"""Estimation of aerodynamic center."""
# This file is part of FAST-OAD_CS23 : A framework for rapid Overall Aircraft Design
# Copyright (C) 2022 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
from stdatm import Atmosphere
[docs]class ComputeAeroCenter(ExplicitComponent):
# TODO: Document equations. Cite sources
"""Aerodynamic center estimation."""
[docs] def setup(self):
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input(
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25", val=np.nan, units="m"
)
self.add_input("data:aerodynamics:wing:cruise:CL_alpha", val=np.nan, units="rad**-1")
self.add_input(
"data:aerodynamics:horizontal_tail:cruise:CL_alpha", val=np.nan, units="rad**-1"
)
self.add_input("data:aerodynamics:horizontal_tail:cruise:downwash_gradient", val=np.nan)
self.add_input("data:aerodynamics:elevator:low_speed:CL_delta", val=np.nan, units="rad**-1")
self.add_input("data:aerodynamics:fuselage:cm_alpha", val=np.nan, units="rad**-1")
self.add_input(
"data:aerodynamics:horizontal_tail:cruise:hinge_moment:CH_alpha",
val=np.nan,
units="rad**-1",
)
self.add_input(
"data:aerodynamics:horizontal_tail:cruise:hinge_moment:CH_delta",
val=np.nan,
units="rad**-1",
)
self.add_input("data:aerodynamics:horizontal_tail:efficiency", val=np.nan)
self.add_input("data:TLAR:v_cruise", val=np.nan, units="m/s")
self.add_input("data:mission:sizing:main_route:cruise:altitude", val=np.nan, units="ft")
self.add_output("data:aerodynamics:cruise:neutral_point:stick_fixed:x")
self.add_output("data:aerodynamics:cruise:neutral_point:stick_free:x")
self.add_output("data:aerodynamics:cruise:neutral_point:free_elevator_factor")
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
l0_wing = inputs["data:geometry:wing:MAC:length"]
lp_ht = inputs["data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25"]
cl_alpha_wing = inputs["data:aerodynamics:wing:cruise:CL_alpha"]
cl_alpha_ht = inputs["data:aerodynamics:horizontal_tail:cruise:CL_alpha"]
downwash_gradient = inputs["data:aerodynamics:horizontal_tail:cruise:downwash_gradient"]
cl_delta_ht = inputs["data:aerodynamics:elevator:low_speed:CL_delta"]
ch_alpha_3d = inputs["data:aerodynamics:horizontal_tail:cruise:hinge_moment:CH_alpha"]
ch_delta_3d = inputs["data:aerodynamics:horizontal_tail:cruise:hinge_moment:CH_delta"]
cm_alpha_fus = inputs["data:aerodynamics:fuselage:cm_alpha"]
tail_efficiency = inputs["data:aerodynamics:horizontal_tail:efficiency"]
v_cruise = inputs["data:TLAR:v_cruise"]
alt_cruise = inputs["data:mission:sizing:main_route:cruise:altitude"]
# TODO: make variable name in computation sequence more english
# FIXME: introduce cm_alpha_wing to the equation (non-symmetrical profile)
x_ca_plane = (tail_efficiency * cl_alpha_ht * lp_ht - cm_alpha_fus * l0_wing) / (
cl_alpha_wing + tail_efficiency * cl_alpha_ht
)
x_aero_center = x_ca_plane / l0_wing + 0.25
outputs["data:aerodynamics:cruise:neutral_point:stick_fixed:x"] = x_aero_center
sos = Atmosphere(alt_cruise).speed_of_sound
mach = v_cruise / sos
beta = np.sqrt(1.0 - mach**2.0)
cl_delta_ht_cruise = cl_delta_ht / beta
# The cl_alpha_ht in the formula for the free_elevator_factor is defined with respect to
# the tail angle of attack, the one we compute is wth respect to the plane so it includes
# downwash, as a consequence we must correct it influence for this specific calculation.
# We will use the formula for elliptical wing as it is well known
downwash_effect = 1.0 - downwash_gradient
cl_alpha_ht_ht = cl_alpha_ht / downwash_effect
free_elevator_factor = 1.0 - (cl_delta_ht_cruise / cl_alpha_ht_ht) * (
ch_alpha_3d / ch_delta_3d
)
outputs["data:aerodynamics:cruise:neutral_point:free_elevator_factor"] = (
free_elevator_factor
)
x_ca_plane_free = (
tail_efficiency * free_elevator_factor * cl_alpha_ht * lp_ht - cm_alpha_fus * l0_wing
) / (cl_alpha_wing + tail_efficiency * free_elevator_factor * cl_alpha_ht)
x_aero_center_free = x_ca_plane_free / l0_wing + 0.25
outputs["data:aerodynamics:cruise:neutral_point:stick_free:x"] = x_aero_center_free