Available online at www.sciencedirect.com SCIENCE DIRECT PROGRESSIN SCIENCES ELSEVIER Progress in Acrospace Sciences 41(2005)143-151 www.elsevier.com/locate/pacrosci Fibre reinforced composites in aircraft construction C.Soutis* Aerospace Engineering.The of Sheffield.Mappin Street.Sheffield SI 3JD.UK Abstract Fibrous h 1.in Nort ahility to shan and tailor their structure to produce mor review odvances using composiesn modemrraft coniructons or Ipolymers,especially ca rced plastics(CFRP)can and mputational simulation of the manufacturing and assembly process as well as the simulation of the performance of the structure. s reserved. Contents Background ..143 4. References.. 150 1.Background ration basis,to military aircraft spoilers.ruddersand doors.With increasing appicti y of on fibre at the o al Aircraft Estab- ras(the osts and the astics)result ishment at Farnborough.UK.in 1964.However,not ing in carbon fibre reinforced plastics (CFRP)compo until the late 1960s did these new composites start to be aluminium and titanium alloys,for primary structures High strength,high modulus carbon fibres are about 376-0421/s.seo dot10.1016fp4 erosc.2005.02.004 2005 Elsevier Ltd.All rights
Progress in Aerospace Sciences 41 (2005) 143–151 Fibre reinforced composites in aircraft construction C. Soutis Aerospace Engineering, The University of Sheffield, Mappin Street, Sheffield S1 3JD, UK Abstract Fibrous composites have found applications in aircraft from the first flight of the Wright Brothers’ Flyer 1, in North Carolina on December 17, 1903, to the plethora of uses now enjoyed by them on both military and civil aircrafts, in addition to more exotic applications on unmanned aerial vehicles (UAVs), space launchers and satellites. Their growing use has risen from their high specific strength and stiffness, when compared to the more conventional materials, and the ability to shape and tailor their structure to produce more aerodynamically efficient structural configurations. In this paper, a review of recent advances using composites in modern aircraft construction is presented and it is argued that fibre reinforced polymers, especially carbon fibre reinforced plastics (CFRP) can and will in the future contribute more than 50% of the structural mass of an aircraft. However, affordability is the key to survival in aerospace manufacturing, whether civil or military, and therefore effort should be devoted to analysis and computational simulation of the manufacturing and assembly process as well as the simulation of the performance of the structure, since they are intimately connected. r 2005 Elsevier Ltd. All rights reserved. Contents 1. Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143 2. Design and analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145 3. Manufacture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 147 4. Applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148 5. Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150 References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150 1. Background The adoption of composite materials as a major contribution to aircraft structures followed from the discovery of carbon fibre at the Royal Aircraft Establishment at Farnborough, UK, in 1964. However, not until the late 1960s did these new composites start to be applied, on a demonstration basis, to military aircraft. Examples of suchdemonstrators were trim tabs, spoilers, rudders and doors. Withincreasing application and experience of their use came improved fibres and matrix materials (thermosets and thermoplastics) resulting in carbon fibre reinforced plastics (CFRP) composites withimproved mechanical properties, allowing them to displace the more conventional materials, aluminium and titanium alloys, for primary structures. High strength, high modulus carbon fibres are about ARTICLE IN PRESS www.elsevier.com/locate/paerosci 0376-0421/$ - see front matter r 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.paerosci.2005.02.004 Tel.: +44 11 42227706; fax: +44 11 42227729. E-mail address: c.soutis@sheffield.ac.uk
14 C.Soutis Proaress in Aerospace Sciences 41 (2005)143-151 56m in diameter and consist of small crystallites of carbon.The rong are weak 的R Fabrics can be wo d plane Young's modulus parallel to the a-axis is 1000 GPa 30GPa.Alignment of the basal plane s of knitting machine.to fibre oSopocsibtke.withcertaing stiff fibre red to the shape of the ity eventu ally speak the cosoPoiasinwoembi.orabopnkmc congated o the fibre during manufacturet setting (epo associated the manufac ropylene. Nylon 6.6 of the is also important since it affects the transverse and shear size to aid bonding to the specified matrix.Wherea propert is pr have a thir tial ide out of gth is.i and a core with important.The aim of the materia pith based exhibit s to a systen alan se the properties erties an lead to imp ved lamina or laminat of cous those of the composites PopcTchcanmpornanticidofibre-malnxnicr proc has t nsile strength (4.5GPa)and in strain to fracture （more than2 PAN- based hi .Thes strongly bonded to the matrix if their high strength an e6is are to high strength(HS.with a modulus of around 230 GPa the interfa A weak interface results in a low stifnes of 4.5 GPa and strength but hi h resistance to fracture.whereas gh-strengr rong intertace pro values of%before fracture.The tensile stresstrain er The selec tion of the ate fibr nd ntal very much on the military aircraft by the characteristics of the interface.In these cases.the oth and high str are desirab relationship betv ween properties and interface charac high-fibre modulus stie nsive experimental evidence ren ector dishes,antennas and their supporting struc are required tun Thermop tic materials are b ovings are the hasic in hich fibr supplicd.a roving being the number of strandsor currently used are the ermosetting epoxics.The matri bundles of filaments wound into a package or creel,the material is the Achilles'heel of the mposite system and ing up the tent
5–6 mm in diameter and consist of small crystallites of ‘turbostratic’ graphite, one of the allotropic forms of carbon. The graphite structure consists of hexagonal layers, in which the bonding is covalent and strong (525 kJ/mol) and there are weak van der Waal forces (o10 kJ/mol) between the layers [1,2]. This means that the basic crystal units are highly anisotropic; the inplane Young’s modulus parallel to the a-axis is approximately 1000 GPa and the Young’s modulus parallel to the c-axis normal to the basal planes is only 30 GPa. Alignment of the basal plane parallel to the fibre axis gives stiff fibres, which, because of the relatively low density of around 2 mg/m3 , have extremely high values of specific stiffness (200 GPa/((mg/m3 )). Imperfections in alignment, introduced during the manufacturing process, result in complex-shaped voids elongated parallel to the fibre axis. These act as stress raisers and points of weakness leading to a reduction in strength properties. Other sources of weakness, which are often associated with the manufacturing method, include surface pits and macro-crystallites. The arrangement of the layer planes in the cross-section of the fibre is also important since it affects the transverse and shear properties of the fibre. Thus, for example, the normal polyacrylonitrile-based (PAN-based) Type I carbon fibres have a thin skin of circumferential layer planes and a core