def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:at25percent:x", val=np.nan, units="m")
self.add_input("data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25", val=np.nan, units="m")
self.add_input("data:geometry:horizontal_tail:area", val=np.nan, units="m**2")
self.add_input("data:geometry:propulsion:nacelle:height", val=np.nan, units="m")
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
self.add_input("data:weight:airframe:landing_gear:main:CG:x", val=np.nan, units="m")
self.add_input("data:weight:aircraft_empty:CG:z", val=np.nan, units="m")
self.add_input("data:weight:propulsion:engine:CG:z", val=np.nan, units="m")
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean", val=np.nan)
self.add_input("data:aerodynamics:aircraft:takeoff:CL_max", val=np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean", val=np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", val=np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CM", val=np.nan)
self.add_input("data:aerodynamics:horizontal_tail:low_speed:CL_alpha_isolated", val=np.nan, units="rad**-1")
self.add_input("data:aerodynamics:horizontal_tail:efficiency", val=np.nan)
self.add_input("takeoff:cl_htp", val=np.nan)
self.add_input("takeoff:cm_wing", val=np.nan)
self.add_input("low_speed:cl_alpha_htp", val=np.nan)
self.add_output("data:handling_qualities:to_rotation_limit:x", units="m")
self.add_output("data:handling_qualities:to_rotation_limit:MAC_position")
self.declare_partials("*", "*", method="fd")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:aerodynamics:aircraft:cruise:CD0", np.nan)
self.add_input("data:aerodynamics:aircraft:cruise:induced_drag_coefficient", np.nan)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
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:holding:fuel", 0.0, 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:thrust_rate", np.nan)
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.add_output("data:mission:sizing:main_route:climb:v_cas", units="m/s")
self.declare_partials(
"*",
[
"data:aerodynamics:aircraft:cruise:CD0",
"data:aerodynamics:aircraft:cruise:induced_drag_coefficient",
"data:geometry:wing:area",
"data:weight:aircraft:MTOW",
"data:mission:sizing:taxi_out:fuel",
"data:mission:sizing:holding:fuel",
"data:mission:sizing:takeoff:fuel",
"data:mission:sizing:initial_climb:fuel",
],
method="fd",
)
class ComputeEngineWeight(ExplicitComponent):
"""
Engine weight estimation calling wrapper
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_output("data:weight:propulsion:engine:mass", units="lb")
self.declare_partials("*", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
b1 = propulsion_model.compute_weight()
outputs["data:weight:propulsion:engine:mass"] = b1
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:mission:sizing:main_route:cruise:altitude", np.nan, units="m")
self.add_input("data:TLAR:cruise_mach", np.nan)
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:aerodynamics:aircraft:cruise:L_D_max", np.nan)
self.add_input("data:geometry:propulsion:engine:count", 2)
self.add_input("settings:mission:sizing:breguet:climb:mass_ratio", 0.97)
self.add_input("settings:mission:sizing:breguet:descent:mass_ratio", 0.98)
self.add_input("settings:mission:sizing:breguet:reserve:mass_ratio", 0.06)
self.add_input("data:TLAR:range", np.nan, units="m")
self.add_input("settings:mission:sizing:breguet:climb_descent_distance", 500.0e3, units="m")
self.add_output("data:mission:sizing:main_route:climb:distance", units="m", ref=1e3)
self.add_output("data:mission:sizing:main_route:cruise:distance", units="m", ref=1e3)
self.add_output("data:mission:sizing:main_route:descent:distance", units="m", ref=1e3)
self.add_output("data:mission:sizing:ZFW", units="kg", ref=1e4)
self.add_output("data:mission:sizing:fuel", units="kg", ref=1e4)
self.add_output("data:mission:sizing:main_route:fuel", units="kg", ref=1e4)
self.add_output("data:mission:sizing:main_route:climb:fuel", units="kg", ref=1e4)
self.add_output("data:mission:sizing:main_route:cruise:fuel", units="kg", ref=1e4)
self.add_output("data:mission:sizing:main_route:descent:fuel", units="kg", ref=1e4)
self.add_output("data:mission:sizing:fuel_reserve", units="kg", ref=1e4)
self.add_output("data:propulsion:SFC", units="kg/s/N", ref=1e-4)
self.add_output("data:propulsion:thrust_rate", lower=0.0, upper=1.0)
self.add_output("data:propulsion:thrust", units="N", ref=1e5)
self.declare_partials("*", "*", method="fd")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
if self.options["taxi_out"]:
self.add_input("data:mission:sizing:taxi_out:thrust_rate", np.nan)
self.add_input("data:mission:sizing:taxi_out:duration",
np.nan,
units="s")
self.add_input("data:mission:sizing:taxi_out:speed",
np.nan,
units="m/s")
self.add_output("data:mission:sizing:taxi_out:fuel", units='kg')
else:
self.add_input("data:mission:sizing:taxi_in:thrust_rate", np.nan)
