# 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 openmdao.api as om
import fastoad.api as oad
from ..constants import SUBMODEL_CL_RUDDER
[docs]@oad.RegisterSubmodel(SUBMODEL_CL_RUDDER, "fastga.submodel.aerodynamics.rudder.roll_moment.legacy")
class ComputeClDeltaRudder(om.ExplicitComponent):
"""
Class to compute the contribution of the rudder to the roll moment coefficient. Depends on
the 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.
Based on :cite:`roskampart6:1985` section 10.3.8
"""
[docs] def initialize(self):
self.options.declare("low_speed_aero", default=False, types=bool)
[docs] def setup(self):
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:root:z", units="m", val=np.nan)
self.add_input(
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25", units="m", val=np.nan
)
self.add_input("data:geometry:vertical_tail:MAC:z", units="m", val=np.nan)
self.add_input("data:geometry:vertical_tail:area", units="m**2", val=np.nan)
self.add_input("data:geometry:fuselage:maximum_height", units="m", val=np.nan)
if self.options["low_speed_aero"]:
self.add_input(
"settings:aerodynamics:reference_flight_conditions:low_speed:AOA",
units="rad",
val=5.0 * np.pi / 180.0,
)
self.add_input(
"data:aerodynamics:rudder:low_speed:Cy_delta_r", val=np.nan, units="rad**-1"
)
self.add_output("data:aerodynamics:rudder:low_speed:Cl_delta_r", units="rad**-1")
else:
self.add_input(
"settings:aerodynamics:reference_flight_conditions:cruise:AOA",
units="rad",
val=1.0 * np.pi / 180.0,
)
self.add_input(
"data:aerodynamics:rudder:cruise:Cy_delta_r", val=np.nan, units="rad**-1"
)
self.add_output("data:aerodynamics:rudder:cruise:Cl_delta_r", units="rad**-1")
self.declare_partials(of="*", wrt="*", method="exact")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
z_v = (
inputs["data:geometry:wing:root:z"]
+ 0.5 * inputs["data:geometry:fuselage:maximum_height"]
+ inputs["data:geometry:vertical_tail:MAC:z"]
)
lp_vt = inputs["data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25"]
wing_span = inputs["data:geometry:wing:span"]
wing_area = inputs["data:geometry:wing:area"]
vt_area = inputs["data:geometry:vertical_tail:area"]
if self.options["low_speed_aero"]:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:low_speed:AOA"]
# Change of reference surface area
cy_delta_r = (
inputs["data:aerodynamics:rudder:low_speed:Cy_delta_r"] * vt_area / wing_area
)
outputs["data:aerodynamics:rudder:low_speed:Cl_delta_r"] = (
cy_delta_r * (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref)) / wing_span
)
else:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:cruise:AOA"]
# Change of reference surface area
cy_delta_r = inputs["data:aerodynamics:rudder:cruise:Cy_delta_r"] * vt_area / wing_area
outputs["data:aerodynamics:rudder:cruise:Cl_delta_r"] = (
cy_delta_r * (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref)) / wing_span
)
[docs] def compute_partials(self, inputs, partials, discrete_inputs=None):
z_v = (
inputs["data:geometry:wing:root:z"]
+ 0.5 * inputs["data:geometry:fuselage:maximum_height"]
+ inputs["data:geometry:vertical_tail:MAC:z"]
)
lp_vt = inputs["data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25"]
wing_span = inputs["data:geometry:wing:span"]
wing_area = inputs["data:geometry:wing:area"]
vt_area = inputs["data:geometry:vertical_tail:area"]
if self.options["low_speed_aero"]:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:low_speed:AOA"]
cy_delta_r = inputs["data:aerodynamics:rudder:low_speed:Cy_delta_r"]
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:aerodynamics:rudder:low_speed:Cy_delta_r",
] = (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref)) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:wing:root:z",
] = cy_delta_r * np.cos(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:fuselage:maximum_height",
] = 0.5 * cy_delta_r * np.cos(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:vertical_tail:MAC:z",
] = cy_delta_r * np.cos(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
] = -cy_delta_r * np.sin(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:vertical_tail:area",
] = (
cy_delta_r
* (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref))
/ wing_span
/ wing_area
)
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:wing:span",
] = (
-cy_delta_r
* (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref))
/ wing_span**2.0
* vt_area
/ wing_area
)
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"data:geometry:wing:area",
] = (
-cy_delta_r
* (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref))
/ wing_span
* vt_area
/ wing_area**2.0
)
partials[
"data:aerodynamics:rudder:low_speed:Cl_delta_r",
"settings:aerodynamics:reference_flight_conditions:low_speed:AOA",
] = (
-cy_delta_r
* (z_v * np.sin(aoa_ref) + lp_vt * np.cos(aoa_ref))
/ wing_span
* vt_area
/ wing_area
)
else:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:cruise:AOA"]
cy_delta_r = inputs["data:aerodynamics:rudder:cruise:Cy_delta_r"]
partials[
"data:aerodynamics:rudder:cruise:Cl_delta_r",
"data:aerodynamics:rudder:cruise:Cy_delta_r",
] = (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref)) / wing_span * vt_area / wing_area
partials["data:aerodynamics:rudder:cruise:Cl_delta_r", "data:geometry:wing:root:z"] = (
(cy_delta_r * np.cos(aoa_ref) / wing_span) * vt_area / wing_area
)
partials[
"data:aerodynamics:rudder:cruise:Cl_delta_r",
"data:geometry:fuselage:maximum_height",
] = 0.5 * cy_delta_r * np.cos(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:cruise:Cl_delta_r",
"data:geometry:vertical_tail:MAC:z",
] = cy_delta_r * np.cos(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:cruise:Cl_delta_r",
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
] = -cy_delta_r * np.sin(aoa_ref) / wing_span * vt_area / wing_area
partials[
"data:aerodynamics:rudder:cruise:Cl_delta_r",
"data:geometry:vertical_tail:area",
] = (
cy_delta_r
* (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref))
/ wing_span
/ wing_area
)
partials["data:aerodynamics:rudder:cruise:Cl_delta_r", "data:geometry:wing:span"] = (
(-cy_delta_r * (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref)) / wing_span**2.0)
* vt_area
/ wing_area
)
partials["data:aerodynamics:rudder:cruise:Cl_delta_r", "data:geometry:wing:area"] = (
-cy_delta_r
* (z_v * np.cos(aoa_ref) - lp_vt * np.sin(aoa_ref))
/ wing_span
* vt_area
/ wing_area**2.0
)
partials[
"data:aerodynamics:rudder:cruise:Cl_delta_r",
"settings:aerodynamics:reference_flight_conditions:cruise:AOA",
] = (
(-cy_delta_r * (z_v * np.sin(aoa_ref) + lp_vt * np.cos(aoa_ref)) / wing_span)
* vt_area
/ wing_area
)