withrandom crystallites. In contrast, some mesophase pith-based fibres exhibit radially oriented layer structures. These different structures result in some significant differences in the properties of the fibres and, of course, those of the composites. Refinements in fibre process technology over the past 20 years have led to considerable improvements in tensile strength(4.5 GPa) and in strain to fracture (more than 2%) for PAN-based fibres. These can now be supplied in three basic forms, high modulus (HM, 380 GPa), intermediate modulus (IM, 290 GPa) and high strength (HS, with a modulus of around 230 GPa and tensile strengthof 4.5 GPa). The more recent developments of the high-strength fibres have led to what are known as high-strain fibres, which have strain values of 2% before fracture. The tensile stress–strain response is elastic up to failure and a large amount of energy is released when the fibres break in a brittle manner. The selection of the appropriate fibre depends very muchon the application. For military aircrafts, both high modulus and high strength are desirable. Satellite applications, in contrast, benefit from use of high-fibre modulus improving stability and stiffness for reflector dishes, antennas and their supporting structures. Rovings are the basic forms in which fibres are supplied, a roving being the number of strands or bundles of filaments wound into a package or creel, the lengthof the roving being up to several kilometres, depending on the package size. Rovings or tows can be woven into fabrics, and a range of fabric constructions are available commercially, suchas plain weave, twills and various satin weave styles, woven witha choice of roving or tow size depending on the weight or density of fabric required. Fabrics can be woven withdifferent kinds of fibre, for example, carbon in the weft and glass in the warp direction, and this increases the range of properties available to the designer. One advantage of fabrics for reinforcing purposes is their ability to drape or conform to curved surfaces without wrinkling. It is now possible, withcertain types of knitting machine, to produce fibre performs tailored to the shape of the eventual component. Generally speaking, however, the more highly convoluted each filament becomes, as at crossover points in woven fabrics, or as loops in knitted fabrics, the lower its reinforcing ability. The fibres are surface treated during manufacture to prepare adhesion with the polymer matrix, whether thermosetting (epoxy, polyester, phenolic, polyimide resins) or thermoplastic (polypropylene, Nylon 6.6, PMMA, PEEK). The fibre surface is roughened by chemical etching and then coated with an appropriate size to aid bonding to the specified matrix. Whereas composite strengthis primarily a function of fibre properties, the ability of the matrix to both support the fibres and provide out-of-plane strength is, in many situations, equally important. The aim of the material supplier is to provide a system witha balanced set of properties. While improvements in fibre and matrix properties can lead to improved lamina or laminate properties, the all-important field of fibre–matrix interface must not be neglected. The load acting on the matrix has to be transferred to the reinforcement via the interface. Thus, fibres must be strongly bonded to the matrix if their high strength and stiffness are to be imparted to the composite. The fracture behaviour is also dependent on the strength of the interface. A weak interface results in a low stiffness and strength but high resistance to fracture, whereas a strong interface produces high stiffness and strength but often a low resistance to fracture, i.e., brittle behaviour. Conflict therefore exists and the designer must select the material most nearly meeting his requirements. Other properties of a composite, suchas resistance to creep, fatigue and environmental degradation, are also affected by the characteristics of the interface. In these cases, the relationship between properties and interface characteristics are generally complex and analytical/numerical models supported by extensive experimental evidence are required. Thermoplastic materials are becoming more available; however, the more conventional matrix materials currently used are thermosetting epoxies. The matrix material is the Achilles’ heel of the composite system and limits the fibre from exhibiting its full potential in terms of laminate properties. The matrix performs a number ARTICLE IN PRESS 144 C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151
C.Soutis Progress in Aerospace Sciences 41 (2005)143-15 of function the fibre ir show that delamination siderable dista ession (providing lateral support).translating the affecting more dramatically the residual strength and fibre properties into the laminat minimising damag stiffness prope es of the co mp site.Another importar m by g plasti PEEK tes is that the Matix-dominated have to be with propo compressive strength) are reduced when the glas and and ethermal be inevitable moisture absorption reduces this temperature more quickly because the lengthy cure schedules for comp tha hmmonctimseuicndingorerealhousae Conve cure at 120-135 an autoclave the with at a higherp erature Sv oughened tolera Z-fibre (carbon. steel or titanium pin go cur ng a e03 The resins mus through the d th z-direction to improve the through lay-up part and have time/ten erature fviscosit forms and the focus is now on affordability The currer suitable fo handling.The characteristi phase is being directed towards affordable processing are in desira ch as non- proc ssing.non-t usually cause fabrication 7.NASA Langley in the USA claims and if this by an 100% damage-tol nt forma cially not desired for a resin transfer moulding(RTM ramme where NCF laminates are pro visc sity f compo et s The composites introduced ogy.which currently results in an expensive solution and reon in 960sand1970 nce produc with a toler e to low mpacts oc urring du Alg ma ure and the 2.Design and analysis enoxy systems ments in this the me y on the use AD laminate oen hole c gths 3-61. facilitate the design pro ess With the introduction of The idea a c hibiting laminated t exhibit an and e the anisotropicprope the gy o ign had to b ieldn hisher-notched compr operties than the mposites should not merely replace the metallic allo (EEK) exceptional EEK t fo ah. a competitor with carbon fibre/epoxies and Al Cuand are not ncountered in the analysi alloys in the aircraft 0 On ol isotropic mat For instance.in a laminate laminates show only an indentation on the impa site connected through their faces shear stresses are deve while in carbon fibre epoxy systems ultrasonic C-scans oped on the faces of each lamina.The transverse stresses
