Induced Drag Reduction of an Airplane Using Local Twist, Panel Method Verification

Abstract

The objective of this work is to modify the wing of the Unmanned Airplane of Ecological Conservation adding wing twist in order to reduce induced drag at a design lift coefficient using a method based on the lifting-line theory, and test such modification using two different three-dimensional panel method codes. The aircraft is a small airplane designed to make daily patrols over petroleum extraction zones. A linear washout distribution between the boom and the wing tip was chosen for the ANCE wing. The optimum twist angle at a design lift coefficient was calculated, resulting in 5.4119 deg. The resultant induced drag reduction for the wing at design lift coefficient is equal to 6.425%. The panel method codes PAN AIR and CMARC were used to study the inviscid aerodynamic flowfield around the aircraft. The results show that the addition of wing twist increases the lift-drag ratio 2.03% and 1.53% respect to the original, when it is computed using PAN AIR and CMARC, respectively, assuming that the wing twist does not change the viscous drag. The lift coefficient at zero angle of attack was reduced by wing twist addition and the lift slope is independent of twist. Nomenclature a m = planform contribution to the coefficients in the infinite series solution to the lifting-line equation b = wingspan b m = washout contribution to the coefficients in the infinite series solution to the lifting-line equation= lift coefficient at zero angle of attack C Lα = lift slope C lα = airfoil section lift slope c = mean aerodynamic chord c c = airfoil section chord e = Oswald efficiency factor e n = two-dimensional free transition criterion k = lift dependent drag factor or induced drag factor, k = (π·e·R A ) -1 (L/D) max = maximum lift-drag ratio n = fixed parameter for two-dimensional free transition criterion R A = wing aspect ratio y to = distance between wing root and the twist start line y, z = spanwise and height coordinates α = angle of attack α 0 = airfoil section zero-lift angle of attack ε Ω = 2 θ = change of variables for the spanwise coordinate κ D = planform contribution to the induced drag factor κ DL = lift-washout contribution to the induced drag factor κ Do = optimum induced drag factor κ DΩ = washout contribution to the induced drag factor Ω = twist angle Ω opt = optimum twist angle ω = normalized washout distribution function