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- 1. OPTIMISATION OF THE DESIGN OF UAV WING J.Alexander, S.Prakash, Dr.BSM.Augustine ABSTRACT This paper presents the research extensively carried out based on thestatic analysis for finding the optimum design of wing structure. The two types ofwings (Rectangular wings) for unmanned airplanes were chosen and designedat the basis of M=0.21 and at M=0.4 respectively. Further it also covers both theaerodynamic and structural designs. The aerodynamics study was achieved byusing the vortex lattice method using XFLR-CFD software and the structuralanalysis was achieved by using the CATIA V5 package consisting of isotropicand composite materials. A wing structure consists of spars, ribs, reinforcements and skin and theywere optimized considering two aspects of weight minimization, and critical loadmaximization. The wing carries an elliptically distributed load along the span.The design variables are based on the positioning of spars and ribs of thestructure. The out come of the research clearly indicates the significant results interms of objective functions obtained through the optimization procedures. W hen composite material is used instead of isotropic material masssaving of 34 % is achieved. The optimum design for each wing was obtainedaccording to the mass, stress and displacement. This could be seen in thepresentation subsequent G: Shear modulus, (pa)Nomenclature ν: Poisson s ratio ρ: Density, (kg/m3)Vn: Normal velocity, (m/s) D.O.F.: Degree of freedomV∞: free stream velocity, (m/s) M.A.C.: Mean aerodynamic chord (m)Θ Flight path angle, (deg.) A/C: AircraftWi: Induced flow component, Gr/Ep: Graphite/epoxy (m/s) H.M Gr/Ep: High modulusα: Angle of attack, (deg.) graphite/epoxyAij: Influence coefficient Sg/Ep: Glass/epoxyLi: Lift of panel i, (N) Wto - total aircraft weightΓ: Strength of vortex (m2/s) Ww :weight of wing structureL: Lift of wing, (N) Wf :weight of fuel stored in wingDi: Drag of panel i, (N)D: Drag of wing, (N) L :length of wingCLcom: Lift coefficient in Lf :length of fuel tank within wingcompressible flow Co :chord length at wing rootCLincom: Lift coefficient in Ct :chord length at wing tipincompressible flow Co :width of fuel tank at wing rootM: Mach numberE: Young s modulus, (pa) Ctf width of fuel tank at Lf structural design of an A/C wing isINTRODUCTION an important factor in th e An Unmanned Air Vehicle (UAV) is performance of the airplanes i.e.an unpiloted aircraft. UAV are obtaining a wing with a highbecoming standard means of stiffness/weight ratio and sustainingcollecting information. The optimum the unexpected loading such as gust and maneuvering situations. This is 1
- 2. accomplished by studying the The shape of the drag force vs.different design parameters required angle of attack is approximatelyto specify the wing geometry. The parabolic. It is desirable for the wingidea of the structural optimization in to have the maximum lift andthe classical sense, has been smallest possible drag.considered to be the minimization of To develop a more accurate model (2)structural mass by varying member for such optimal design studies ,sizes or shell thickness of a model the Structure of the wing thatin which the geometry remains combines the composite(11) andunchanged. Therefore many studies isotropic materials and compare thehave been made during the last wing with the wing that design onlyyears to find the structural from isotropic material in order tooptimization. obtain high strength/weight ratio forOne of the major loading is finding optimum design.aerodynamic loading. Aerodynamiccharacteristics of the UAV (7) vary Types of load acting on thewith certain parameters like the aircraft wingangle of attack and others. 1.Aerodynamic loadingExperimental works on UAVs have 2.Structural loadingbeen conducted in many places with 3.Fuel W eightvarious aerofoil profiles but enough The optimization chain consistswork with the Computational Fluid of:Dynamics (CFD) analysis is not 1 Aerodynamic gridavailable that much till now. Thepresent work contains mainly CFD(8) generationanalysis to determine the flow 2 CFD Analysis by usingpattern a n d th e aerodynamic XFLR softwarecharacteristics of an UAV. The 3 Parametric CFD modelshape of an aircraft is designed to generationmake the airflow through the surface 4 Structural grid generationto produce a lifting force in mostefficient manner. In addition to the 5 FEM Analysis andlift, a force directly opposing the structure sizingmotion of the wing through the