of functions amongst which are stabilising the fibre in compression (providing lateral support), translating the fibre properties into the laminate, minimising damage due to impact by exhibiting plastic deformation and providing out-of-plane properties to the laminate. Matrix-dominated properties (interlaminar strength, compressive strength) are reduced when the glass transition temperature is exceeded whereas with a dry laminate this is close to the cure temperature, the inevitable moisture absorption reduces this temperature and hence limits the application of most high-temperature-cure thermoset epoxy composites to less than 120 1C. Conventional epoxy aerospace resins are designed to cure at 120–135 1C or 180 1C usually in an autoclave or close cavity tool at pressures up to 8 bar, occasionally with a post cure at a higher temperature. Systems intended for high-temperature applications may undergo curing at temperatures up to 350 1C. The resins must have a room temperature life beyond the time it takes to lay-up a part and have time/temperature/viscosity suitable for handling. The resultant resin characteristics are normally a compromise between certain desirable characteristics. For example, improved damage tolerance performance usually causes a reduction in hot–wet compression properties and if this is attained by an increased thermoplastic content then the resin viscosity can increase significantly. Increased viscosity is especially not desired for a resin transfer moulding (RTM) resin where a viscosity of 50 cPs or less is often required, but toughness may also be imparted by the fabric structure suchas a stitched non-crimped fabric (NCF). The first generation of composites introduced to aircraft construction in the 1960s and 1970s employed brittle epoxy resin systems leading to laminated structures witha poor tolerance to low-energy impact caused by runway debris thrown up by aircraft wheels or the impacts occurring during manufacture and subsequent servicing operation. Although the newer toughened epoxy systems provide improvements in this respect, they are still not as damage-tolerant as thermoplastic materials. A measure of damage tolerance is the laminate compression after impact (CAI) and the laminate open hole compressive (OHC) strengths [3–6]. The ideal solution is to provide a composite exhibiting equal OHC and CAI strengths and while the thermoplastics are tougher they have not capitalised on this by yielding higher-notched compression properties than the thermoset epoxy composites. Polyetheretherketone (PEEK) is a relatively costly thermoplastic with good mechanical properties. Carbon fibre reinforced PEEK is a competitor withcarbon fibre/epoxies and Al–Cu and Al–Li alloys in the aircraft industry. On impact, at relatively low energies (5–10 J) carbon fibre–PEEK laminates show only an indentation on the impact site while in carbon fibre–epoxy systems ultrasonic C-scans show that delamination extends a considerable distance affecting more dramatically the residual strength and stiffness properties of the composite. Another important advantage of carbon fibre–PEEK composites is that they possess unlimited shelf-life at ambient temperature; the fabricator does not have to be concerned with proportioning and mixing resins, hardeners and accelerators as with thermosets; and the reversible thermal behaviour of thermoplastics means that components can be fabricated more quickly because the lengthy cure schedules for thermosets, sometimes extending over several hours, are eliminated. It can be seen that in an effort to improve the through-the-thickness strength properties and impact resistance, the composites industry has moved away from brittle resins and progressed to thermoplastic resins, toughened epoxies, through damage-tolerant methodology, Z-fibre (carbon, steel or titanium pins driven through the z-direction to improve the throughthe-thickness properties), stitched fabrics, stitched performs and the focus is now on affordability. The current phase is being directed towards affordable processing methods such as non-autoclave processing, non-thermal electron beam curing by radiation and cost effective fabrication [7]. NASA Langley in the USA claims a 100% improvement in damage-tolerant performance withstitched fabrics relative to conventional materials (ref. Advanced Composites Technology, ACT, programme where NCF laminates are processed by resin film infusion (RFI). It is essential that if composites were to become affordable they must change their basic processes to get away from pre-preg material technology, which currently results in an expensive solution and hence product. However, autoclaved continuous fibre composites will still dominate the high levels of structural efficiency required. 2. Design and analysis Aircraft design from the 1940s has been based primarily on the use of aluminium alloys and as such an enormous amount of data and experience exists to facilitate the design process. With the introduction of laminated composites that exhibit anisotropic properties, the methodology of design had to be reviewed and in many cases replaced. It is accepted that designs in composites should not merely replace the metallic alloy but should take advantage of exceptional composite properties if the most efficient designs are to evolve. Of course, the design should account for through-thethickness effects that are not encountered in the analysis of isotropic materials. For instance, in a laminated structure, since the layers (laminae) are elastically connected through their faces, shear stresses are developed on the faces of each lamina. The transverse stresses ARTICLE IN PRESS C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151 145
146 composite.For example.a brittle polymero her in The laminate stacking quence can significantly the interlaminar norm loot or an f such a large effect. Fracture in that the fatigue strength of a(仕l5/±45)，boron fibre/ catastrophically without warning.but tends to be cis about 175 MP sy from ion to co pull-ou by changing the stacking sequence and thus accounts for matrix debonding and fibre rupt imate analytical mcthods and numerical appr such differenc hnite element (FE) mechani sms but as said earlier it is not yet possible to the inter edg the fat bolted oints (a comple three-di al 3-D of the co osite.In contrast to homogeneou rial blem and help to in which fatigue lure by the initiatio e and kincmatic bo The lay-up geometry of a composite strongly affects damage modes. including fibre/matrix de-bonding but also crack n and hbre Ir aiti to the resence of stres concentrators The selection of nage develops throughout the bulk of the and leads to controlled if optimum residual strength [15-20 toughness is to be material Although thes omplexities(free edge impac shear strengths damage, fatigu life p diction)leng with fracture in metal the mas avings and improvements in aerodynami rch into the ture behaviot sites is ir ncy that resu inite element analysis is al vet nts in stric for predicting the toughness of all composite We are data and using mode not able yet to d csign with c he interfa The key is usin tly it d to In metallic and plastic materials,even relatively brittle Works)in St.Louis.USA. more than 6 months te sipat d in non-e ess and lost in mo with a handful of craze formation in a polymer.In composites.the more attractive [7 nte The majority of aircraft co ha esin)For example if the fibr x ho tal fabric tech nts i may run thr h both the fib and ma dynamic eft can be c ned by moving to doubl without cur the pro of the separate component toughnes On the othe mould tools allow the shar to be tailored to meet the hand,if the bond is weak the crack path be mes very required pertormanc targets at various points in rk of y to ta
