C. Soutis/ Progress in Aerospace Sciences 41(2005)143-1. of functions amongst which are stabilising the fibre in show that delamination extends a considera ble distance compression (providing lateral support), translating the affecting more dramatically the residual strength and fibre properties into the laminate, minimising damage stiffness properties of the composite. Another important due to impact by exhibiting plastic deformation an advantage of carbon fibre-PEEK composites is that they providing out-of-plane properties to the laminate. possess unlimited shelf-life at ambient temperature; the Matrix-dominated properties (interlaminar strength, fabricator does not have to be concerned with propo compressive strength) are reduced when the glass tioning and mixing resins, hardeners and accelerators as transition temperature is exceeded whereas with a dr with thermosets: and the reversible thermal behaviour of laminate this is close to the cure temperature, the thermoplastics means that components can be fabricated nevitable moisture absorption reduces this temperature more quickly because the lengthy cure schedules for and hence limits the application of most high-tempera thermosets sometimes extending over several hours are ture-cure thermoset epoxy composites to less than eliminated 120°C It can be seen that in an effort to improve the Conventional epoxy aerospace resins are designed to through-the-thickness strength properties and impact cure at 120-135C or 180C usually in an autoclave or resistance, the composites industry has moved away lose cavity tool at pressures up to 8 bar, occasionally from brittle resins and progressed to thermoplast th a post cure at a higher temperature. System intended for high-temperature applications may under- methodology, Z-fibre (carbon, steel or titanium pins go curing at temperatures up to 350C. The resins must driven through the z-direction to improve the through have a room temperature life beyond the time it takes to the-thickness properties), stitched fabrics, stitched per lay-up a part and have time/temperature/viscosity forms and the focus is now on affordability. The current uitable for handling. The resultant resin characteristics phase is being directed towards affordable processing are normally a compromise between certain desirable methods such as non-autoclave processing, non-thermal characteristics. For example, improved damage toler- electron beam curing by radiation and cost effective ance performance usually causes a reduction in hot-wet fabrication [7. NASA Langley in the USa claims a compression properties and if this is attained by an 100% improvement in damage-tolerant performance increased thermoplastic content then the resin viscosity with stitched fabrics relative to conventional materials can increase significantly. Increased viscosity is espe ref. Advanced Composites Technology, ACT, pro- cially not desired for a resin transfer moulding(RTM) gramme where NCF laminates are processed by resin resin where a viscosity of 50 cPs or less is often required, film infusion(RFD). It is essential that if composites were but toughness may also be imparted by the fabric to become affordable they must change their basic structure such as a stitched non-crimped fabric (NCF). processes to get away from pre-preg material technol- The first generation of composites introduced to ogy, which currently results in an expensive solution and aircraft construction in the 1960s and 1970s employed hence product. However, autoclaved continuous fibre brittle epoxy resin systems leading to laminated struc- composites will still dominate the high levels of ures with a poor tolerance structural efficiency required by runway debris thrown up by aircraft wheels or the impacts occurring during manufacture and subsequent servicing operation. Although the newer toughened 2. Design and analysis epoxy systems provide improvements in this respect, they are still not as damage-tolerant as thermoplastic Aircraft design from the 1940s has been based materials. a measure of damage tolerance is the rimarily on the use of aluminium alloys and as such laminate compression after impact (CAD and the an enormous amount of data and experience exists to laminate open hole compressive(OHC) strengths [3-6]. facilitate the design process. With the introduction of The ideal solution is to provide a composite exhibiting laminated composites that exhibit anisotropic proper equal OHC and CAl strengths and while the thermo- ties, the methodology of design had to be reviewed and plastics are tougher they have not capitalised on this by in many cases replaced. It is accepted that designs in yielding higher-notched compression properties than the composites should not merely replace the metallic alloy thermoset epoxy composites. Polyetheretherketone but should take advantage of exceptional composite (PEEK) is a relatively costly thermoplastic with goo properties if the most efficient designs are to evolve. Of mechanical properties. Carbon fibre reinforced PEEK course, the design should account for through-the- a competitor with carbon fibre/epoxies and Al-Cu and thickness effects that are not encountered in the analysis AFLi alloys in the aircraft industry. On impact, at of isotropic materials. For instance, in a laminated relatively low energies (5-10J)carbon fibre-PEEK structure, since the layers (laminae) are elastically laminates show only an indentation on the impact site connected through their faces, shear stresses are devel- while in carbon fibre- epoxy systems ultrasonic C-scans oped on the faces of each lamina. The transverse stressesof 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