air isalways present, which is called a AIM AND SCOPE OF THEdrag force. The angle between the PRESENT INVESTIGATIONrelative wind and the chord line isthe angle of attack of the aerofoil. A light rectangular wing of an UAVThe lift and drag forces developed with NACA0012 airfoil is developedby an aircraft will vary with the using CATIA software and the staticchange of angle of attack. The cross analysis of the wing by CATIA V5sectional shape obtained by the package has been presented andintersection of the wing with the used to determine the optimumperpendicular plane is called an design for the wing. This operationaerofoil. Here NACA 0012(10) involves using the vortex latticesymmetric aerofoil profiles have method to obtain the aerodynamicbeen used for the present research load(lift) for rectangle wing. Thiswork. The lift force increases aerodynamic result (lift) isalmost linearly with the angle of combined(4) with the fuel weight andattack until a maximum value is structural weight of the wing, and it isreached where upon the wing is said used to simulate the wing loading onto stall. the wings during the static analysis. 2
- 3. Some of the parameters of aerofoil The whole wing is divided into manyand properties of air have been kept numbers of panels and each panel isconstant and some have been considered as a bound vortex. Thevaried. The flow of air over the required strength of the bound vortexaerofoil is varied as per requirement. on each panel is to be calculated byThe chord length of the aerofoil is applying a surface flow boundary450mm. (10)The free stream airflow condition. The equation used is thehas been kept 60 m/s and the effect usual condition of zero flow normalof the temperature in the study has to the surface. For each panel thebeen neglected. The density of air condition is applied at the 3/4 chord(ρo)=1.22 kg/m3 , operating pressure position along the centre line of the(Po)=0.101 MPa (1.01 bar) and panel(9). The normal velocity is madeabsolute viscosity up of a free stream component and(μ)=1.789x10-5kg/m-s.The Reynolds an induced flow component, asNumber has been considered as shown in Fig.(2).variable.METHOD OF ANALYSISAERODYNAMICSThe aircraft motion through theatmosphere produces forces calledthe aerodynamic forces mainly thelift and drag forces and the wing isthe major source of this aerodynamic α V Panel iforce. The wing is that surface ofaircraft which supports the aircraft by Fig 2 : Velocity component of themeans of the dynamic reaction on panel i on the wingthe air producing pressuredistribution. This pressuredistribution on the wing changes with Vn = 0 = Sinθ+W i --------------------(1)different wing angles of attack andflight condition. The induced component is a function of strength of all vortex panels on the Methods of calculating aerodynamicload (Lift) wing. Assuming small angles1.Prandtl’s classical lifting line theory Sinθ=sin(α-β)= (α-β)=-dz/dx--------(2)2.Vortex lattice numerical method3.Panel method Wi= Гj ---------------------(3)In this research work vortex latticemethod is used Гj= -Vsin(θ) -----------------(4)Vortex lattice method(8) Where θ is flight path angle Aij is the influence coefficient which represents the induced flow on panel I due to the vortex on panel j A solution for the strength of the vortex lines on each panel is found by solving the matrix of equations the lift coefficient for the wing at a given angle of attack will be obtained by integrating the panel liftFig1: Wing panels distribution. The lift on a particular panel can be found using the Kutta law. 3
- 4. Li = ρV∞ Гj 2k Lift of panel i x = position along wing ka = lift profileL= Lift of wing ---------(4) coefficientThe down wash velocity induced ona panel can be calculated once thestrength of the wing loading isknown. The variation between localflow angles of the panel and the freestream velocity can be found. Figure 3: Derivation of Liftconsequence of this down wash flowis that the direction of action of eachpanels lift vector is rotated relative to We can determine the total lift bythe free stream direction. The local integrating across the length of thelift vectors are rotated backward and wing:hence give rise to a lift induced drag. Within the notebook interface weBy integrating the component of define ql(x) and calculate its integralpanel lift coefficient that acts parallel (Figure 5). W e incorporate mathto the free stream across the span equations, descriptive text, andthen the induced drag coefficient can images into our calculations tobe