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-tempera￾ture-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 under￾go 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 toler￾ance 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 espe￾cially 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 struc￾tures 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 thermo￾plastics 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 propor￾tioning 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 through￾the-thickness properties), stitched fabrics, stitched per￾forms 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, pro￾gramme 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 technol￾ogy, 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 proper￾ties, 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-the￾thickness 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 devel￾oped on the faces of each lamina. The transverse stresses ARTICLE IN PRESS C. Soutis / Progress in Aerospace Sciences 41 (2005) 143–151 145