# This file is part of FAST-OAD_CS23 : A framework for rapid Overall Aircraft Design
# Copyright (C) 2024 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
[docs]class AlphaRatio(om.ExplicitComponent):
[docs] def initialize(self):
self.options.declare("number_of_points", types=int, default=250)
[docs] def setup(self):
n = self.options["number_of_points"]
self.add_input("total_pressure_41", units="Pa", shape=n, val=np.nan)
self.add_input("total_pressure_45", units="Pa", shape=n, val=np.nan)
self.add_input(
"data:propulsion:turboprop:design_point:turbine_entry_temperature",
np.nan,
units="K",
)
self.add_input(
"total_temperature_45",
units="degK",
val=np.full(n, np.nan),
shape=n,
)
self.add_output("data:propulsion:turboprop:design_point:alpha", val=np.full(n, 0.8))
self.add_output("data:propulsion:turboprop:design_point:alpha_p", val=np.full(n, 0.3))
self.declare_partials(
of="data:propulsion:turboprop:design_point:alpha",
wrt=[
"data:propulsion:turboprop:design_point:turbine_entry_temperature",
"total_temperature_45",
],
method="exact",
)
self.declare_partials(
of="data:propulsion:turboprop:design_point:alpha_p",
wrt=[
"total_pressure_41",
"total_pressure_45",
],
method="exact",
)
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
total_temperature_41 = inputs[
"data:propulsion:turboprop:design_point:turbine_entry_temperature"
]
total_temperature_45 = inputs["total_temperature_45"]
total_pressure_45 = inputs["total_pressure_45"]
total_pressure_41 = inputs["total_pressure_41"]
outputs["data:propulsion:turboprop:design_point:alpha"] = (
total_temperature_45 / total_temperature_41
)
outputs["data:propulsion:turboprop:design_point:alpha_p"] = (
total_pressure_45 / total_pressure_41
)
[docs] def compute_partials(self, inputs, partials, discrete_inputs=None):
total_temperature_41 = inputs[
"data:propulsion:turboprop:design_point:turbine_entry_temperature"
]
total_temperature_45 = inputs["total_temperature_45"]
total_pressure_45 = inputs["total_pressure_45"]
total_pressure_41 = inputs["total_pressure_41"]
partials["data:propulsion:turboprop:design_point:alpha", "total_temperature_45"] = np.diag(
1.0 / total_temperature_41
)
partials[
"data:propulsion:turboprop:design_point:alpha",
"data:propulsion:turboprop:design_point:turbine_entry_temperature",
] = -total_temperature_45 / total_temperature_41**2.0
partials["data:propulsion:turboprop:design_point:alpha_p", "total_pressure_45"] = np.diag(
1.0 / total_pressure_41
)
partials["data:propulsion:turboprop:design_point:alpha_p", "total_pressure_41"] = -np.diag(
total_pressure_45 / total_pressure_41**2.0
)