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 1994259manual 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 construc￾tion. Another interesting relatively new field of develop￾ment 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 produc￾tion. 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 proces￾sing costs, impact damage and damage tolerance [22–24], repair and inspection [25–28], dimensional tolerance, size effects on strength [29,30] and conserva￾tism 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 afford￾able. References [1] Kelly A, editor. Concise encyclopaedia of composite materials. New York: Pergamon; 1994. [2] Hull D. An introduction to composite materials. Cam￾bridge: Cambridge University Press; 1987. [3] Soutis C, Fleck NA. Static compression failure of carbon fibre T800/924C composite plate withsingle hole. J Compos Mater 1990;24(5):536–58. [4] Soutis C, Fleck NA, SmithPA. Failure prediction technique for compression loaded carbon fibre-epoxy laminates withopen holes. J Compos Mater 1991;25(11):1476–98. [5] Souti C, Curtis PT, Fleck NA. Compressive failure of notched carbon fibre composites. Proc R Soc Lond A 1993;440:241–56. [6] Soutis C, Filiou C, Pateau V. Strengthprediction of CFRP plates witha hole under biaxial compression-tension. AIAA J 2000;38(1):110–4. [7] Aerospace composite structures in the USA. Report for the international technology service (Overseas Missions Unit) of the DTI, UK, 1999. [8] Matthews FL, Rawlings RD. Composite materials: en￾gineering and science. New York: Chapman & Hall; 1994. [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: in￾plane 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