"""Simple module for climb computation."""
# 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 os
import logging
import time
import numpy as np
import openmdao.api as om
from scipy.constants import g
from scipy.interpolate import interp1d
# noinspection PyProtectedMember
from fastoad.module_management._bundle_loader import BundleLoader
import fastoad.api as oad
from fastoad.constants import EngineSetting
from stdatm import Atmosphere
from fastga.models.performances.mission.takeoff import SAFETY_HEIGHT
from ..dynamic_equilibrium import DynamicEquilibrium
from ..constants import SUBMODEL_CLIMB, SUBMODEL_CLIMB_SPEED
_LOGGER = logging.getLogger(__name__)
POINTS_NB_CLIMB = 100
MAX_CALCULATION_TIME = 15 # time in seconds
oad.RegisterSubmodel.active_models[SUBMODEL_CLIMB] = (
"fastga.submodel.performances.mission.climb.legacy"
)
oad.RegisterSubmodel.active_models[SUBMODEL_CLIMB_SPEED] = (
"fastga.submodel.performances.mission.climb_speed.legacy"
)
[docs]@oad.RegisterSubmodel(SUBMODEL_CLIMB, "fastga.submodel.performances.mission.climb.legacy")
class ComputeClimb(DynamicEquilibrium):
"""
Compute the fuel consumption on climb segment with constant VCAS and a climb rate which
varies linearly with the altitude.
The hypothesis of small alpha/gamma angles is done.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
[docs] def initialize(self):
super().initialize()
self.options.declare("propulsion_id", default="", types=str)
[docs] def setup(self):
super().setup()
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:aerodynamics:aircraft:cruise:CD0", np.nan)
self.add_input("data:aerodynamics:wing:cruise:induced_drag_coefficient", np.nan)
self.add_input("data:aerodynamics:horizontal_tail:cruise:induced_drag_coefficient", np.nan)
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:mission:sizing:taxi_out:fuel", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:fuel", np.nan, units="kg")
self.add_input("data:mission:sizing:initial_climb:fuel", np.nan, units="kg")
self.add_input(
"data:mission:sizing:main_route:climb:climb_rate:sea_level", val=np.nan, units="m/s"
)
self.add_input(
"data:mission:sizing:main_route:climb:climb_rate:cruise_level", val=np.nan, units="m/s"
)
self.add_input("data:mission:sizing:main_route:climb:v_cas", val=np.nan, units="m/s")
self.add_output("data:mission:sizing:main_route:climb:fuel", units="kg")
self.add_output("data:mission:sizing:main_route:climb:distance", units="m")
self.add_output("data:mission:sizing:main_route:climb:duration", units="s")
self.declare_partials("*", "*", method="fd")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
# Delete previous .csv results
if self.options["out_file"] != "":
# noinspection PyBroadException
try:
os.remove(self.options["out_file"])
except OSError:
_LOGGER.info("Failed to remove %s file!", self.options["out_file"])
propulsion_model = self._engine_wrapper.get_model(inputs)
wing_area = inputs["data:geometry:wing:area"]
cruise_altitude = inputs["data:mission:sizing:main_route:cruise:altitude"]
mtow = inputs["data:weight:aircraft:MTOW"]
m_to = inputs["data:mission:sizing:taxi_out:fuel"]
m_tk = inputs["data:mission:sizing:takeoff:fuel"]
m_ic = inputs["data:mission:sizing:initial_climb:fuel"]
v_cas = inputs["data:mission:sizing:main_route:climb:v_cas"]
climb_rate_sl = float(inputs["data:mission:sizing:main_route:climb:climb_rate:sea_level"])
climb_rate_cl = float(
inputs["data:mission:sizing:main_route:climb:climb_rate:cruise_level"]
)
# Define initial conditions
t_start = time.time()
altitude_t = SAFETY_HEIGHT # conversion to m
distance_t = 0.0
time_t = 0.0
mass_t = mtow - (m_to + m_tk + m_ic)
mass_fuel_t = 0.0
previous_step = ()
self.flight_points = []
# Calculate constant speed (cos(gamma)~1) and corresponding climb angle
atm = Atmosphere(altitude_t, altitude_in_feet=False)
atm.calibrated_airspeed = v_cas
# Define specific time step ~POINTS_NB_CLIMB points for calculation (with ground conditions)
time_step = ((cruise_altitude - SAFETY_HEIGHT) / climb_rate_sl) / float(POINTS_NB_CLIMB)
