"""
Compressibility drag computation.
"""
# 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 CdCompressibility(ExplicitComponent):
"""
Computation of drag increment due to compressibility effects.
Formula from §4.2.4 of :cite:`supaero:2014`. This formula can be used for aircraft
before year 2000.
Earlier aircraft have more optimized wing profiles that are expected to limit the
compressibility drag below 2 drag counts. Until a better model can be provided, the
variable `tuning:aerodynamics:aircraft:cruise:CD:compressibility:characteristic_mach_increment`
allows to move the characteristic Mach number, thus moving the CD divergence to higher
Mach numbers.
"""
[docs] def setup(self):
self.add_input("data:TLAR:cruise_mach", val=np.nan)
self.add_input("data:aerodynamics:aircraft:cruise:CL", shape_by_conn=True, val=np.nan)
self.add_input("data:geometry:wing:sweep_25", units="deg", val=np.nan)
self.add_input("data:geometry:wing:thickness_ratio", val=np.nan)
self.add_input("tuning:aerodynamics:aircraft:cruise:CD:compressibility:max_value", val=0.5)
self.add_input(
"tuning:aerodynamics:aircraft:cruise:CD:compressibility:characteristic_mach_increment",
val=0.0,
)
self.add_output(
"data:aerodynamics:aircraft:cruise:CD:compressibility",
copy_shape="data:aerodynamics:aircraft:cruise:CL",
)
[docs] def setup_partials(self):
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs):
cl = inputs["data:aerodynamics:aircraft:cruise:CL"]
m = inputs["data:TLAR:cruise_mach"]
max_cd_comp = inputs["tuning:aerodynamics:aircraft:cruise:CD:compressibility:max_value"]
delta_m_charac_0 = inputs[
"tuning:aerodynamics:aircraft:cruise:CD:compressibility:characteristic_mach_increment"
]
sweep_angle = inputs["data:geometry:wing:sweep_25"]
thickness_ratio = inputs["data:geometry:wing:thickness_ratio"]
# Computation of characteristic Mach for 28° sweep and 0.12 of relative thickness
m_charac_comp_0 = (
-0.5 * np.maximum(0.35, cl) ** 2
+ 0.35 * np.maximum(0.35, cl)
+ 0.765
+ delta_m_charac_0
)
# Computation of characteristic Mach for actual sweep angle and relative thickness
m_charac_comp = (
m_charac_comp_0 * np.cos(np.radians(28)) + 0.12 - thickness_ratio
) / np.cos(np.radians(sweep_angle))
cd_comp = np.minimum(max_cd_comp, 0.002 * np.exp(42.58 * (m - m_charac_comp)))
outputs["data:aerodynamics:aircraft:cruise:CD:compressibility"] = cd_comp