def compute_from(self, start: FlightPoint) -> pd.DataFrame:
self.complete_flight_point(
start) # needed to ensure all speed values are computed.
if self.target.altitude is not None:
if isinstance(self.target.altitude, str):
# Target altitude will be modified along the process, so we keep track
# of the original order in target CL, that is not used otherwise.
self.target.CL = self.target.altitude # pylint: disable=invalid-name
# let's put a numerical, negative value in self.target.altitude to
# ensure there will be no problem in self._get_distance_to_target()
self.target.altitude = -1000.0
self.interrupt_if_getting_further_from_target = False
else:
# Target altitude is fixed, back to original settings (in case
# this instance is used more than once)
self.target.CL = None
self.interrupt_if_getting_further_from_target = True
atm = AtmosphereSI(start.altitude)
if self.target.equivalent_airspeed == self.CONSTANT_VALUE:
atm.equivalent_airspeed = start.equivalent_airspeed
start.true_airspeed = atm.true_airspeed
elif self.target.mach == self.CONSTANT_VALUE:
atm.mach = start.mach
start.true_airspeed = atm.true_airspeed
return super().compute_from(start)
def test_EngineSet():
"""Tests behaviour of FuelEngineSet"""
engine = DummyEngine(1.2e5, 1.0e-5)
# input = thrust rate
thrust_rate = np.linspace(0.0, 1.0, 20)
for engine_count in [1, 2, 3, 4]:
engine_set = FuelEngineSet(engine, engine_count)
flight_point = FlightPoint(mach=0.0,
altitude=0.0,
engine_setting=1,
thrust_rate=thrust_rate)
engine_set.compute_flight_points(flight_point)
assert_allclose(flight_point.sfc, engine.max_sfc * thrust_rate)
assert_allclose(flight_point.thrust,
engine.max_thrust * engine_count * thrust_rate)
# input = thrust
thrust = np.linspace(0.0, engine.max_thrust, 30)
for engine_count in [1, 2, 3, 4]:
engine_set = FuelEngineSet(engine, engine_count)
flight_point = FlightPoint(mach=0.0,
altitude=0.0,
engine_setting=1,
thrust=thrust)
engine_set.compute_flight_points(flight_point)
assert_allclose(flight_point.thrust_rate,
thrust / engine_count / engine.max_thrust)
assert_allclose(flight_point.sfc,
engine.max_sfc * flight_point.thrust_rate)
def _climb_to_altitude_and_cruise(
self,
start: FlightPoint,
cruise_altitude: float,
climb_segment: AltitudeChangeSegment,
cruise_segment: CruiseSegment,
):
"""
Climbs up to cruise_altitude and cruise, while ensuring final ground_distance is
equal to self.target.ground_distance.
:param start:
:param cruise_altitude:
:param climb_segment:
:param cruise_segment:
:return:
"""
climb_segment.target = FlightPoint(
altitude=cruise_altitude,
mach=cruise_segment.target.mach,
true_airspeed=cruise_segment.target.true_airspeed,
equivalent_airspeed=cruise_segment.target.equivalent_airspeed,
)
climb_points = climb_segment.compute_from(start)
cruise_start = FlightPoint.create(climb_points.iloc[-1])
cruise_segment.target.ground_distance = (self.target.ground_distance -
cruise_start.ground_distance)
cruise_points = cruise_segment.compute_from(cruise_start)
return pd.concat([climb_points, cruise_points]).reset_index(drop=True)
def complete_flight_point(self, flight_point: FlightPoint):
"""
Computes data for provided flight point.
Assumes that it is already defined for time, altitude, mass,
ground distance and speed (TAS, EAS, or Mach).
