2D NACA 0012 Airfoil with Xfoil

Computational Fluid Dynamics (CFD), Xfoil

Description

Calculate the lift and drag for NACA 0012 airfoil at given Reynolds and Mach number.

Reference

  1. 2DN00: 2D NACA 0012 Airfoil Validation Case

Analysis type

Isotherm, incompressible

Tools used

  • Xfoil

Geometry

Figure 1: NACA 0012
Figure 1: NACA 0012

Boundary conditions

Isotherm, incompressible flow

  • Re = 6 * 106
  • Mach = 0.15

Fluid properties

Dry air at standard seal level:

  • ρ0 = 1.225 kg/m³
  • η0 = 1.80 * 10-5 kg/m/s

Computation target

The quantities of interest for comparison are as follows:

  • Lift coefficient (CL) vs. angle of attack (alpha) (for 0 < alpha < amax)
  • Drag coefficient (CD) vs. CL for 0 < alpha < amax)
  • Surface pressure coefficient (Cp) vs. x/c (for alpha = 0, 10, 15)
  • Surface skin friction coefficient (Cf) vs. x/c (for alpha = 0, 10, 15

where amax is the maximum angle of attack at which the CFD yields steady-state results with no force oscillations.

Results

Lift coefficient (CL) vs. angle of attack (alpha)
Figure 2: Lift coefficient (CL) vs. angle of attack (alpha)
Lift coefficient (CL) vs. Drag coefficient (CD)
Figure 3: Lift coefficient (CL) vs. Drag coefficient (CD)
Figure 3: Surface pressure coefficient (Cp) vs. x/c for alpha = 0
Figure 4: Surface pressure coefficient (Cp) vs. x/c for alpha = 0
Figure 4: Surface pressure coefficient (Cp) vs. x/c for alpha = 10
Figure 5: Surface pressure coefficient (Cp) vs. x/c for alpha = 10
Figure 5: Surface pressure coefficient (Cp) vs. x/c for alpha = 15
Figure 6: Surface pressure coefficient (Cp) vs. x/c for alpha = 15

Conclusion

Xfoil results are comparable with experimental results.

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