"""Estimation of the wing profile drag."""
# 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
import fastoad.api as oad
from openmdao.core.explicitcomponent import ExplicitComponent
from fastga.models.geometry.profiles.get_profile import get_profile
from ..constants import SUBMODEL_CD0_WING
[docs]@oad.RegisterSubmodel(SUBMODEL_CD0_WING, "fastga.submodel.aerodynamics.wing.cd0.legacy")
class Cd0Wing(ExplicitComponent):
"""
Profile drag estimation for the wing
Based on : Gudmundsson, Snorri. General aviation aircraft design: Applied Methods and
Procedures. Butterworth-Heinemann, 2013.
"""
[docs] def initialize(self):
self.options.declare("low_speed_aero", default=False, types=bool)
self.options.declare("airfoil_folder_path", default=None, types=str, allow_none=True)
self.options.declare(
"wing_airfoil_file", default="naca23012.af", types=str, allow_none=True
)
[docs] def setup(self):
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:fuselage:maximum_width", val=np.nan, units="m")
self.add_input("data:geometry:wing:root:y", val=np.nan, units="m")
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:wing:sweep_25", val=np.nan, units="deg")
self.add_input("data:geometry:wing:wet_area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:thickness_ratio", val=np.nan)
self.add_input("data:geometry:propeller:diameter", val=np.nan, units="m")
self.add_input("data:geometry:propulsion:engine:layout", val=np.nan)
if self.options["low_speed_aero"]:
self.add_input("data:aerodynamics:low_speed:mach", val=np.nan)
self.add_input("data:aerodynamics:low_speed:unit_reynolds", val=np.nan, units="m**-1")
self.add_output("data:aerodynamics:wing:low_speed:CD0")
else:
self.add_input("data:aerodynamics:cruise:mach", val=np.nan)
self.add_input("data:aerodynamics:cruise:unit_reynolds", val=np.nan, units="m**-1")
self.add_output("data:aerodynamics:wing:cruise:CD0")
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
l2_wing = inputs["data:geometry:wing:root:chord"]
l4_wing = inputs["data:geometry:wing:tip:chord"]
y1_wing = inputs["data:geometry:fuselage:maximum_width"] / 2.0
y2_wing = inputs["data:geometry:wing:root:y"]
span = inputs["data:geometry:wing:span"]
sweep_25 = inputs["data:geometry:wing:sweep_25"]
wet_area_wing = inputs["data:geometry:wing:wet_area"]
wing_area = inputs["data:geometry:wing:area"]
thickness = inputs["data:geometry:wing:thickness_ratio"]
prop_dia = inputs["data:geometry:propeller:diameter"]
engine_layout = inputs["data:geometry:propulsion:engine:layout"]
if self.options["low_speed_aero"]:
mach = inputs["data:aerodynamics:low_speed:mach"]
unit_reynolds = inputs["data:aerodynamics:low_speed:unit_reynolds"]
else:
mach = inputs["data:aerodynamics:cruise:mach"]
unit_reynolds = inputs["data:aerodynamics:cruise:unit_reynolds"]
# Sear max thickness position ratio
profile = get_profile(
airfoil_folder_path=self.options["airfoil_folder_path"],
file_name=self.options["wing_airfoil_file"],
)
relative_thickness = profile.get_relative_thickness()
index = int(
np.where(relative_thickness["thickness"] == np.max(relative_thickness["thickness"]))[0]
)
x_t_max = relative_thickness["x"][index]
# Root: 45% NLF
x_trans = 0.45
x0_turbulent = 36.9 * x_trans**0.625 * (1 / (unit_reynolds * l2_wing)) ** 0.375
cf_root = 0.074 / (unit_reynolds * l2_wing) ** 0.2 * (1 - (x_trans - x0_turbulent)) ** 0.8
# Tip: 55% NLF
x_trans = 0.55
x0_turbulent = 36.9 * x_trans**0.625 * (1 / (unit_reynolds * l4_wing)) ** 0.375
cf_tip = 0.074 / (unit_reynolds * l4_wing) ** 0.2 * (1 - (x_trans - x0_turbulent)) ** 0.8
if engine_layout == 1.0:
# Wing fully turbulent behind the propeller
cf_turbulent = 0.074 / (unit_reynolds * l2_wing) ** 0.2
cf_wing = (
cf_turbulent * prop_dia
+ (cf_root + cf_tip) / 2.0 * (span / 2.0 - (y1_wing + prop_dia))
) / (span / 2.0 - y1_wing)
else:
# Global
cf_wing = (
cf_root * (y2_wing - y1_wing) + 0.5 * (span / 2.0 - y2_wing) * (cf_root + cf_tip)
) / (span / 2.0 - y1_wing)
form_factor = 1 + 0.6 / x_t_max * thickness + 100 * thickness**4
if mach > 0.2:
form_factor = form_factor * 1.34 * mach**0.18 * (np.cos(sweep_25 * np.pi / 180)) ** 0.28
cd0_wing = form_factor * cf_wing * wet_area_wing / wing_area
if self.options["low_speed_aero"]:
outputs["data:aerodynamics:wing:low_speed:CD0"] = cd0_wing
else:
outputs["data:aerodynamics:wing:cruise:CD0"] = cd0_wing