Start work on optimized multiprocessing random a/c gen. & eval.HEADmaster

1 files changed, 312 insertions, 0 deletions

diff --git a/aircraftstudio/creator/wing.py b/aircraftstudio/creator/wing.py
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+"""
+The wing.py module contains class definitions for and various components
+we add to an airfoil (spars, stringers, and ribs).
+
+Classes:
+    Airfoil: instantiated with class method to provide coordinates to heirs.
+    Spar: inherits from Airfoil.
+    Stringer: also inherits from Airfoil.
+
+Functions:
+    plot_geom(airfoil): generates a 2D plot of the airfoil & any components.
+"""
+
+import logging
+import numpy as np
+from math import sin, cos, atan
+import bisect as bi
+import matplotlib.pyplot as plt
+
+from aircraftstudio.creator import base
+import resources.materials as mt
+
+
+class Airfoil(base.Component):
+    """This class represents a single NACA airfoil.
+
+    The coordinates are saved as two np.arrays
+    for the x- and z-coordinates. The coordinates start at
+    the leading edge, travel over the airfoil's upper edge,
+    then loop back to the leading edge via the lower edge.
+
+    This method was chosen for easier future exports
+    to 3D CAD packages like SolidWorks, which can import such
+    geometry as coordinates written in a CSV file.
+    """
+    def __init__(self,
+                 parent,
+                 name,
+                 chord=68,
+                 semi_span=150,
+                 material=mt.aluminium):
+        super().__init__(parent, name)
+        parent.wing = self
+        if chord > 20:
+            self.chord = chord
+        else:
+            self.chord = 20
+            logging.debug('Chord too small, using minimum value of 20.')
+        parent
+        self.semi_span = semi_span
+        self.material = material
+        self.spars = []
+        self.stringers = []
+
+    def add_naca(self, naca_num=2412):
+        """Generate surface geometry for a NACA airfoil.
+
+        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
+        """
+        self.naca_num = naca_num
+        # Variables extracted from naca_num argument passed to the function
+        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 camber z-coordinate from 1 'x' along the airfoil chord.
+            """
+            z_c = float()
+            if 0 <= x < p_c:
+                z_c = (m / (p**2)) * (2 * p * (x / self.chord) -
+                                      (x / self.chord)**2)
+            elif p_c <= x <= self.chord:
+                z_c = (m /
+                       ((1 - p)**2)) * ((1 - 2 * p) + 2 * p *
+                                        (x / self.chord) - (x / self.chord)**2)
+            return (z_c * self.chord)
+
+        def get_thickness(x):
+            """Return thickness from 1 'x' along the airfoil chord."""
+            x = 0 if x < 0 else x
+            z_t = 5 * t * self.chord * (+0.2969 *
+                                        (x / self.chord)**0.5 - 0.1260 *
+                                        (x / self.chord)**1 - 0.3516 *
+                                        (x / self.chord)**2 + 0.2843 *
+                                        (x / self.chord)**3 - 0.1015 *
+                                        (x / self.chord)**4)
+            return z_t
+
+        def get_theta(x):
+            dz_c = float()
+            if 0 <= x < p_c:
+                dz_c = ((2 * m) / p**2) * (p - x / self.chord)
+            elif p_c <= x <= self.chord:
+                dz_c = (2 * m) / ((1 - p)**2) * (p - x / self.chord)
+
+            theta = atan(dz_c)
+            return theta
+
+        def get_coord_u(x):
+            x = x - get_thickness(x) * sin(get_theta(x))
+            z = get_camber(x) + get_thickness(x) * cos(get_theta(x))
+            return (x, z)
+
+        def get_coord_l(x):
+            x = x + get_thickness(x) * sin(get_theta(x))
+            z = get_camber(x) - get_thickness(x) * cos(get_theta(x))
+            return (x, z)
+
+        # Densify x-coordinates 10 times for first 1/4 chord length
+        x_chord_25_percent = round(self.chord / 4)
+        x_chord = [i / 10 for i in range(x_chord_25_percent * 10)]
+        x_chord.extend(i for i in range(x_chord_25_percent, self.chord + 1))
+        # Generate our airfoil skin geometry from previous sub-functions
+        self.x_c = np.array([])
+        self.z_c = np.array([])
+        # Upper surface and camber line
+        for x in x_chord:
+            self.x_c = np.append(self.x_c, x)
+            self.z_c = np.append(self.z_c, get_camber(x))
+            self.x = np.append(self.x, get_coord_u(x)[0])
+            self.z = np.append(self.z, get_coord_u(x)[1])
+        # Lower surface
+        for x in x_chord[::-1]:
+            self.x = np.append(self.x, get_coord_l(x)[0])
+            self.z = np.append(self.z, get_coord_l(x)[1])
+        return None
+
+
+class Spar(base.Component):
+    """Contains a single spar's data."""
+    def __init__(self, parent, name, loc_percent=0.30, material=mt.aluminium):
+        """Set spar location as percent of chord length."""
+        super().__init__(parent, name)
+        parent.spars.append(self)
+        self.material = material
+        self.cap_area = float()
+        # bi.bisect_left: returns index of first value in parent.x > loc
+        # This ensures that spar geom intersects with airfoil geom.