:param flight_point: the flight point that will be completed in-place
"""
flight_point.engine_setting = self.engine_setting
self._complete_speed_values(flight_point)
atm = AtmosphereSI(flight_point.altitude)
reference_force = 0.5 * atm.density * flight_point.true_airspeed**2 * self.reference_area
if self.polar:
flight_point.CL = flight_point.mass * g / reference_force
flight_point.CD = self.polar.cd(flight_point.CL)
else:
flight_point.CL = flight_point.CD = 0.0
flight_point.drag = flight_point.CD * reference_force
self._compute_propulsion(flight_point)
flight_point.slope_angle, flight_point.acceleration = self._get_gamma_and_acceleration(
flight_point.mass, flight_point.drag, flight_point.thrust)
def test_breguet_cruise(polar):
propulsion = FuelEngineSet(DummyEngine(0.5e5, 1.0e-5), 2)
segment = BreguetCruiseSegment(
target=FlightPoint(ground_distance=5.0e5),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
engine_setting=EngineSetting.
CRUISE, # The engine model does not use this setting
)
flight_points = segment.compute_from(
FlightPoint(time=10000.0, mass=70000.0, altitude=10000.0, mach=0.78))
print_dataframe(flight_points)
first_point = flight_points.iloc[0]
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with CruiseSegment and 0.05s as
# time step
assert_allclose(first_point.altitude, 10000.0)
assert_allclose(first_point.true_airspeed, 233.6, atol=0.1)
assert first_point.engine_setting == EngineSetting.CRUISE
assert_allclose(last_point.ground_distance, 500000.0)
assert_allclose(last_point.altitude, 10000.0)
assert_allclose(last_point.time, 12141.0, rtol=1e-2)
assert_allclose(last_point.true_airspeed, 233.6, atol=0.1)
assert_allclose(last_point.mass, 69568.0, rtol=1e-4)
assert last_point.engine_setting == EngineSetting.CRUISE
def test_climb_fixed_altitude_at_constant_EAS(polar):
propulsion = FuelEngineSet(DummyEngine(1.0e5, 1.0e-5), 2)
flight_points = AltitudeChangeSegment(
target=FlightPoint(altitude=10000.0, equivalent_airspeed="constant"),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
thrust_rate=1.0,
time_step=2.0,
engine_setting=EngineSetting.
CRUISE, # The engine model does not use this setting
).compute_from(
FlightPoint(altitude=5000.0, mass=70000.0, equivalent_airspeed=100.0))
first_point = flight_points.iloc[0]
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.01s as time step
assert_allclose(last_point.altitude, 10000.0)
assert_allclose(last_point.equivalent_airspeed, 100.0)
assert_allclose(last_point.time, 145.2, rtol=1e-2)
assert_allclose(first_point.true_airspeed, 129.0, atol=0.1)
assert_allclose(last_point.true_airspeed, 172.3, atol=0.1)
assert_allclose(last_point.mass, 69710.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 20915.0, rtol=1e-3)
assert last_point.engine_setting == EngineSetting.CRUISE
def _compute_breguet(self, inputs, outputs):
"""
Computes mission using simple Breguet formula at altitude==100m and Mach 0.1
(but max L/D ratio is assumed anyway)
Useful only for initiating the computation.
