# --- # jupyter: # jupytext: # default_lexer: ipython3 # text_representation: # extension: .py # format_name: percent # format_version: '1.3' # jupytext_version: 1.19.4 # kernelspec: # display_name: Python 3 (ipykernel) # language: python # name: python3 # --- # %% [markdown] # # 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 # %% [markdown] # ## 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" # %% [markdown] # ### 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 # %% [markdown] # ### 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() # %% [markdown] # ### 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() # %% [markdown] # ### 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}") # %% [markdown] # ### Close the modeler # %% modeler.close() # %% [markdown] # ## 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" # %% [markdown] # ### 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() # %% [markdown] # ### 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() # %% [markdown] # ### 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}") # %% [markdown] # ### Close the modeler # %% modeler.close()