Welcome to AeroPy’s, an easy to use aerodynamic tool

AeroPy is an library for calculating aerodynamic properties. The main feature of this library is the Python interface with XFOIL. The main objective of this library is to use XFOIL via Python iteratively in a total of 4 lines total (one line for most uses). Through this interface coupling with other softwares (Abaqus, Ansys, etc) is possible and iterative processes (optimization, design sensitivity) are also possible. For a thorough explanation please check the documentation and the tutorials.

Aerodynamic tools

• xfoil_module: contains all the functions relating to XFOIL.
• aero_module: contains all functions related to aerodynamics, but not related to XFOIL.
• filehandling: constains functions to read and output in several formats

Geometric tools

• geometry: tools to generate and modify airfoils
• 2D_CST: contains libraries to develop 2D shapes using the Class/Shape Transformation equations
• 3D_CST: contains libraries to develop 2D shapes using the Class/Shape Transformation equations
• morphing: functions to generate structurally consistent morphed configurations

Contents:

• 1. Xfoil Module Documentation
• 2. Aerodynamic Module Documentation
• 3. Airfoil Module Documentation
• 4. Tutorial

To Do

• Include asymmetric wing
• Create airfoil and wing classes

Recommended Collaborations

Please use and adapt this library to your needs. There are several functionalities I wished to implement, but did not have the time. Hence I am strongly recommending the following collaborations:

• Airfoil generator with a GUI
• Atmospheric module (use the library already available in aero_module)
• Extend aero_module for wings with non-constant cross sections

Indices and tables

• Index
• Module Index
• Search Page

Examples

A simple example demonstrating how to use AeroPy to calculate lift, drag and moment coefficients for a NACA0012 airfoil at angle of attack of 12 degrees.

from aeropy.xfoil_module import find_coefficients
find_coefficients(airfoil='naca0012'alpha=12.)
>>> {'CM': 0.0134, 'CL': 1.2453, 'Top_Xtr': 0.0194, 'CD': 0.01934, 'CDp': 0.01379, 'alpha': 12.0, 'Bot_Xtr': 1.0}

Another example showing how to use AeroPy to calculate pressure coefficients for a NACA0012 in viscous flow at Reynolds = 10000000 and angle of attack of 12 degrees

from aeropy.xfoil_module import find_pressure_coefficients
find_pressure_coefficients(airfoil='naca0012', Reynolds = 1e6, alpha=12.,NACA=True)
>>> 'y': [0.00126, ..., -0.00126], 'x': [1.0, ..., 1.0], 'Cp': [0.41501, ..., 0.41501]}