found. clearly document our work.Di=ρV∞Гi2kSin(αi) Through integration, we find thatDrag of panel i Lift=πL2ka/4 ---------------(7)D= Drag of the wing--(5) W e determine ka by equating the lift expression that we just calculated with the lift expressed in terms of theThe induced flow angle (αi) aircraft’s load factor. In aircraftrepresents the amount of rotation of design, load factor is the ratio of liftthe lift vector backward and must becalculated from the velocities to total aircraft weight:induced on the bound vortex of the n=Lift/W to ----------------(8)panel by other panels and the free Load factor equals 1 during straightstream. and level flight, and is greater than 1Pitching moment about the wing root when an aircraft is climbing or duringloading edge can be calculated by other maneuvers where lift exceedssumming the panel lift multiplied by a aircraft weight.moment arm which extends in the x- Equating two lift expressions,direction from the loading edge of Wton/2= πL2ka/4 -------------------(9)the wing to the centre of the boundvortex for the panel. and solve for the unknown kaANALYTICAL MODELING OF term. Our analysis assumes that liftAIRCRAFT WING LOADS forces are concentrated on the two wings of the aircraft, which is whyLift the left-hand side of the equation isWe assume an elliptical distribution divided by 2. W e do not consider liftfor lift across the length of the wing, on the fuselage or other surfacesresulting in the following expressionforlift profile(4) ---------(6)where L = length of wing L i ft= ∫ dx-----10) 4
- 5. Plugging ka into our original ql(x)expression, we obtain an expressionfor lift: An analytical model like this helpsus understand how variousparameters affect lift. For example,we see that lift is directly proportionalto load factor (n) and that for a loadfactor of 1, the maximum lift , occursat the wing root (x=0). Weight of Fuel Stored in Wing Weight of Wing StructureWe assume that the load caused by We define the load from the weightthe weight of the wing structure is of the fuel stored in the wing as aproportional to chord length (the piecewise function where load iswidth of the wing), which is highest zero when x > Lf. We assume thatat the wing base (Co) and tapers off this load is proportional to the widthtoward the wing tip (Ct). Therefore, of the fuel tank, which is at itsthe load profile can be expressed as maximum at the base of the wing. W e define qw(x) and integrate it (Cof) and tapers off as we approachacross the length of the wing to the tip of the fuel storage tank (Ctf).calculate the total load from the wing We derive qf(x) in the same way thatstructure: we derived qw(x), resulting in an We then equate this structural load equation of the same form:equation with the structural loadexpressed in terms of load factorand weight of the wing structure(Wws)and solve for kw. Plugging kw into our original qw(x) ---------(12)expression, we obtain an analyticalexpression for load due to weight ofthe wing structure. Figure 5: showing the weight ofFigure 4: showing the weight of fuelwing structure Total Load We calculate total load by adding the three individual load components. This analytical model gives a clear view of how aircraft weight and geometry parameters affect total load. 5
- 6. Results of the theoretical aerodynamic analysis by CFD- XFLR. Table (1) shows the main characteristics of the wings. The aerodynamic results (lift) are used to simulate the wing loading on the MODELING AND ANALYSIS wings during the static stress analysis. Table1: Main characteristics of theAerodynamic model is developed wingsand analyzed by using CFD-XFLR Characteristic Rectanglesoftware and the structural model is winggenerated by CATIA V5 and is Wing span (m) 3.35analysed by using ansys. Gross area (m2) 1.53CFD ANALYSIS Aspect ratio 7.33 Taper ratio 1The wing is designed with single Tip chord (m) 0.45aerofoil profile. The aerofoil number Root chord (m) 0.45is NACA 0012 which is a symmetric M. A. C. (m) 0.45aerofoil. Sweep angle of 0 leading edge (deg.) Section profile NACA 0012 Table 2: Calculated polar through CFD analysis XFLR5_v4.17 polar for: NACA 0012 Mach = 0.2 Re = 1.000 e 6 A OA Cl Cd Cm 1 0.11 0.0051 0.001Fig6 :Cl Vs Alpha Curve At 3 0.32 0.0058 0.0022various Reynolds number 5 0.54 0.0078 0.0054 7 0.82 0.01 0.0084 9 0.992 0.013 -0.0084 11 1.15 0.0177 0.0012 13 1.29 0.023 0.0102 15 1.33 0.035 0.021 17 1.1801 0.086 0.0093 The aerodynamic results are taken for a range of wing incidences Figs.(7) and and only the maximum wing loadings are applied to the wings to serve wing stress conditionsFig 7: Lift Distribution over the wing 6
