"""Computation of CL characteristics at low speed."""
# 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
import openmdao.api as om
from fastoad.module_management.service_registry import RegisterSubmodel
from ..constants import SERVICE_LOW_SPEED_CL_AOA
[docs]@RegisterSubmodel(SERVICE_LOW_SPEED_CL_AOA, "fastoad.submodel.aerodynamics.low_speed.AoA.legacy")
class ComputeAerodynamicsLowSpeed(om.ExplicitComponent):
"""
Computes CL gradient and CL at low speed.
CL gradient from :cite:`raymer:1999` Eq 12.6
"""
# TODO: complete source
[docs] def setup(self):
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)
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")
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)
self.add_output("data:aerodynamics:aircraft:takeoff:CL_alpha", units="1/rad")
self.add_output("data:aerodynamics:aircraft:takeoff:CL0_clean")
[docs] def setup_partials(self):
self.declare_partials("*", "*", 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"]
l2_wing = inputs["data:geometry:wing:root:chord"]
l4_wing = inputs["data:geometry:wing:tip:chord"]
el_ext = inputs["data:geometry:wing:tip:thickness_ratio"]
sweep_25 = inputs["data:geometry:wing:sweep_25"]
wing_area = inputs["data:geometry:wing:area"]
mach = 0.2
beta = np.sqrt(1 - mach ** 2)
d_f = np.sqrt(width_max * height_max)
fuselage_lift_factor = 1.07 * (1 + d_f / span) ** 2
lambda_wing_eff = lambda_wing * (1 + 1.9 * l4_wing * el_ext / span)
cl_alpha_wing_low = (
2
* np.pi
* lambda_wing_eff
/ (
2
+ np.sqrt(
4
+ lambda_wing_eff ** 2
* beta ** 2
/ 0.95 ** 2
* (1 + (np.tan(sweep_25 / 180.0 * np.pi)) ** 2 / beta ** 2)
)
)
* (wing_area - l2_wing * width_max)
/ wing_area
* fuselage_lift_factor
)
outputs["data:aerodynamics:aircraft:takeoff:CL_alpha"] = cl_alpha_wing_low
outputs["data:aerodynamics:aircraft:takeoff:CL0_clean"] = 0.2 # FIXME: hard-coded value