From 588c34a3d595fcad5e93b8d4893f1098ce64d046 Mon Sep 17 00:00:00 2001 From: blendoit Date: Mon, 30 Sep 2019 18:42:34 -0700 Subject: First commit! Changed coordinate lists into numpy arrays. --- .gitignore | 4 + README.org | 7 + creator/fuselage.py | 0 creator/propulsion.py | 0 creator/wing.py | 375 +++++++++++++++++++++++ evaluator.py | 284 ++++++++++++++++++ example_airfoil.py | 82 +++++ generator.py | 64 ++++ gui.py | 81 +++++ resources/materials.py | 8 + wing_scripts/eye_beam_example.m | 70 +++++ wing_scripts/get_dp.m | 4 + wing_scripts/get_ds.m | 20 ++ wing_scripts/get_int.m | 35 +++ wing_scripts/get_z.m | 34 +++ wing_scripts/my_progress.m | 459 ++++++++++++++++++++++++++++ wing_scripts/stringersBeamExample.m | 47 +++ wing_scripts/wingAnalysis_190422.m | 579 ++++++++++++++++++++++++++++++++++++ 18 files changed, 2153 insertions(+) create mode 100644 .gitignore create mode 100644 README.org create mode 100644 creator/fuselage.py create mode 100644 creator/propulsion.py create mode 100644 creator/wing.py create mode 100644 evaluator.py create mode 100644 example_airfoil.py create mode 100644 generator.py create mode 100644 gui.py create mode 100644 resources/materials.py create mode 100644 wing_scripts/eye_beam_example.m create mode 100644 wing_scripts/get_dp.m create mode 100644 wing_scripts/get_ds.m create mode 100644 wing_scripts/get_int.m create mode 100644 wing_scripts/get_z.m create mode 100644 wing_scripts/my_progress.m create mode 100644 wing_scripts/stringersBeamExample.m create mode 100644 wing_scripts/wingAnalysis_190422.m diff --git a/.gitignore b/.gitignore new file mode 100644 index 0000000..c796902 --- /dev/null +++ b/.gitignore @@ -0,0 +1,4 @@ +# .gitignore +**/__pycache__/ +**/log.txt +save/ \ No newline at end of file diff --git a/README.org b/README.org new file mode 100644 index 0000000..caa04a9 --- /dev/null +++ b/README.org @@ -0,0 +1,7 @@ +#+TITLE: UCLA MAE 154B +#+SUBTITLE: Spring 2019 Final Project + +This program enables the creation of NACA airfoils; +the analysis of the airfoil's structural properties; +the optimization via genetic algorithm of a population of airfoils; +With the final objective of designing a lightweight FAR 23 compliant airfoil. diff --git a/creator/fuselage.py b/creator/fuselage.py new file mode 100644 index 0000000..e69de29 diff --git a/creator/propulsion.py b/creator/propulsion.py new file mode 100644 index 0000000..e69de29 diff --git a/creator/wing.py b/creator/wing.py new file mode 100644 index 0000000..4988cb5 --- /dev/null +++ b/creator/wing.py @@ -0,0 +1,375 @@ +""" +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 sys +import os.path +import logging +import numpy as np +from math import sin, cos, atan +import bisect as bi +import matplotlib.pyplot as plt + +logging.basicConfig(filename='log.txt', + level=logging.DEBUG, + format='%(asctime)s - %(levelname)s - %(message)s') + + +class Component: + """Basic component providing coordinates and tools.""" + + # TODO: define defaults in separate module + def __init__(self): + self.x = np.array([]) + self.z = np.array([]) + self.material = str() + self.mass = float() + + def set_material(self, material): + """Set the component bulk material.""" + self.material = material + + def info_print(self, round): + """Print all the component's coordinates to the terminal.""" + name = f' CREATOR DATA FOR {str(self).upper()} ' + num_of_dashes = len(name) + print(num_of_dashes * '-') + print(name) + for k, v in self.