:param inputs: OpenMDAO input vector
:param outputs: OpenMDAO output vector
"""
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:engine:count"]
)
high_speed_polar = self._get_initial_polar(inputs)
distance = np.asscalar(
np.sum(self._mission_wrapper.get_route_ranges(inputs, self.options["mission_name"]))
)
altitude = 100.0
cruise_mach = 0.1
breguet = BreguetCruiseSegment(
FlightPoint(ground_distance=distance),
propulsion=propulsion_model,
polar=high_speed_polar,
use_max_lift_drag_ratio=True,
)
start_point = FlightPoint(
mass=inputs[self._mission_vars.TOW.value], altitude=altitude, mach=cruise_mach
)
flight_points = breguet.compute_from(start_point)
end_point = FlightPoint.create(flight_points.iloc[-1])
outputs[self._mission_vars.NEEDED_BLOCK_FUEL.value] = start_point.mass - end_point.mass
def test_optimal_cruise(polar):
propulsion = FuelEngineSet(DummyEngine(0.5e5, 1.0e-5), 2)
segment = OptimalCruiseSegment(
target=FlightPoint(ground_distance=5.0e5),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
engine_setting=EngineSetting.CRUISE,
)
flight_points = segment.compute_from(
FlightPoint(mass=70000.0, time=1000.0, ground_distance=1e5, mach=0.78))
print_dataframe(flight_points)
first_point = flight_points.iloc[0]
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.05s as time step
assert_allclose(first_point.altitude, 9156.0, atol=1.0)
assert_allclose(first_point.true_airspeed, 236.4, atol=0.1)
assert_allclose(first_point.CL, polar.optimal_cl)
assert_allclose(first_point.CL, polar.optimal_cl)
assert first_point.engine_setting == EngineSetting.CRUISE
assert_allclose(last_point.ground_distance, 600000.0)
assert_allclose(last_point.CL, polar.optimal_cl)
assert_allclose(last_point.altitude, 9196.0, atol=1.0)
assert_allclose(last_point.time, 3115.0, rtol=1e-2)
assert_allclose(last_point.true_airspeed, 236.3, atol=0.1)
assert_allclose(last_point.mass, 69577.0, rtol=1e-4)
assert last_point.engine_setting == EngineSetting.CRUISE
def test_climb_fixed_altitude_at_constant_TAS(polar):
propulsion = FuelEngineSet(DummyEngine(1.0e5, 1.0e-5), 2)
# initialisation then change instance attributes
segment = AltitudeChangeSegment(
target=FlightPoint(altitude=10000.0),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
engine_setting=EngineSetting.CLIMB,
) # not constant TAS order, as it is the default
segment.thrust_rate = 1.0
segment.time_step = 2.0
flight_points = segment.compute_from(
FlightPoint(altitude=5000.0, mass=70000.0,
true_airspeed=150.0)) # Test with dict
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.01s as time step
assert_allclose(last_point.altitude, 10000.0)
assert_allclose(last_point.true_airspeed, 150.0)
assert_allclose(last_point.time, 143.5, rtol=1e-2)
assert_allclose(last_point.mass, 69713.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 20943.0, rtol=1e-3)
assert last_point.engine_setting == EngineSetting.CLIMB
def compute_next_flight_point(self, flight_points: List[FlightPoint],
time_step: float) -> FlightPoint:
"""
Computes time, altitude, speed, mass and ground distance of next flight point.
:param flight_points: previous flight points
:param time_step: time step for computing next point
:return: the computed next flight point
"""
start = flight_points[0]
previous = flight_points[-1]
next_point = FlightPoint()
next_point.mass = previous.mass - self.propulsion.get_consumed_mass(
previous, time_step)
next_point.time = previous.time + time_step
next_point.ground_distance = (
previous.ground_distance +
previous.true_airspeed * time_step * np.cos(previous.slope_angle))
self._compute_next_altitude(next_point, previous)
if self.target.true_airspeed == self.CONSTANT_VALUE:
next_point.true_airspeed = previous.true_airspeed
elif self.target.equivalent_airspeed == self.CONSTANT_VALUE:
next_point.equivalent_airspeed = start.equivalent_airspeed
elif self.target.mach == self.CONSTANT_VALUE:
next_point.mach = start.mach
else:
next_point.true_airspeed = previous.true_airspeed + time_step * previous.acceleration
# The naming is not done in complete_flight_point for not naming the start point
next_point.name = self.name
return next_point
def compute(self, inputs: Vector, outputs: Vector,
start_flight_point: FlightPoint) -> pd.DataFrame:
"""
To be used during compute() of an OpenMDAO component.
Builds the mission from input file, and computes it. `outputs` vector is
filled with duration, burned fuel and covered ground distance for each
part of the flight.
