NACA0012 2D Airfoil

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Introduction

This is an example of steady-state incompressible flow analysis of a two-dimensional subsonic airfoil. The NACA0012 airfoil flow validation problem uses the conditions from the website below.

https://turbmodels.larc.nasa.gov/naca0012_val.html

Use the mesh created in the BaramMesh tutorial.

The flow conditions are as follows

  • solver : buoyantSimpleNFoam
  • turbulence model : $SST$ $k-\omega$
  • velocity : 51.4815 m/s
  • Mach No. : 0.15
  • Reynolds No. : 6E+06
  • Angle of Attack : 15 degree
  • farfield pressure : 0 Pa

Start BaramFlow

Run the program and select [New Case] from the launcher. In the launcher, select [Pressure-based] for [Solver Type] and [None] for [Multiphase Model].

Use the given polyMesh folder. In the top tab, click [File]-[Load Mesh]-[OpenFOAM] in that order and select the polyMesh folder.

General

For this example, we’ll use default values for all.

Models

For this example, we’ll use SST k-$\omega$ model for turbulence.

Materials

The material properties of the fluid are set as follows from the velocity and Reynolds number conditions.

  • density : 1
  • viscosity : 8.58E-06

Boundary Conditions

Set the boundary conditions as follows

  • Hex6_1_xMin, Hex6_1_xMax, Hex6_1_yMin, Hex6_1_yMax : Free Stream
    • Specification Method : AOA and AOS
    • Drag direction : (1 0 0)
    • Lift direction : (0 1 0)
    • Angle of Attack : 15
    • Speed : 51.4815
    • Turbulence specification method : Intensity and Viscosity Ratio
    • Turbulence intensity : 1
    • Turbulence viscosity ratio : 1
    • It is convenient to set only one boundary and use the [Copy] function at the bottom
  • Hex6_1_zMin : Empty
  • wing_surface : Wall(No Slip)

Reference Values

  • Area, Length : 1
  • Density : 1
  • Pressure : 0
  • Velocity : 51.4815

Numerical Conditions

  • Pressure-Velocity Coupling Scheme : SIMPLE
  • Discretization Schemes : Second Order Upwind for flow and turbulence
  • Convergence Criteria : 0 for pressure

Monitor

Select [Add]-[Forces] and set as follows

  • Flow Direction
    • Specification Method : AOA and AOA
    • Direction for AOA=0, AOS=0 : (1 0 0) for drag, (0 1 0) for lift
    • Angle of Attack : 15
    • Sideslip Angle : 0
  • Center of Rotation : (0 0 0)
  • Boundaries : wing_surface

Initialization

For the initial condition, select Hex6_1_xMin for [Compute from] and the corresponding boundary condition value will be set automatically.

Click the [Initialize] button at the bottom. After that, click the [File] – [Save] button on the menu to save.

Run

Set [Run Conditions] as follows and click [Start Calculation] button, then simulation starts.

  • Number of Iterations : 1000
  • Save Interval : 1000

When the calculation starts, a residual graph is plotted as shown below.

Post-processing

Click the [External tools]-[ParaView] button in the menu to launch ParaView.

Select U and you will see the following distribution.