"""Computation of CL characteristics at low speed."""
# This file is part of FAST-OAD_CS25
# Copyright (C) 2023 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 fastoad.api as oad
import numpy as np
import openmdao.api as om
from ..constants import SERVICE_CL_ALPHA
[docs]
@oad.RegisterSubmodel(SERVICE_CL_ALPHA, "fastoad.submodel.aerodynamics.CLalpha.legacy")
class ComputeCLAlpha(om.ExplicitComponent):
"""
Computes CL gradient.
CL gradient from :cite:`raymer:1999` Eq 12.6
"""
[docs]
def initialize(self):
self.options.declare("low_speed_aero", default=False, types=bool)
[docs]
def setup(self):
if not self.options["low_speed_aero"]:
self.add_input("data:TLAR:cruise_mach", val=np.nan, units="unitless")
self.add_input("data:geometry:fuselage:maximum_width", val=np.nan, units="m")
self.add_input("data:geometry:fuselage:maximum_height", val=np.nan, units="m")
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:wing:aspect_ratio", val=np.nan, units="unitless")
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="rad")
self.add_input("data:geometry:wing:root:chord", val=np.nan, units="m")
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:tip:thickness_ratio", val=np.nan, units="unitless")
if self.options["low_speed_aero"]:
self.add_output("data:aerodynamics:aircraft:low_speed:CL_alpha", units="1/rad")
else:
self.add_output("data:aerodynamics:aircraft:high_speed:CL_alpha", units="1/rad")
[docs]
def setup_partials(self):
if self.options["low_speed_aero"]:
self.declare_partials("data:aerodynamics:aircraft:low_speed:CL_alpha", "*", method="fd")
else:
self.declare_partials(
"data:aerodynamics:aircraft:high_speed:CL_alpha", "*", method="fd"
)
[docs]
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
width_max = inputs["data:geometry:fuselage:maximum_width"]
height_max = inputs["data:geometry:fuselage:maximum_height"]
span = inputs["data:geometry:wing:span"]
lambda_wing = inputs["data:geometry:wing:aspect_ratio"]
root_chord = inputs["data:geometry:wing:root:chord"]
tip_chord = inputs["data:geometry:wing:tip:chord"]
tip_thickness_ratio = inputs["data:geometry:wing:tip:thickness_ratio"]
sweep_25 = inputs["data:geometry:wing:sweep_25"]
wing_area = inputs["data:geometry:wing:area"]
if self.options["low_speed_aero"]:
mach = 0.2
else:
mach = inputs["data:TLAR:cruise_mach"]
beta = np.sqrt(1.0 - mach**2)
d_f = np.sqrt(width_max * height_max)
fuselage_lift_factor = 1.07 * (1.0 + d_f / span) ** 2
lambda_wing_eff = lambda_wing * (1.0 + 1.9 * tip_chord * tip_thickness_ratio / span)
cl_alpha_wing = (
2.0
* np.pi
* lambda_wing_eff
/ (
2.0
+ np.sqrt(
4.0
+ lambda_wing_eff**2
* beta**2
/ 0.9025 # equals 0.95**2
* (1.0 + (np.tan(sweep_25)) ** 2 / beta**2)
)
)
* (wing_area - root_chord * width_max)
/ wing_area
* fuselage_lift_factor
)
if self.options["low_speed_aero"]:
outputs["data:aerodynamics:aircraft:low_speed:CL_alpha"] = cl_alpha_wing
else:
outputs["data:aerodynamics:aircraft:high_speed:CL_alpha"] = cl_alpha_wing