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# This file is part of Marius Peter's airfoil analysis package (this program).
#
# This program is free software: you can redistribute it and/or modify
# it under the terms of the GNU General Public License as published by
# the Free Software Foundation, either version 3 of the License, or
# (at your option) any later version.
#
# This program is distributed in the hope that it will be useful,
# but WITHOUT ANY WARRANTY; without even the implied warranty of
# MERCHANTABILITY or FITNESS FOR A PARTICULAR PURPOSE. See the
# GNU General Public License for more details.
#
# You should have received a copy of the GNU General Public License
# along with this program. If not, see <https://www.gnu.org/licenses/>.
import sys
import os.path
import numpy as np
from math import sin, cos, tan, atan, sqrt, ceil
import bisect as bi
import matplotlib.pyplot as plt
import matplotlib as mpl
from mpl_toolkits.mplot3d import Axes3D
# This variable is required for main.py constant wing dimensions
# to be passed to inheriting classes (Airfoil, Spar, Stringer, Rib).
# This way, we don't have to redeclare our coordinates as parameters for
# our spars, stringers and ribs. This makes for more elegant code.
global parent
class Coordinates:
"""
All airfoil components need the following:
Parameters:
* Component material
* Coordinates relative to the chord & semi-span.
Methods:
* Print component coordinates
* Save component coordinates to file specified in main.py
So, all component classes inherit from class Coordinates.
"""
def __init__(self, chord, semi_span):
# Global dimensions
self.chord = chord
if chord < 10:
self.chord = 10
self.semi_span = semi_span
# Component material
self.material = str()
# Upper coordinates
self.x_u = []
self.y_u = []
# Lower coordinates
self.x_l = []
self.y_l = []
# Coordinates x_u, y_u, x_l, y_l packed in single list
self.coord = []
# The airfoil components know the Coordinates instance's coords
global parent
parent = self
def __str__(self):
return type(self).__name__
def print_coord(self, round):
"""
Print all the component's coordinates to the terminal.
This function's output is piped to the 'save_coord' function below.
"""
print('============================')
print('Component:', str(self))
print('Chord length:', self.chord)
print('Semi-span:', self.semi_span)
print('============================')
print('x_u the upper x-coordinates:\n', np.around(self.x_u, round))
print('y_u the upper y-coordinates:\n', np.around(self.y_u, round))
print('x_l the lower x-coordinates:\n', np.around(self.x_l, round))
print('y_l the lower y-coordinates:\n', np.around(self.y_l, round))
# print('\n')
return None
def save_coord(self, save_dir_path):
"""
Save all the object's coordinates (must be full path).
"""
file_name = str(self)
full_path = os.path.join(save_dir_path, file_name + '.txt')
file = open(full_path, 'w')
sys.stdout = file
self.print_coord(4)
return None
def pack_coord(self):
self.coord.append(self.x_u)
self.coord.append(self.y_u)
self.coord.append(self.x_l)
self.coord.append(self.y_l)
class Airfoil(Coordinates):
"""This class enables the creation of a single NACA airfoil."""
def __init__(self):
global parent
# Run 'Coordinates' super class init method with same chord & 1/2 span.
super().__init__(parent.chord, parent.semi_span)
# NACA number
self.naca_num = int()
# Mean camber line
self.x_c = [] # Contains only integers from 0 to self.chord
self.y_c = [] # Contains floats
# Thickness
self.y_t = []
# dy_c / d_x
self.dy_c = []
# Theta
self.theta = []
def add_naca(self, naca_num):
"""
This function generates geometry for our chosen NACA airfoil shape.
The nested functions perform the required steps to generate geometry,
and can be called to solve the geometry y-coordinate for any 'x' input.
Equation coefficients were retrieved from Wikipedia.org.
Parameters:
naca_num: 4-digit NACA wing
Return:
None
"""
# Variables extracted from 'naca_num' argument passed to the function
self.naca_num = naca_num
m = int(str(naca_num)[0]) / 100
p = int(str(naca_num)[1]) / 10
t = int(str(naca_num)[2:]) / 100
# x-coordinate of maximum camber
p_c = p * self.chord
def get_camber(x):
"""
Returns 1 camber y-coordinate from 1 'x' along the airfoil chord.