self.add_input("data:mission:sizing:taxi_in:duration",
np.nan,
units="s")
self.add_input("data:mission:sizing:taxi_in:speed",
np.nan,
units="m/s")
self.add_output("data:mission:sizing:taxi_in:fuel", units='kg')
self.declare_partials("*", "*", method="fd")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:weight:aircraft:CG:aft:MAC_position", val=np.nan)
self.add_input("data:aerodynamics:fuselage:cruise:CnBeta", val=np.nan)
self.add_input("data:aerodynamics:vertical_tail:cruise:CL_alpha",
val=np.nan,
units="rad**-1")
self.add_input("data:TLAR:v_approach", val=np.nan, units="m/s")
self.add_input(
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input("data:geometry:propulsion:nacelle:wet_area",
val=np.nan,
units="m**2")
self.add_input("data:geometry:propulsion:nacelle:y",
val=np.nan,
units="m")
self.add_output("data:geometry:vertical_tail:area",
val=2.5,
units="m**2")
self.declare_partials(
"*",
"*",
method="fd",
)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_alpha", np.nan, units="rad**-1")
self.add_input("data:aerodynamics:aircraft:low_speed:CD0", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CD", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:induced_drag_coefficient", np.nan)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
self.add_input("data:geometry:wing:span", np.nan, units="m")
self.add_input("data:geometry:landing_gear:height", np.nan, units="m")
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:thrust_rate", np.nan)
self.add_input("data:mission:sizing:takeoff:friction_coefficient_no_brake", np.nan)
self.add_input("vr:speed", np.nan, units='m/s')
self.add_input("v2:angle", np.nan, units='rad')
self.add_output("data:mission:sizing:takeoff:VR", units='m/s')
self.add_output("data:mission:sizing:takeoff:VLOF", units='m/s')
self.add_output("data:mission:sizing:takeoff:V2", units='m/s')
self.add_output("data:mission:sizing:takeoff:TOFL", units='m')
self.add_output("data:mission:sizing:takeoff:duration", units='s')
self.add_output("data:mission:sizing:takeoff:fuel", units='kg')
self.add_output("data:mission:sizing:initial_climb:fuel", units='kg')
self.declare_partials("*", "*", method="fd")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_output("data:weight:propulsion:engine_oil:mass", units="lb")
class _compute_taxi(om.ExplicitComponent):
"""
Compute the fuel consumption for taxi based on speed and duration.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
self.options.declare("taxi_out", default=True, types=bool)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
if self.options["taxi_out"]:
self.add_input("data:mission:sizing:taxi_out:thrust_rate", np.nan)
self.add_input("data:mission:sizing:taxi_out:duration", np.nan, units="s")
self.add_input("data:mission:sizing:taxi_out:speed", np.nan, units="m/s")
self.add_output("data:mission:sizing:taxi_out:fuel", units='kg')
else:
self.add_input("data:mission:sizing:taxi_in:thrust_rate", np.nan)
self.add_input("data:mission:sizing:taxi_in:duration", np.nan, units="s")
self.add_input("data:mission:sizing:taxi_in:speed", np.nan, units="m/s")
self.add_output("data:mission:sizing:taxi_in:fuel", units='kg')
self.declare_partials("*", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
if self.options["taxi_out"]:
thrust_rate = inputs["data:mission:sizing:taxi_out:thrust_rate"]
duration = inputs["data:mission:sizing:taxi_out:duration"]
mach = inputs["data:mission:sizing:taxi_out:speed"]/Atmosphere(0.0).speed_of_sound
else:
thrust_rate = inputs["data:mission:sizing:taxi_in:thrust_rate"]
duration = inputs["data:mission:sizing:taxi_in:duration"]
mach = inputs["data:mission:sizing:taxi_in:speed"] / Atmosphere(0.0).speed_of_sound
# FIXME: no specific settings for taxi (to be changed in fastoad\constants.py)
flight_point = FlightPoint(
mach=mach, altitude=0.0, engine_setting=EngineSetting.TAKEOFF,
thrust_rate=thrust_rate
)
propulsion_model.compute_flight_points(flight_point)
fuel_mass = propulsion_model.get_consumed_mass(flight_point, duration)
if self.options["taxi_out"]:
outputs["data:mission:sizing:taxi_out:fuel"] = fuel_mass
else:
outputs["data:mission:sizing:taxi_in:fuel"] = fuel_mass
class ComputeUnusableFuelWeight(ExplicitComponent):
"""
Weight estimation for motor oil
Based on : Wells, Douglas P., Bryce L. Horvath, and Linwood A. McCullers. "The Flight Optimization System Weights
Estimation Method." (2017). Equation 121
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="ft**2")
self.add_input("data:weight:aircraft:MFW", val=np.nan, units="lb")
self.add_output("data:weight:propulsion:unusable_fuel:mass",
units="lb")
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
n_eng = inputs["data:geometry:propulsion:count"]
wing_area = inputs["data:geometry:wing:area"]
mfw = inputs["data:weight:aircraft:MFW"]
n_tank = 2.0
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), n_eng)
flight_point = FlightPoint(
mach=0.0,
altitude=0.0,
engine_setting=EngineSetting.TAKEOFF,