__dict__.items(): + if type(v) != list: + print('{}:\n'.format(k), v) + print(num_of_dashes * '-') + for k, v in self.__dict__.items(): + if type(v) == list: + print('{}:\n'.format(k), np.around(v, round)) + return None + + def info_save(self, save_path, number): + """Save all the object's coordinates (must be full path).""" + file_name = f'{str(self).lower()}_{number}.txt' + full_path = os.path.join(save_path, file_name) + try: + with open(full_path, 'w') as sys.stdout: + self.info_print(6) + # This line required to reset behavior of sys.stdout + sys.stdout = sys.__stdout__ + logging.debug(f'Successfully wrote to file {full_path}') + except IOError: + print(f'Unable to write {file_name} to specified directory.\n', + 'Was the full path passed to the function?') + return None + + +class Airfoil(Component): + """This class represents a single NACA airfoil. + + The coordinates are saved as two lists + 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. + """ + + # TODO: default values in separate module + def __init__(self, chord, semi_span, material): + super().__init__() + # self.x = np.array([]) + # self.z = np.array([]) + # self.chord = chord + """Create airfoil from its chord and semi-span.""" + self.chord = chord if chord > 20 else 20 + if chord <= 20: + logging.debug('Chord too small, using minimum value of 20.') + self.semi_span = semi_span + self.material = material + self.naca_num = int() + + def __str__(self): + return type(self).__name__ + + def add_naca(self, naca_num): + """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(Component): + """Contains a single spar's data.""" + def __init__(self, airfoil, loc_percent, material): + """Set spar location as percent of chord length.""" + super().__init__() + super().set_material(material) + self.cap_area = float() + loc = loc_percent * airfoil.chord + # bi.bisect_left: returns index of first value in airfoil.x > loc + # This ensures that spar geom intersects with airfoil geom. + # Spar upper coordinates + spar_u = bi.bisect_left(airfoil.x, loc) - 1 + self.x = np.append(self.x, airfoil.x[spar_u]) + self.z = np.append(self.z, airfoil.z[spar_u]) + # Spar lower coordinates + spar_l = bi.bisect_left(airfoil.x[::-1], loc) + self.x = np.append(self.x, airfoil.x[-spar_l]) + self.z = np.append(self.z, airfoil.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(Component): + """Contains the coordinates of all stringers.""" + def __init__(self): + super().__init__() + self.x_start = [] + self.x_end = [] + self.z_start = [] + self.z_end = [] + self.diameter = float() + self.area = float() + + def add_coord(self, airfoil, spars, 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: + airfoil_coord: packed airfoil coordinates + spar_coord: packed spar coordinates + 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 + """ + + # Find distance between leading edge and first upper stringer + interval = spars.x[0][0] / (stringer_u_1 + 1) + # initialise first self.stringer_x 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 > x + i = bi.bisect_left(airfoil.x, x) + self.x.append(airfoil.x[i]) + self.z.append(airfoil.z[i]) + x += interval + # Add upper stringers from first spar until last spar + # TODO: stringer placement if only one spar is created + interval = (airfoil.spar.x[-1][0] - + airfoil.spar.x[0][0]) / (stringer_u_2 + 1) + x = interval + airfoil.spar.x[0][0] + for _ in range(0, stringer_u_2): + i = bi.bisect_left(airfoil.x, x) + self.x.append(airfoil.x[i]) + self.z.append(airfoil.z[i]) + x += interval + + # Find distance between leading edge and first lower stringer + interval = airfoil.spar.x[0][1] / (stringer_l_1 + 1) + x = interval + # Add lower stringers from leading edge until first spar. + for _ in range(0, stringer_l_1): + i = bi.bisect_left(airfoil.x[::-1], x) + self.x.append(airfoil.x[-i]) + self.z.append(airfoil.z[-i]) + x += interval + # Add lower stringers from first spar until last spar + interval = (airfoil.spar.x[-1][1] - + airfoil.spar.x[0][1]) / (stringer_l_2 + 1) + x = interval + airfoil.spar.x[0][1] + for _ in range(0, stringer_l_2): + i = bi.bisect_left(airfoil.x[::-1], x) + self.x.append(airfoil.x[-i]) + self.z.append(airfoil.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): + super().info_print(round) + print('Stringer Area:\n', np.around(self.area, round)) + return None + + +def plot_geom(airfoil, spars, stringers): + """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') + + # Plot spars + try: + for spar in spars: + x = (spar.x) + y = (spar.z) + ax.plot(x, y, '-', color='y', linewidth='4') + except AttributeError: + print('No spars to plot.') + # Plot stringers + try: + for _ in range(0, len(airfoil.stringer.x)): + x = airfoil.stringer.x[_] + y = airfoil.stringer.z[_] + ax.plot(x, y, '.', color='y', markersize=12) + except AttributeError: + print('No stringers to plot.') + + # Graph formatting + # plot_bound = np.amax(airfoil.x) + ax.set( + title='NACA ' + str(airfoil.naca_num) + ' airfoil', + xlabel='X axis', + # xlim=[-0.10 * plot_bound, 1.10 * plot_bound], + ylabel='Z axis') + # ylim=[-(1.10 * plot_bound / 2), (1.10 * plot_bound / 2)]) + + 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 + + +def main(): + return None + + +if __name__ == '__main__': + main() diff --git a/evaluator.py b/evaluator.py new file mode 100644 index 0000000..85325a9 --- /dev/null +++ b/evaluator.py @@ -0,0 +1,284 @@ +""" +The evaluator.py module contains a single Evaluator class, +which knows all the attributes of a specified Airfoil instance, +and contains functions to analyse the airfoil's geometrical +& structural properties. +""" + +import sys +import os.path +import numpy as np +from math import sqrt +import matplotlib.pyplot as plt + + +class Evaluator: + """Performs structural evaluations for the airfoil passed as argument.""" + def __init__(self, airfoil): + # Evaluator knows all geometrical info from evaluated airfoil + self.airfoil = airfoil + self.spar = airfoil.spar + self.stringer = airfoil.stringer + # Global dimensions + self.chord = airfoil.chord + self.semi_span = airfoil.semi_span + # Mass & spanwise distribution + self.mass_total = float(airfoil.mass + airfoil.spar.mass + + airfoil.stringer.mass) + self.mass_dist = [] + # Lift + self.lift_rectangular = [] + self.lift_elliptical = [] + self.lift_total = [] + # Drag + self.drag = [] + # centroid + self.centroid = [] + # Inertia terms: + self.I_ = {'x': 0, 'z': 0, 'xz': 0} + + def __str__(self): + return type(self).__name__ + + def info_print(self, round): + """Print all the component's evaluated data to the terminal.""" + name = ' EVALUATOR DATA FOR {} '.format(str(self).upper()) + num_of_dashes = len(name) + print(num_of_dashes * '-') + print(name) + for k, v in self.__dict__.items(): + if type(v) != list: + print('{}:\n'.format(k), v) + print(num_of_dashes * '-') + for k, v in self.