"""
Computation of Oswald coefficient
"""
# 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 OswaldCoefficient(ExplicitComponent):
# TODO: Document equations. Cite sources (M. Nita and D. Scholz)
# FIXME: output the real Oswald coefficient (coef_e instead of coef_k)
""" Computes Oswald efficiency number """
[docs] def initialize(self):
self.options.declare("low_speed_aero", default=False, types=bool)
[docs] def setup(self):
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:fuselage:maximum_height", val=np.nan, units="m")
self.add_input("data:geometry:fuselage:maximum_width", val=np.nan, units="m")
self.add_input("data:geometry:wing:root: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:sweep_25", val=np.nan, units="deg")
if self.options["low_speed_aero"]:
self.add_input("data:aerodynamics:aircraft:takeoff:mach", val=np.nan)
self.add_output("data:aerodynamics:aircraft:low_speed:induced_drag_coefficient")
else:
self.add_input("data:TLAR:cruise_mach", val=np.nan)
self.add_output("data:aerodynamics:aircraft:cruise:induced_drag_coefficient")
[docs] def setup_partials(self):
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs):
wing_area = inputs["data:geometry:wing:area"]
span = inputs["data:geometry:wing:span"] / math.cos(5.0 / 180 * math.pi)
height_fus = inputs["data:geometry:fuselage:maximum_height"]
width_fus = inputs["data:geometry:fuselage:maximum_width"]
l2_wing = inputs["data:geometry:wing:root:chord"]
l4_wing = inputs["data:geometry:wing:tip:chord"]
sweep_25 = inputs["data:geometry:wing:sweep_25"]
if self.options["low_speed_aero"]:
mach = inputs["data:aerodynamics:aircraft:takeoff:mach"]
else:
mach = inputs["data:TLAR:cruise_mach"]
aspect_ratio = span ** 2 / wing_area
df = math.sqrt(width_fus * height_fus)
lamda = l4_wing / l2_wing
delta_lamda = -0.357 + 0.45 * math.exp(0.0375 * sweep_25 / 180.0 * math.pi)
lamda = lamda - delta_lamda
f_lamda = (
0.0524 * lamda ** 4 - 0.15 * lamda ** 3 + 0.1659 * lamda ** 2 - 0.0706 * lamda + 0.0119
)
e_theory = 1 / (1 + f_lamda * aspect_ratio)
if mach <= 0.4:
ke_m = 1.0
else:
ke_m = -0.001521 * ((mach - 0.05) / 0.3 - 1) ** 10.82 + 1
ke_f = 1 - 2 * (df / span) ** 2
coef_e = e_theory * ke_f * ke_m * 0.9
coef_k = 1.0 / (math.pi * aspect_ratio * coef_e)
if self.options["low_speed_aero"]:
outputs["data:aerodynamics:aircraft:low_speed:induced_drag_coefficient"] = coef_k
else:
outputs["data:aerodynamics:aircraft:cruise:induced_drag_coefficient"] = coef_k