"""
Computation of wing area
"""
# 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 scipy.constants import g
from fastoad.module_management.constants import ModelDomain
from fastoad.module_management.service_registry import RegisterOpenMDAOSystem
[docs]@RegisterOpenMDAOSystem("fastoad.loop.wing_area", domain=ModelDomain.OTHER)
class ComputeWingArea(om.Group):
"""
Computes needed wing area for:
- having enough lift at required approach speed
- being able to load enough fuel to achieve the sizing mission
"""
[docs] def setup(self):
self.add_subsystem("wing_area", _ComputeWingArea(), promotes=["*"])
self.add_subsystem("constraints", _ComputeWingAreaConstraints(), promotes=["*"])
class _ComputeWingArea(om.ExplicitComponent):
""" Computation of wing area from needed approach speed and mission fuel """
def setup(self):
self.add_input("data:geometry:wing:aspect_ratio", val=np.nan)
self.add_input("data:geometry:wing:root:thickness_ratio", val=np.nan)
self.add_input("data:geometry:wing:tip:thickness_ratio", val=np.nan)
self.add_input("data:weight:aircraft:sizing_block_fuel", val=np.nan, units="kg")
self.add_input("data:TLAR:approach_speed", val=np.nan, units="m/s")
self.add_input("data:weight:aircraft:MLW", val=np.nan, units="kg")
self.add_input("data:aerodynamics:aircraft:landing:CL_max", val=np.nan)
self.add_output("data:geometry:wing:area", val=100.0, units="m**2")
def setup_partials(self):
self.declare_partials("data:geometry:wing:area", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
lambda_wing = inputs["data:geometry:wing:aspect_ratio"]
root_thickness_ratio = inputs["data:geometry:wing:root:thickness_ratio"]
tip_thickness_ratio = inputs["data:geometry:wing:tip:thickness_ratio"]
mfw_mission = inputs["data:weight:aircraft:sizing_block_fuel"]
wing_area_mission = (
max(1000.0, mfw_mission - 1570.0)
/ 224
/ lambda_wing ** -0.4
/ (0.6 * root_thickness_ratio + 0.4 * tip_thickness_ratio)
) ** (1.0 / 1.5)
approach_speed = inputs["data:TLAR:approach_speed"]
mlw = inputs["data:weight:aircraft:MLW"]
max_CL = inputs["data:aerodynamics:aircraft:landing:CL_max"]
wing_area_approach = 2 * mlw * g / ((approach_speed / 1.23) ** 2) / (1.225 * max_CL)
outputs["data:geometry:wing:area"] = np.nanmax([wing_area_mission, wing_area_approach])
class _ComputeWingAreaConstraints(om.ExplicitComponent):
def setup(self):
self.add_input("data:weight:aircraft:sizing_block_fuel", val=np.nan, units="kg")
self.add_input("data:weight:aircraft:MFW", val=np.nan, units="kg")
self.add_input("data:TLAR:approach_speed", val=np.nan, units="m/s")
self.add_input("data:weight:aircraft:MLW", val=np.nan, units="kg")
self.add_input("data:aerodynamics:aircraft:landing:CL_max", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_output("data:weight:aircraft:additional_fuel_capacity", units="kg")
self.add_output("data:aerodynamics:aircraft:landing:additional_CL_capacity")
def setup_partials(self):
self.declare_partials(
"data:weight:aircraft:additional_fuel_capacity",
["data:weight:aircraft:MFW", "data:weight:aircraft:sizing_block_fuel"],
method="fd",
)
self.declare_partials(
"data:aerodynamics:aircraft:landing:additional_CL_capacity",
[
"data:TLAR:approach_speed",
"data:weight:aircraft:MLW",
"data:aerodynamics:aircraft:landing:CL_max",
"data:geometry:wing:area",
],
method="fd",
)
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
mfw = inputs["data:weight:aircraft:MFW"]
mission_fuel = inputs["data:weight:aircraft:sizing_block_fuel"]
v_approach = inputs["data:TLAR:approach_speed"]
cl_max = inputs["data:aerodynamics:aircraft:landing:CL_max"]
mlw = inputs["data:weight:aircraft:MLW"]
wing_area = inputs["data:geometry:wing:area"]
outputs["data:weight:aircraft:additional_fuel_capacity"] = mfw - mission_fuel
outputs["data:aerodynamics:aircraft:landing:additional_CL_capacity"] = cl_max - mlw * g / (
0.5 * 1.225 * (v_approach / 1.23) ** 2 * wing_area
)