Ansys ICEM CFD – Import Points NACA Airfoil 4412

Written by cfd.ninja

March 10, 2020

Bookstore

Source: Wikipedia

The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA). The shape of the NACA airfoils is described using a series of digits following the word “NACA”. The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties.

The NACA four-digit wing sections define the profile by:[1]

  1. First digit describing maximum camber as percentage of the chord.
  2. Second digit describing the distance of maximum camber from the airfoil leading edge in tenths of the chord.
  3. Last two digits describing maximum thickness of the airfoil as percent of the chord.[2]

For example, the NACA 2412 airfoil has a maximum camber of 2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the chord.

The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long.

Equation for a symmetrical 4-digit NACA airfoil[edit]

Plot of a NACA 0015 foil generated from formula

The formula for the shape of a NACA 00xx foil, with “x” being replaced by the percentage of thickness to chord, is[3]

{\displaystyle y_{t}=5t\left[0.2969{\sqrt {x}}-0.1260x-0.3516x^{2}+0.2843x^{3}-0.1015x^{4}\right],}

[4][5]

where:

x is the position along the chord from 0 to 1.00 (0 to 100%),

{\displaystyle y_{t}}

 is the half thickness at a given value of x (centerline to surface),

t is the maximum thickness as a fraction of the chord (so t gives the last two digits in the NACA 4-digit denomination divided by 100).

Note that in this equation, at x/c = 1 (the trailing edge of the airfoil), the thickness is not quite zero. If a zero-thickness trailing edge is required, for example for computational work, one of the coefficients should be modified such that they sum to zero. Modifying the last coefficient (i.e. to −0.1036) will result in the smallest change to the overall shape of the airfoil. The leading edge approximates a cylinder with a radius of

{\displaystyle r=1.1019{\frac {t^{2}}{c}}.}

[6]

Now the coordinates 

{\displaystyle (x_{U},y_{U})}

 of the upper airfoil surface and 

{\displaystyle (x_{L},y_{L})}

 of the lower airfoil surface are

{\displaystyle x_{U}=x_{L}=x,\quad y_{U}=+y_{t},\quad y_{L}=-y_{t}.}

Symmetrical 4-digit series airfoils by default have maximum thickness at 30% of the chord from the leading edge.

Equation for a cambered 4-digit NACA airfoil[edit]

Plot of a NACA 2412 foil. The camber line is shown in red, and the thickness – or the symmetrical airfoil 0012 – is shown in purple.

The simplest asymmetric foils are the NACA 4-digit series foils, which use the same formula as that used to generate the 00xx symmetric foils, but with the line of mean camber bent. The formula used to calculate the mean camber line is[3]

{\displaystyle y_{c}={\begin{cases}{\dfrac {m}{p^{2}}}\left(2p\left({\dfrac {x}{c}}\right)-\left({\dfrac {x}{c}}\right)^{2}\right),&0\leq x\leq pc,\\{\dfrac {m}{(1-p)^{2}}}\left((1-2p)+2p\left({\dfrac {x}{c}}\right)-\left({\dfrac {x}{c}}\right)^{2}\right),&pc\leq x\leq c,\end{cases}}}

where

m is the maximum camber (100 m is the first of the four digits),
p is the location of maximum camber (10 p is the second digit in the NACA xxxx description).

For this cambered airfoil, because the thickness needs to be applied perpendicular to the camber line, the coordinates 

{\displaystyle (x_{U},y_{U})}

 and 

{\displaystyle (x_{L},y_{L})}

, of respectively the upper and lower airfoil surface, become[7]

{\displaystyle {\begin{aligned}x_{U}&=x-y_{t}\,\sin \theta ,&y_{U}&=y_{c}+y_{t}\,\cos \theta ,\\x_{L}&=x+y_{t}\,\sin \theta ,&y_{L}&=y_{c}-y_{t}\,\cos \theta ,\end{aligned}}}

where

{\displaystyle \theta =\arctan {\frac {dy_{c}}{dx}},}

{\displaystyle {\frac {dy_{c}}{dx}}={\begin{cases}{\dfrac {2m}{p^{2}}}\left(p-{\dfrac {x}{c}}\right),&0\leq x\leq pc,\\{\dfrac {2m}{(1-p)^{2}}}\left(p-{\dfrac {x}{c}}\right),&pc\leq x\leq c.\end{cases}}}

In this first tutorial, you will learn how to import points (NACA Airfoil 4412) from Excel to Ansys ICEM CFD.

ICEM is a powerful tool for generating meshing.

Do you want to learn Ansys Meshing?

Related Articles

Ansys Meshing – Mesh Copy Control

Ansys Meshing – Mesh Copy Control

The Mesh Copy control enables you copy mesh from one body to another. This option can be used to reduce the mesh setup time for repetitive bodies/parts. Association to CAD is maintained after performing mesh copy.

Mesh controls are scoped only to the source anchor body. When the mesh is generated, the source anchor body is meshed and the mesh is then copied to targets.

Ansys Meshing – Pinch

Ansys Meshing – Pinch

The Pinch feature lets you remove small features (such as short edges and narrow regions) at the mesh level in order to generate better quality elements around those features.

Ansys Meshing – Mesh Types (Hexa, Prism, Polyhedral)

Ansys Meshing – Mesh Types (Hexa, Prism, Polyhedral)

When geometries are complex or the range of length scales of the flow is large, a triangular/tetrahedral mesh can be created with far fewer cells than the equivalent mesh consisting of quadrilateral/hexahedral elements.

Ansys Meshing – Hexahedral Mesh (Pipe)

Ansys Meshing – Hexahedral Mesh (Pipe)

The MultiZone mesh method provides automatic decomposition of geometry into mapped (structured/sweepable) regions and free (unstructured) regions.

Stay Up to Date With The Latest News & Updates

Help us keep growing

CFD.NINJA is financed with its own resources, if you want to support us we will be grateful.

Join Our Newsletter

Subscribe to receive emails with detailed information related to the CFD.

Follow Us

Subscribe to our social networks to receive notifications about our new tutorials

Pin It on Pinterest

Shares

Share This

Share this post with your friends!