"""
x_c = x
y_c = float()
if 0 <= x < p_c:
y_c = (m / (p**2)) * (2 * p * (x / self.chord) -
(x / self.chord)**2)
elif p_c <= x <= self.chord:
y_c = (m /
((1 - p)**2)) * ((1 - 2 * p) + 2 * p *
(x / self.chord) - (x / self.chord)**2)
else:
print('x-coordinate for camber is out of bounds. '
'Check that 0 < x <= chord.')
return (x_c, y_c * self.chord)
def get_thickness(x):
"""
Returns thickness from 1 'x' along the airfoil chord.
"""
y_t = float()
if 0 <= x <= self.chord:
y_t = 5 * t * self.chord * (0.2969 * sqrt(x / self.chord) -
0.1260 *
(x / self.chord) - 0.3516 *
(x / self.chord)**2 + 0.2843 *
(x / self.chord)**3 - 0.1015 *
(x / self.chord)**4)
else:
print('x-coordinate for thickness is out of bounds. '
'Check that 0 < x <= chord.')
return y_t
def get_dy_c(x):
"""
Returns dy_c/dx from 1 'x' along the airfoil chord.
"""
dy_c = float()
if 0 <= x < p_c:
dy_c = ((2 * m) / p**2) * (p - x / self.chord)
elif p_c <= x <= self.chord:
dy_c = (2 * m) / ((1 - p)**2) * (p - x / self.chord)
return dy_c
def get_theta(dy_c):
theta = atan(dy_c)
return theta
def get_upper_coordinates(x):
x_u = float()
y_u = float()
if 0 <= x < self.chord:
x_u = x - self.y_t[x] * sin(self.theta[x])
y_u = self.y_c[x] + self.y_t[x] * cos(self.theta[x])
elif x == self.chord:
x_u = x - self.y_t[x] * sin(self.theta[x])
y_u = 0 # Make upper curve finish at y = 0
return (x_u, y_u)
def get_lower_coordinates(x):
x_l = float()
y_l = float()
if 0 <= x < self.chord:
x_l = (x + self.y_t[x] * sin(self.theta[x]))
y_l = (self.y_c[x] - self.y_t[x] * cos(self.theta[x]))
elif x == self.chord:
x_l = (x + self.y_t[x] * sin(self.theta[x]))
y_l = 0 # Make lower curve finish at y = 0
return (x_l, y_l)
# Generate all our wing geometries from previous sub-functions
for x in range(0, self.chord + 1):
self.x_c.append(get_camber(x)[0])
self.y_c.append(get_camber(x)[1])
self.y_t.append(get_thickness(x))
self.dy_c.append(get_dy_c(x))
self.theta.append(get_theta(self.dy_c[x]))
self.x_u.append(get_upper_coordinates(x)[0])
self.y_u.append(get_upper_coordinates(x)[1])
self.x_l.append(get_lower_coordinates(x)[0])
self.y_l.append(get_lower_coordinates(x)[1])
super().pack_coord()
return None
class Spar(Coordinates):
"""Contains a single spar's location."""
global parent
def __init__(self):
super().__init__(parent.chord, parent.semi_span)
def add(self, airfoil_coord, spar_x):
"""
Add a single spar at the % chord location given to function.
Parameters:
coordinates: provided by Airfoil.coordinates[x_u, y_u, x_l, y_l].
material: spar's material. Assumes homogeneous material.
spar_x: spar's location as a % of total chord length.
Return:
None
"""
# Airfoil surface coordinates
# unpacked from 'coordinates' (list of lists in 'Coordinates').
x_u = airfoil_coord[0]
y_u = airfoil_coord[1]
x_l = airfoil_coord[2]
y_l = airfoil_coord[3]
# Scaled spar location with regards to chord
loc = spar_x * self.chord
# bisect_left: returns index of first value in x_u > loc.
# This ensures that the spar coordinates intersect with airfoil surface.
spar_x_u = bi.bisect_left(x_u, loc) # index of spar's x_u
spar_x_l = bi.bisect_left(x_l, loc) # index of spar's x_l
# These x and y coordinates are assigned to the spar, NOT airfoil.
self.x_u.append(x_u[spar_x_u])
self.y_u.append(y_u[spar_x_u])
self.x_l.append(x_l[spar_x_l])
self.y_l.append(y_l[spar_x_l])
super().pack_coord()
return None
class Stringer(Coordinates):
"""Contains the coordinates of all stringers."""
global parent
def __init__(self):
super().__init__(parent.chord, parent.semi_span)
def add(self, airfoil_coord, spar_coord, stringer_u_1, stringer_u_2,
stringer_l_1, stringer_l_2):
"""
Add equally distributed stringers to four airfoil locations
(upper nose, lower nose, upper surface, lower surface).
