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External Aerodynamics: Parametric modeling of of an airfoil using blockMeshDict in OpenFOAM

The construction of OpenFOAM cases for flow over an aerofoil can be done in many ways. For example, the tutorial section demonstrates two methods. The first method uses simpleFoam to demonstrate steady state and incompressible flow over the aerofoil. This method focuses on the overall methodology and a mesh generated in other application has been used. The second method as described in section compressible / sonicFoam / ras which uses a mesh created in STAR and utility star3ToFoam to convert the mesh to the OpenFOAM format. Yet another method using the data points of the airfoil and GMSH is explained in *this tutorial.*

This page demonstrates a method which can be used to generate mesh for flow over an aerofoil when the points defining the cross-section of the aerofoil is known. The aerofoil shape is created using arc and spline utility of blockMesh. The entire domain will look like as shown below.

Using the boundary conditions defined in tutorial case incompressible / simpleFoam / airFoil2D, the pressure contour generated out simulation is as per screeshot shown below.

The blockmeshDict file is here. The complete case data is here.

Check the effect of parameters on shape of the NACA airfoil:

- R = Resultant force acting on the wing
- N = Component of R acting perpendicular to the chord
- A = Component of R acting parallel to the chord
- Lift L = Component of R acting perpendicular to the relative wind
- Drag D = Component of R acting parallel to the relative wind direction
- Angle of Attack, AoA: α = angle between the chord line and relative wind direction.
- L = N × cosα - A × sinα, D = N × sinα + A × cosα
- A
_{WING}(also designated as 'S' in many textbooks), Wing (projected) area: area of the wing responsible for lift - Freestream density: ρ
_{∞}= density of air flow over wings, Freestream velocity: V_{∞}= velocity of air flow over wings - C
_{L}= Lift Coefficient, C_{D}= Drag coefficient - L = C
_{L}× ½ρ_{∞}V_{∞}^{2}A_{WING}, D = C_{D}× ½ρ_{∞}V_{∞}^{2}A_{WING} - At level flight: lift = weigth of aircraft (W), drag = thrust of the engine
- Wing loading, q = W / A
_{WING} - Drag coefficient = [C
_{D_i}] induced drag (drag caused by shape responsible for generation of lift, includes that due to trailing edge vortex downstream the lifting surface) + [C_{D_o}] zero-lift drag (drag due to shapes not contributing to generation of lift such as fuselage, cockpit...) - Drag polar equation: C
_{D}= C_{D_o}+ K.C_{L}^{2}where K = induced drag correcton factor and lift-dependent drag coefficient C_{D_i}= K.C_{L}^{2} - Steady level flight: Thrust (T) = Drag (D), Weight (W) = Lift (L). Thus: W = L = L/D × D = L/D × T
- Similarly, D = L × D/L = W × D/L = W × C
_{D}/C_{L}. Thus, drag D would be minimum at maximum value of the ratio of C_{D}/C_{L} - m: chordwise location for maximum ordinate of airfoil on camber line
- p: maximum ordinate of 2-digit camber line
- ζ = x/c, M = m/c, P = p/c, a NACA airfoil can be either of the form y/c = a
_{0}ζ^{1/2}+ a_{1}ζ + a_{2}ζ^{2}+ a_{3}ζ^{3}+ a_{4}ζ^{4}or y/c = M/P^{2}(2Pζ - ζ^{2}) for 0≤ζ≤P, y/c = M/(1-P)^{2}(1-2P+2Pζ-ζ^{2}) for P≤ζ≤1. - C
_{MLE}: Pitching moment about leading edge - C
_{M1/4}: Pitching moment about quarter-chord point

The following integration formula will be helpful in calculation of lift coefficient of airfoils.

- Pitching moment is positive when it tends to raise the nose of the aircraft upward. If M
_{a}= pitching moment about a known distance 'a' from the leading edge 'LE' then: M_{LE}= M_{a}- L × a × cosα - D × a × sinα ...[I] Similarly, pitching moment about a point away from LE by distance 'x': M_{LE}= M_{x}- L · x · cosα - D · x · sinα ...[II] - Subtracting equation [I] from [II]: M
_{a}= M_{x}+ (L × cosα + D × sinα) (a - x) - C
_{Mx}= C_{Ma}- (C_{L}× cosα + C_{D}× sinα) (a - x)/c ... [III] - C
_{MLE}= C_{Ma}- (C_{L}× cosα + C_{D}× sinα) (a - 0)/c ... [IV] - [III] - [IV]: C
_{Mx}= C_{MLE}+ (C_{L}× cosα + C_{D}× sinα) x/c

C_{MAC} = C_{Ma} + (C_{L} × cosα + C_{D} × sinα) (x_{AC} - a)/c

C_{MAC} = C_{Ma} + C_{L} × cosα (x_{AC} - a)/c

Problem Types:

- Type-a: Flow is isentropic thoughout, given Mach number at exit and area ratio calculate mass flow rate, temperature...
- Type-b: Given area ratio and operating pressure ratio, find out location of shock.
- Type-c: Area ratio and Mach number at exit is given, calculate operating pressure ratio.
- Type-d: Given location of shock and Ae/A*, find operating pressure ratio.
- Type-e: Design exit Mach number and reservoir pressure specified, calculate critical pressure ratios.
- Type-f: Given inlet conditions, exit//throat area ratio and Mach number before shock, find exit pressure, mass flow rate.

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