# 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 ..figure_digitization import FigureDigitization
from ...constants import SUBMODEL_CL_P_HT
[docs]@oad.RegisterSubmodel(
SUBMODEL_CL_P_HT, "fastga.submodel.aerodynamics.horizontal_tail.roll_moment_roll_rate.legacy"
)
class ComputeClRollRateHorizontalTail(FigureDigitization):
"""
Class to compute the contribution of the horizontal tail to the roll moment coefficient due
to roll rate (roll damping). Depends on the lift coefficient of the horizontal tail,
hence on the reference angle of attack, so the same remark as in ..compute_cy_yaw_rate.py
holds. The convention from :cite:`roskampart6:1985` are used, meaning that for lateral
derivative, the reference length is the wing span. Another important point is that,
for the derivative with respect to yaw and roll, the rotation speed are made dimensionless by
multiplying them by the wing span and dividing them by 2 times the airspeed.
Based on :cite:`roskampart6:1985` section 10.2.6
"""
[docs] def initialize(self):
self.options.declare("low_speed_aero", default=False, types=bool)
[docs] def setup(self):
self.add_input("data:geometry:horizontal_tail:aspect_ratio", val=np.nan)
self.add_input("data:geometry:horizontal_tail:taper_ratio", val=np.nan)
self.add_input("data:geometry:horizontal_tail:sweep_25", val=np.nan, units="rad")
self.add_input("data:geometry:horizontal_tail:dihedral", val=0.0, units="deg")
self.add_input("data:geometry:horizontal_tail:z:from_wingMAC25", val=np.nan, units="m")
self.add_input("data:geometry:wing:root:z", val=np.nan, units="m")
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:horizontal_tail:span", val=np.nan, units="m")
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:horizontal_tail:area", val=np.nan, units="m**2")
self.add_input(
"data:aerodynamics:horizontal_tail:airfoil:CL_alpha", val=np.nan, units="rad**-1"
)
ls_tag = "low_speed" if self.options["low_speed_aero"] else "cruise"
ref_aoa = 5.0 if self.options["low_speed_aero"] else 1.0
self.add_input(
"settings:aerodynamics:reference_flight_conditions:" + ls_tag + ":AOA",
units="rad",
val=ref_aoa * np.pi / 180.0,
)
self.add_input("data:aerodynamics:" + ls_tag + ":mach", val=np.nan)
self.add_input("data:aerodynamics:horizontal_tail:" + ls_tag + ":CL0", val=np.nan)
self.add_input(
"data:aerodynamics:horizontal_tail:" + ls_tag + ":CL_alpha", val=np.nan, units="rad**-1"
)
self.add_input("data:aerodynamics:horizontal_tail:" + ls_tag + ":CD0", val=np.nan)
self.add_output("data:aerodynamics:horizontal_tail:" + ls_tag + ":Cl_p", units="rad**-1")
self.declare_partials(of="*", wrt="*", method="fd")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
ht_ar = inputs["data:geometry:horizontal_tail:aspect_ratio"]
wing_span = inputs["data:geometry:wing:span"]
ht_span = inputs["data:geometry:horizontal_tail:span"]
wing_area = inputs["data:geometry:wing:area"]
ht_area = inputs["data:geometry:horizontal_tail:area"]
ht_taper_ratio = inputs["data:geometry:horizontal_tail:taper_ratio"]
ht_sweep_25 = inputs["data:geometry:horizontal_tail:sweep_25"] # In rad !!!
ht_dihedral = inputs["data:geometry:horizontal_tail:dihedral"] # In rad
if float(inputs["data:geometry:horizontal_tail:z:from_wingMAC25"]) == 0.0:
z2_ht = 0.0 # Aligned with the fuselage centerline
else:
z2_ht = (
inputs["data:geometry:wing:root:z"]
- inputs["data:geometry:horizontal_tail:z:from_wingMAC25"]
)
cl_alpha_airfoil = inputs["data:aerodynamics:horizontal_tail:airfoil:CL_alpha"]
ls_tag = "low_speed" if self.options["low_speed_aero"] else "cruise"
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:" + ls_tag + ":AOA"]
mach = inputs["data:aerodynamics:" + ls_tag + ":mach"]
cl_0_ht = inputs["data:aerodynamics:horizontal_tail:" + ls_tag + ":CL0"]
cl_alpha_ht = inputs["data:aerodynamics:horizontal_tail:" + ls_tag + ":CL_alpha"]
cd0_ht = inputs["data:aerodynamics:horizontal_tail:" + ls_tag + ":CD0"]
k = cl_alpha_airfoil / (2.0 * np.pi)
beta = np.sqrt(1.0 - mach**2.0)
roll_damping_parameter = self.cl_p_roll_damping_parameter(
ht_taper_ratio, ht_ar, mach, ht_sweep_25, k
)
cl_ht = (cl_0_ht + cl_alpha_ht * aoa_ref) * wing_area / ht_area
cd0_ht = cd0_ht * wing_area / ht_area
dihedral_effect = (
1.0
- 4.0 * z2_ht / ht_span * np.sin(ht_dihedral)
+ 12.0 * (z2_ht / ht_span) ** 2.0 * np.sin(ht_dihedral) ** 2.0
)
roll_damping_due_to_induced_drag_parameter = self.cl_p_cdi_roll_damping(ht_sweep_25, ht_ar)
roll_damping_due_to_induced_drag = (
roll_damping_due_to_induced_drag_parameter * cl_ht**2.0 - 0.125 * cd0_ht
)
# The assumption we make on the lift curve makes the second term equal to 1.0
cl_p_h = (
(
roll_damping_parameter * k / beta * 1.0 * dihedral_effect
+ roll_damping_due_to_induced_drag
)
* (ht_area / wing_area)
* (ht_span / wing_span) ** 2.0
)
outputs["data:aerodynamics:horizontal_tail:" + ls_tag + ":Cl_p"] = cl_p_h