thrust_rate=1.0) # with engine_setting as EngineSetting
propulsion_model.compute_flight_points(flight_point)
sl_thrust_newton = float(flight_point.thrust)
sl_thrust_lbs = sl_thrust_newton / lbf
sl_thrust_lbs_per_engine = sl_thrust_lbs / n_eng
b3 = 11.5 * n_eng * sl_thrust_lbs_per_engine ** 0.2 + \
0.07 * wing_area + \
1.6 * n_tank * mfw ** 0.28
outputs["data:weight:propulsion:unusable_fuel:mass"] = b3
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:cabin:luggage:mass_max",
val=np.nan,
units="kg")
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:geometry:propulsion:count", val=np.nan)
self.add_input("data:geometry:cabin:seats:passenger:NPAX_max",
val=np.nan)
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:wing:MAC:at25percent:x",
val=np.nan,
units="m")
self.add_input("data:geometry:fuselage:front_length",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:pilot:length",
val=np.nan,
units="m")
self.add_input("data:geometry:fuselage:PAX_length",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:count_by_row",
val=np.nan)
self.add_input("data:geometry:cabin:seats:passenger:length",
val=np.nan,
units="m")
self.add_input("data:weight:payload:rear_fret:CG:x",
val=np.nan,
units="m")
self.add_input("data:weight:aircraft_empty:CG:x",
val=np.nan,
units="m")
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
self.add_input("data:weight:aircraft_empty:mass",
val=np.nan,
units="kg")
self.add_input("data:weight:propulsion:unusable_fuel:mass",
val=np.nan,
units="kg")
self.add_input("data:weight:propulsion:tank:CG:x",
val=np.nan,
units="m")
self.add_input("data:weight:aircraft:MFW", val=np.nan, units="kg")
self.add_output(
"data:weight:aircraft:CG:flight_condition:max:MAC_position")
self.add_output(
"data:weight:aircraft:CG:flight_condition:min:MAC_position")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:TLAR:NPAX", val=np.nan)
self.add_input("data:geometry:cabin:seats:pilot:length",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:pilot:width",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:length",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:width",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:count_by_row",
val=np.nan)
self.add_input("data:geometry:cabin:aisle_width",
val=np.nan,
units="m")
self.add_input("data:geometry:propulsion:layout", val=np.nan)
self.add_input("data:geometry:wing:MAC:at25percent:x",
val=np.nan,
units="m")
self.add_input(
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input(
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input("data:geometry:horizontal_tail:MAC:length",
val=np.nan,
units="m")
self.add_input("data:geometry:vertical_tail:MAC:length",
val=np.nan,
units="m")
self.add_output("data:geometry:cabin:NPAX")
self.add_output("data:geometry:fuselage:length", units="m")
self.add_output("data:geometry:fuselage:maximum_width", units="m")
self.add_output("data:geometry:fuselage:maximum_height", units="m")
self.add_output("data:geometry:fuselage:front_length", units="m")
self.add_output("data:geometry:fuselage:rear_length", units="m")
self.add_output("data:geometry:fuselage:PAX_length", units="m")
self.add_output("data:geometry:cabin:length", units="m")
self.add_output("data:geometry:fuselage:wet_area", units="m**2")
self.add_output("data:geometry:fuselage:luggage_length", units="m")
self.declare_partials(
"*", "*",
method="fd") # FIXME: declare proper partials without int values
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="ft**2")
self.add_input("data:weight:aircraft:MFW", val=np.nan, units="lb")
self.add_output("data:weight:propulsion:unusable_fuel:mass",
units="lb")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("settings:weight:aircraft:CG:range", val=0.3)
self.add_input("data:mission:sizing:takeoff:thrust_rate", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:MAC:at25percent:x",
val=np.nan,
units="m")
self.add_input(
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:propulsion:nacelle:height",
val=np.nan,
units="m")
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
self.add_input("data:weight:aircraft:MLW", val=np.nan, units="kg")
self.add_input("data:weight:aircraft:CG:aft:x", val=np.nan, units="m")
self.add_input("data:weight:airframe:landing_gear:main:CG:x",
val=np.nan,
units="m")
self.add_input("data:aerodynamics:aircraft:low_speed:CL0_clean",
val=np.nan)
self.add_input("data:aerodynamics:aircraft:landing:CL_max", val=np.nan)
self.add_input("data:aerodynamics:aircraft:takeoff:CL_max", val=np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean",
val=np.nan)
self.add_input("data:aerodynamics:flaps:landing:CL", val=np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", val=np.nan)
self.add_input("data:aerodynamics:flaps:landing:CM", val=np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CM", val=np.nan)
self.add_input("data:aerodynamics:horizontal_tail:low_speed:CL_alpha",
val=np.nan,
