18 TH INTERNATIONAL CONFERENCE ON COMPOSITE MATERIALS COMMERCIAL TRANSPORT AIRCRAFT COMPOSITE WING BOX TRADE STUDY W. Kang 1* , I.S. Hwang 1 1 Aeronautical System Division, Korea Aerospace Research Institute, Daejeon, Korea * Corresponding author(wgkang@kari.re.kr) Keywords : composite, transport airplane, trade study, Finite element, HyperSizer, weight cost optimization 1 Introduction introduced to the conventional airframe Recently composite materials are widely used in constructions. In the trade study, weight and cost the building of transport airplanes. The first between conventional concept and new technologies significant application of composite materials in are main items. In this view, composite wingbox commercial transports was an all composite rudder concepts are estimated for the 90 seats turboprop for A300/310 by Airbus in 1983. In 1985, Airbus transport airplane in this work. The wing of 90 seats introduced a composite vertical tail fin in the same turboprop transport is designed as an upper wing airplanes. With the success of A300/310, Airbus type. The wing consists of a center wing box, two introduced a full composite tail structure for A320. outer wingbox at the right and left sides, leading The composite weight of A320 is totaled to 15% of edge and trailing edge, inboard and outboard flaps, structure weight. In the late of 1970, NASA and and ailerons. The wingbox consists of front and rear major airframe companies such as Boeing, Lockheed, spars, upper and lower cover panels and ribs. For the MD started the ACEE program. The main goal of first work of this development, 3D digital model of this program was to reduce airframe structure weight the wing airframe was constructed using CATIA. by using composite materials. With the ACEE Then, aerodynamic, inertia and control loads were program, the empennage of B737 was replaced by calculated for the initial sizing using 2D panel composite materials, MD developed a full composite methods and point mass models of 3D CATIA wing for commercial transport, and Lockheed model. Based on this 3D CAD model, FE analysis designed new composite vertical tail and aileron for model was constructed. Almost all structures L1011. In the US, the most significant use of including skin panels, ribs, spars, leading edge, composites in commercial transports has been on the trailing edge and control surfaces were modeled in B777. Composite structures make up 10 percent of 2D shell elements by MSC/PATRAN while some the structure weight of the B777. The empennage, skin stringers were modeled in 1D element. floor beams, flaps and outboard aileron of B777 Although the material properties are not different were developed by composite materials. Composite along spanwise direction in this FE model, elements materials are used for fuselage and wing structures were divided to some groups to be used in the for recently developed commercial transports by HyperSizer program. Airbus and Boeing. The composite weight The internal loads which were calculated by the FE percentage of A350 and B787 will be more than analysis were used for the detail sizing for each 50%. Both airplanes adopted composite materials for structure components. For this purpose, HyperSizer wing box and fuselage structures. program was used for structure sizing and concept proofs. This commercial program provides structurally optimized results for the panel and beam 2 Trade Study Procedure components based on the initial FE model In the early stage of airplane development, trade studies on many possible structure concepts are main 3 Material Selection job for structure engineers. Structure engineer team studies the possibility of adapting the state of art With technological improvements in the material technologies for their new aircraft. New concepts are properties, or introductions of new airframe
fabrication methods, airframe designer should defined by specific number should be described choose major material concepts for the main in details for the final selection. structural components. For the material selection, selection criteria should be defined first. Table 2. Material properties definition Young Table 1. Material selection criteria Yield ’s Densi Fractu Criteria Design Consideration Material Stren modul ty re Price (lb/in 3 s gth us toughn ($/lbs) (10 6 (ksi) ) ess Static Strength Limit load capability psi) Stiffness Deflection requirement Al 2024 47 10.5 0.101 33.7 2~3 Crack initiation, fracture Al 7075 71 10.3 0.101 26.4 2~3 Fatigue toughness Al-Li 63.2 10.9 0.094 35 5~10 Damage Crack growth rate, critical crack 2198 Tolerance length(residual strength) AlMgSc 46 10.7 0.096 5~10 Ko8242 Bird strike, hail, operation Impact damage mishap 50~10 CFRP 140 22 0.065 2~5 0 Crashworthiness Plasticity, design capability GFRP 60 5 0.056 10 15~25 Density, minimum gage, design Weight allowable Some criteria such as producibility are not easy to be represented by definite numbers in the Corrosion Galvanic, stress corrosion early stage. Such criteria can be briefly summarized as some distinctive features such as Commercial Availability, Lead Producibility Times, Fabrication Alternatives weldability, creep form capability, or (welding) machining/ATL capabilities. Maintainability Repair methods availability Table 3. Al-Li material characteristic description Raw material, fabrication & Pro Cons Cost assembly - - Low densities High price : TRL Including past experiences (2.55~2.58) 5$/lbs - - High elastic Fast crack The selection criteria can be different for each modulus growth in - structural component and different areas for Excellent fatigue compression each component. For leading edge of wing, and cryogenic loading area - impact damage by bird strike is quite more strength and Consistent import than other criteria. Weight & cost are the toughness quality is not key criteria for material selection for all properties, guaranteed - materials. To proceed material selection, Superior fatigue material properties should be defined by number crack growth resistance for each criteria. The criteria which cannot be
For the rib structure, wh which support compression Brief description of candidate ma material can be loads between upper and nd lower skin panels, full helpful for final decision, as the f e final call will composite and metal were c e constructed and compared be done by collaboration between en many design in the cost and weight. In t In the composite rib design, functions, even including sup supply chains, optimal rib spacing is qui quite important for overall manufacturing and marketing als also. With the weight and cost optimiz ization. Ribs are usually definition of selection criteria ria and each attached skin panel and s spar via metal clip with properties for every candidate, the he final material mechanical fasteners. Diff ifferent rib-panel and spar for each structural componen nents can be fastening concepts are evalu aluated. determined. Through this work, th he overall weight and manufacturing cost were es estimated for each concepts Table 5. Final material selectio tion result and the optimal structure c concept was suggested for Component Material Comments the overall wing airframe co construction. Conventional Welding for Fuselage Aluminum lower panel Wing CFRP Metal rib Empennage CFRP Thermoplastic Th GFRP for Removables CFRP im impact critical area 4 Full Composite Wing of the Comm mmercial Transport Aircraft Various concepts were considered an and validated for Fig.1.Regional Turbopro rop Wing Configuration. the wing box design in the preliminary ary design phase. For the skin panel and stringer stru structure, several different stringer concepts for the sk skin panel were Table 6. Various skin p panel stringer concepts considered such as T-shaped, C-sha haped and blade type stringers. The optimal stringer ger spacing was Design Structural Concept another design variable to be opti ptimized by the suitability efficiency HyperSizer program. In the skin n panel design, I blade Buckling of f f flat Symmetric, stringer run-out is quite important. In In our design, we plate with a f a free Buckling of flat chose three piece wingbox. In this is concept, outer edge Better r for plate with a free and center wing box joint design is o is one of the key lower stringers ers edge issues in the early wingbox configu figuration set-up. Wing bending moment is the design l n loads for upper and lower wing skin panels. T blade Better for low lower Symmetric, In addition, two different composi osite spars were stringers Buckling of flat considered; one is the one piec iece spar with plate with a free cocured/cobonded stiffeners and the the other is the edge three pieces spar with mechanic nically fastened stiffeners. Many point loads, such as as engine inertia J Section Better for Asymmetric, load, wing-fuselage mating loads, are are introduced in metallic Complex the winbox through the spars. Concen centrated loads in compression composite structure come to a catastro strophic failure in flexural torsion overall structure. Several loads distr istributing design modes concepts are introduced and evaluated ted. 3
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