stepped lap Fig.9.2 Schematic illustration of several types of bonded joint. aluminum alloy fuselage component by the extensive use of bonded construction. The demonstrated weight-saving was about 15%,with a 20%saving in cost. Lap joints relying solely on adhesive bonding,although structurally very attractive,are not generally used by major aircraft manufacturers in primary structural applications(such as fuselage splice joints)because of concerns with long-term environmental durability.These concerns stem from some early poor service experience with the environmental durability of adhesive bonds,resulting from the use of inadequate pre-bonding surface treatments and ambient-curing adhesives
JOINING OF COMPOSITE STRUCTURES 295 Double overlap Single overlap Scarf Double scarf ~===---~====~::::.~ ..... ,/ Stepped lap ZZZZZZZZ-'--ZZZZ_ZZZZZ__'_=_~ Double-stepped lap Fig. 9.2 Schematic illustration of several types of bonded joint. aluminum alloy fuselage component by the extensive use of bonded construction. The demonstrated weight-saving was about 15%, with a 20% saving in cost. Lap joints relying solely on adhesive bonding, although structurally very attractive, are not generally used by major aircraft manufacturers in primary structural applications (such as fuselage splice joints) because of concerns with long-term environmental durability. These concerns stem from some early poor service experience with the environmental durability of adhesive bonds, resulting from the use of inadequate pre-bonding surface treatments and ambient-curing adhesives
296 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES SCARF-AND STEPPED-LAP JOINTS FAILURES OUTSIDE SHEAR FAILURES FAILURES SHOWN REPRESENT THE BEST POSSIBLE FROM TAPERED-LAP JOINT EFFICIENT DESIGN FOR EACH GEOMETRY LAMINATE STRENGTH OUTSIDE JOINT DOUBLE-LAP JOINT PEEL FAILURES SINGLE-LAP JOINT BENDING OF AOHFRENDS DUE TO ECCENTRIC LOAD PATH ADHEREND THICKNESS Fig.9.3 Load-carrying capacity of adhesive joints.Taken from Ref.3. 9.3.2 Design/Analysis of Bonded Lap Joints Reviews of analytical procedures for joints involving composites are provided in Refs.5 and 6.Hart-Smith undertook comprehensive analytical studies7-9 on adhesive joints,particularly advanced fiber composite to composite and composite to metal joints.His studies,based on the earlier approaches,cover the important aspect of non-linear (elastic/plastic)deformation in the adhesive. The stress level for joint (adhesive)failure is determined by shear strain to failure of the adhesive(y+yp)in the bondline;the design aim being that this stress level should well exceed adherend strength.Peel stresses are avoided by careful design rather than considered as a potential failure mode. Several earlier attempts were made to represent non-linear behavior in the adhesive assuming realistic shear stress/shear strain behavior,but they were too
296 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES O Z o SCAPJ:- AND STEPPED-LAP JOINTS i , , J" FAILURES SHOWN REPRESENT THE BEST POSSIBLE FROM EFFICIENT DI~SION FOR EACH GEOMETRY TAPERED-!~P JOINT ' DOUBLE-LAP JOINT II . , '! PEEL FAILURES SINGLE-LAP JOINT i BENDING OF AOHFRENDS DUE TO ECCENTRIC LOAD PATH ADHEREND THICKNESS t Fig. 9.3 Load-carrying capacity of adhesive joints. Taken from Ref. 3. 9.3.2 Design~Analysis of Bonded Lap Joints Reviews of analytical procedures for joints involving composites are provided in Refs. 5 and 6. Hart-Smith undertook comprehensive analytical studies 7-9 on adhesive joints, particularly advanced fiber composite to composite and composite to metal joints. His studies, based on the earlier approaches, cover the important aspect of non-linear (elastic/plastic) deformation in the adhesive. The stress level for joint (adhesive) failure is determined by shear strain to failure of the adhesive (% + "yp) in the bondline; the design aim being that this stress level should well exceed adherend strength. Peel stresses are avoided by careful design rather than considered as a potential failure mode. Several earlier attempts were made to represent non-linear behavior in the adhesive assuming realistic shear stress/shear strain behavior, but they were too