units="rad**-1")
self.add_input("landing:cl_ht", val=np.nan)
self.add_input("takeoff:cl_ht", val=np.nan)
self.add_input("landing:cm_wing", val=np.nan)
self.add_input("takeoff:cm_wing", val=np.nan)
self.add_input("low_speed:cl_alpha_ht", val=np.nan)
self.add_output("data:geometry:horizontal_tail:area",
val=4.0,
units="m**2")
self.declare_partials(
"*", "*",
method="fd") # FIXME: write partial avoiding discrete parameters
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
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_output("data:TLAR:v_max_sl", units="m/s")
self.declare_partials("*", "*", method="fd")
class ComputeOilWeight(ExplicitComponent):
"""
Weight estimation for motor oil
Based on : Wells, Douglas P., Bryce L. Horvath, and Linwood A. McCullers. "The Flight Optimization System Weights
Estimation Method." (2017). Equation 123
Not used since already included in the engine installed weight but left there in case
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_output("data:weight:propulsion:engine_oil:mass", units="lb")
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
n_eng = inputs["data:geometry:propulsion:count"]
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), n_eng)
flight_point = FlightPoint(
mach=0.0,
altitude=0.0,
engine_setting=EngineSetting.TAKEOFF,
thrust_rate=1.0) # with engine_setting as EngineSetting
propulsion_model.compute_flight_points(flight_point)
# This should give the UNINSTALLED weight
sl_thrust_newton = float(flight_point.thrust)
sl_thrust_lbs = sl_thrust_newton / lbf
b1_2 = 0.082 * n_eng * sl_thrust_lbs**0.65
outputs["data:weight:propulsion:engine_oil:mass"] = b1_2
class ComputeEngineWeight(ExplicitComponent):
"""
Engine weight estimation calling wrapper
Based on : Raymer Daniel. Aircraft Design: A Conceptual Approach. AIAA
Education Series 1996 for installed engine weight, table 15.2
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_output("data:weight:propulsion:engine:mass", units="lb")
self.declare_partials("*", "*", method="fd")
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs),
inputs["data:geometry:propulsion:count"])
# This should give the UNINSTALLED weight
uninstalled_engine_weight = propulsion_model.compute_weight()
b1 = 1.4 * uninstalled_engine_weight
outputs["data:weight:propulsion:engine:mass"] = b1
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("data:geometry:propulsion:layout", val=np.nan)
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
if self.options["low_speed_aero"]:
self.add_input("data:aerodynamics:low_speed:mach", val=np.nan)
self.add_input("data:aerodynamics:low_speed:unit_reynolds",
val=np.nan)
self.add_output("data:aerodynamics:nacelles:low_speed:CD0")
else:
self.add_input("data:aerodynamics:cruise:mach", val=np.nan)
self.add_input("data:aerodynamics:cruise:unit_reynolds",
val=np.nan)
self.add_output("data:aerodynamics:nacelles:cruise:CD0")
self.declare_partials("*", "*", method="fd")
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:layout", val=np.nan)
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:propulsion:y_ratio", val=np.nan)
self.add_input("data:geometry:fuselage:maximum_width",
val=np.nan,
units="m")
self.add_output("data:geometry:propulsion:nacelle:length", units="m")
self.add_output("data:geometry:propulsion:nacelle:height", units="m")
self.add_output("data:geometry:propulsion:nacelle:width", units="m")
self.add_output("data:geometry:propulsion:nacelle:wet_area",
units="m**2")
self.add_output("data:geometry:landing_gear:height", units="m")
self.add_output("data:geometry:propulsion:nacelle:y", units="m")
self.declare_partials("*", "*", method="fd")
def setup(self):
super().setup()
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
self.add_input("data:weight:aircraft:MLW", val=np.nan, units="kg")
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input(
"data:aerodynamics:wing:low_speed:induced_drag_coefficient",
val=np.nan)
self.add_input(
"data:aerodynamics:horizontal_tail:low_speed:induced_drag_coefficient",
val=np.nan)
self.add_input("data:aerodynamics:aircraft:landing:CL_max", val=np.nan)
self.add_input("data:aerodynamics:aircraft:low_speed:CD0", val=np.nan)
self.add_output("data:handling_qualities:balked_landing_limit:x",
val=4.0,
units="m")
self.add_output(
"data:handling_qualities:balked_landing_limit:MAC_position",
val=np.nan)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self._mission_input = self.options["mission_file_path"]
if not self._mission_input:
with path(resources, "sizing_mission.yml") as mission_input_file:
self._mission_input = MissionDefinition(mission_input_file)
self._mission = MissionWrapper(self._mission_input)
self._mission.setup(self)
self.add_input("data:geometry:propulsion:engine:count", 2)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
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:altitude",
np.nan,
units="m")
