"""
Estimation of vertical tail 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
[docs]class ComputeVTcg(ExplicitComponent):
# TODO: Document equations. Cite sources
""" Vertical tail center of gravity estimation """
[docs] def setup(self):
self.add_input("data:geometry:vertical_tail:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:vertical_tail:root:chord", val=np.nan, units="m")
self.add_input("data:geometry:vertical_tail:tip:chord", val=np.nan, units="m")
self.add_input(
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25", val=np.nan, units="m"
)
self.add_input("data:geometry:vertical_tail:MAC:at25percent:x:local", val=np.nan, units="m")
self.add_input("data:geometry:vertical_tail:sweep_25", val=np.nan, units="deg")
self.add_input("data:geometry:vertical_tail:span", 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:vertical_tail:CG:x", units="m")
[docs] def setup_partials(self):
self.declare_partials("data:weight:airframe:vertical_tail:CG:x", "*", method="fd")
[docs] def compute(self, inputs, outputs):
root_chord = inputs["data:geometry:vertical_tail:root:chord"]
tip_chord = inputs["data:geometry:vertical_tail:tip:chord"]
lp_vt = inputs["data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25"]
mac_vt = inputs["data:geometry:vertical_tail:MAC:length"]
fa_length = inputs["data:geometry:wing:MAC:at25percent:x"]
x0_vt = inputs["data:geometry:vertical_tail:MAC:at25percent:x:local"]
sweep_25_vt = inputs["data:geometry:vertical_tail:sweep_25"]
b_v = inputs["data:geometry:vertical_tail:span"]
tmp = root_chord * 0.25 + b_v * math.tan(sweep_25_vt / 180.0 * math.pi) - tip_chord * 0.25
l_cg_vt = (1 - 0.55) * (root_chord - tip_chord) + tip_chord
x_cg_vt = 0.42 * l_cg_vt + 0.55 * tmp
x_cg_vt_absolute = lp_vt + fa_length - 0.25 * mac_vt + (x_cg_vt - x0_vt)
outputs["data:weight:airframe:vertical_tail:CG:x"] = x_cg_vt_absolute