# 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_CY_P
[docs]@oad.RegisterSubmodel(SUBMODEL_CY_P, "fastga.submodel.aerodynamics.aircraft.cy_roll_rate.legacy")
class ComputeCyRollRateAircraft(om.ExplicitComponent):
"""
Computation of the increase in side force due to a roll rate. Assumes that only the
contribution of the VTP is relevant for this coefficient. Same remark on the dependency on
the AoA of this coefficient as in .compute_cy_yaw_rate. The convention from
:cite:`roskampart6:1985` are used, meaning 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:wing:span", val=np.nan, units="m")
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: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:vertical_tail:low_speed:Cy_beta", val=np.nan, units="rad**-1"
)
self.add_output("data:aerodynamics:aircraft:low_speed:Cy_p", 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:vertical_tail:cruise:Cy_beta", val=np.nan, units="rad**-1"
)
self.add_output("data:aerodynamics:aircraft:cruise:Cy_p", 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"]
if self.options["low_speed_aero"]:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:low_speed:AOA"]
cy_beta_vt = inputs["data:aerodynamics:vertical_tail:low_speed:Cy_beta"]
outputs["data:aerodynamics:aircraft:low_speed:Cy_p"] = (
2.0 * cy_beta_vt * (-lp_vt * np.sin(aoa_ref) + z_v * np.cos(aoa_ref)) / wing_span
)
else:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:cruise:AOA"]
cy_beta_vt = inputs["data:aerodynamics:vertical_tail:cruise:Cy_beta"]
outputs["data:aerodynamics:aircraft:cruise:Cy_p"] = (
2.0 * cy_beta_vt * (-lp_vt * np.sin(aoa_ref) + z_v * np.cos(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"]
if self.options["low_speed_aero"]:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:low_speed:AOA"]
cy_beta_vt = inputs["data:aerodynamics:vertical_tail:low_speed:Cy_beta"]
partials[
"data:aerodynamics:aircraft:low_speed:Cy_p",
"data:aerodynamics:vertical_tail:low_speed:Cy_beta",
] = 2.0 * (-lp_vt * np.sin(aoa_ref) + z_v * np.cos(aoa_ref)) / wing_span
partials["data:aerodynamics:aircraft:low_speed:Cy_p", "data:geometry:wing:root:z"] = (
2.0 * cy_beta_vt * np.cos(aoa_ref) / wing_span
)
partials[
"data:aerodynamics:aircraft:low_speed:Cy_p",
"data:geometry:fuselage:maximum_height",
] = cy_beta_vt * np.cos(aoa_ref) / wing_span
partials[
"data:aerodynamics:aircraft:low_speed:Cy_p",
"data:geometry:vertical_tail:MAC:z",
] = 2.0 * cy_beta_vt * np.cos(aoa_ref) / wing_span
partials[
"data:aerodynamics:aircraft:low_speed:Cy_p",
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
] = -2.0 * cy_beta_vt * np.sin(aoa_ref) / wing_span
partials["data:aerodynamics:aircraft:low_speed:Cy_p", "data:geometry:wing:span"] = (
-2.0
* cy_beta_vt
* (-lp_vt * np.sin(aoa_ref) + z_v * np.cos(aoa_ref))
/ wing_span**2.0
)
partials[
"data:aerodynamics:aircraft:low_speed:Cy_p",
"settings:aerodynamics:reference_flight_conditions:low_speed:AOA",
] = -2.0 * cy_beta_vt * (lp_vt * np.cos(aoa_ref) + z_v * np.sin(aoa_ref)) / wing_span
else:
aoa_ref = inputs["settings:aerodynamics:reference_flight_conditions:cruise:AOA"]
cy_beta_vt = inputs["data:aerodynamics:vertical_tail:cruise:Cy_beta"]
partials[
"data:aerodynamics:aircraft:cruise:Cy_p",
"data:aerodynamics:vertical_tail:cruise:Cy_beta",
] = 2.0 * (-lp_vt * np.sin(aoa_ref) + z_v * np.cos(aoa_ref)) / wing_span
partials["data:aerodynamics:aircraft:cruise:Cy_p", "data:geometry:wing:root:z"] = (
2.0 * cy_beta_vt * np.cos(aoa_ref) / wing_span
)
partials[
"data:aerodynamics:aircraft:cruise:Cy_p",
"data:geometry:fuselage:maximum_height",
] = cy_beta_vt * np.cos(aoa_ref) / wing_span
partials[
"data:aerodynamics:aircraft:cruise:Cy_p",
"data:geometry:vertical_tail:MAC:z",
] = 2.0 * cy_beta_vt * np.cos(aoa_ref) / wing_span
partials[
"data:aerodynamics:aircraft:cruise:Cy_p",
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
] = -2.0 * cy_beta_vt * np.sin(aoa_ref) / wing_span
partials["data:aerodynamics:aircraft:cruise:Cy_p", "data:geometry:wing:span"] = (
-2.0
* cy_beta_vt
* (-lp_vt * np.sin(aoa_ref) + z_v * np.cos(aoa_ref))
/ wing_span**2.0
)
partials[
"data:aerodynamics:aircraft:cruise:Cy_p",
"settings:aerodynamics:reference_flight_conditions:cruise:AOA",
] = -2.0 * cy_beta_vt * (lp_vt * np.cos(aoa_ref) + z_v * np.sin(aoa_ref)) / wing_span