self.add_input("data:mission:sizing:takeoff:fuel", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:V2", np.nan, units="m/s")
self.add_output("data:mission:sizing:ZFW", units="kg")
self.declare_partials(["*"], ["*"])
@staticmethod
def get_model(inputs) -> IPropulsion:
engine_params = {
"max_power": inputs["data:propulsion:IC_engine:max_power"],
"design_altitude": inputs["data:mission:sizing:main_route:cruise:altitude"],
"design_speed": inputs["data:TLAR:v_cruise"],
"fuel_type": inputs["data:propulsion:IC_engine:fuel_type"],
"strokes_nb": inputs["data:propulsion:IC_engine:strokes_nb"],
"prop_layout": inputs["data:geometry:propulsion:layout"]
}
return DummyEngine(**engine_params)
BundleLoader().context.install_bundle(__name__).start()
def test_update_vt_area():
""" Tests computation of the vertical tail area """
# Research independent input value in .xml file
ivc = get_indep_var_comp(list_inputs(UpdateVTArea(propulsion_id=ENGINE_WRAPPER)), __file__, XML_FILE)
ivc.add_output("data:weight:aircraft:CG:aft:MAC_position", 0.364924)
ivc.add_output("data:aerodynamics:fuselage:cruise:CnBeta", -0.0599)
# Run problem and check obtained value(s) is/(are) correct
register_wrappers()
problem = run_system(UpdateVTArea(propulsion_id=ENGINE_WRAPPER), ivc)
vt_area = problem.get_val("data:geometry:vertical_tail:area", units="m**2")
assert vt_area == pytest.approx(2.44, abs=1e-2) # old-version obtained value 2.4m²
class ComputeNacelleGeometry(om.ExplicitComponent):
# TODO: Document equations. Cite sources
""" Nacelle and pylon geometry estimation """
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:wing:span", val=np.nan, units="m")
self.add_input("data:geometry:propulsion:y_ratio", val=np.nan)
self.add_input("data:geometry:fuselage:maximum_width",
val=np.nan,
units="m")
self.add_output("data:geometry:propulsion:nacelle:length", units="m")
self.add_output("data:geometry:propulsion:nacelle:height", units="m")
self.add_output("data:geometry:propulsion:nacelle:width", units="m")
self.add_output("data:geometry:propulsion:nacelle:wet_area",
units="m**2")
self.add_output("data:geometry:propulsion:propeller:depth", units="m")
self.add_output("data:geometry:propulsion:propeller:diameter",
units="m")
self.add_output("data:geometry:landing_gear:height", units="m")
self.add_output("data:geometry:propulsion:nacelle:y", units="m")
self.declare_partials("*", "*", method="fd")
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), 1.0)
prop_layout = inputs["data:geometry:propulsion:layout"]
span = inputs["data:geometry:wing:span"]
y_ratio = inputs["data:geometry:propulsion:y_ratio"]
b_f = inputs["data:geometry:fuselage:maximum_width"]
nac_height, nac_width, nac_length, nac_wet_area, prop_dia, prop_depth = propulsion_model.compute_dimensions(
)
if prop_layout == 1.0:
y_nacelle = y_ratio * span / 2
elif prop_layout == 2.0:
y_nacelle = b_f / 2 + 0.8 * nac_width
elif prop_layout == 3.0:
y_nacelle = 0.0
else:
y_nacelle = y_ratio * span / 2
warnings.warn(
'Propulsion layout {} not implemented in model, replaced by layout 1!'
.format(prop_layout))
lg_height = 0.41 * prop_dia
outputs["data:geometry:propulsion:nacelle:length"] = nac_length
outputs["data:geometry:propulsion:nacelle:height"] = nac_height
outputs["data:geometry:propulsion:nacelle:width"] = nac_width
outputs["data:geometry:propulsion:nacelle:wet_area"] = nac_wet_area
outputs["data:geometry:propulsion:propeller:depth"] = prop_depth
outputs["data:geometry:propulsion:propeller:diameter"] = prop_dia
outputs["data:geometry:landing_gear:height"] = lg_height
outputs["data:geometry:propulsion:nacelle:y"] = y_nacelle
class ComputeVh(om.ExplicitComponent):
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
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_output("data:TLAR:v_max_sl", units="kn")
self.declare_partials("*", "*", method="fd")
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
# The maximum Sea Level flight velocity is computed using a method which finds for which speed
# the thrust required for flight (drag) is equal to the thrust available
design_mass = inputs["data:weight:aircraft:MTOW"]
Vh = self.max_speed(inputs, 0.0, design_mass)
outputs["data:TLAR:v_max_sl"] = Vh
def max_speed(self, inputs, altitude, mass):
# noinspection PyTypeChecker
roots = optimize.fsolve(self.delta_axial_load,
300.0,
args=(inputs, altitude, mass))[0]
return np.max(roots[roots > 0.0])
def delta_axial_load(self, air_speed, inputs, altitude, mass):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs),
inputs["data:geometry:propulsion:count"])
wing_area = inputs["data:geometry:wing:area"]
cd0 = inputs["data:aerodynamics:aircraft:cruise:CD0"]
coef_k = inputs[
"data:aerodynamics:wing:cruise:induced_drag_coefficient"]
# Get the available thrust from propulsion system
atm = Atmosphere(altitude, altitude_in_feet=False)
flight_point = FlightPoint(mach=air_speed / atm.speed_of_sound,
altitude=altitude,
engine_setting=EngineSetting.TAKEOFF,
