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Applied: Create a NACA 4-digit airfoil

NACA airfoils are a series of airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA). They are a standardized system of airfoil shapes that are defined by a series of digits. The digits, which indicate the shape of the airfoil, are used to create the airfoil shape.

Each digit in the NACA airfoil number has a specific meaning:

• The first digit defines the maximum camber as a percentage of the chord length.

• The second digit defines the position of the maximum camber as a percentage of the chord length.

• The last two digits define the maximum thickness of the airfoil as a percentage of the chord length.

To fully understand the previous definitions, it is important to know that the chord length is the length of the airfoil from the leading edge to the trailing edge. The camber is the curvature of the airfoil, and the thickness is the distance between the upper and lower surfaces.

Symmetric airfoils have a camber of 0% and consequently, the first two digits of the NACA number are 0. For example, the NACA 0012 airfoil is a symmetric airfoil with a maximum thickness of 12%.

Define the NACA 4-digit airfoil equation

The following code uses the equation for a NACA 4-digit airfoil to create a set of points that define the airfoil shape. These points are Point2D objects in PyAnsys Geometry.

from typing import List, Union

import numpy as np

from ansys.geometry.core.math import Point2D

def naca_airfoil_4digits(number: Union[int, str], n_points: int = 200) -> List[Point2D]:
    """
    Generate a NACA 4-digits airfoil.

    Parameters
    ----------
    number : int or str
        NACA 4-digit number.
    n_points : int
        Number of points to generate the airfoil. The default is ``200``.
        Number of points in the upper side of the airfoil.
        The total number of points is ``2 * n_points - 1``.

    Returns
    -------
    List[Point2D]
        List of points that define the airfoil.
    """
    # Check if the number is a string
    if isinstance(number, str):
        number = int(number)

    # Calculate the NACA parameters
    m = number // 1000 * 0.01
    p = number // 100 % 10 * 0.1
    t = number % 100 * 0.01

    # Generate the airfoil
    points = []
    for i in range(n_points):

        # Make it a exponential distribution so the points are more concentrated
        # near the leading edge
        x = (1 - np.cos(i / (n_points - 1) * np.pi)) / 2

        # Check if it is a symmetric airfoil
        if p == 0 and m == 0:
            # Camber line is zero in this case
            yc = 0
            dyc_dx = 0
        else:
            # Compute the camber line
            if x < p:
                yc = m / p**2 * (2 * p * x - x**2)
                dyc_dx = 2 * m / p**2 * (p - x)
            else:
                yc = m / (1 - p) ** 2 * ((1 - 2 * p) + 2 * p * x - x**2)
                dyc_dx = 2 * m / (1 - p) ** 2 * (p - x)

        # Compute the thickness
        yt = 5 * t * (0.2969 * x**0.5
                      - 0.1260 * x
                      - 0.3516 * x**2
                      + 0.2843 * x**3
                      - 0.1015 * x**4)

        # Compute the angle
        theta = np.arctan(dyc_dx)

        # Compute the points (upper and lower side of the airfoil)
        xu = x - yt * np.sin(theta)
        yu = yc + yt * np.cos(theta)
        xl = x + yt * np.sin(theta)
        yl = yc - yt * np.cos(theta)

        # Append the points
        points.append(Point2D([xu, yu]))
        points.insert(0, Point2D([xl, yl]))

        # Remove the first point since it is repeated
        if i == 0:
            points.pop(0)

    return points

Example of a symmetric airfoil: NACA 0012

Now that the function for generating a NACA 4-digit airfoil is generated, this code creates a NACA 0012 airfoil, which is symmetric. This airfoil has a maximum thickness of 12%. The NACA number is a constant.

NACA_AIRFOIL = "0012"

Required imports

Before you start creating the airfoil points, you must import the necessary modules to create the airfoil using PyAnsys Geometry.

from ansys.geometry.core import launch_modeler
from ansys.geometry.core.sketch import Sketch

Generate the airfoil points

Using the function defined previously, you generate the points that define the NACA 0012 airfoil. Create a Sketch object and add the points to it. Then, approximate the airfoil using straight lines between the points.

# Create a sketch
sketch = Sketch()

# Generate the points of the airfoil
points = naca_airfoil_4digits(NACA_AIRFOIL)

# Create the segments of the airfoil
for i in range(len(points) - 1):
    sketch.segment(points[i], points[i + 1])

# Close the airfoil
sketch.segment(points[-1], points[0])

# Plot the airfoil
sketch.plot()

Create the 3D airfoil

Once the Sketch object is created, you create a 3D airfoil. For this operation, you must create a modeler object, create a design object, and extrude the sketch.

# Launch the modeler
modeler = launch_modeler()

# Create the design
design = modeler.create_design(f"NACA_Airfoil_{NACA_AIRFOIL}")

# Extrude the airfoil
design.extrude_sketch("Airfoil", sketch, 1)

# Plot the design
design.plot()

Save the design

Finally, save the design to a file. This file can be used in other applications or imported into a simulation software. This code saves the design as an FMD file, which can then be imported into Ansys Fluent.

# Save the design
file = design.export_to_fmd()
print(f"Design saved to {file}")

Design saved to C:\Users\ansys\actions-runner\_work\pyansys-geometry\pyansys-geometry\doc\source\examples\04_applied\NACA_Airfoil_0012.fmd

Close the modeler

modeler.close()

Example of a cambered airfoil: NACA 6412

This code creates a NACA 6412 airfoil, which is cambered. This airfoil has a maximum camber of 6% and a maximum thickness of 12%. After overriding the NACA number, the code generates the airfoil points.

NACA_AIRFOIL = "6412"

Generate the airfoil points

As before, you generate the points that define the NACA 6412 airfoil. Create a Sketch object and add the points to it. Then, approximate the airfoil using straight lines.

# Create a sketch
sketch = Sketch()

# Generate the points of the airfoil
points = naca_airfoil_4digits(NACA_AIRFOIL)

# Create the segments of the airfoil
for i in range(len(points) - 1):
    sketch.segment(points[i], points[i + 1])

# Close the airfoil
sketch.segment(points[-1], points[0])

# Plot the airfoil
sketch.plot()

Create the 3D airfoil

# Launch the modeler
modeler = launch_modeler()

# Create the design
design = modeler.create_design(f"NACA_Airfoil_{NACA_AIRFOIL}")

# Extrude the airfoil
design.extrude_sketch("Airfoil", sketch, 1)

# Plot the design
design.plot()

Save the design

In this case, the design is saved as an SCDOCX file.

# Save the design
file = design.export_to_scdocx()
print(f"Design saved to {file}")

Design saved to C:\Users\ansys\actions-runner\_work\pyansys-geometry\pyansys-geometry\doc\source\examples\04_applied\NACA_Airfoil_6412.scdocx

Close the modeler

modeler.close()

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