ðsz; txz; tyzÞ thus produced can be quite large near a free boundary (free edge, cut-out, an open hole) and may influence the failure of the laminate [8,9]. The laminate stacking sequence can significantly influence the magnitude of the interlaminar normal and shear stresses, and thus the stacking sequence of plies can be important to a designer. It has been reported that the fatigue strength of a ð15= 45Þs boron fibre/ epoxy laminate is about 175MPa lower than a ð45= 15Þs laminate of the same system [10]. The interlaminar normal stress, szz, changes from tension to compression by changing the stacking sequence and thus accounts for the difference in strengths. In this case, progressive delamination is the failure mode in fatigue. Approximate analytical methods and numerical approaches suchas finite difference and finite element (FE) techniques [9] can be used to analyse the interlaminar stress distributions near free edges, open holes and bolted joints (a complex three-dimensional, 3-D, problem) [11–13], and help to identify the optimum fibre orientation and laminate stacking sequence for the given loading and kinematic boundary conditions. The lay-up geometry of a composite strongly affects not only crack initiation but also crack propagation, with the result that some laminates appear highly notchsensitive whereas others are totally insensitive to the presence of stress concentrators [5]. The selection of fibres and resins, the manner in which they are combined in the lay-up and the quality of the manufactured composite, must all be carefully controlled if optimum toughness is to be achieved. Furthermore, materials requirements for highest tensile and shear strengths of laminates are often incompatible withrequirements for highest toughness. Compared with fracture in metals, research into the fracture behaviour of composites is in its infancy. Much of the necessary theoretical framework is not yet fully developed and there is no simple recipe for predicting the toughness of all composites. We are not able yet to design withcertainty the structure of any composite so as to produce the optimum combination of strengthand toughness. In metallic and plastic materials, even relatively brittle ones, energy is dissipated in non-elastic deformation mechanisms in the region of the crack tip. This energy is lost in moving dislocations in metal and viscoelastic flow or craze formation in a polymer. In composites, the fibres interfere with crack growth, but their effect depends on how strongly they are bonded to the matrix (resin). For example, if the fibre/matrix bond is strong, the crack may run through both the fibre and matrix without deviation, in which case the composite toughness would be low and approximately equal to the sum of the separate component toughness. On the other hand, if the bond is weak the crack path becomes very complex and many separate damage mechanisms may then contribute to the overall fracture work of the composite. For example, a brittle polymer or epoxy resin witha fracture energy G ffi 0:1 kJ m2 and brittle glass fibres with G ffi 0:01 kJ m2 can be combined together in composites some of which have energies of up to 100 kJ m2 . For an explanation of sucha large effect, we must look beyond simple addition. Fracture in composite materials seldom occurs catastrophically without warning, but tends to be progressive, withsubstantial damage widely dispersed through the material. Tensile loading can produce matrix cracking, fibre bridging, fibre pull-out, fibre/ matrix debonding and fibre rupture, which provide extra toughness and delay failure. The fracture behaviour of the composite can be reasonably well explained in terms of some summation of the contributions from these mechanisms but as said earlier it is not yet possible to design a laminated composite to have a given toughness. Another important modelling issue is the fatigue life of the composite. In contrast to homogeneous materials, in which fatigue failure generally occurs by the initiation and propagation of a single crack, the fatigue process in composite materials is very complex and involves several damage modes, including fibre/matrix de-bonding, matrix cracking, delamination and fibre fracture (tensile or compressive failure in the form of fibre microbuckling or kinking) [14]. By a combination of these processes, widespread damage develops throughout the bulk of the composite and leads to a permanent degradation in mechanical properties, notably laminate stiffness and residual strength [15–20]. Although these complexities (free edge effects, impact damage, joints, fatigue life prediction) lengthen the design process, they are more than compensated for by the mass savings and improvements in aerodynamic efficiency that result. The finite element analysis is also a crucial component, and the biggest time-saving strides have been in the user-friendly developments in creating the data and interpreting the results using modern sophisticated graphical user interfaces. The key is using parametric software to generate the geometry and the meshes. Apparently it used to take Boeing (Phantom Works) in St. Louis, USA, more than 6 months to perform the initial FE stiffness and strength analysis for a complete aircraft and this now takes less than 3 weeks witha handful of engineers, so composites can become more attractive [7]. The majority of aircraft control-lift surfaces produced has a single degree of curvature due to limitation of metal fabrication techniques. Improvements in aerodynamic efficiency can be obtained by moving to double curvature allowing, for example, the production of variable camber, twisted wings. Composites and modern mould tools allow the shape to be tailored to meet the required performance targets at various points in the flying envelope. A further benefit is the ability to tailor the aeroelasticity of the surface to further improve the ARTICLE IN PRESS 146 C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151
C.Soutis Progress in Aerospace Sciences 41 (2005)143-151 147 aerodynamic performance.This tailoring can involve intensive hand lay-up techniques to those requiring high adopting laminate configurations that allow the cross- capital investment in automatic tape layers (ATLs). coupling of flexure and torsion such that wing twist can Tape-laying machines operating under numerical con- result from bending and vice versa.FE analysis allows trol are currently limited in production applications to this process of aeroelastic tailoring,along with strength flat lay-up and significant effort is being directed by and dynamic stiffness (flutter)requirements to be machine manufacturers at overcoming these problems performed automatically with a minimum of post- associated with laying on contoured surfaces.The width analysis engineering yielding a minimum mass solution. of UD tape applied varies considerably from about Early composite designs were replicas of those that 150mm down to a single tow for complex structures. employed metallic materials,and as a result the high The cost of machinery is high and deposition rates low. material cost and man-hour-intensive laminate produc- In 1988,the first Cincinnati tape layer was installed in tion jeopardised their acceptance.This was compounded the Phantom Works and in 1995 a seven-axis Ingersol by the increase in assembly costs due to initial difficulties fibre placement machine was installed.This gave the of machining and hole production.The cost is directly capability to steer fibres within an envelope of 40ft x proportional to the number of parts in the assembly and, 20ft with a 32-tow capability.An overwing panel had as a consequence,designs