thrust_rate=1.0)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
# Get the necessary thrust to overcome
cl = (mass * g) / (0.5 * atm.density * wing_area * air_speed**2.0)
cd = cd0 + coef_k * cl**2.0
drag = 0.5 * atm.density * wing_area * cd * air_speed**2.0
return thrust - drag
class ComputeFuselageGeometryCabinSizing(ExplicitComponent):
# TODO: Document equations. Cite sources
""" Geometry of fuselage part A - Cabin (Commercial) estimation """
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:TLAR:NPAX", val=np.nan)
self.add_input("data:geometry:cabin:seats:pilot:length",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:pilot:width",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:length",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:width",
val=np.nan,
units="m")
self.add_input("data:geometry:cabin:seats:passenger:count_by_row",
val=np.nan)
self.add_input("data:geometry:cabin:aisle_width",
val=np.nan,
units="m")
self.add_input("data:geometry:propulsion:layout", val=np.nan)
self.add_input("data:geometry:wing:MAC:at25percent:x",
val=np.nan,
units="m")
self.add_input(
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input(
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input("data:geometry:horizontal_tail:MAC:length",
val=np.nan,
units="m")
self.add_input("data:geometry:vertical_tail:MAC:length",
val=np.nan,
units="m")
self.add_output("data:geometry:cabin:NPAX")
self.add_output("data:geometry:fuselage:length", units="m")
self.add_output("data:geometry:fuselage:maximum_width", units="m")
self.add_output("data:geometry:fuselage:maximum_height", units="m")
self.add_output("data:geometry:fuselage:front_length", units="m")
self.add_output("data:geometry:fuselage:rear_length", units="m")
self.add_output("data:geometry:fuselage:PAX_length", units="m")
self.add_output("data:geometry:cabin:length", units="m")
self.add_output("data:geometry:fuselage:wet_area", units="m**2")
self.add_output("data:geometry:fuselage:luggage_length", units="m")
self.declare_partials(
"*", "*",
method="fd") # FIXME: declare proper partials without int values
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), 1.0)
npax = inputs["data:TLAR:NPAX"]
l_pilot_seats = inputs["data:geometry:cabin:seats:pilot:length"]
w_pilot_seats = inputs["data:geometry:cabin:seats:pilot:width"]
l_pass_seats = inputs["data:geometry:cabin:seats:passenger:length"]
w_pass_seats = inputs["data:geometry:cabin:seats:passenger:width"]
seats_p_row = inputs[
"data:geometry:cabin:seats:passenger:count_by_row"]
w_aisle = inputs["data:geometry:cabin:aisle_width"]
prop_layout = inputs["data:geometry:propulsion:layout"]
fa_length = inputs["data:geometry:wing:MAC:at25percent:x"]
ht_lp = inputs[
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25"]
vt_lp = inputs[
"data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25"]
ht_length = inputs["data:geometry:horizontal_tail:MAC:length"]
vt_length = inputs["data:geometry:vertical_tail:MAC:length"]
# Length of instrument panel
l_instr = 0.7
# Length of pax cabin
# noinspection PyBroadException
try:
npax_1 = math.ceil(npax / seats_p_row) * seats_p_row
except:
npax_1 = npax
n_rows = npax_1 / seats_p_row
lpax = l_pilot_seats + n_rows * l_pass_seats
# Cabin width considered is for side by side seats
wcabin = max(2 * w_pilot_seats, seats_p_row * w_pass_seats + w_aisle)
r_i = wcabin / 2
radius = 1.06 * r_i
# Cylindrical fuselage
b_f = 2 * radius
# 0.14m is the distance between both lobe centers of the fuselage
h_f = b_f + 0.14
# Luggage length
l_lug = npax_1 * 0.20 / (math.pi * radius**2)
# Cabin total length
cabin_length = l_instr + lpax + l_lug
# Calculate nose length
if prop_layout == 3.0: # engine located in nose
_, _, propulsion_length, _ = propulsion_model.compute_dimensions()
lav = propulsion_length
else:
lav = 1.7 * h_f
# Calculate fuselage length
fus_length = fa_length + max(ht_lp + 0.75 * ht_length,
vt_lp + 0.75 * vt_length)
lar = fus_length - (lav + cabin_length)
# Calculate wet area
fus_dia = math.sqrt(b_f * h_f) # equivalent diameter of the fuselage
cyl_length = fus_length - lav - lar
wet_area_nose = 2.45 * fus_dia * lav
wet_area_cyl = 3.1416 * fus_dia * cyl_length
wet_area_tail = 2.3 * fus_dia * lar
wet_area_fus = (wet_area_nose + wet_area_cyl + wet_area_tail)
outputs["data:geometry:cabin:NPAX"] = npax_1
outputs["data:geometry:fuselage:length"] = fus_length
outputs["data:geometry:fuselage:maximum_width"] = b_f
outputs["data:geometry:fuselage:maximum_height"] = h_f
outputs["data:geometry:fuselage:front_length"] = lav
outputs["data:geometry:fuselage:rear_length"] = lar
outputs["data:geometry:fuselage:PAX_length"] = lpax
outputs["data:geometry:cabin:length"] = cabin_length
outputs["data:geometry:fuselage:wet_area"] = wet_area_fus
outputs["data:geometry:fuselage:luggage_length"] = l_lug
class _simulate_takeoff(om.ExplicitComponent):
"""
Simulate take-off from 0m/s speed to safety height using input VR.