__dict__.items(): + if type(v) == list: + print('{}:\n'.format(k), np.around(v, round)) + return None + + def info_save(self, save_path, number): + """Save all the object's coordinates (must be full path).""" + file_name = 'airfoil_{}_eval.txt'.format(number) + full_path = os.path.join(save_path, file_name) + try: + with open(full_path, 'w') as sys.stdout: + self.info_print(6) + # This line required to reset behavior of sys.stdout + sys.stdout = sys.__stdout__ + print('Successfully wrote to file {}'.format(full_path)) + except IOError: + print( + 'Unable to write {} to specified directory.\n'.format( + file_name), 'Was the full path passed to the function?') + return None + + # All these functions take integer arguments and return lists. + + def get_lift_rectangular(self, lift): + L_prime = [lift / (self.semi_span * 2) for x in range(self.semi_span)] + return L_prime + + def get_lift_elliptical(self, L_0): + L_prime = [ + L_0 / (self.semi_span * 2) * sqrt(1 - (y / self.semi_span)**2) + for y in range(self.semi_span) + ] + return L_prime + + def get_lift_total(self): + F_z = [(self.lift_rectangular[_] + self.lift_elliptical[_]) / 2 + for _ in range(len(self.lift_rectangular))] + return F_z + + def get_mass_distribution(self, total_mass): + F_z = [total_mass / self.semi_span for x in range(0, self.semi_span)] + return F_z + + def get_drag(self, drag): + # Transform semi-span integer into list + semi_span = [x for x in range(0, self.semi_span)] + + # Drag increases after 80% of the semi_span + cutoff = round(0.8 * self.semi_span) + + # Drag increases by 25% after 80% of the semi_span + F_x = [drag for x in semi_span[0:cutoff]] + F_x.extend([1.25 * drag for x in semi_span[cutoff:]]) + return F_x + + def get_centroid(self): + """Return the coordinates of the centroid.""" + stringer_area = self.stringer.area + cap_area = self.spar.cap_area + + caps_x = [value for spar in self.spar.x for value in spar] + caps_z = [value for spar in self.spar.z for value in spar] + stringers_x = self.stringer.x + stringers_z = self.stringer.z + + denominator = float( + len(caps_x) * cap_area + len(stringers_x) * stringer_area) + + centroid_x = float( + sum([x * cap_area for x in caps_x]) + + sum([x * stringer_area for x in stringers_x])) + centroid_x = centroid_x / denominator + + centroid_z = float( + sum([z * cap_area for z in caps_z]) + + sum([z * stringer_area for z in stringers_z])) + centroid_z = centroid_z / denominator + + return (centroid_x, centroid_z) + + def get_inertia_terms(self): + """Obtain all inertia terms.""" + stringer_area = self.stringer.area + cap_area = self.spar.cap_area + + # Adds upper and lower components' coordinates to list + x_stringers = self.stringer.x + z_stringers = self.stringer.z + x_spars = self.spar.x[:][0] + self.spar.x[:][1] + z_spars = self.spar.z[:][0] + self.spar.z[:][1] + stringer_count = range(len(x_stringers)) + spar_count = range(len(self.spar.x)) + + # I_x is the sum of the contributions of the spar caps and stringers + # TODO: replace list indices with dictionary value + I_x = sum([ + cap_area * (z_spars[i] - self.centroid[1])**2 for i in spar_count + ]) + I_x += sum([ + stringer_area * (z_stringers[i] - self.centroid[1])**2 + for i in stringer_count + ]) + + I_z = sum([ + cap_area * (x_spars[i] - self.centroid[0])**2 for i in spar_count + ]) + I_z += sum([ + stringer_area * (x_stringers[i] - self.centroid[0])**2 + for i in stringer_count + ]) + + I_xz = sum([ + cap_area * (x_spars[i] - self.centroid[0]) * + (z_spars[i] - self.centroid[1]) for i in spar_count + ]) + I_xz += sum([ + stringer_area * (x_stringers[i] - self.centroid[0]) * + (z_stringers[i] - self.centroid[1]) for i in stringer_count + ]) + return (I_x, I_z, I_xz) + + def get_dx(self, component): + return [x - self.centroid[0] for x in component.x_start] + + def get_dz(self, component): + return [x - self.centroid[1] for x in component.x_start] + + def get_dP(self, xDist, zDist, V_x, V_z, area): + I_x = self.I_['x'] + I_z = self.I_['z'] + I_xz = self.I_['xz'] + denom = float(I_x * I_z - I_xz**2) + z = float() + for _ in range(len(xDist)): + z += float(-area * xDist[_] * (I_x * V_x - I_xz * V_z) / denom - + area * zDist[_] * (I_z * V_z - I_xz * V_x) / denom) + return z + + def analysis(self, V_x, V_z): + """Perform all analysis calculations and store in class instance.""" + self.drag = self.get_drag(10) + self.lift_rectangular = self.get_lift_rectangular(13.7) + self.lift_elliptical = self.get_lift_elliptical(15) + self.lift_total = self.get_lift_total() + self.mass_dist = self.get_mass_distribution(self.mass_total) + self.centroid = self.get_centroid() + self.I_['x'] = self.get_inertia_terms()[0] + self.I_['z'] = self.get_inertia_terms()[1] + self.I_['xz'] = self.get_inertia_terms()[2] + spar_dx = self.get_dx(self.spar) + spar_dz = self.get_dz(self.spar) + self.spar.dP_x = self.get_dP(spar_dx, spar_dz, V_x, 0, + self.spar.cap_area) + self.spar.dP_z = self.get_dP(spar_dx, spar_dz, 0, V_z, + self.spar.cap_area) + return None + + +def plot_geom(evaluator): + """This function plots analysis results over the airfoil's geometry.""" + # Plot chord + x_chord = [0, evaluator.chord] + y_chord = [0, 0] + plt.plot(x_chord, y_chord, linewidth='1') + # Plot quarter chord + plt.plot(evaluator.chord / 4, + 0, + '.', + color='g', + markersize=24, + label='Quarter-chord') + # Plot airfoil surfaces + x = [0.98 * x for x in evaluator.airfoil.x] + y = [0.98 * z for z in evaluator.airfoil.z] + plt.fill(x, y, color='w', linewidth='1', fill=False) + x = [1.02 * x for x in evaluator.airfoil.x] + y = [1.02 * z for z in evaluator.airfoil.z] + plt.fill(x, y, color='b', linewidth='1', fill=False) + + # Plot spars + try: + for _ in range(len(evaluator.spar.x)): + x = (evaluator.spar.x[_]) + y = (evaluator.spar.z[_]) + plt.plot(x, y, '-', color='b') + except AttributeError: + print('No spars to plot.') + # Plot stringers + try: + for _ in range(0, len(evaluator.stringer.x)): + x = evaluator.stringer.x[_] + y = evaluator.stringer.z[_] + plt.plot(x, y, '.', color='y', markersize=12) + except AttributeError: + print('No stringers to plot.') + + # Plot centroid + x = evaluator.centroid[0] + y = evaluator.centroid[1] + plt.plot(x, y, '.', color='r', markersize=24, label='centroid') + + # Graph formatting + plt.xlabel('X axis') + plt.ylabel('Z axis') + + plot_bound = max(evaluator.airfoil.x) + plt.xlim(-0.10 * plot_bound, 1.10 * plot_bound) + plt.ylim(-(1.10 * plot_bound / 2), (1.10 * plot_bound / 2)) + plt.gca().set_aspect('equal', adjustable='box') + plt.gca().legend() + plt.grid(axis='both', linestyle=':', linewidth=1) + plt.show() + return None + + +def plot_lift(evaluator): + x = range(evaluator.semi_span) + y_1 = evaluator.lift_rectangular + y_2 = evaluator.lift_elliptical + y_3 = evaluator.lift_total + plt.plot(x, y_1, '.', color='b', markersize=4, label='Rectangular