JOINING OF COMPOSITE STRUCTURES 297 Table 9.2 Computer Programs Developed by Dr John Hart-Smith for Stress Analysis of Bonded Joints Joint to be analyzed Program Joint to be analyzed Program Single-lap joint: A4EA Double-lap joint:Elastic A4EB Joint strengths and efficiencies adherend and elastic/plastic in non-dimensional form. adhesive. Deals only with identical Can deal with unbalanced adherends.Three failure cases joints and allows for thermal are considered:a)adherend mismatch between adherends. bending,b)adhesive shear, Provides ratio of maximum to and c)adhesive peel. average shear strength and non-dimensionalized joint strength. Scarf Joint: A4EE Step-lap joint: A4EG Elastic adherend and elastic/ Elastic adherend and elastic/ plastic adhesive. plastic adhesive. Provides a)shear stress Provides a)shear stress distribution along the joint b) distribution along the joint, displacement of inner and b)displacement of inner and outer adherends,and c) outer adherends,and c) potential joint strength. potential joint strength. Step-lap joint: A4EI Elastic adherend and elastic/ plastic adhesive. Similar to A4EG but more comprehensive;allows for variations in adhesive thickness and adhesive defects.Bond width can also be varied. complex for most analytical approaches.However,as discussed later,Hart-Smith shows that a simple elastic/ideally plastic formulation gives similar results to more realistic representations of adhesive behavior,providing the strain energy density in shear in the adhesive (area under the stress-strain curve)is comparable to that expected for the real curve. As a major part of these studies,software programs were developed for the analysis of double-overlap and the other types of joint discussed here;these are listed in Table 9.2.Similar programs are available through the Engineering Sciences Data Unit (ESDU)9 and proprietary programs have been developed by manufacturers. Inevitably,many of the complications in real joints are neglected or inadequately dealt with in these relatively simple studies.These include:
Table 9.2 JOINING OF COMPOSITE STRUCTURES Computer Programs Developed by Dr John Hart-Smith for Stress Analysis of Bonded Joints 297 Joint to be analyzed Program Joint to be analyzed Program Single-lap joint: A4EA Double-lap joint: Elastic A4EB Joint strengths and efficiencies adherend and elastic/plastic in non-dimensional form. adhesive. Deals only with identical Can deal with unbalanced adherends. Three failure cases joints and allows for thermal are considered: a) adherend mismatch between adherends. bending, b) adhesive shear, Provides ratio of maximum to and c) adhesive peel. average shear strength and non-dimensionalized joint strength. A4EE Step-lap joint: A4EG Elastic adherend and elastic/ plastic adhesive. Provides a) shear stress distribution along the joint, b) displacement of inner and outer adherends, and c) potential joint strength. Scarf Joint: Elastic adherend and elastic/ plastic adhesive. Provides a) shear stress distribution along the joint b) displacement of inner and outer adherends, and c) potential joint strength. Step-lap joint: Elastic adherend and elastic/ plastic adhesive. Similar to A4EG but more comprehensive; allows for variations in adhesive thickness and adhesive defects. Bond width can also be varied. A4EI complex for most analytical approaches. However, as discussed later, Hart-Smith shows that a simple elastic/ideally plastic formulation gives similar results to more realistic representations of adhesive behavior, providing the strain energy density in shear in the adhesive (area under the stress-strain curve) is comparable to that expected for the real curve. As a major part of these studies, software programs were developed for the analysis of double-overlap and the other types of joint discussed here; these are listed in Table 9.2. Similar programs are available through the Engineering Sciences Data Unit (ESDU)9 and proprietary programs have been developed by manufacturers. Inevitably, many of the complications in real joints are neglected or inadequately dealt with in these relatively simple studies. These include:
298 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Influence of flaws in the form of local porosity,local disbonds,etc. Adhesive thickness variations .Through-thickness variation of shear stresses Through-thickness stresses .Stress-free state at the ends of the adhesive .Highly beneficial effect of adhesive spew,excess adhesive that forms a fillet at the edges of the joint True shear stress/shear strain behavior Most of these complexities are best modelled using finite element procedures.For example,the simple analytical procedures for lap joints mentioned here predict that the maximum shear stress occurs at the free ends of the overlap.However, because the end of the overlap is a free surface,the principle of complimentary shears is violated because the horizontal shear force at the ends cannot be balanced by a vertical shear force.In reality,therefore,the stress along the bond line right at the edge must fall to zero.More realistic stress analysis-using the finite element approach shows that this is the case-shear stress falls rapidly to zero over a distance of the order of the adherend thickness;these observations are confirmed by direct experimental observations.However,the shear stress distribution along the bond line and magnitude of the maximum stress predicted by the simple analytical procedures turns out to be approximately correct.Similar observations have been made concerning normal or peel stresses. A further considerable complication,difficult to handle even with finite element methods,is the time dependency or viscoelastic (and viscoplastic) behavior of adhesives. 