- 7. load distance (m) N Wto = 108 kg 0.25 73.39 Ww = 14.17kg 0.5 36.352 Wf = 32 kg 0.75 17.885 L = 3.35 m 1 17.471 Lf = 2.3 m 1.5 16.235 Co = 0 .45 m 2 14.349 Ct = 0.45 m 2.5 4.59 Co = 0.275 m 3 -0.594 Ctf = 0.275 m 3.35 -8.763Table 3: Load acting at variousportion of the w ing Distance from leading edge in m α Fig 9 Load distribution over the wingFig 8) Alpha Vs Cl Curves at STRUCTURAL STATIC ANALYSISMach no 0.2 Structural analysis is probably theCALCULATION OF WING most common application of theLOADING finite element method. The static analysis is used to determinePlot of individual load displacements, stresses, etc. undercomponents and total load along static loading conditions. In this workthe length of the wing. the structural static analysis was we see that lift is the largest achieved by using the CATIA V5contributor to total load and that the package in order to obtain stressmaximum load occurs at the end of and displacement distribution in thethe fuel tank. Fuel load also wings by using isotropic andcontributes significantly to the total composite material which are usedload, while the weight of the wing is commonly these days especially forthe smallest contributor. unmanned A/C. The CATIA V5While it is useful to visualize wing program has many finite elementloads, what really concerns us are analysis capabilities, ranging from athe shear force and bending simple, linear, static analysis to amoments resulting from these loads. complex, non linear, transientWe need to determine whether dynamic analysis. A typical CATIAworst- case bending moments ex- V5 analysis has following distinctperienced by the wing are within steps:design limits i) Build the model. ii) Apply loads 7
- 8. iii) Obtain the solution. iv) Review the results.MaterialsHistorically, aluminum materialshave been the primary material foraircraft and space craft construction.Today, structural weight andstiffness requirements haveexceeded th e capability ofconventional aluminum, and high-performance payloads have Fig 10.CATIA WING MODELdemanded extreme thermo-elasticstability in the aircraft design Optimization Procedureenvironment. During the past The wing structural design problemdecades, advanced composite is composed into two levels in amaterials have been increasingly hierarchical structure at the firstaccepted for aircraft and aerospace level, the wing configuration isstructural materials by numerous completely made up of isotropicdevelopments and flight applications. material and the following designComposite materials are those parameters were investigated suchcontaining more than one bonded as number of stiffeners, skinmaterial, each with different thickness, cross sectional area ofstructural properties. stiffener and the type of material.For the sake of simplicity two types The second level wing substructureof materials have been used for the (spars, stiffener) is made of isotropicoptimization procedure.One is material and the skin panel is madeisotropic material that wing of composite material. Various typesconfiguration is completely made up of composite materials are alsoof Aluminum or titanium (7075-T6, considered. Then from the resultsadv. Aluminum, Ti6A 14V). Another we take the optimum design .one is composite material that wingconfiguration is Graphite/Epoxy or S- RESULTS AND DISCUSSIONSglass/Epoxy.The material propertiesare shown in Table (4) The structural analysis was achieved (12) by using the CATIA V5 package in Table 4: Material propertiesIsotropic 7075-T6 Adv Ti6A14V order to obtain stress andmaterial Aluminium displacement distributions in theE(N/m2) 71E+9 82.7E+9 110.32E+9 wings by using isotropic and γ 0.33 0.318 0.29Yield 470E+6 424E+6 598E+6 composite material to find thestress(N/m2) optimum design. The work involves 2800 2906.4 4429 the investigation effects of changingΡ(Kg/m3) the skin shell thickness (0.001- 0.003m), the stiffener beam areaComposite H M Gr/Ep Gr/Ep Sg/Ep (44mm2-81mm2), adding ribs spanmaterial wisely (3-5) adding stringers chordE1(Gpa) 137.9 145 55 wisely (0-5),. The optimum designE2=E3 (Gpa) 14.48 10 16γ 12= γ 12= γ 12 0.21 0.25 0.28 parameter was achieved, where theG12=G23=G13 5.86 4 .8 7 .6 skin thickness (0.001m) for rectangle(Gpa) wing and the stiffener (3×5) . TheΡ(Kg/m3) 1743.84 1580 1593.42 optimum isotropic material is (7075- T6) . Also the optimum cross section. area of stiffener equal to (44 mm2) 8