Fuel burn is supposed negligible : mass = MTOW.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_alpha", np.nan, units="rad**-1")
self.add_input("data:aerodynamics:aircraft:low_speed:CD0", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CD", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:induced_drag_coefficient", np.nan)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
self.add_input("data:geometry:wing:span", np.nan, units="m")
self.add_input("data:geometry:landing_gear:height", np.nan, units="m")
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:thrust_rate", np.nan)
self.add_input("data:mission:sizing:takeoff:friction_coefficient_no_brake", np.nan)
self.add_input("vr:speed", np.nan, units='m/s')
self.add_input("v2:angle", np.nan, units='rad')
self.add_output("data:mission:sizing:takeoff:VR", units='m/s')
self.add_output("data:mission:sizing:takeoff:VLOF", units='m/s')
self.add_output("data:mission:sizing:takeoff:V2", units='m/s')
self.add_output("data:mission:sizing:takeoff:TOFL", units='m')
self.add_output("data:mission:sizing:takeoff:duration", units='s')
self.add_output("data:mission:sizing:takeoff:fuel", units='kg')
self.add_output("data:mission:sizing:initial_climb:fuel", units='kg')
self.declare_partials("*", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
cl_max_clean = inputs["data:aerodynamics:wing:low_speed:CL_max_clean"]
cl0 = inputs["data:aerodynamics:wing:low_speed:CL0_clean"] + inputs["data:aerodynamics:flaps:takeoff:CL"]
cl_alpha = inputs["data:aerodynamics:wing:low_speed:CL_alpha"]
cd0 = inputs["data:aerodynamics:aircraft:low_speed:CD0"] + inputs["data:aerodynamics:flaps:takeoff:CD"]
coef_k = inputs["data:aerodynamics:wing:low_speed:induced_drag_coefficient"]
wing_area = inputs["data:geometry:wing:area"]
wing_span = inputs["data:geometry:wing:span"]
lg_height = inputs["data:geometry:landing_gear:height"]
mtow = inputs["data:weight:aircraft:MTOW"]
thrust_rate = inputs["data:mission:sizing:takeoff:thrust_rate"]
friction_coeff = inputs["data:mission:sizing:takeoff:friction_coefficient_no_brake"]
alpha_v2 = float(inputs["v2:angle"])
# Define ground factor effect on Drag
k_ground = lambda altitude: (
33. * ((lg_height + altitude) / wing_span) ** 1.5
/ (1. + 33. * ((lg_height + altitude) / wing_span) ** 1.5)
)
# Determine rotation speed from regulation CS23.51
vs1 = math.sqrt((mtow * g) / (0.5 * Atmosphere(0).density * wing_area * cl_max_clean))
if inputs["data:geometry:propulsion:count"] == 1.0:
k = 1.0
else:
k = 1.1
vr = max(k * vs1, float(inputs["vr:speed"]))
# Start calculation of flight from null speed to 35ft high
alpha_t = 0.0
gamma_t = 0.0
v_t = 0.0
altitude_t = 0.0
distance_t = 0.0
mass_fuel1_t = 0.0
mass_fuel2_t = 0.0
time_t = 0.0
vloff = 0.0
climb = False
while altitude_t < SAFETY_HEIGHT:
# Estimation of thrust
atm = Atmosphere(altitude_t, altitude_in_feet=False)
flight_point = FlightPoint(
mach=max(v_t, vr) / atm.speed_of_sound, altitude=altitude_t, engine_setting=EngineSetting.TAKEOFF,
thrust_rate=thrust_rate
)
# FIXME: (speed increased to vr to have feasible consumptions)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
# Calculate lift and drag
cl = cl0 + cl_alpha * alpha_t
lift = 0.5 * atm.density * wing_area * cl * v_t ** 2
cd = cd0 + k_ground(altitude_t) * coef_k * cl ** 2
drag = 0.5 * atm.density * wing_area * cd * v_t ** 2
# Check if lift-off condition reached
if ((lift + thrust * math.sin(alpha_t) - mtow * g * math.cos(gamma_t)) >= 0.0) and not climb:
climb = True
vloff = v_t
# Calculate acceleration on x/z air axis
if climb:
acc_z = (lift + thrust * math.sin(alpha_t) - mtow * g * math.cos(gamma_t)) / mtow
acc_x = (thrust * math.cos(alpha_t) - mtow * g * math.sin(gamma_t) - drag) / mtow
else:
friction = (mtow * g - lift - thrust * math.sin(alpha_t)) * friction_coeff
acc_z = 0.0
acc_x = (thrust * math.cos(alpha_t) - drag - friction) / mtow
# Calculate gamma change and new speed
delta_gamma = math.atan((acc_z * TIME_STEP) / (v_t + acc_x * TIME_STEP))
v_t_new = math.sqrt((acc_z * TIME_STEP) ** 2 + (v_t + acc_x * TIME_STEP) ** 2)
# Trapezoidal integration on distance/altitude
delta_altitude = (v_t_new * math.sin(gamma_t + delta_gamma) + v_t * math.sin(gamma_t)) / 2 * TIME_STEP
delta_distance = (v_t_new * math.cos(gamma_t + delta_gamma) + v_t * math.cos(gamma_t)) / 2 * TIME_STEP
# Update temporal values
if v_t >= vr:
alpha_t = min(alpha_v2, alpha_t + ALPHA_RATE * TIME_STEP)
gamma_t = gamma_t + delta_gamma
altitude_t = altitude_t + delta_altitude
if not climb:
mass_fuel1_t += propulsion_model.get_consumed_mass(flight_point, TIME_STEP)
distance_t = distance_t + delta_distance
time_t = time_t + TIME_STEP
else:
mass_fuel2_t += propulsion_model.get_consumed_mass(flight_point, TIME_STEP)
time_t = time_t + TIME_STEP
v_t = v_t_new
outputs["data:mission:sizing:takeoff:VR"] = vr
outputs["data:mission:sizing:takeoff:VLOF"] = vloff
outputs["data:mission:sizing:takeoff:V2"] = v_t
outputs["data:mission:sizing:takeoff:TOFL"] = distance_t
outputs["data:mission:sizing:takeoff:duration"] = time_t
outputs["data:mission:sizing:takeoff:fuel"] = mass_fuel1_t
outputs["data:mission:sizing:initial_climb:fuel"] = mass_fuel2_t
class _vr_from_v2(om.ExplicitComponent):
"""
Search VR for given lift-off conditions by doing reverted simulation.