lift') + plt.plot(x, y_2, '.', color='g', markersize=4, label='Elliptical lift') + plt.plot(x, y_3, '.', color='r', markersize=4, label='Total lift') + + # Graph formatting + plt.xlabel('Semi-span location') + plt.ylabel('Lift') + + plt.gca().legend() + plt.grid(axis='both', linestyle=':', linewidth=1) + plt.show() + return None diff --git a/example_airfoil.py b/example_airfoil.py new file mode 100644 index 0000000..292c1e9 --- /dev/null +++ b/example_airfoil.py @@ -0,0 +1,82 @@ +"""This example illustrates the usage of creator, evaluator and generator. + +All the steps of airfoil creation & evaluation are detailed here; +furthermore, the generator.py module contains certain presets +(default airfoils). + +Create an airfoil; +Evaluate an airfoil; +Generate a population of airfoils & optimize. +""" + +from resources import materials as mt +from creator import wing, fuselage, propulsion +# from evaluator import +# from generator import + +import time +start_time = time.time() + +# Airfoil dimensions (in) +NACA_NUM = 2412 + +# Thicknesses +SPAR_THICKNESS = 0.4 +SKIN_THICKNESS = 0.1 + +# Component masses (lbs) +AIRFOIL_MASS = 10 +SPAR_MASS = 10 +STRINGER_MASS = 5 + +# Area (sqin) +SPAR_CAP_AREA = 0.3 +STRINGER_AREA = 0.1 + +# Amount of stringers +TOP_STRINGERS = 6 +BOTTOM_STRINGERS = 4 +NOSE_TOP_STRINGERS = 3 +NOSE_BOTTOM_STRINGERS = 5 + +SAVE_PATH = '/home/blendux/Projects/Aircraft_Studio/save' + +# Create airfoil instance +af = wing.Airfoil(68, 150, mt.aluminium) +af.add_naca(NACA_NUM) +# af.info_print(2) +af.info_save(SAVE_PATH, 'foo_name') + +# Create spar instances +af.spar1 = wing.Spar(af, 0.23, mt.aluminium) +af.spar2 = wing.Spar(af, 0.57, mt.aluminium) +# af.spar1.info_print(2) +# af.spar2.info_print(2) +af.spar1.info_save(SAVE_PATH, 'spar1') +af.spar2.info_save(SAVE_PATH, 'spar2') + +# # Create stringer instance +# af.stringer = wing.Stringer() +# # Compute the stringer coordinates from their quantity in each zone +# af.stringer.add_coord(af, [af.spar1, af.spar2], NOSE_TOP_STRINGERS, TOP_STRINGERS, +# NOSE_BOTTOM_STRINGERS, BOTTOM_STRINGERS) +# af.stringer.add_area(STRINGER_AREA) +# af.stringer.add_webs(SKIN_THICKNESS) +# af.stringer.info_print(2) +# af.stringer.info_save(SAVE_PATH, 'foo_name') + +# Plot components with matplotlib +wing.plot_geom(af, [af.spar1, af.spar2], None) + +# Evaluator object contains airfoil analysis results. +# eval = evaluator.Evaluator(af) +# The analysis is performed in the evaluator.py module. +# eval.analysis(1, 1) +# eval.info_print(2) +# eval.info_save(SAVE_PATH, 'foo_name') +# evaluator.plot_geom(eval) +# evaluator.plot_lift(eval) + +# Final execution time +final_time = time.time() - start_time +print(f"--- {round(final_time, 4)}s seconds ---") diff --git a/generator.py b/generator.py new file mode 100644 index 0000000..0213828 --- /dev/null +++ b/generator.py @@ -0,0 +1,64 @@ +# 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 . +""" +The generator.py module contains a single Population class, +which represents a collection of randomized airfoils. +""" + +from tools import creator + + +def default_airfoil(): + """Generate the default airfoil.""" + airfoil = creator.Airfoil.from_dimensions(100, 200) + airfoil.add_naca(2412) + airfoil.add_mass(10) + + airfoil.spar = creator.Spar() + airfoil.spar.add_coord(airfoil, 0.23) + airfoil.spar.add_coord(airfoil, 0.57) + airfoil.spar.add_spar_caps(0.3) + airfoil.spar.add_mass(10) + airfoil.spar.add_webs(0.4) + + airfoil.stringer = creator.Stringer() + airfoil.stringer.add_coord(airfoil, 3, 6, 5, 4) + airfoil.stringer.add_area(0.1) + airfoil.stringer.add_mass(5) + airfoil.stringer.add_webs(0.1) + + return airfoil + + +class Population(creator.Airfoil): + """Collection of random airfoils.""" + + def __init__(self, size): + af = creator.Airfoil + # print(af) + self.size = size + self.gen_number = 0 # incremented for every generation + + def mutate(self, prob_mt): + """Randomly mutate the genes of prob_mt % of the population.""" + + def crossover(self, prob_cx): + """Combine the genes of prob_cx % of the population.""" + + def reproduce(self, prob_rp): + """Pass on the genes of the fittest prob_rp % of the population.""" + + def fitness(): + """Rate the fitness of an individual on a relative scale (0-100)""" diff --git a/gui.py b/gui.py new file mode 100644 index 0000000..2eea281 --- /dev/null +++ b/gui.py @@ -0,0 +1,81 @@ +from tools import creator, evaluator, generator +# import creator +# import evaluator +# import generator +import tkinter as tk +import tkinter.ttk as ttk + +from matplotlib.backends.backend_tkagg import ( + FigureCanvasTkAgg, NavigationToolbar2Tk) + + +class MainWindow(tk.Frame): + """Main editor window.""" + + def __init__(self, *args, **kwargs): + tk.Frame.__init__(self, *args, **kwargs) + root = tk.Tk() + root.wm_title('MAE 154B - Airfoil Design, Evaluation, Optimization') + + # self.button = tk.Button(self, text="Create new window", + # command=self.create_window) + # self.button.pack(side="top") + frame_1 = ttk.Frame(root) + l_naca, e_naca = new_field(frame_1, 'naca') + l_chord, e_chord = new_field(frame_1, 'chord') + l_semi_span, e_semi_span = new_field(frame_1, 'semi_span') + af = generator.default_airfoil() + # Graph window + frame_2 = ttk.Frame(root) + fig, ax = creator.plot_geom(af, False) + plot = FigureCanvasTkAgg(fig, frame_2) + # plot.draw() + toolbar = NavigationToolbar2Tk(plot, frame_2) + # toolbar.update() + + l_naca.grid(row=0, column=0) + e_naca.grid(row=0, column=1, padx=4) + # b_naca.grid(row=0, column=2) + l_chord.grid(row=1, column=0) + e_chord.grid(row=1, column=1, padx=4) + l_semi_span.grid(row=2, column=0, padx=4) + e_semi_span.grid(row=2, column=1, padx=4) + frame_1.pack(side=tk.LEFT) + # Graph window + plot.get_tk_widget().pack(expand=1, fill=tk.BOTH) + toolbar.pack() + frame_2.pack(side=tk.LEFT) + + def create_window(self): + self.counter += 1 + window = tk.Toplevel(self) + window.wm_title("Window #%s" % self.counter) + label = tk.Label(window, text="This is window #%s" % self.counter) + label.pack(side="top", fill="both", expand=True, padx=100, pady=100) + + +def new_field(parent, name): + """Add a new user input field.""" + + label = ttk.Label(parent, text=name) + entry = ttk.Entry(parent) + return label, entry + + +def set_naca(name): + naca_num = name.get() + print(naca_num) + + +def set_chord(name): + chord = name.get() + print(chord) + + +def set_semi_span(name): + semi_span = name.get() + print(semi_span) + + +# plot.get_tk_widget().pack() +MainWindow().mainloop() diff --git a/resources/materials.py b/resources/materials.py new file mode 100644 index 0000000..480b518 --- /dev/null +++ b/resources/materials.py @@ -0,0 +1,8 @@ +aluminium = { + "name": "aluminium", + "category": "metal", + "density": 2.70, + "mod_young": 70E9, + "mod_shear": 26E9, + "mod_bulk": 76E9 +} diff --git a/wing_scripts/eye_beam_example.m b/wing_scripts/eye_beam_example.m