+        loc = loc_percent * parent.chord
+        # Spar upper coordinates
+        spar_u = bi.bisect_left(parent.x, loc) - 1
+        self.x = np.append(self.x, parent.x[spar_u])
+        self.z = np.append(self.z, parent.z[spar_u])
+        # Spar lower coordinates
+        spar_l = bi.bisect_left(parent.x[::-1], loc)
+        self.x = np.append(self.x, parent.x[-spar_l])
+        self.z = np.append(self.z, parent.z[-spar_l])
+        return None
+
+    def set_cap_area(self, cap_area):
+        self.cap_area = cap_area
+        return None
+
+    def set_mass(self, mass):
+        self.mass = mass
+        return None
+
+
+class Stringer(base.Component):
+    """Contains the coordinates of all stringers."""
+    def __init__(self,
+                 parent,
+                 name,
+                 den_u_1=4,
+                 den_u_2=4,
+                 den_l_1=4,
+                 den_l_2=4):
+        """Add equally distributed stringers to four airfoil locations
+        (upper nose, lower nose, upper surface, lower surface).
+
+        den_u_1: upper nose number of stringers
+        den_u_2: upper surface number of stringers
+        den_l_1: lower nose number of stringers
+        den_l_2: lower surface number of stringers
+        """
+        super().__init__(parent, name)
+        parent.stringers = self
+        self.x_start = []
+        self.x_end = []
+        self.z_start = []
+        self.z_end = []
+        self.diameter = float()
+        self.area = float()
+
+        # Find distance between leading edge and first upper stringer
+        # interval = self.parent.spars[0].x[0] / (den_u_1 + 1)
+        interval = 2
+        # initialise first self.stringer_x at first interval
+        x = interval
+        # Add upper stringers from leading edge until first spar.
+        for _ in range(0, den_u_1):
+            # Index of the first value of airfoil.x > x
+            i = bi.bisect_left(self.parent.x, x)
+            self.x = np.append(self.x, self.parent.x[i])
+            self.z = np.append(self.z, self.parent.z[i])
+            x += interval
+            # Add upper stringers from first spar until last spar
+        interval = (self.parent.spars[-1].x[0] -
+                    self.parent.spars[0].x[0]) / (den_u_2 + 1)
+        x = interval + self.parent.spars[0].x[0]
+        for _ in range(0, den_u_2):
+            i = bi.bisect_left(self.parent.x, x)
+            self.x = np.append(self.x, self.parent.x[i])
+            self.z = np.append(self.z, self.parent.z[i])
+            x += interval
+
+        # Find distance between leading edge and first lower stringer
+        interval = self.parent.spars[0].x[1] / (den_l_1 + 1)
+        x = interval
+        # Add lower stringers from leading edge until first spar.
+        for _ in range(0, den_l_1):
+            i = bi.bisect_left(self.parent.x[::-1], x)
+            self.x = np.append(self.x, self.parent.x[-i])
+            self.z = np.append(self.z, self.parent.z[-i])
+            x += interval
+            # Add lower stringers from first spar until last spar
+        interval = (self.parent.spars[-1].x[1] -
+                    self.parent.spars[0].x[1]) / (den_l_2 + 1)
+        x = interval + self.parent.spars[0].x[1]
+        for _ in range(0, den_l_2):
+            i = bi.bisect_left(self.parent.x[::-1], x)
+            self.x = np.append(self.x, self.parent.x[-i])
+            self.z = np.append(self.z, self.parent.z[-i])
+            x += interval
+        return None
+
+    def add_area(self, area):
+        self.area = area
+        return None
+
+    def add_mass(self, mass):
+        self.mass = len(self.x) * mass + len(self.x) * mass
+        return None
+
+    def add_webs(self, thickness):
+        """Add webs to stringers."""
+        for _ in range(len(self.x) // 2):
+            self.x_start.append(self.x[_])
+            self.x_end.append(self.x[_ + 1])
+            self.z_start.append(self.z[_])
+            self.z_end.append(self.z[_ + 1])
+            self.thickness = thickness
+        return None
+
+    def info_print(self, round=2):
+        super().info_print(round)
+        print('Stringer Area:\n', np.around(self.area, round))
+        return None
+
+
+def plot_geom(airfoil):
+    """This function plots the airfoil's + sub-components' geometry."""
+    fig, ax = plt.subplots()
+
+    # Plot chord
+    x = [0, airfoil.chord]
+    y = [0, 0]
+    ax.plot(x, y, linewidth='1')
+    # Plot quarter chord
+    ax.plot(airfoil.chord / 4,
+            0,
+            '.',
+            color='g',
+            markersize=24,
+            label='Quarter-chord')
+    # Plot mean camber line
+    ax.plot(airfoil.x_c,
+            airfoil.z_c,
+            '-.',
+            color='r',
+            linewidth='2',
+            label='Mean camber line')
+    # Plot airfoil surfaces
+    ax.plot(airfoil.x, airfoil.z, color='b', linewidth='1')
+
+    try:  # Plot spars
+        for spar in airfoil.spars:
+            x = (spar.x)
+            y = (spar.z)
+            ax.plot(x, y, '-', color='y', linewidth='4')
+    except AttributeError:
+        print('No spars to plot.')
+    try:  # Plot stringers
+        for i in range(len(airfoil.stringers.x)):
+            x = airfoil.stringers.x[i]
+            y = airfoil.stringers.z[i]
+            ax.plot(x, y, '.', color='y', markersize=12)
+    except AttributeError:
+        print('No stringers to plot.')
+
+    ax.set(title='NACA ' + str(airfoil.naca_num) + ' airfoil',
+           xlabel='X axis',
+           ylabel='Z axis')
+
+    plt.grid(axis='both', linestyle=':', linewidth=1)
+    plt.gca().set_aspect('equal', adjustable='box')
+    plt.gca().legend(bbox_to_anchor=(1, 1),
+                     bbox_transform=plt.gcf().transFigure)
+    plt.show()
+    return fig, ax