TH 18 INTERNATIONAL CONFERENCE ON COMPOSITE MATERIALS

COMMERCIAL TRANSPORT COMPOSITE WING BOX TRADE STUDY

W. Kang 1*, I.S. Hwang 1 1 Aeronautical System Division, Korea Aerospace Research Institute, Daejeon, Korea * Corresponding author( [email protected] )

Keywords : composite, transport airplane, trade study, Finite element, HyperSizer, weight cost optimization 1 Introduction introduced to the conventional 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 concepts are estimated for the 90 seats turboprop for A300/310 by 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 , inboard and outboard flaps, structure weight. In the late of 1970, NASA and and . 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 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 for model was constructed. Almost all structures L1011. In the US, the most significant use of including skin panels, ribs, spars, , 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 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

Table 1. Material selection criteria Young Yield ’s Densi Fractu Criteria Design Consideration Material Stren modul ty re Price s gth us (lb/in 3 toughn ($/lbs) Static Strength Limit load capability (ksi) (10 6 ) ess psi) Stiffness Deflection requirement Al 2024 47 10.5 0.101 33.7 2~3 Crack initiation, fracture Fatigue Al 7075 71 10.3 0.101 26.4 2~3 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 Bird strike, hail, operation Ko8242 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 Corrosion Galvanic, stress corrosion to be represented by definite numbers in the early stage. Such criteria can be briefly Commercial Availability, Lead summarized as some distinctive features such as 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 & Cost Pro Cons 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 for each criteria. The criteria which cannot be resistance

Brief description of candidate material can be For the structure, which support compression helpful for final decision, as the final call will loads between upper and lower skin panels, full be done by collaboration between many design composite and metal were constructed and compared functions, even including supply chains, in the cost and weight . In the composite rib design, manufacturing and marketing also. With the optimal rib spacing is quite important for overall weight and cost optimiza tion. Ribs are usually definition of selection criteria and each attached skin panel and via metal clip with properties for every candidate, the final material mechanical fasteners. Different rib -panel and spar for each structural components can be fastening concepts are evaluated. determined. Through this work, t he overall weight and manufacturing cost were estimated for each concepts Table 5. Final material selection result and the optimal structure concept was suggested for Component Material Comments the overall wing airframe construction. Conventional Welding for Fuselage Aluminum lower panel Wing CFRP Metal rib Empennage CFRP Thermoplastic GFRP for Removables CFRP impact critical area

4 Full Composite Wing of the Commercial Transport Aircraft Various concepts were considered and validated for the wing box design in the preliminary design phase . Fig.1. Regional Turboprop Wing Configuration . For the skin panel and stringer structure, several different stringer concepts for the skin panel were considered such as T-shaped, C-shaped and blade Table 6. Various skin panel stringer concepts type stringers. The optimal stringer spacing was Design Structural Concept another design variable to be optimized by the suitability efficiency HyperSizer program. In the skin panel design, stringer run-out is qui te important. In our design, we I blade Buckling of flat Symmetric, chose three piece wingbox. In this concept, outer plate with a free Buckling of flat edge Better for plate with a free and center wing box joint design is one of the key issues in the early wingbox configuration set -up. lower stringers edge Wing bending moment is the design loads for upper and lower wing skin panels. T blade Better for lower Symmetric, In addition, two different composite spars were stringers Buckling of flat considered; one is the one piece spar with plate with a free cocured/cobonded stiffeners and the other is the edge three pieces spar with mechanically fastened stiffe ners. Many point loads, such as engine inertia J Section Better for Asymmetric, load, wing-fu selage mating loads, are introduced in metallic Complex the winbox through the spars. Concentrated loads in compression composite structure come to a catastrophic failure in flexural torsion overall structure. Several loads distributing design modes concepts are introduced and evaluated.

3 Table 7 . Two different spar concepts wing has 13% weight reduction potential also. Structural These final weight estimation can be used for Concept Design suitability efficiency fabrication cost for all wing structures. With the weight & cost estimation for metal and Single May require Carries composite wing, the final material for wing and piece spar fasteners to material in each wing component can be decided. reinforce bonds at lower loaded

highly loaded regions Table 8. Weight estimation for metal/composite areas wing Composite 3 piece Pre-cured stiffener No advantage Sample Part Metal Composite weight spar is co-bonded to of composite potential spar web spar material by Upper skin 29.92 22.63 24% web & spar cab is mechanical Lower skin 24.11 20.91 13% mechanically fastening

joined Acknowledgement

This work was supported by the Technology 5 Weight Estimation Development Program for Aeros pace Parts and The final structural concept should be Components of MKE. Authors are gratefully approved by weight & cost comparison to acknowledging the support by the MKE of conventional metallic wing concept. Each Korea. component of full composite wing is sized to the design ultimate loads. Design ultimate loads References are applied to FEM model of full composite wing. Internal loads for each component are [1] B. A. Byers, R. L. Stoecklin “Preliminary Design of calculated by FEM analysis. Detail sizing is Graphite Composite Wing Panels for Commercial Transport Aircraft ”. NASA -CR-159150, 1980. carried for structural elements for metal & [2] Michael Karal “AST Composite Wing Program – composite wing. By using 3D cad program, the Executive Summary”. NASA/CR -2001-210650 initial weights for each element are estimated. [3] Ram C. Madan “Composite Transport Wing To verify the weight potential of composite Technology Development ”. NASA-CR-178409, wing, HyperSizer is used for weight estima tion 1988. of upper and lower skin panels. Uniaxial [4] I. F. Sakata, R. B. Ostrom “Study on Utilization of Advanced Composites in Commercial Aircraft Wing stiffened panel family with T-shaped stringer is Structures”. NASA-CR-1453811, 1978 . selected for optimization process. T-shaped stringers are bonded to skins as defined previous chapter. Panel height, flange width , spacing are optimized main va riables. Final panel weight for metal and composite wing is calculated for upper and lower panels.

6 Conclusion The result in table 8 shows that composite upper skin panel has 24% weight reduction potential to metal wing concepts. The lower