and manufacture techniques been manufactured where it was able to steer around had to be modified to integrate parts,thereby reducing cut-outs.Collaboration with DASA on global optimisa- the number of associated fasteners.A number of tion software was to be completed at the end of 1998. avenues are available for reducing the parts count, This software is claimed to have produced a 13%weight amongst which are the use of integrally stiffened saving.Other applications include an engine cowling structures,co-curing or co-bonding of substructures door,ducting with a complex structure,FA18 E/F and onto lift surfaces such as wings and stabilisers and the T45 horizontal stabiliser skins.Its capacity was extended use of honeycomb sandwich panels.Hand lay-up to take a 6-in wide tape and Boeing 777 has been techniques and conventional assembly results in manu- converted from hand lay-up to fibre placed (back to facturing costs 60%higher than the datum and only back then split)spars with a saving $5000 per set.Bell with the progressive introduction of automated lay-up Textron has a 10-axis Ingersol,contoured automatic and advanced assembly techniques composites compete tape laying machine for the B609 skin lay-up,which is with their metallic counterparts.Also,the introduction placing a 6-in wide T300 tape onto an inner mould line of virtual reality and virtual manufacturing will play an Invar tool with pre-installed hat stringers.Fibre place- enormous role in further reducing the overall cost.The ment and filament winding technologies are also being use of virtual reality models in engineering prior to used to manufacture components for the V22 [7]. manufacture to identify potential problems is relatively Once the component is laid-up on,the mould is new but has already demonstrated great potential.Bell enclosed in a flexible bag tailored approximately to the Textron in the USA made a significant use of IT during desired shape and the assembly is enclosed usually in an the product definition phase (for the V22 Osprey Tilt- autoclave,a pressure vessel designed to contain a gas at rotor,Fig.1)to ensure 'right first time'approach. pressures generally up to 1.5 MPa and fitted with a Other manufacturing tools that can reduce produc- means of raising the internal temperature to that tion cost and make composites more attractive are required to cure the resin.The flexible bag is first Virtual Fabrication(creating parts from raw materials), evacuated,thereby removing trapped air and organic Virtual Assembly (creation of assembly from parts), vapours from the composite,after which the chamber is Virtual Factory (evaluation of the shop floor).Virtual pressurised to provide additional consolidation during manufacturing validates the product definition and cure.The process produces structures of low porosity, optimises the product cost;it reduces rework and less than 1%and high mechanical integrity.Large improves learning. autoclaves have been installed in the aircraft industry capable of housing complete wing or tail sections. Alternatively.low-cost non-autoclave processing 3.Manufacture methods [21]can be used like vacuum moulding (VM), RTM,Fig.2,vacuum-assisted RTM (VARTM)and The largest proportion of carbon fibre composites RFI.The vacuum moulding process makes use of used on primary class-one structures is fabricated by atmospheric pressure to consolidate the material while placing layer upon layer of unidirectional (UD)material curing.thereby obviating the need for an autoclave or a to the designer's requirement in terms of ply profile and hydraulic press.The laminate in the form of pre- fibre orientation.On less critical items.woven fabrics impregnated fibres or fabric is placed on a single mould very often replace the prime unidirectional form.A surface and is overlaid by a flexible membrane,which is number of techniques have been developed for the sealed around the edges of the mould by a suitable accurate placement of the material,ranging from labour clamping arrangement.The space between the mould
aerodynamic performance. This tailoring can involve adopting laminate configurations that allow the crosscoupling of flexure and torsion suchthat wing twist can result from bending and vice versa. FE analysis allows this process of aeroelastic tailoring, along with strength and dynamic stiffness (flutter) requirements to be performed automatically witha minimum of postanalysis engineering yielding a minimum mass solution. Early composite designs were replicas of those that employed metallic materials, and as a result the high material cost and man-hour-intensive laminate production jeopardised their acceptance. This was compounded by the increase in assembly costs due to initial difficulties of machining and hole production. The cost is directly proportional to the number of parts in the assembly and, as a consequence, designs and manufacture techniques had to be modified to integrate parts, thereby reducing the number of associated fasteners. A number of avenues are available for reducing the parts count, amongst which are the use of integrally stiffened structures, co-curing or co-bonding of substructures onto lift surfaces suchas wings and stabilisers and the use of honeycomb sandwich panels. Hand lay-up techniques and conventional assembly results in manufacturing costs 60% higher than the datum and only withthe progressive introduction of automated lay-up and advanced assembly techniques composites compete with their metallic counterparts. Also, the introduction of virtual reality and virtual manufacturing will play an enormous role in further reducing the overall cost. The use of virtual reality models in engineering prior to manufacture to identify potential problems is relatively new but has already demonstrated great potential. Bell Textron in the USA made a significant use of IT during the product definition phase (for the V22 Osprey Tiltrotor, Fig. 1) to ensure ‘right first time’ approach. Other manufacturing tools that can reduce production cost and make composites more attractive are Virtual Fabrication (creating parts from raw materials), Virtual Assembly (creation of assembly from parts), Virtual Factory (evaluation of the shop floor). Virtual manufacturing validates the product definition and optimises the product cost; it reduces rework and improves learning. 