Parameters:
stringer_u_1: upper nose number of stringers
stringer_u_2: upper surface number of stringers
stringer_l_1: lower nose number of stringers
stringer_l_2: lower surface number of stringers
Returns:
None
"""
# Airfoil surface coordinates
# unpacked from 'coordinates' (list of lists in 'Coordinates').
airfoil_x_u = airfoil_coord[0]
airfoil_y_u = airfoil_coord[1]
airfoil_x_l = airfoil_coord[2]
airfoil_y_l = airfoil_coord[3]
# Spar coordinates
# unpacked from 'coordinates' (list of lists in 'Coordinates').
try:
spar_x_u = spar_coord[0]
spar_y_u = spar_coord[1]
spar_x_l = spar_coord[2]
spar_y_l = spar_coord[3]
except:
print('Unable to initialize stringers. Were spars created?')
# Find distance between leading edge and first upper stringer
interval = spar_x_u[0] / (stringer_u_1 + 1)
# initialise first self.stringer_x_u at first interval
x = interval
# Add upper stringers from leading edge until first spar.
for _ in range(0, stringer_u_1):
# Index of the first value of airfoil_x_u > x
index = bi.bisect_left(airfoil_x_u, x)
self.x_u.append(airfoil_x_u[index])
self.y_u.append(airfoil_y_u[index])
x += interval
# Add upper stringers from first spar until last spar
interval = (spar_x_u[-1] - spar_x_u[0]) / (stringer_u_2 + 1)
x = interval + spar_x_u[0]
for _ in range(0, stringer_u_2):
index = bi.bisect_left(airfoil_x_u, x)
self.x_u.append(airfoil_x_u[index])
self.y_u.append(airfoil_y_u[index])
x += interval
# Find distance between leading edge and first lower stringer
interval = spar_x_l[0] / (stringer_l_1 + 1)
x = interval
# Add lower stringers from leading edge until first spar.
for _ in range(0, stringer_l_1):
index = bi.bisect_left(airfoil_x_l, x)
self.x_l.append(airfoil_x_l[index])
self.y_l.append(airfoil_y_l[index])
x += interval
# Add lower stringers from first spar until last spar
interval = (spar_x_l[-1] - spar_x_l[0]) / (stringer_l_2 + 1)
x = interval + spar_x_l[0]
for _ in range(0, stringer_l_2):
index = bi.bisect_left(airfoil_x_l, x)
self.x_l.append(airfoil_x_l[index])
self.y_l.append(airfoil_y_l[index])
x += interval
super().pack_coord()
return None
def plot(airfoil, spar, stringer):
"""This function plots the elements passed as arguments."""
print('Plotting airfoil.')
# Plot chord
x_chord = [0, airfoil.chord]
y_chord = [0, 0]
plt.plot(x_chord, y_chord, linewidth='1')
# Plot mean camber line
plt.plot(airfoil.x_c,
airfoil.y_c,
'-.',
color='r',
linewidth='2',
label='mean camber line')
# Plot upper surface
plt.plot(airfoil.x_u, airfoil.y_u, '', color='b', linewidth='1')
# Plot lower surface
plt.plot(airfoil.x_l, airfoil.y_l, '', color='b', linewidth='1')
# Plot spars
try:
for _ in range(0, len(spar.x_u)):
x = (spar.x_u[_], spar.x_l[_])
y = (spar.y_u[_], spar.y_l[_])
plt.plot(x, y, '.-', color='b')
plt.legend()
except:
print('Did not plot spars. Were they added?')
# Plot stringers
try:
# Upper stringers
for _ in range(0, len(stringer.x_u)):
x = stringer.x_u[_]
y = stringer.y_u[_]
plt.plot(x, y, '.', color='y')
# Lower stringers
for _ in range(0, len(stringer.x_l)):
x = stringer.x_l[_]
y = stringer.y_l[_]
plt.plot(x, y, '.', color='y')
except:
print('Unable to plot stringers. Were they created?')
# Graph formatting
plt.gcf().set_size_inches(9, 2.2)
plt.xlabel('X axis')
plt.ylabel('Y axis')
# plt.gcf().set_size_inches(self.chord, max(self.y_u) - min(self.y_l))
plt.grid(axis='both', linestyle=':', linewidth=1)
plt.show()
return None
def main():
return None
if __name__ == '__main__':
main()
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