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