3. Manufacture The largest proportion of carbon fibre composites used on primary class-one structures is fabricated by placing layer upon layer of unidirectional (UD) material to the designer’s requirement in terms of ply profile and fibre orientation. On less critical items, woven fabrics very often replace the prime unidirectional form. A number of techniques have been developed for the accurate placement of the material, ranging from labour intensive hand lay-up techniques to those requiring high capital investment in automatic tape layers (ATLs). Tape-laying machines operating under numerical control are currently limited in production applications to flat lay-up and significant effort is being directed by machine manufacturers at overcoming these problems associated withlaying on contoured surfaces. The width of UD tape applied varies considerably from about 150 mm down to a single tow for complex structures. The cost of machinery is high and deposition rates low. In 1988, the first Cincinnati tape layer was installed in the Phantom Works and in 1995 a seven-axis Ingersol fibre placement machine was installed. This gave the capability to steer fibres within an envelope of 40 ft 20 ft witha 32-tow capability. An overwing panel had been manufactured where it was able to steer around cut-outs. Collaboration withDASA on global optimisation software was to be completed at the end of 1998. This software is claimed to have produced a 13% weight saving. Other applications include an engine cowling door, ducting witha complex structure, FA18 E/F and T45 horizontal stabiliser skins. Its capacity was extended to take a 6-in wide tape and Boeing 777 has been converted from hand lay-up to fibre placed (back to back then split) spars with a saving $5000 per set. Bell Textron has a 10-axis Ingersol, contoured automatic tape laying machine for the B609 skin lay-up, which is placing a 6-in wide T300 tape onto an inner mould line Invar tool withpre-installed hat stringers. Fibre placement and filament winding technologies are also being used to manufacture components for the V22 [7]. Once the component is laid-up on, the mould is enclosed in a flexible bag tailored approximately to the desired shape and the assembly is enclosed usually in an autoclave, a pressure vessel designed to contain a gas at pressures generally up to 1.5 MPa and fitted witha means of raising the internal temperature to that required to cure the resin. The flexible bag is first evacuated, thereby removing trapped air and organic vapours from the composite, after which the chamber is pressurised to provide additional consolidation during cure. The process produces structures of low porosity, less than 1% and high mechanical integrity. Large autoclaves have been installed in the aircraft industry capable of housing complete wing or tail sections. Alternatively, low-cost non-autoclave processing methods [21] can be used like vacuum moulding (VM), RTM, Fig. 2, vacuum-assisted RTM (VARTM) and RFI. The vacuum moulding process makes use of atmospheric pressure to consolidate the material while curing, thereby obviating the need for an autoclave or a hydraulic press. The laminate in the form of preimpregnated fibres or fabric is placed on a single mould surface and is overlaid by a flexible membrane, which is sealed around the edges of the mould by a suitable clamping arrangement. The space between the mould ARTICLE IN PRESS C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151 147
148 C.Soutis Progress in Aerospace Sciences 41 (2005)143-151 variations [21].In traditional prepreg technology,the resin has already infiltrated the fibres and processing mainly removes air and volatiles.consolidates and cures. RTM in its simplest form involves a fabric preform being placed in an enclosed cavity and resin forced into the mould to fill the gaps under pressure and cure.The RFI method utilises precast resin tiles with thickness ranging from 0.125 to 0.25 in.This approach reduces the number of consumables used,but is very process- sensitive relying on the resin being of sufficiently low permeability to fully impregnate the fabric before cure advances too far.The use of an autoclave or press to apply pressure varies.The RFI process is being applied within the Advanced Composites Technology (ACT) Programme in conjunction with traditional autoclave Fig.1.V22-Osprey tilt-rotor plane (courtesy of Bell Textron, processing.Heat is the energy source to activate the USA). resin cure,but some resin systems can be activated by radiation.Wright Paterson claim that thermal oven processing could save 90%of autoclave processing time and energy and hence 50%cost.There is also a radiation curing process developed jointly by NASA and Advanced Composites Group (ACG)and of innovative electron beam cured structures being devel- oped by Foster Miller,Lockheed Martin and Oakridge National Laboratories in the USA [7]. The vacuum-assisted RTM is a liquid resin infusion process and is currently considered by the aircraft industry to be the favoured low-cost manufacturing process for the future.It is an autoclave-free process that Fig.2.Aircraft wing rib element produced by RTM. has been identified as reducing the cost of component processing.It is reported that dimensional tolerance and mass measurements are comparable with stitched RFI and membrane is then evacuated and the vacuum is autoclave panels.A conventional blade stiffened test maintained until the resin has cured.Quite large,thin panel (3ft x 2ft with 4-in high blades 0.5in thick)has shell mouldings can be made in this way at low cost.The been manufactured recently at NASA by using the majority of systems suitable for vacuum-only processing VARTM method,achieving a reasonable quality. are cured at 60-120C and then postcured typically at Further cost reduction when manufacturing with 180C to fully developed properties.In 1991,the composites will be achieved by reducing the assembly evaluation of this method started at the Phantom Works cost,by moving away from fastening (drilling of using the resin system LTM10 (low-temperature mould- thousands of holes followed by fastener insertion and ing)and they created a small allowables database for sealing)towards bonding and to assembly with less or their X36 fighter research aircraft study.In 1996. no expensive jigging.Bell Textron among others are McDonnel Douglas characterised LTM45 EL for the building and developing a number of structures(for the Joint Strike Force(JSF)prototype and generated design V22 and B609)where they are applying state-of-the-art allowable data.In 1998,Boeing also produced LTM45 composites technology/processes to achieve a unitised EL data.LTM10 applications demonstrated for com- approach to manufacturing and assembly.There are of plex parts with a 140F cure under vacuum include a course significant certification challenges with an adhe- serpent inlet duct.A box using LTM10 was shown at the sively bonded joint for a primary aircraft structure 1998 Farnborough Airshow.A research programme at application that need to be addressed. NASA Langley is looking at the development of 180C material properties using low-temperature curing resins. The main advantages of LTM systems are the potential 4.Applications to use autoclave free cures,the use of cheaper tooling and reduced springback of parts. In the pioneering days of flight,aircraft structures RTM and RFI are the predominant curing processes were composite being fabricated largely of wood being developed today of which there are several (natural composite),wire and fabric.Aluminium alloys