:param inputs: the input vector of the OpenMDAO component
:param outputs: the output vector of the OpenMDAO component
:param start_flight_point: the starting flight point just after takeoff
:return: a pandas DataFrame where columns names match fields of
:class:`~fastoad.model_base.flight_point.FlightPoint`
"""
mission = self.build(inputs, self.mission_name)
def _compute_vars(name_root, start: FlightPoint, end: FlightPoint):
"""Computes duration, burned fuel and covered distance."""
if name_root + ":duration" in outputs:
outputs[name_root + ":duration"] = end.time - start.time
if name_root + ":fuel" in outputs:
outputs[name_root + ":fuel"] = start.mass - end.mass
if name_root + ":distance" in outputs:
outputs[
name_root +
":distance"] = end.ground_distance - start.ground_distance
if not start_flight_point.name:
start_flight_point.name = mission.flight_sequence[0].name
current_flight_point = start_flight_point
flight_points = mission.compute_from(start_flight_point)
for part in mission.flight_sequence:
var_name_root = "data:mission:%s" % part.name
part_end = FlightPoint.create(
flight_points.loc[flight_points.name.str.startswith(
part.name)].iloc[-1])
_compute_vars(var_name_root, current_flight_point, part_end)
if isinstance(part, FlightSequence):
# In case of a route, outputs are computed for each phase in the route
phase_start = current_flight_point
for phase in part.flight_sequence:
phase_points = flight_points.loc[flight_points.name ==
phase.name]
if len(phase_points) > 0:
phase_end = FlightPoint.create(phase_points.iloc[-1])
var_name_root = "data:mission:%s" % phase.name
_compute_vars(var_name_root, phase_start, phase_end)
phase_start = phase_end
current_flight_point = part_end
# Outputs for the whole mission
var_name_root = "data:mission:%s" % mission.name
_compute_vars(var_name_root, start_flight_point, current_flight_point)
return flight_points
def compute_flight_points(self, flight_point: FlightPoint):
if flight_point.thrust_is_regulated or flight_point.thrust_rate is None:
flight_point.thrust_rate = flight_point.thrust / self.max_thrust
else:
flight_point.thrust = self.max_thrust * flight_point.thrust_rate
flight_point.sfc = self.max_sfc * (1.0 + flight_point.thrust_rate) / 2.0
def test_compute_flight_points():
engine = RubberEngine(5, 30, 1500, 1, 0.95,
10000) # f0=1 so that output is simply fc/f0
# Test with scalars
flight_point = FlightPoint(
mach=0,
altitude=0,
engine_setting=EngineSetting.TAKEOFF,
thrust_rate=0.8) # with engine_setting as EngineSetting
engine.compute_flight_points(flight_point)
np.testing.assert_allclose(flight_point.thrust, 0.9553 * 0.8, rtol=1e-4)
np.testing.assert_allclose(flight_point.sfc, 0.99210e-5, rtol=1e-4)
flight_point = FlightPoint(mach=0.3,
altitude=0,
engine_setting=EngineSetting.CLIMB.value,
thrust=0.35677) # with engine_setting as int
engine.compute_flight_points(flight_point)
np.testing.assert_allclose(flight_point.thrust_rate, 0.5, rtol=1e-4)
np.testing.assert_allclose(flight_point.sfc, 1.3496e-5, rtol=1e-4)