- 9. Figures(12), (13) show the Von-Table (5) shows the Von-Mises Mises stress distribution in casestress and (Uy) results for rectangle of using isotropic and compositewing when adding the ribs span materials for rectangular wing, inwisely, increasing the number of ribs both cases the root Von-Misesspan wisely gives small reduction variation increases through outeffect on the Von-Mises stress and the (0.2-0.35) chord zone.displacement distribution.Table (5) Variation of vonmisses stress due to variationof no of ribs No. of ribs 3 4 5 properties V.M. stress 129 129 130 (M Pa) UY(mm) 0.022 0.022 0.022 Stress ratio 29.57 29.57 30.00 % Fig11: Effect of von mises stress Mass (Kg) 1.991 2.01 2.11 in changing the compositeYield/weight 8.02 7.86 7.72 material* 106( 1/m2 )Weight/Area 38.27 38.98 39.75 ( N/m2 )Table (6) Result of isotropicAnd composite materials Material Aluminu Graphite/ properties m allay Epoxy 7075 -T6 V.M. stress 103 109 (MPa) Fig:12 Contours of Von misses UY(m) 0.026 0.027 stress(Pa) distribution for Stress ratio 21.91 23.19 isotropic w ing % Mass (Kg) 5.58 2.01Yield/weight 5.96 7.76*106(1/m2 )Weight/Area 51.48 39.56 (N/m2 )Table (6) shows comparison of theresults between the isotropic andcomposite material. The goalcomparison is to emphasize that thewing total mass is reduced and Von-Mises stress increases whencomposite material is used insteadof isotropic material. 9
- 10. Fig:17 Contours of Von misses isotropic and composite materials instress(Pa) distribution for design the parts of wing. Thereforecomposite wing the major and general observations and conclusions for this work can beFigures shows the displacement listed below:distribution of Isotropic and 1. From the Aerodynamic results itComposite wings can be noticed that lift coefficient increases with the increase of Mach number at the same angle of attacks due to the compressibility. 2. From the structural results it is noticed that increasing the skin thickness is considered as an important factor in reducing the stress and displacement levels compared to the increase of, cross section area of beam and number of layers. 3. Increasing the ribs span wisely does not change noticeable the Von-Fig13 : Contours of Displacement Misses stress distribution.distribution of isotropic wing 4. Also it is found that the mass saves (34%) when composite material is used instead of isotropic material wing. 5. It is desirable to use composite materials in design of the skin of a wing References. 1. N. G. R., Iyengar, and S. P., Joshi, “Optimal Design of Antisymmetric Laminated Composite Plates”, Journal of Aircraft, Vol. 23, No. 5, May 1986.Fig :14: Contours of 2. R. A., Confield, R. V., Granhi, andDisplacement distribution of V. B., Venkayya, “Optimum DesignComposite wing of Structures with Multiple Constraints”, AIAA Journal, Vol. 26,5.SUMMARY AND No. 1, January 1988, p. 78.CONCLUSSION 3. An example of preliminaryThe static analysis of the wing by design of Jet plane E .GCATIA V5 package has been .Tulaburkara,A.Venkatraman,presented and used to determine the V.Ganesh 2007optimum design for the wing. This 4. Dan Doharty, Analyticaloperation involves using the vortex modeling of aircraft wing usinglattice method to obtain the matlab, Matlab digestaerodynamic result at (M=0.2) for 5.Thomas Chamberlain, Methods ofrectangle wing. This aerodynamic calculating aerodynamic load onresult (lift) is used to simulate the aircraft structureswing loading on the wings during the 6Tobias F.W underlich .”static analysis and also using Multidisciplinary w ing design and 10
- 11. optimization for transport aircraft” 10” Design analysis of UAV usingDRL Institute of aeronautics and flow NACA 0012 Airfoil’ Alimul Rajib1,Technology Sep 2008. Bhuiyan Shameem Mahmud Ebna7. Shawn E.Gano,John E.Renaud ” Hai1 and Md Abdus Salam2Optimized unmanned aerial 1Department of Mechanicalvehicle with wing Morphing”, , Engineering, Military Institute ofAIAA Journal 2002“. Science and Technology, Dhaka-8.”Aerodynamics of 3-D lifting 1216, Bangladesh..surfaces using vortex lattice 11.A.Kao Mechanics of compositemethod” NASA Journal Nov 1998. materials .9. J. J., Bertin, and M. S., Smith, 12. P. J., Rol, D. N., Marris, and D.“Aerdynamics for Engineering P.,Schrage,“Combined AerodynamicStudents”, Department of and Structural Optimization of aAeronautical Engineering, University High-Speed Civil Transport Wing”,of Sydney, 2000. School of Aerospace Engineering, American Institute of Aeronautic and Astronautic, Inc., 1995. 11

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