and membrane is then evacuated and the vacuum is maintained until the resin has cured. Quite large, thin shell mouldings can be made in this way at low cost. The majority of systems suitable for vacuum-only processing are cured at 60–120 1C and then postcured typically at 180 1C to fully developed properties. In 1991, the evaluation of this method started at the Phantom Works using the resin system LTM10 (low-temperature moulding) and they created a small allowables database for their X36 fighter research aircraft study. In 1996, McDonnel Douglas characterised LTM45 EL for the Joint Strike Force (JSF) prototype and generated design allowable data. In 1998, Boeing also produced LTM45 EL data. LTM10 applications demonstrated for complex parts witha 140 F cure under vacuum include a serpent inlet duct. A box using LTM10 was shown at the 1998 FarnboroughAirshow. A researchprogramme at NASA Langley is looking at the development of 180 1C material properties using low-temperature curing resins. The main advantages of LTM systems are the potential to use autoclave free cures, the use of cheaper tooling and reduced springback of parts. RTM and RFI are the predominant curing processes being developed today of which there are several variations [21]. In traditional prepreg technology, the resin has already infiltrated the fibres and processing mainly removes air and volatiles, consolidates and cures. RTM in its simplest form involves a fabric preform being placed in an enclosed cavity and resin forced into the mould to fill the gaps under pressure and cure. The RFI method utilises precast resin tiles with thickness ranging from 0.125 to 0.25 in. This approach reduces the number of consumables used, but is very processsensitive relying on the resin being of sufficiently low permeability to fully impregnate the fabric before cure advances too far. The use of an autoclave or press to apply pressure varies. The RFI process is being applied within the Advanced Composites Technology (ACT) Programme in conjunction withtraditional autoclave processing. Heat is the energy source to activate the resin cure, but some resin systems can be activated by radiation. Wright Paterson claim that thermal oven processing could save 90% of autoclave processing time and energy and hence 50% cost. There is also a radiation curing process developed jointly by NASA and Advanced Composites Group (ACG) and of innovative electron beam cured structures being developed by Foster Miller, Lockheed Martin and Oakridge National Laboratories in the USA [7]. The vacuum-assisted RTM is a liquid resin infusion process and is currently considered by the aircraft industry to be the favoured low-cost manufacturing process for the future. It is an autoclave-free process that has been identified as reducing the cost of component processing. It is reported that dimensional tolerance and mass measurements are comparable withstitched RFI autoclave panels. A conventional blade stiffened test panel (3 ft 2 ft with 4-in high blades 0.5 in thick) has been manufactured recently at NASA by using the VARTM method, achieving a reasonable quality. Further cost reduction when manufacturing with composites will be achieved by reducing the assembly cost, by moving away from fastening (drilling of thousands of holes followed by fastener insertion and sealing) towards bonding and to assembly withless or no expensive jigging. Bell Textron among others are building and developing a number of structures (for the V22 and B609) where they are applying state-of-the-art composites technology/processes to achieve a unitised approachto manufacturing and assembly. There are of course significant certification challenges with an adhesively bonded joint for a primary aircraft structure application that need to be addressed. 4. Applications In the pioneering days of flight, aircraft structures were composite being fabricated largely of wood (natural composite), wire and fabric. Aluminium alloys ARTICLE IN PRESS Fig. 2. Aircraft wing rib element produced by RTM. Fig. 1. V22-Osprey tilt-rotor plane (courtesy of Bell Textron, USA). 148 C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151
C.Soutis Progress in Aerospace Sciences 41 (2005)143-151 Advanced composites airframe proo (ACAP to the recent time when they were able to achieve a 20%reduction in (DH98 plywoo 4 V22)are V? point design with wood.The whic was only possible in composites at low enou s (from 1960s onwards) phenolic resin incorpo ating The skins of the V22 wing are I-stiffened with co-bonded regarded as the nt civil aird nd ut it is to fh on replacing the secondary structure with fibrous Over 60%of the whoe vehicleweigh is composites whe th have th in GR he nd of the osite coup a ction but is mo dom of its greater tolerance to environ the compos primary erodynamic and structural member fabricated 400 A1 used.In addition.the CFRP fin box cor only 9 -A380 compared with 2076 parts A320h the the horizontal stabiliser in addition to the plethora of leading to a an indicat tion of the b fit of such w has beene stimated that I kg weight reduction saves ove are Fig.3.Airbus A380(courtesy of Airbus S.A.S in the A 380 suner iumbo airliner that is cu tly built by the new 550-seat A380.Fig.3. d the laminate is laid up.and melts when heat is applie system (GLARE)will be used for the A380 fuselage sites have been used in Bell helic ters (Dallas Fort Worth.USA)since the 1980s following their Fig.4.Eurofighter-Typhoon
took over in the 1930s and have dominated the industry to the recent time. Wooden structures did however persist until world war II and the de Havilland mosquito aircraft (DH98) constructed of a plywood–balsa–plywood sandwich laminate probably represents the high point of engineering design withwood. The DH91 Albatross airliner in 1937 was moulded as a ply–balsa– ply sandwichconstruction and the Spitfire fuselage in 1940 was designed and built of Gordon Aerolite material that was a phenolic resin incorporating untwisted flax fibres that could be regarded as the precursor of modern fibre reinforced plastics. Current civil aircraft applications have concentrated on replacing the secondary structure with fibrous composites where the reinforcement media have either been carbon, glass, Kevlar or hybrids of these. The matrix material, a thermosetting epoxy system is either a 125 or 180 1C curing system withthe latter becoming dominant because of its greater tolerance to environmental degradation. Typical examples of the extensive application of composites in this manner are the Boeing 757, 767 and 777 and from Europe the Airbus A310, A320, A330 and A340 airliners. The A310 carries a vertical stabiliser (8.3 m high by 7.8 m wide at the base) a primary aerodynamic and structural member fabricated in its entirety from carbon composite (now £10-20/kg for large tow HS fibre) witha total weight saving of almost 400 kg when compared with the Al alloy unit previously used. In addition, the CFRP fin box comprises only 95 parts excluding fasteners, compared with2076 parts in the metal unit, thus making it easier to produce. The A320 has extended the use of composites to the horizontal stabiliser in addition to the plethora of panels and secondary control surfaces leading to a weight saving of 800 kg over Al alloy skin construction. As an indication of the benefit of such weight saving it has been estimated that 1 kg weight reduction saves over 2900 l of fuel per year. Larger amounts of FRPs are used in the bigger A330, A340 models and of course in the A380 super jumbo airliner that is currently built by the Airbus consortium. GKN Aerospace Services at Cowes, Isle of Wight, UK, is committed to produce some 70 fixed trailing edge panels for the wings of each of the new 550-seat A380, Fig. 3. The wing trailing edge panels are made of glass and CFRPs using a new RFI method, in which resin film, interleaved between glass and carbon fabric layers, when the laminate is laid up, and melts when heat is applied. Melted low-viscosity resin migrates easily through the thickness of the laminate where it cures to form the final component. A hybrid aluminium/glass reinforced plastic system (GLARE) will be used for the A380 fuselage crown that results in reduced weight, increased damage tolerance and improved fatigue life. Composites have been used in