"""
Estimation of tanks center of gravity
"""
# This file is part of FAST-OAD : A framework for rapid Overall Aircraft Design
# Copyright (C) 2021 ONERA & ISAE-SUPAERO
# FAST 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 math
import numpy as np
from openmdao.core.explicitcomponent import ExplicitComponent
from scipy import interpolate
from fastoad.models.geometry.profiles import get_profile
[docs]class ComputeTanksCG(ExplicitComponent):
# TODO: Document equations. Cite sources
""" Tanks center of gravity estimation """
[docs] def initialize(self):
self.options.declare("ratio", default=0.2, types=float)
self.ratio = self.options["ratio"]
[docs] def setup(self):
self.add_input("data:geometry:wing:spar_ratio:front:root", val=np.nan)
self.add_input("data:geometry:wing:spar_ratio:front:kink", val=np.nan)
self.add_input("data:geometry:wing:spar_ratio:front:tip", val=np.nan)
self.add_input("data:geometry:wing:spar_ratio:rear:root", val=np.nan)
self.add_input("data:geometry:wing:spar_ratio:rear:kink", val=np.nan)
self.add_input("data:geometry:wing:spar_ratio:rear:tip", val=np.nan)
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:leading_edge:x:local", val=np.nan, units="m")
self.add_input("data:geometry:wing:root:chord", val=np.nan, units="m")
self.add_input("data:geometry:wing:kink:chord", val=np.nan, units="m")
self.add_input("data:geometry:wing:tip:chord", val=np.nan, units="m")
self.add_input("data:geometry:wing:root:y", val=np.nan, units="m")
self.add_input("data:geometry:wing:kink:leading_edge:x:local", val=np.nan, units="m")
self.add_input("data:geometry:wing:kink:y", val=np.nan, units="m")
self.add_input("data:geometry:wing:tip:leading_edge:x:local", val=np.nan, units="m")
self.add_input("data:geometry:wing:tip:y", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:at25percent:x", val=np.nan, units="m")
self.add_input("data:geometry:fuselage:maximum_width", val=np.nan, units="m")
self.add_output("data:weight:fuel_tank:CG:x", units="m")
[docs] def setup_partials(self):
self.declare_partials("data:weight:fuel_tank:CG:x", "*", method="fd")
[docs] def compute(self, inputs, outputs):
# TODO: decompose into a function to make the code more clear
front_spar_ratio_root = inputs["data:geometry:wing:spar_ratio:front:root"]
front_spar_ratio_kink = inputs["data:geometry:wing:spar_ratio:front:kink"]
front_spar_ratio_tip = inputs["data:geometry:wing:spar_ratio:front:tip"]
rear_spar_ratio_root = inputs["data:geometry:wing:spar_ratio:rear:root"]
rear_spar_ratio_kink = inputs["data:geometry:wing:spar_ratio:rear:kink"]
rear_spar_ratio_tip = inputs["data:geometry:wing:spar_ratio:rear:tip"]
l0_wing = inputs["data:geometry:wing:MAC:length"]
x0_wing = inputs["data:geometry:wing:MAC:leading_edge:x:local"]
l2_wing = inputs["data:geometry:wing:root:chord"]
l3_wing = inputs["data:geometry:wing:kink:chord"]
l4_wing = inputs["data:geometry:wing:tip:chord"]
y2_wing = inputs["data:geometry:wing:root:y"]
x3_wing = inputs["data:geometry:wing:kink:leading_edge:x:local"]
y3_wing = inputs["data:geometry:wing:kink:y"]
y4_wing = inputs["data:geometry:wing:tip:y"]
x4_wing = inputs["data:geometry:wing:tip:leading_edge:x:local"]
fa_length = inputs["data:geometry:wing:MAC:at25percent:x"]
width_max = inputs["data:geometry:fuselage:maximum_width"]
profile = get_profile("airfoil_f_15_15.txt", chord_length=1.0)
height_root_front, height_root_rear = self._get_thickness(
profile, l2_wing, [front_spar_ratio_root, rear_spar_ratio_root]
)
profile = get_profile("airfoil_f_15_12.txt", chord_length=1.0)
height_kink_front, height_kink_rear = self._get_thickness(
profile, l3_wing, [front_spar_ratio_root, rear_spar_ratio_root]
)
profile = get_profile("airfoil_f_15_11.txt", chord_length=1.0)
height_tip_front, height_tip_rear = self._get_thickness(
profile, l4_wing, [front_spar_ratio_root, rear_spar_ratio_root]
)
l_root = l2_wing * (rear_spar_ratio_root - front_spar_ratio_root)
l_kink = l3_wing * (rear_spar_ratio_kink - front_spar_ratio_kink)
l_tip = l4_wing * (rear_spar_ratio_tip - front_spar_ratio_tip)
s_root = (height_root_front + height_root_rear) * l_root / 2
s_kink = (height_kink_front + height_kink_rear) * l_kink / 2
s_tip = (height_tip_front + height_tip_rear) * l_tip / 2
vol_central = s_root * width_max