# Test full arrays
# 2D arrays are used, where first line is for thrust rates, and second line
# is for thrust values.
# As thrust rates and thrust values match, thrust rate results are 2 equal
# lines and so are thrust value results.
machs = [0, 0.3, 0.3, 0.8, 0.8]
altitudes = [0, 0, 0, 10000, 13000]
thrust_rates = [0.8, 0.5, 0.5, 0.4, 0.7]
thrusts = [0.9553 * 0.8, 0.38851, 0.35677, 0.09680, 0.11273]
engine_settings = [
EngineSetting.TAKEOFF,
EngineSetting.TAKEOFF,
EngineSetting.CLIMB,
EngineSetting.IDLE,
EngineSetting.CRUISE.value,
] # mix EngineSetting with integers
expected_sfc = [0.99210e-5, 1.3496e-5, 1.3496e-5, 1.8386e-5, 1.5957e-5]
flight_points = pd.DataFrame()
flight_points["mach"] = machs + machs
flight_points["altitude"] = altitudes + altitudes
flight_points["engine_setting"] = engine_settings + engine_settings
flight_points["thrust_is_regulated"] = [False] * 5 + [True] * 5
flight_points["thrust_rate"] = thrust_rates + [0.0] * 5
flight_points["thrust"] = [0.0] * 5 + thrusts
engine.compute_flight_points(flight_points)
np.testing.assert_allclose(flight_points.sfc,
expected_sfc + expected_sfc,
rtol=1e-4)
np.testing.assert_allclose(flight_points.thrust_rate,
thrust_rates + thrust_rates,
rtol=1e-4)
np.testing.assert_allclose(flight_points.thrust,
thrusts + thrusts,
rtol=1e-4)
def test_dummy_reserve():
dummy_reserve = DummyTransitionSegment(
target=FlightPoint(altitude=0.0, mach=0.0), reserve_mass_ratio=0.1
)
flight_points = dummy_reserve.compute_from(FlightPoint(altitude=0.0, mach=0.0, mass=55.0e3))
assert_allclose(flight_points.mass, [55.0e3, 55.0e3, 50.0e3])
assert_allclose(flight_points.altitude, [0.0, 0.0, 0.0])
assert_allclose(flight_points.ground_distance, [0.0, 0.0, 0.0])
assert_allclose(flight_points.mach, [0.0, 0.0, 0.0])
assert_allclose(flight_points.true_airspeed, [0.0, 0.0, 0.0], rtol=1.0e-4)
def test_ranged_route(low_speed_polar, high_speed_polar, cleanup):
engine = DummyEngine(1e5, 2.0e-5)
propulsion = FuelEngineSet(engine, 2)
total_distance = 2.0e6
kwargs = dict(propulsion=propulsion, reference_area=120.0)
initial_climb = InitialClimbPhase(**kwargs,
polar=low_speed_polar,
thrust_rate=1.0,
name="initial_climb",
time_step=0.2)
climb = ClimbPhase(
**kwargs,
polar=high_speed_polar,
thrust_rate=0.8,
target_altitude="mach",
maximum_mach=0.78,
name="climb",
time_step=5.0,
)
cruise = CruiseSegment(
**kwargs,
target=FlightPoint(ground_distance=0.0),
polar=high_speed_polar,
engine_setting=EngineSetting.CRUISE,
name=FlightPhase.CRUISE.value,
)
descent = DescentPhase(
**kwargs,
polar=high_speed_polar,
thrust_rate=0.05,
target_altitude=1500.0 * foot,
name=FlightPhase.DESCENT.value,
time_step=5.0,
)
flight_calculator = RangedRoute([initial_climb, climb], cruise, [descent],
total_distance)
assert flight_calculator.cruise_speed == ("mach", 0.78)
start = FlightPoint(true_airspeed=150.0 * knot,
altitude=100.0 * foot,
mass=70000.0,
ground_distance=100000.0)
flight_points = flight_calculator.compute_from(start)
# plot_flight(flight_points, "test_ranged_flight.png")
assert_allclose(
flight_points.iloc[-1].ground_distance,
total_distance + start.ground_distance,
atol=flight_calculator.distance_accuracy,
)
def test_acceleration_not_enough_thrust(polar):
propulsion = FuelEngineSet(DummyEngine(0.5e5, 1.0e-5), 2)
segment = SpeedChangeSegment(
target=FlightPoint(true_airspeed=250.0),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
thrust_rate=0.1,
