"""
Estimation of aerodynamic center
"""
# 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 ComputeAeroCenter(ExplicitComponent):
# TODO: Document equations. Cite sources
""" Aerodynamic center estimation """
[docs] def setup(self):
self.add_input("data:geometry:wing:MAC:leading_edge:x:local", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:wing:root:virtual_chord", val=np.nan, units="m")
self.add_input("data:geometry:fuselage:maximum_width", val=np.nan, units="m")
self.add_input("data:geometry:fuselage:length", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:at25percent:x", val=np.nan, units="m")
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:horizontal_tail:area", val=np.nan, units="m**2")
self.add_input(
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25", val=np.nan, units="m"
)
self.add_input("data:aerodynamics:aircraft:cruise:CL_alpha", val=np.nan)
self.add_input("data:aerodynamics:horizontal_tail:cruise:CL_alpha", val=np.nan)
self.add_output("data:aerodynamics:cruise:neutral_point:x")
[docs] def setup_partials(self):
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs):
x0_wing = inputs["data:geometry:wing:MAC:leading_edge:x:local"]
l0_wing = inputs["data:geometry:wing:MAC:length"]
l1_wing = inputs["data:geometry:wing:root:virtual_chord"]
width_max = inputs["data:geometry:fuselage:maximum_width"]
fa_length = inputs["data:geometry:wing:MAC:at25percent:x"]
fus_length = inputs["data:geometry:fuselage:length"]
wing_area = inputs["data:geometry:wing:area"]
s_h = inputs["data:geometry:horizontal_tail:area"]
lp_ht = inputs["data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25"]
cl_alpha_wing = inputs["data:aerodynamics:aircraft:cruise:CL_alpha"]
cl_alpha_ht = inputs["data:aerodynamics:horizontal_tail:cruise:CL_alpha"]
# TODO: make variable name is computation sequence more english
x0_25 = fa_length - 0.25 * l0_wing - x0_wing + 0.25 * l1_wing
ratio_x025 = x0_25 / fus_length
# fitting result of Raymer book, figure 16.14
k_h = 0.01222 - 7.40541e-4 * ratio_x025 * 100 + 2.1956e-5 * (ratio_x025 * 100) ** 2
# equation from Raymer book, eqn 16.22
cm_alpha_fus = k_h * width_max ** 2 * fus_length / (l0_wing * wing_area) * 180.0 / np.pi
x_ca_plane = (
cl_alpha_wing * fa_length / l0_wing
- cm_alpha_fus
+ cl_alpha_ht * (1 - 0.4) * 0.9 * s_h / wing_area * (lp_ht + fa_length) / l0_wing
) / (cl_alpha_wing + cl_alpha_ht * (1 - 0.4) * 0.9 * s_h / wing_area)
x_aero_center = x_ca_plane - fa_length / l0_wing + 0.25
outputs["data:aerodynamics:cruise:neutral_point:x"] = x_aero_center