9.3.3 Models for Adhesive Stress/Strain Behavior For analysis of stress distribution in the joint,a model for the shear stress/ strain behavior of the adhesive is required.The simplest model assumes that the adhesive is strained only within its linear elastic range.This model may be adequate if fatigue is a major concern and the primary aim is to avoid plastic cycling of the adhesive;then the stresses must not be allowed to exceed Tp. However,use of the elastic model is overly conservative for assessing the static strength of a joint,particularly if it is bonded with a highly ductile adhesive. To account for plastic deformation,the actual stress/strain behavior must be modelled.In computer-based approaches such as the finite element method,the stress/strain curve can be closely modelled using the actual constitutive rela- tionship.However,for analytical approaches,much simpler models are needed. Figure 9.4 shows stress/strain behavior for a typical ductile adhesive and the models of this behavior used for joint analysis by Hart-Smith.The intuitive simple non-linear model is the bilinear characteristic because this most closely approximates to the real curve.However,even use of this simple model is mathematically complex,greatly limiting the cases that can be analyzed to produce closed-form solutions
298 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES • Influence of flaws in the form of local porosity, local disbonds, etc. • Adhesive thickness variations • Through-thickness variation of shear stresses • Through-thickness stresses • Stress-free state at the ends of the adhesive • Highly beneficial effect of adhesive spew, excess adhesive that forms a fillet at the edges of the joint • True shear stress/shear strain behavior Most of these complexities are best modelled using finite element procedures. For example, the simple analytical procedures for lap joints mentioned here predict that the maximum shear stress occurs at the free ends of the overlap. However, because the end of the overlap is a free surface, the principle of complimentary shears is violated because the horizontal shear force at the ends cannot be balanced by a vertical shear force. In reality, therefore, the stress along the bond line right at the edge must fall to zero. More realistic stress analysis--using the finite element approach 1° shows that this is the case--shear stress falls rapidly to zero over a distance of the order of the adherend thickness; these observations are confirmed by direct experimental observations. However, the shear stress distribution along the bond line and magnitude of the maximum stress predicted by the simple analytical procedures turns out to be approximately correct. Similar observations have been made concerning normal or peel stresses. A further considerable complication, difficult to handle even with finite element methods, is the time dependency or viscoelastic (and viscoplastic) behavior of adhesives. 9.3.3 Models for Adhesive Stress/Strain Behavior For analysis of stress distribution in the joint, a model for the shear stress/ strain behavior of the adhesive is required. The simplest model assumes that the adhesive is strained only within its linear elastic range. This model may be adequate if fatigue is a major concem and the primary aim is to avoid plastic cycling of the adhesive; then the stresses must not be allowed to exceed "rp. However, use of the elastic model is overly conservative for assessing the static strength of a joint, particularly if it is bonded with a highly ductile adhesive. To account for plastic deformation, the actual stress/strain behavior must be modelled. In computer-based approaches such as the finite element method, the stress/strain curve can be closely modelled using the actual constitutive relationship. However, for analytical approaches, much simpler models are needed. Figure 9.4 shows stress/strain behavior for a typical ductile adhesive and the models of this behavior used for joint analysis by Hart-Smith. 11 The intuitive simple non-linear model is the bilinear characteristic because this most closely approximates to the real curve. However, even use of this simple model is mathematically complex, greatly limiting the cases that can be analyzed to produce closed-form solutions