)
assert len(
segment.compute_from(
FlightPoint(altitude=5000.0, true_airspeed=150.0, mass=70000.0)))
def test_dummy_descent_with_reserve():
dummy_descent_reserve = DummyTransitionSegment(
target=FlightPoint(altitude=0.0, mach=0.0, ground_distance=500.0e3),
mass_ratio=0.9,
reserve_mass_ratio=0.08,
)
flight_points = dummy_descent_reserve.compute_from(
FlightPoint(altitude=9.0e3, mass=60.0e3, mach=0.8))
assert_allclose(flight_points.mass, [60.0e3, 54.0e3, 50.0e3])
assert_allclose(flight_points.altitude, [9.0e3, 0.0, 0.0])
assert_allclose(flight_points.ground_distance, [0.0, 500.0e3, 500.0e3])
assert_allclose(flight_points.mach, [0.8, 0.0, 0.0])
assert_allclose(flight_points.true_airspeed, [243.04, 0.0, 0.0],
rtol=1.0e-4)
def test_climb_not_enough_thrust(polar):
propulsion = FuelEngineSet(DummyEngine(1.0e5, 1.0e-5), 2)
segment = AltitudeChangeSegment(
target=FlightPoint(altitude=10000.0, true_airspeed="constant"),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
thrust_rate=0.1,
)
assert (len(
segment.compute_from(
FlightPoint(altitude=5000.0, true_airspeed=150.0,
mass=70000.0))) == 1)
def test_dummy_climb():
dummy_climb = DummyTransitionSegment(
target=FlightPoint(altitude=9.0e3, mach=0.8, ground_distance=400.0e3), mass_ratio=0.8
)
flight_points = dummy_climb.compute_from(
FlightPoint(altitude=0.0, mass=100.0e3, mach=0.0, ground_distance=100.0e3)
)
assert_allclose(flight_points.mass, [100.0e3, 80.0e3])
assert_allclose(flight_points.altitude, [0.0, 9.0e3])
assert_allclose(flight_points.ground_distance, [100.0e3, 500.0e3])
assert_allclose(flight_points.mach, [0.0, 0.8])
assert_allclose(flight_points.true_airspeed, [0.0, 243.04], rtol=1.0e-4)
def compute_from(self, start: FlightPoint) -> pd.DataFrame:
parts = []
part_start = start
for part in self.flight_sequence:
flight_points = part.compute_from(part_start)
if len(parts) > 0 and len(flight_points) > 1:
# First point of the segment is omitted, as it is the last of previous segment.
#
# But sometimes (especially in the case of simplistic segments), the new first
# point may contain more information than the previous last one. In such case,
# it is interesting to complete the previous last one.
last_flight_points = parts[-1]
last_index = last_flight_points.index[-1]
for name in flight_points.columns:
value = last_flight_points.loc[last_index, name]
if not value:
last_flight_points.loc[last_index,
name] = flight_points.loc[0,
name]
parts.append(flight_points.iloc[1:])
else:
# But it is kept if the computed segment is the first one.
parts.append(flight_points)
part_start = FlightPoint.create(flight_points.iloc[-1])
if parts:
return pd.concat(parts).reset_index(drop=True)
def test_hold(polar):
propulsion = FuelEngineSet(DummyEngine(0.5e5, 2.0e-5), 2)
segment = HoldSegment(target=FlightPoint(time=3000.0),
propulsion=propulsion,
reference_area=120.0,
polar=polar)
flight_points = segment.compute_from(
FlightPoint(altitude=500.0, equivalent_airspeed=250.0, mass=60000.0), )
last_point = flight_points.iloc[-1]
assert_allclose(last_point.time, 3000.0)
assert_allclose(last_point.altitude, 500.0)
assert_allclose(last_point.equivalent_airspeed, 250.0, atol=0.1)
assert_allclose(last_point.mass, 57776.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 768323.0, rtol=1.0e-3)
def flight_sequence(self) -> List[Union[IFlightPart, str]]:
return [
AltitudeChangeSegment(