new file mode 100644 index 0000000..70b4d92 --- /dev/null +++ b/wing_scripts/eye_beam_example.m @@ -0,0 +1,70 @@ +% Bending/Shear stress example +close all; + +length = 20; % in +force = 10000; %lbs + +%eye-beam dimensions + +max_width = 4; % in +min_width = 1; % in +y_max = 4; % in +center_y = 2; % in + + +%max bending moment at the root... + +M = force*length; + +I = min_width*(2*center_y)^3/12 + 2*( max_width*(y_max-center_y)^3/12 + ... + max_width*(y_max-center_y)*((y_max+center_y)/2)^2); + +sigma_max = M * y_max / I; + + +% solve for shear stress distribution +% V / (I * t) * int(y*da) + +% Point 1: evaluated at location just before thickness changes from 4 to 1 in +tempCoeff = force / (I * max_width); +int_y_da = ((y_max+center_y)/2) * max_width*(y_max-center_y); +shear_1 = tempCoeff*int_y_da; + +% Point 2: evaluated at location just after thickness changes from 4 to 1 in +tempCoeff = force / (I * min_width); +shear_2 = tempCoeff*int_y_da; + + +% Point 3: evaluated at center of beam +tempCoeff = force / (I * min_width); +int_y_da = (center_y/2) * min_width*center_y; +shear_3 = shear_2+tempCoeff*int_y_da; + +%evaluating continous integral for width of 4.. +int_y_da_4 = force / (I * max_width)*4*(y_max^2/2 - (center_y:.1:y_max).^2/2); + +%evaluating continous integral for width of 1.. +int_y_da_1 = shear_2 + force / (I * min_width)*1*(center_y^2/2 - (0:.1:center_y).^2/2); + +figure; grid on; hold on;set(gcf,'color',[1 1 1]); +plot(int_y_da_4,center_y:.1:y_max,'linewidth',2) +plot(int_y_da_1,0:.1:center_y,'linewidth',2) +plot(int_y_da_1,0:-.1:-center_y,'linewidth',2) +plot(int_y_da_4,-center_y:-.1:-y_max,'linewidth',2) +plot([shear_1 shear_2],[center_y center_y],'linewidth',2) +plot([shear_1 shear_2],[-center_y -center_y],'linewidth',2) + +plot(shear_1,center_y,'o') +plot(shear_2,center_y,'o') +plot(shear_3,0,'o') +plot(shear_2,-center_y,'o') +plot(shear_1,-center_y,'o') + +xlabel('shear stress (lb/in^2)','fontsize',16,'fontweight','bold');ylabel('Distance from Center (in)','fontsize',16,'fontweight','bold') +set(gca,'FontSize',16,'fontweight','bold'); + + +figure; grid on; hold on;set(gcf,'color',[1 1 1]); +plot([0 4 4 2.5 2.5 4 4 0 0 1.5 1.5 0 0],[4 4 2 2 -2 -2 -4 -4 -2 -2 2 2 4],'linewidth',2) + + diff --git a/wing_scripts/get_dp.m b/wing_scripts/get_dp.m new file mode 100644 index 0000000..2a3281d --- /dev/null +++ b/wing_scripts/get_dp.m @@ -0,0 +1,4 @@ +function z = get_dp(xDist,zDist,Vx,Vz,Ix,Iz,Ixz,A) + +denom = (Ix*Iz-Ixz^2); +z = -A*xDist*(Ix*Vx-Ixz*Vz)/denom - A*zDist*(Iz*Vz-Ixz*Vx)/denom; diff --git a/wing_scripts/get_ds.m b/wing_scripts/get_ds.m new file mode 100644 index 0000000..2f0eb9d --- /dev/null +++ b/wing_scripts/get_ds.m @@ -0,0 +1,20 @@ +function ds = get_ds(xi,xf,u) + +dist = 0; +numSteps = 10; +dx = (xf-xi)/numSteps; +z0 = get_z(xi,u); +x0 = xi; +for i=1:10 + tempX = x0+dx; + if tempX > 0 + tempZ = get_z(tempX,u); + else + tempZ = 0; + end + dist = dist + (dx^2+(tempZ-z0)^2)^.5; + z0 = tempZ; + x0 = tempX; +end + +ds =dist; \ No newline at end of file diff --git a/wing_scripts/get_int.m b/wing_scripts/get_int.m new file mode 100644 index 0000000..edbfda3 --- /dev/null +++ b/wing_scripts/get_int.m @@ -0,0 +1,35 @@ +function z = get_int(xi,xf,u) + +M = 0.02; +P = 0.4; +T = 0.12; +a0 = 0.2969; +a1 = -0.126; +a2 = -0.3516; +a3 = 0.2843; +a4 = -0.1015; + + +%evaluate the integral of camber line, depending on xi and xf related to P + +if xf