Bell helicopters (Dallas Fort Worth, USA) since the 1980s following their Advanced Composites Airframe Programme (ACAP) when they were able to achieve a 20% reduction in weight on metallic airframes. All blades on their newer vehicles (412, 407, 427, 214, 609, OH58D, V22) are all composite. The V22 Osprey tilt-rotor has an allcomposite wing, chosen for its stiffness critical design, which was only possible in composites at low enough weight. Early demonstrators (from 1960s onwards) did not meet expectations until composites were available. The skins of the V22 wing are I-stiffened with co-bonded spars and bolted on ribs (the civil 609 version will use bonded ribs). The pylon support spindle is currently filament wound but it is planned to fibre place this part. Over 60% of the whole vehicle weight is carbon composite, plus a further 12% in GRP. The V22 uses tape laying, hand lay-up and filament winding for most of the composite construction but is moving to fibre placement for the 609 civil version [7]. Mechanical fastening features heavily in the composite structure, some 3000 on eachside of the wing, is introduced by ARTICLE IN PRESS Fig. 3. Airbus A380 (courtesy of Airbus S.A.S). Fig. 4. Eurofighter-Typhoon. C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151 149
150 C.Soutis Proaress in Aerospace Sciences 41 (2005)143-151 manual drilling with templates. ft the has been References parts (including fasteners).weight reduced by 20% (260kg) nth the e [Kelly A edito ew York:Pergamon;199 ently being designed or built in the USA USF or F. materials.Cam 35)and Europe(EFA.Fig.4)contain in the region of Uniw sity Press: 198 in the fibre T800/924C agility of the aircraft would be lost if this amount of Fleck NA. mith PA.Failure composite material was not used because of the the mat in the paper have been implemented in their construc on of CFRI concept that requires the designer to achieve the smallest possible rada (RCS),to e in the USA.Report for tial c constant change of radius of the airfra sets is much easier to form in ☒Matthews FL Rawlings RD nposite n (RAM) atis C.Finit mate nals and structures 5.Summary composite materials.London 107 The application of carbon fibre has developed from yCP.SoutisC.Exper Rtal and small-scale technology demonstrators in the 1970s to aminates.RaeS autJ1998:1021018 the price of carbon fibre has dropped to less tha :66143- applications such that the mass and part reduction.complex shape manufacture. utis C.Fleck NA.Smith PA.Com atigu ge ges re ricting their use are material and [15]Zhang J.Fa tisC.Analysis of costs.impact damage and damage e1992.235291-8 24 repair and 9.30 associated with uncertaintics about relatively new transverse ply cracks.Composites and sometimes variable materials p re here e to stay in terms savings can be achieved.For weight savings approaching 40 ng 1.Sou C.Fan J.Strai rat 30e with local 1994259
manual drilling withtemplates, but they are looking towards the use of automated drilling and probably involving water jet cutting. Other examples where composites will be extensively applied are the future military cargo Airbus A400M and the tail of the C17 (USA). A 62 ft C-17 tail demonstrator has been successfully completed yielding 4300 fewer parts (including fasteners), weight reduced by 20% (260 kg) and cost by 50% compared withthe existing metal tail. Without exception all agile fighter aircraft currently being designed or built in the USA (JSF or F- 35) and Europe (EFA, Fig. 4) contain in the region of 40% of composites in the structural mass, covering some 70% of the surface area of the aircraft. The essential agility of the aircraft would be lost if this amount of composite material was not used because of the consequential mass increase. Many of the materials, processes and manufacturing methods discussed earlier in the paper have been implemented in their construction. Another interesting relatively new field of development in the military aircraft sphere is that of ‘stealth’, a concept that requires the designer to achieve the smallest possible radar cross-section (RCS), to reduce the chances of early detection by defending radar sets. The essential compound curvature of the airframe with a constant change of radius is much easier to form in composites than in metal while radar absorbent material (RAM) can be effectively produced in composites. 5. Summary The application of carbon fibre has developed from small-scale technology demonstrators in the 1970s to large structures today. From being a very expensive exotic material when first developed relatively few years ago, the price of carbon fibre has dropped to less than £10/kg, which has increased applications such that the aerospace market accounts for only 20% of all production. The main advantages provided by CFRP include mass and part reduction, complex shape manufacture, reduced scrap, improved fatigue life, design optimisation and generally improved corrosion resistance. The main challenges restricting their use are material and processing costs, impact damage and damage tolerance [22–24], repair and inspection [25–28], dimensional tolerance, size effects on strength [29,30] and conservatism associated withuncertainties about relatively new and sometimes variable materials. Carbon fibre composites are here to stay in terms of future aircraft construction, since significant weight savings can be achieved. For secondary structures, weight savings approaching 40% are feasible by using composites instead of light metal alloys, while for primary structures, suchas wings and fuselages, 20% is more realistic. These figures can always improve but innovation is key to making composites more affordable. References [1] Kelly A, editor. 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[9] Matthews FL, Davies GAO, Hitchings D, Soutis C. Finite element modelling of composite materials and structures. Woodhead Publishing Ltd; 2000. [10] Jones RM. Mechanics of composite materials. London: Taylor & Francis; 1975. [11] Andreasson N, Mackinlay CP, Soutis C. Experimental and numerical failure analysis of bolted joints in CFRP woven laminates. RaeS, Aeronaut J 1998;102(1018):445–50. [12] Andreasson N, Mackinlay CP, Soutis C. Tensile behaviour of bolted joints in low temperature cure CFRP woven laminates. Adv Comp Lett 1997;6(6):143–8. [13] Hu FZ, Soutis C, Edge EC. Interlaminar stresses in composite laminates witha circular hole. Compos Struct 1997;37(2):223–32. [14] Soutis C, Fleck NA, SmithPA. Compression fatigue behaviour of notched carbon fibre-epoxy laminates. Int J Fatigue 1991;13(4):303–12. [15] Zhang J, Fan J, Soutis C. Analysis of multiple matrix cracking in [7ym/90n]s composite laminates. Part I: inplane stiffness properties. Composites 1992;23(5):291–8. [16] Zhang J, Fan J, Soutis C. Analysis of multiple matrix cracking in [7ym/90n]s composite laminates. Part II: development of transverse ply cracks. Composites 1992;23(5):299–304. [17] Zhang J, Soutis C, Fan J. Effects of matrix cracking and hygrothermal stresses on the strain energy release rate for edge delamination in composite laminates. Composites 1994;25(1):27–35. [18] Zhang J, Soutis C, Fan J. Strain energy release rate associated withlocal delamination in cracked composite laminates. Composites 1994;25(9):851–62. ARTICLE IN PRESS 150 C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151
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