vol_side_inner = (
(s_root + s_kink + math.sqrt(s_root * s_kink)) * (y3_wing - y2_wing) * 2 / 3
)
# Assume in the region outward 0.8 of the span, no tank installed
ratio_1 = y4_wing * self.ratio / (y4_wing - y3_wing)
s_real_tip = (ratio_1 * (s_kink ** 0.5 - s_tip ** 0.5) + s_tip ** 0.5) ** 2
vol_side_out = (
(s_kink + s_real_tip + math.sqrt(s_kink * s_real_tip))
* (y4_wing - y3_wing)
* (1 - ratio_1)
* 2
/ 3
)
# vol_side = vol_side_inner + vol_side_out
# assume only 80% of the total volume can be used for fuel, based on Raymer(0.85),
# for central tank, the ratio is 0.92 fuel density is 785kg/m3
# total_mass = (vol_side * 0.8 + vol_central * 0.92) * 785
#######################################################################
# CG calculation of tank
x_cg_central = (
l_root
/ 3
* (height_root_front + 2 * height_root_rear)
/ (height_root_front + height_root_rear)
+ l2_wing * front_spar_ratio_root
)
x_cg_central_absolute = fa_length - 0.25 * l0_wing - x0_wing + x_cg_central
y_side_inner_cg = (
(y3_wing - y2_wing)
/ 4
* (s_root + 3 * s_kink + 2 * math.sqrt(s_root * s_kink))
/ (s_root + s_kink + math.sqrt(s_root * s_kink))
)
height_side_inner_cg_front = (y_side_inner_cg / (y3_wing - y2_wing)) * (
height_kink_front - height_root_front
) + height_root_front
height_side_inner_cg_rear = (y_side_inner_cg / (y3_wing - y2_wing)) * (
height_kink_rear - height_root_rear
) + height_root_rear
l_side_inner_cg = (y_side_inner_cg / (y3_wing - y2_wing)) * (l_kink - l_root) + l_root
x_side_inner_front = (y_side_inner_cg / (y3_wing - y2_wing)) * (
x3_wing + l3_wing * front_spar_ratio_kink - l2_wing * front_spar_ratio_root
) + l2_wing * front_spar_ratio_root
x_cg_side_inner = x_side_inner_front + l_side_inner_cg / 3 * (
height_side_inner_cg_front + 2 * height_side_inner_cg_rear
) / (height_side_inner_cg_front + height_side_inner_cg_rear)
x_cg_side_inner_absolute = fa_length - 0.25 * l0_wing - x0_wing + x_cg_side_inner
y4_tank = y4_wing * (1 - self.ratio)
x4_tank = (x4_wing - x3_wing) * (1 - y4_wing * self.ratio / (y4_wing - y3_wing)) + x3_wing
l4_tank = l4_wing + (l3_wing - l4_wing) * ratio_1
height_tank_front = height_tip_front + (height_kink_front - height_tip_front) * ratio_1
height_tank_rear = height_tip_rear + (height_kink_rear - height_tip_rear) * ratio_1
l_tank_tip = l_tip + (l_kink - l_tip) * ratio_1
front_spar_ratio_tank = (
front_spar_ratio_tip + (front_spar_ratio_kink - front_spar_ratio_tip) * ratio_1
)
y_side_out_cg = (
(y4_tank - y3_wing)
/ 4
* (s_kink + 3 * s_real_tip + 2 * math.sqrt(s_kink * s_real_tip))
/ (s_kink + s_real_tip + math.sqrt(s_kink * s_real_tip))
)
height_side_out_cg_front = (y_side_out_cg / (y4_tank - y3_wing)) * (
height_tank_front - height_kink_front
) + height_kink_front
height_side_out_cg_rear = (y_side_out_cg / (y4_tank - y3_wing)) * (
height_tank_rear - height_kink_rear
) + height_kink_rear
l_side_out_cg = (y_side_out_cg / (y4_tank - y3_wing)) * (l_tank_tip - l_kink) + l_kink
x_side_out_front = (
(y_side_out_cg / (y4_tank - y3_wing))
* (
x4_tank
+ l4_tank * front_spar_ratio_tank
- x3_wing
- l3_wing * front_spar_ratio_kink
)
+ x3_wing
+ l3_wing * front_spar_ratio_kink
)
x_cg_side_out = x_side_out_front + l_side_out_cg / 3 * (
height_side_out_cg_front + 2 * height_side_out_cg_rear
) / (height_side_out_cg_front + height_side_out_cg_rear)
x_cg_side_out_absolute = fa_length - 0.25 * l0_wing - x0_wing + x_cg_side_out
x_cg_tank = (
vol_side_out * 0.8 * x_cg_side_out_absolute
+ vol_side_inner * 0.8 * x_cg_side_inner_absolute
+ vol_central * 0.92 * x_cg_central_absolute
) / (vol_side_out * 0.8 + vol_side_inner * 0.8 + vol_central * 0.92)
# x_cg_tank=(vol_side_out*0.8*x_cg_side_out+vol_side_inner
# *0.8*x_cg_side_inner+vol_central*0.8*x_cg_central)/
# (vol_side_out*0.8+vol_side_inner*0.8+vol_central*0.8)
outputs["data:weight:fuel_tank:CG:x"] = x_cg_tank
@staticmethod
def _get_thickness(profile, l2_wing, relative_x):
"""
:param profile:
:param l2_wing:
:param relative_x:
:return: profile thickness at provide value(s) of relative_x
"""
relative_thickness = profile.get_relative_thickness()
relative_x_scale = relative_thickness.x
thickness = relative_thickness.thickness * l2_wing
return interpolate.interp1d(relative_x_scale, thickness)(relative_x)