# NACA0012 airfoil

### From CFD-Wiki

(→Lift Curve) |
|||

Line 6: | Line 6: | ||

[[Image:Naca0012 mesh final.JPG]] | [[Image:Naca0012 mesh final.JPG]] | ||

<br><br> | <br><br> | ||

- | The mesh is a 30,000 cell C-grid. The width off the first cell at the foil boundary is 0.02 mm. At Re = 3e6 this results in a wall y+ = 1.3 ± 0.4 . The mesh shown is for an | + | The mesh is a 30,000 cell C-grid. The width off the first cell at the foil boundary is 0.02 mm. At Re = 3e6 and zero angle of attack, this results in a wall y+ = 1.3 ± 0.4, which is low enough for the turbulence model to resolve the sub layer. The mesh shown is for an angle of attack of 6 degrees. |

<br> | <br> | ||

== Drag Coefficient == | == Drag Coefficient == | ||

Line 12: | Line 12: | ||

[[Image:Naca0012 cd tripwire.JPG]] | [[Image:Naca0012 cd tripwire.JPG]] | ||

<br><br> | <br><br> | ||

- | The drag coefficient at zero Angle of Attack depends on the Reynold's number. The experimental data is for an airfoil with a trip wire | + | The drag coefficient at zero Angle of Attack depends on the Reynold's number. The experimental data is for an airfoil with a trip wire, which forces the experimental boundary layer to be completely turbulent.[1] This corresponds to the Fluent model, which has an active turbulence model over the complete airfoil. Note that the calculated drag coefficient is somewhat higher than the experimental one. Possibly, the modeled boundary layer is turbulent from the beginning, while in reality the trip wire is not at the very leading edge of the foil. |

+ | |||

+ | <br> | ||

+ | Modeling the NACA 0012 airfoil without a trip wire is more complicated, since Fluent itself is unable to predict the point along the chord where the transition from a laminar to a turbulent boundary layer takes place. One option is to manually set this transition point at Re = 5.5e6. | ||

<br> | <br> | ||

== Lift Curve == | == Lift Curve == | ||

Line 18: | Line 21: | ||

[[Image:Naca0012 lift curve2.JPG]] | [[Image:Naca0012 lift curve2.JPG]] | ||

<br><br> | <br><br> | ||

- | The lift coefficient depends on the | + | The lift coefficient depends on the angle of attack. For Re = 2e6 I compare the lift coefficient to experimental results.[2] |

<br> | <br> | ||

## Revision as of 18:55, 10 November 2009

## Contents |

## Introduction

The NACA 0012 airfoil is widely used. The simple geometry and the large amount of wind tunnel data provide an excellent 2D validation case. For this case I use the Spalart-Allmaras turbulence model.

## Mesh

The mesh is a 30,000 cell C-grid. The width off the first cell at the foil boundary is 0.02 mm. At Re = 3e6 and zero angle of attack, this results in a wall y+ = 1.3 ± 0.4, which is low enough for the turbulence model to resolve the sub layer. The mesh shown is for an angle of attack of 6 degrees.

## Drag Coefficient

The drag coefficient at zero Angle of Attack depends on the Reynold's number. The experimental data is for an airfoil with a trip wire, which forces the experimental boundary layer to be completely turbulent.[1] This corresponds to the Fluent model, which has an active turbulence model over the complete airfoil. Note that the calculated drag coefficient is somewhat higher than the experimental one. Possibly, the modeled boundary layer is turbulent from the beginning, while in reality the trip wire is not at the very leading edge of the foil.

Modeling the NACA 0012 airfoil without a trip wire is more complicated, since Fluent itself is unable to predict the point along the chord where the transition from a laminar to a turbulent boundary layer takes place. One option is to manually set this transition point at Re = 5.5e6.

## Lift Curve

The lift coefficient depends on the angle of attack. For Re = 2e6 I compare the lift coefficient to experimental results.[2]

## Lift Curve Slope

The initial slope of the lift curve depends on the Reynold's number. Here I compare the lift curve slope to experimental results.[1]

## References

1. W. J. McCroskey, *A Critical Assessment of Wind Tunnel Results for the NACA 0012 Airfoil*, NASA Technical Memorandum 10001 9 (1987)

2. L. Lazauskus, NACA 0012 Lift Data