target=FlightPoint(equivalent_airspeed="constant", altitude=400.0 * foot),
engine_setting=EngineSetting.TAKEOFF,
**self.segment_kwargs,
),
SpeedChangeSegment(
target=FlightPoint(equivalent_airspeed=250.0 * knot),
engine_setting=EngineSetting.TAKEOFF,
**self.segment_kwargs,
),
AltitudeChangeSegment(
target=FlightPoint(equivalent_airspeed="constant", altitude=1500.0 * foot),
engine_setting=EngineSetting.TAKEOFF,
**self.segment_kwargs,
),
]
def test_taxi():
propulsion = FuelEngineSet(DummyEngine(0.5e5, 1.0e-5), 2)
segment = TaxiSegment(target=FlightPoint(time=500.0),
propulsion=propulsion,
thrust_rate=0.1)
flight_points = segment.compute_from(
FlightPoint(altitude=10.0,
true_airspeed=10.0,
mass=50000.0,
time=10000.0), )
last_point = flight_points.iloc[-1]
assert_allclose(last_point.altitude, 10.0, atol=1.0)
assert_allclose(last_point.time, 10500.0, rtol=1e-2)
assert_allclose(last_point.true_airspeed, 10.0, atol=0.1)
assert_allclose(last_point.mass, 49973.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 5000.0)
def test_climb_and_cruise_at_optimal_flight_level(polar):
propulsion = FuelEngineSet(DummyEngine(0.5e5, 3.0e-5), 2)
reference_area = 120.0
segment = ClimbAndCruiseSegment(
target=FlightPoint(
ground_distance=10.0e6,
altitude=AltitudeChangeSegment.OPTIMAL_FLIGHT_LEVEL),
propulsion=propulsion,
reference_area=reference_area,
polar=polar,
engine_setting=EngineSetting.CRUISE,
climb_segment=AltitudeChangeSegment(
target=FlightPoint(),
propulsion=propulsion,
reference_area=reference_area,
polar=polar,
thrust_rate=0.9,
engine_setting=EngineSetting.CLIMB,
),
)
flight_points = segment.compute_from(
FlightPoint(mass=70000.0,
altitude=8000.0,
mach=0.78,
ground_distance=1.0e6))
first_point = flight_points.iloc[0]
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 1.0s as time step
assert_allclose(first_point.altitude, 8000.0)
assert_allclose(first_point.thrust_rate, 0.9)
assert_allclose(first_point.true_airspeed, 240.3, atol=0.1)
assert first_point.engine_setting == EngineSetting.CLIMB
assert_allclose(flight_points.mach, 0.78)
assert_allclose(last_point.altitude, 9753.6)
assert_allclose(last_point.ground_distance, 11.0e6)
assert_allclose(last_point.time, 42659.0, rtol=1e-3)
assert_allclose(last_point.true_airspeed, 234.4, atol=0.1)
assert_allclose(last_point.mass, 48874.0, rtol=1e-4)
assert last_point.engine_setting == EngineSetting.CRUISE
def test_descent_to_fixed_altitude_at_constant_EAS(polar):
propulsion = FuelEngineSet(DummyEngine(1.0e5, 1.0e-5), 2)
flight_points = AltitudeChangeSegment(
target=FlightPoint(altitude=5000.0, equivalent_airspeed="constant"),
propulsion=propulsion,
reference_area=100.0,
polar=polar,
thrust_rate=0.1,
time_step=2.0,
).compute_from(FlightPoint(altitude=10000.0, equivalent_airspeed=200.0, mass=70000.0))
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.01s as time step
assert_allclose(last_point.altitude, 5000.0)
assert_allclose(last_point.equivalent_airspeed, 200.0)
assert_allclose(last_point.time, 821.4, rtol=1e-2)
assert_allclose(last_point.mass, 69910.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 243155.0, rtol=1e-3)
def flight_sequence(self) -> List[Union[IFlightPart, str]]:
self.segment_kwargs["engine_setting"] = EngineSetting.IDLE
return [
AltitudeChangeSegment(
target=FlightPoint(equivalent_airspeed=300.0 * knot, mach="constant"),
**self.segment_kwargs,
),
AltitudeChangeSegment(
target=FlightPoint(altitude=10000.0 * foot, equivalent_airspeed="constant"),
