"""
The evaluator.py module contains a single Evaluator class,
which knows all the attributes of a specified Aircraft 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
import logging

logging.basicConfig(filename='log_eval.txt',
                    level=logging.DEBUG,
                    format='%(asctime)s - %(levelname)s - %(message)s')


class Evaluator:
    """Performs structural evaluations on aircrafts.
    Individual aircrafts must claim an Evaluator object as parent."""
    def __init__(self, name):
        self.name = name
        self.aircrafts = []
        self.results = []

        self.I_ = {'x': 0, 'z': 0, 'xz': 0}

    def get_lift_rectangular(aircraft, lift=50):
        # L_prime = [
        #     lift / (aircraft.wing.semi_span * 2)
        #     for x in range(aircraft.wing.semi_span)
        # ]
        return lift

    def get_lift_elliptical(aircraft, L_0=3.2):
        # L_prime = [
        #     L_0 / (aircraft.wing.semi_span * 2) *
        #     sqrt(1 - (y / aircraft.wing.semi_span)**2)
        #     for y in range(aircraft.wing.semi_span)
        # ]
        return L_0

    def get_lift_total(self, aircraft):
        F_z = 100
        # F_z = [(self.get_lift_rectangular() + self.get_lift_elliptical() / 2
        #        for _ in range(len(aircraft.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(aircraft, drag):
        # Transform semi-span integer into list
        semi_span = [x for x in range(0, aircraft.wing.semi_span)]

        # Drag increases after 80% of the semi_span
        cutoff = round(0.8 * aircraft.wing.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(aircraft):
        """Return the coordinates of the centroid."""
        stringer_area = aircraft.stringer.area
        cap_area = aircraft.spar.cap_area

        caps_x = [value for spar in aircraft.spar.x for value in spar]
        caps_z = [value for spar in aircraft.spar.z for value in spar]
        stringers_x = aircraft.stringer.x
        stringers_z = aircraft.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):
        """Perform all analysis calculations and store in self.results."""
        for aircraft in self.aircrafts:
            results = {"Lift": 400, "Drag": 20, "Centroid": [0.2, 4.5]}
            self.results.append(results)
        #     results = {
        #         "Lift": self.get_lift_total(aircraft),
        #         "Drag": self.get_drag(aircraft.wing),
        #         "Centroid": self.get_centroid(aircraft.wing)
        #     }
        return results

    # 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)
    #     print("yayyyyy")
    #     return None

    # print(f"Analysis results for {aircraft.name}:\n", results)
    # self.results = self.get_lift_total(aircraft)

    # 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 tree_print(self):
        """Print a list of subcomponents."""
        name = f"    COMPONENT TREE FOR {[_.name for _ in self.aircrafts]}    "
        num_of_dashes = len(name)
        print(num_of_dashes * '-')
        print(name)
        for aircraft in self.aircrafts:
            print(".")
            print(f"`-- {aircraft}")
            print(f"    |--{aircraft.wing}")
            for spar in aircraft.wing.spars[:-1]:
                print(f"    |   |-- {spar}")
            print(f"    |   `-- {aircraft.wing.spars[-1]}")
            print(f"    |-- {aircraft.fuselage}")
            print(f"    `-- {aircraft.propulsion}")
        print(num_of_dashes * '-')
        return None

    def tree_save(self,
                  save_path='/home/blendux/Projects/Aircraft_Studio/save'):
        """Save the evaluator's tree to a file."""
        file_name = f"{self.name}_tree.txt"
        full_path = os.path.join(save_path, file_name)
        with open(full_path, 'w') as f:
            try:
                for aircraft in self.aircrafts:
                    f.write(".\n")
                    f.write(f"`-- {aircraft}\n")
                    f.write(f"    |--{aircraft.wing}\n")
                    for spar in aircraft.wing.spars[:-1]:
                        f.write(f"    |   |-- {spar}\n")
                    f.write(f"    |   `-- {aircraft.wing.spars[-1]}\n")
                    f.write(f"    |-- {aircraft.fuselage}\n")
                    f.write(f"    `-- {aircraft.propulsion}\n")
                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
        try:
            with open(full_path, 'w') as 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

    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


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