"""
Estimation of wing 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 numpy as np
from openmdao.core.explicitcomponent import ExplicitComponent
[docs]class ComputeWingCG(ExplicitComponent):
# TODO: Document equations. Cite sources
""" Wing center of gravity estimation """
[docs] def setup(self):
self.add_input("data:geometry:wing:kink:span_ratio", val=np.nan)
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:span", 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_output("data:weight:airframe:wing:CG:x", units="m")
[docs] def setup_partials(self):
self.declare_partials("data:weight:airframe:wing:CG:x", "*", method="fd")
[docs] def compute(self, inputs, outputs):
wing_break = inputs["data:geometry:wing:kink:span_ratio"]
front_spar_ratio_root = inputs["data:geometry:wing:spar_ratio:front:root"]
front_spar_ratio_middle = 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_middle = inputs["data:geometry:wing:spar_ratio:rear:kink"]
rear_spar_ratio_tip = inputs["data:geometry:wing:spar_ratio:rear:tip"]
span = inputs["data:geometry:wing:span"]
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"]
# TODO: make this constant an option
if wing_break >= 0.35:
y_cg = span / 2 * 0.35
l_cg = (y3_wing - y_cg) / (y3_wing - y2_wing) * (l2_wing - l3_wing) + l3_wing
front_spar_cg = (y3_wing - y_cg) / (y3_wing - y2_wing) * (
l2_wing * front_spar_ratio_root - l3_wing * front_spar_ratio_middle
) + l3_wing * front_spar_ratio_middle
rear_spar_cg = (y3_wing - y_cg) / (y3_wing - y2_wing) * (
l2_wing * rear_spar_ratio_root - l3_wing * rear_spar_ratio_middle
) + l3_wing * rear_spar_ratio_middle
x_cg = (
(y_cg - y2_wing) / (y3_wing - y2_wing) * x3_wing
+ front_spar_cg
+ (l_cg - front_spar_cg - rear_spar_cg) * 0.7
)
elif wing_break < 0.35:
y_cg = span / 2 * 0.35
l_cg = (y4_wing - y_cg) / (y4_wing - y3_wing) * (l3_wing - l4_wing) + l4_wing
front_spar_cg = (y4_wing - y_cg) / (y4_wing - y3_wing) * (
l3_wing * front_spar_ratio_middle - l4_wing * front_spar_ratio_tip
) + l4_wing * front_spar_ratio_tip
rear_spar_cg = (y4_wing - y_cg) / (y4_wing - y3_wing) * (
l3_wing * rear_spar_ratio_middle - l4_wing * rear_spar_ratio_tip
) + l4_wing * rear_spar_ratio_tip
x_cg = (
(y_cg - y3_wing) / (y4_wing - y3_wing) * x4_wing
+ front_spar_cg
+ (l_cg - front_spar_cg - rear_spar_cg) * 0.7
)
x_cg_absolute = fa_length - 0.25 * x0_wing + (x_cg - x0_wing)
outputs["data:weight:airframe:wing:CG:x"] = x_cg_absolute