**self.segment_kwargs,
),
SpeedChangeSegment(
target=FlightPoint(equivalent_airspeed=250.0 * knot), **self.segment_kwargs,
),
AltitudeChangeSegment(
target=FlightPoint(altitude=self.target_altitude, equivalent_airspeed="constant"),
**self.segment_kwargs,
),
]
def test_acceleration_to_mach(polar):
propulsion = FuelEngineSet(DummyEngine(0.5e5, 1.0e-5), 2)
flight_points = SpeedChangeSegment(
target=FlightPoint(mach=0.8),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
thrust_rate=1.0,
time_step=0.2,
).compute_from(
FlightPoint(altitude=1000.0, true_airspeed=150.0, mass=70000.0))
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.01s as time step
assert_allclose(last_point.time, 138.0, rtol=1e-2)
assert_allclose(last_point.altitude, 1000.0)
assert_allclose(last_point.mach, 0.8, atol=1e-5)
assert_allclose(last_point.mass, 69862.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 29470.0, rtol=1e-3)
def test_climb_optimal_flight_level_at_fixed_mach(polar):
propulsion = FuelEngineSet(DummyEngine(1.0e5, 1.0e-5), 2)
flight_points = AltitudeChangeSegment(
target=FlightPoint(altitude=AltitudeChangeSegment.OPTIMAL_FLIGHT_LEVEL, mach="constant"),
propulsion=propulsion,
reference_area=120.0,
polar=polar,
thrust_rate=1.0,
time_step=2.0,
).compute_from(FlightPoint(altitude=5000.0, mach=0.82, mass=70000.0))
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.01s as time step
assert_allclose(flight_points.mach, 0.82)
assert_allclose(last_point.altitude / foot, 32000.0, atol=0.1)
assert_allclose(last_point.time, 77.5, rtol=1e-2)
assert_allclose(last_point.true_airspeed, 246.44, rtol=1e-4)
assert_allclose(last_point.mass, 69843.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 19179.0, rtol=1e-3)
def _compute_taxi_out(self, inputs, outputs, propulsion_model):
"""
Computes the taxi-out segment.
"""
start_of_taxi_out = FlightPoint(
altitude=inputs[self._mission_vars.TAKEOFF_ALTITUDE.value],
true_airspeed=0.0,
# start mass is irrelevant here as long it does not get negative during computation.
mass=inputs[self._mission_vars.TOW.value],
)
taxi_segment = TaxiSegment(
target=FlightPoint(time=inputs[self._mission_vars.TAXI_OUT_DURATION.value],),
thrust_rate=inputs[self._mission_vars.TAXI_OUT_THRUST_RATE.value],
propulsion=propulsion_model,
)
flight_points = taxi_segment.compute_from(start_of_taxi_out)
end_of_taxi_out = flight_points.iloc[-1]
outputs[self._mission_vars.TAXI_OUT_FUEL.value] = (
start_of_taxi_out.mass - end_of_taxi_out.mass
)
def test_descent_to_fixed_EAS_at_constant_mach(polar):
propulsion = FuelEngineSet(DummyEngine(1.0e5, 1.0e-5), 2)
flight_points = AltitudeChangeSegment(
target=FlightPoint(equivalent_airspeed=150.0, mach="constant"),
propulsion=propulsion,
reference_area=100.0,
polar=polar,
thrust_rate=0.1,
# time_step=5.0, # we use default time step
).compute_from(FlightPoint(altitude=10000.0, mass=70000.0, mach=0.78))
last_point = flight_points.iloc[-1]
# Note: reference values are obtained by running the process with 0.01s as time step
assert_allclose(last_point.equivalent_airspeed, 150.0)
assert_allclose(last_point.mach, 0.78)
assert_allclose(last_point.time, 343.6, rtol=1e-2)
assert_allclose(last_point.altitude, 8654.0, atol=1.0)
assert_allclose(last_point.true_airspeed, 238.1, atol=0.1)
assert_allclose(last_point.mass, 69962.0, rtol=1e-4)
assert_allclose(last_point.ground_distance, 81042.0, rtol=1e-3)