while altitude_t < cruise_altitude:
flight_point = oad.FlightPoint(
altitude=altitude_t,
time=time_t,
ground_distance=distance_t,
engine_setting=EngineSetting.CLIMB,
thrust_is_regulated=True,
mass=mass_t,
name="sizing:main_route:climb",
)
climb_rate = interp1d([0.0, float(cruise_altitude)], [climb_rate_sl, climb_rate_cl])(
altitude_t
)
self.complete_flight_point(flight_point, v_cas=v_cas, climb_rate=climb_rate)
# Calculate dynamic pressure
atm = Atmosphere(altitude_t, altitude_in_feet=False)
atm.calibrated_airspeed = v_cas
v_tas = atm.true_airspeed
atm_1 = Atmosphere(altitude_t + 1.0, altitude_in_feet=False)
atm_1.calibrated_airspeed = v_cas
dv_tas_dh = atm_1.true_airspeed - v_tas
dvx_dt = dv_tas_dh * v_tas * np.sin(flight_point.gamma)
dynamic_pressure = 0.5 * atm.density * v_tas**2
# Find equilibrium
previous_step = self.dynamic_equilibrium(
inputs,
flight_point.gamma,
dynamic_pressure,
dvx_dt,
0.0,
mass_t,
"none",
previous_step[0:2],
)
flight_point.thrust = float(previous_step[1])
# Compute consumption
propulsion_model.compute_flight_points(flight_point)
if flight_point.thrust_rate > 1.0:
_LOGGER.warning("Thrust rate is above 1.0, value clipped at 1.0")
# Save results
self.compute_flight_point_drag(
flight_point=flight_point, equilibrium_result=previous_step, wing_area=wing_area
)
self.add_flight_point(flight_point=flight_point, equilibrium_result=previous_step)
consumed_mass_1s = propulsion_model.get_consumed_mass(flight_point, 1.0)
# Calculate distance variation (earth axis)
v_z = v_tas * np.sin(flight_point.gamma)
v_x = v_tas * np.cos(flight_point.gamma)
time_step = min(time_step, (cruise_altitude - altitude_t) / v_z)
altitude_t += v_z * time_step
distance_t += v_x * time_step
# Estimate mass evolution and update time
mass_fuel_t += consumed_mass_1s * time_step
mass_t = mass_t - consumed_mass_1s * time_step
time_t += time_step
# Check calculation duration
if (time.time() - t_start) > MAX_CALCULATION_TIME:
raise Exception(
"Time calculation duration for climb phase [%f s] exceeded!"
% MAX_CALCULATION_TIME
)
# Save mission
if self.options["out_file"] != "":
self.save_csv()
outputs["data:mission:sizing:main_route:climb:fuel"] = mass_fuel_t
outputs["data:mission:sizing:main_route:climb:distance"] = distance_t
outputs["data:mission:sizing:main_route:climb:duration"] = time_t
[docs]@oad.RegisterSubmodel(
SUBMODEL_CLIMB_SPEED, "fastga.submodel.performances.mission.climb_speed.legacy"
)
class ComputeClimbSpeed(om.ExplicitComponent):
[docs] def setup(self):
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:aerodynamics:aircraft:cruise:CD0", val=np.nan)
self.add_input("data:aerodynamics:wing:cruise:induced_drag_coefficient", val=np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean", val=np.nan)
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:mission:sizing:taxi_out:fuel", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:fuel", np.nan, units="kg")
self.add_input("data:mission:sizing:initial_climb:fuel", np.nan, units="kg")
self.add_output("data:mission:sizing:main_route:climb:v_cas", units="m/s")
[docs] def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
altitude_t = SAFETY_HEIGHT # conversion to m
cd0 = inputs["data:aerodynamics:aircraft:cruise:CD0"]
coeff_k_wing = inputs["data:aerodynamics:wing:cruise:induced_drag_coefficient"]
c_l_max_clean = inputs["data:aerodynamics:wing:low_speed:CL_max_clean"]
wing_area = inputs["data:geometry:wing:area"]
mtow = inputs["data:weight:aircraft:MTOW"]
m_to = inputs["data:mission:sizing:taxi_out:fuel"]
m_tk = inputs["data:mission:sizing:takeoff:fuel"]
m_ic = inputs["data:mission:sizing:initial_climb:fuel"]
mass_t = mtow - (m_to + m_tk + m_ic)
c_l = np.sqrt(3 * cd0 / coeff_k_wing)
atm = Atmosphere(altitude_t, altitude_in_feet=False)
vs1 = np.sqrt((mass_t * g) / (0.5 * atm.density * wing_area * c_l_max_clean))
v_cas = max(np.sqrt((mass_t * g) / (0.5 * atm.density * wing_area * c_l)), 1.3 * vs1)
outputs["data:mission:sizing:main_route:climb:v_cas"] = v_cas