"""Computation of lift and drag increment due to high-lift devices."""
# 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/>.
from typing import Union
import numpy as np
import fastoad.api as oad
from .figure_digitization import FigureDigitization
from ..constants import SUBMODEL_DELTA_ELEVATOR
[docs]@oad.RegisterSubmodel(SUBMODEL_DELTA_ELEVATOR, "fastga.submodel.aerodynamics.elevator.delta.legacy")
class ComputeDeltaElevator(FigureDigitization):
"""
Provides lift and drag increments due to high-lift devices.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self.phase = None
[docs] def setup(self):
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:geometry:horizontal_tail:sweep_25", val=np.nan, units="rad")
self.add_input("data:geometry:horizontal_tail:elevator_chord_ratio", val=np.nan)
self.add_input("data:geometry:horizontal_tail:thickness_ratio", val=np.nan)
self.add_input("data:aerodynamics:low_speed:mach", val=np.nan)
self.add_input(
"data:aerodynamics:horizontal_tail:airfoil:CL_alpha", val=np.nan, units="rad**-1"
)
self.add_input("data:mission:sizing:landing:elevator_angle", val=np.nan, units="deg")
self.add_output("data:aerodynamics:elevator:low_speed:CL_delta", units="rad**-1")
self.add_output("data:aerodynamics:elevator:low_speed:CD_delta", units="rad**-2")
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
wing_area = inputs["data:geometry:wing:area"]
htp_area = inputs["data:geometry:horizontal_tail:area"]
elevator_chord_ratio = inputs["data:geometry:horizontal_tail:elevator_chord_ratio"]
# Computes elevator contribution during low speed operations (for different deflection
# angle)
outputs["data:aerodynamics:elevator:low_speed:CL_delta"] = self._get_elevator_delta_cl(
inputs,
25.0,
) # get derivative for 25° angle assuming it is linear when <= to 25 degree,
# derivative wrt to the wing, multiplies the deflection angle squared
outputs["data:aerodynamics:elevator:low_speed:CD_delta"] = (
self.delta_cd_plain_flap(
float(elevator_chord_ratio),
abs(float(inputs["data:mission:sizing:landing:elevator_angle"])),
)
/ (abs(inputs["data:mission:sizing:landing:elevator_angle"]) * np.pi / 180.0) ** 2.0
* np.cos(inputs["data:geometry:horizontal_tail:sweep_25"])
* htp_area
/ wing_area
)
def _get_elevator_delta_cl(
self, inputs, elevator_angle: Union[float, np.array]
) -> Union[float, np.array]:
"""
Computes the elevator lift increment as a plain flap following the method presented in
Roskam part 6, section 8.1.2.1.a.
:param elevator_angle: elevator angle (in Degree).
:return: lift coefficient derivative.
"""
ht_area = inputs["data:geometry:horizontal_tail:area"]
wing_area = inputs["data:geometry:wing:area"]
elevator_chord_ratio = inputs["data:geometry:horizontal_tail:elevator_chord_ratio"]
htp_thickness_ratio = inputs["data:geometry:horizontal_tail:thickness_ratio"]
cl_alpha_airfoil_ht = inputs["data:aerodynamics:horizontal_tail:airfoil:CL_alpha"]
# Elevator (plain flap). Default: maximum deflection (25deg)
cl_delta_theory = self.cl_delta_theory_plain_flap(
float(htp_thickness_ratio), float(elevator_chord_ratio)
)
k = self.k_prime_plain_flap(abs(float(elevator_angle)), float(elevator_chord_ratio))
k_cl_delta = self.k_cl_delta_plain_flap(
float(htp_thickness_ratio), float(cl_alpha_airfoil_ht), float(elevator_chord_ratio)
)
cl_alpha_elev = (cl_delta_theory * k * k_cl_delta) * ht_area / wing_area
cl_alpha_elev *= 0.9 # Correction for the central fuselage part (no elevator there)
return cl_alpha_elev