The error introduced comes from acceleration acc(t)~acc(t+dt) => v(t-dt)~V(t)-acc(t)*dt.
Time step has been reduced by 1/5 to limit integration error.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_alpha", np.nan, units="rad**-1")
self.add_input("data:aerodynamics:aircraft:low_speed:CD0", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CD", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:induced_drag_coefficient", np.nan)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
self.add_input("data:geometry:wing:span", np.nan, units="m")
self.add_input("data:geometry:landing_gear:height", np.nan, units="m")
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:thrust_rate", np.nan)
self.add_input("data:mission:sizing:takeoff:friction_coefficient_no_brake", np.nan)
self.add_input("vloff:speed", np.nan, units='m/s')
self.add_input("vloff:angle", np.nan, units='rad')
self.add_output("vr:speed", units='m/s')
self.declare_partials("*", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
cl0 = inputs["data:aerodynamics:wing:low_speed:CL0_clean"] + inputs["data:aerodynamics:flaps:takeoff:CL"]
cl_alpha = inputs["data:aerodynamics:wing:low_speed:CL_alpha"]
cd0 = inputs["data:aerodynamics:aircraft:low_speed:CD0"] + inputs["data:aerodynamics:flaps:takeoff:CD"]
coef_k = inputs["data:aerodynamics:wing:low_speed:induced_drag_coefficient"]
wing_area = inputs["data:geometry:wing:area"]
wing_span = inputs["data:geometry:wing:span"]
lg_height = inputs["data:geometry:landing_gear:height"]
mtow = inputs["data:weight:aircraft:MTOW"]
thrust_rate = inputs["data:mission:sizing:takeoff:thrust_rate"]
friction_coeff = inputs["data:mission:sizing:takeoff:friction_coefficient_no_brake"]
v_t = float(inputs["vloff:speed"])
alpha_t = float(inputs["vloff:angle"])
# Define ground factor effect on Drag
k_ground = 33. * (lg_height / wing_span) ** 1.5 / (1. + 33. * (lg_height / wing_span) ** 1.5)
# Start reverted calculation of flight from lift-off to 0° alpha angle
atm = Atmosphere(0.0)
while (alpha_t != 0.0) and (v_t != 0.0):
# Estimation of thrust
flight_point = FlightPoint(
mach=v_t / atm.speed_of_sound, altitude=0.0, engine_setting=EngineSetting.TAKEOFF,
thrust_rate=thrust_rate
)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
# Calculate lift and drag
cl = cl0 + cl_alpha * alpha_t
lift = 0.5 * atm.density * wing_area * cl * v_t ** 2
cd = cd0 + k_ground * coef_k * cl ** 2
drag = 0.5 * atm.density * wing_area * cd * v_t ** 2
# Calculate rolling resistance load
friction = (mtow * g - lift - thrust * math.sin(alpha_t)) * friction_coeff
# Calculate acceleration
acc_x = (thrust * math.cos(alpha_t) - drag - friction) / mtow
# Speed and angle update (feedback)
dt = min(TIME_STEP / 5, alpha_t / ALPHA_RATE, v_t / acc_x)
v_t = v_t - acc_x * dt
alpha_t = alpha_t - ALPHA_RATE * dt
outputs["vr:speed"] = v_t
class _vloff_from_v2(om.ExplicitComponent):
"""
Search alpha-angle<=alpha(v2) at which Vloff is operated such that
aircraft reaches v>=v2 speed @ safety height with imposed rotation speed.
Fuel burn is neglected : mass = MTOW.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_alpha", np.nan, units="rad**-1")
self.add_input("data:aerodynamics:aircraft:low_speed:CD0", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CD", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:induced_drag_coefficient", np.nan)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
self.add_input("data:geometry:wing:span", np.nan, units="m")
self.add_input("data:geometry:landing_gear:height", np.nan, units="m")
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_input("data:mission:sizing:takeoff:thrust_rate", np.nan)
self.add_input("v2:speed", np.nan, units='m/s')
self.add_input("v2:angle", np.nan, units='rad')
self.add_output("vloff:speed", units='m/s')
self.add_output("vloff:angle", units='rad')
self.declare_partials("*", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
cl0 = inputs["data:aerodynamics:wing:low_speed:CL0_clean"] + inputs["data:aerodynamics:flaps:takeoff:CL"]
cl_alpha = inputs["data:aerodynamics:wing:low_speed:CL_alpha"]
cd0 = inputs["data:aerodynamics:aircraft:low_speed:CD0"] + inputs["data:aerodynamics:flaps:takeoff:CD"]
coef_k = inputs["data:aerodynamics:wing:low_speed:induced_drag_coefficient"]
wing_area = inputs["data:geometry:wing:area"]
wing_span = inputs["data:geometry:wing:span"]
lg_height = inputs["data:geometry:landing_gear:height"]
mtow = inputs["data:weight:aircraft:MTOW"]
thrust_rate = inputs["data:mission:sizing:takeoff:thrust_rate"]
v2_target = float(inputs["v2:speed"])
alpha_v2 = float(inputs["v2:angle"])
# Define ground factor effect on Drag
k_ground = lambda altitude: (
33. * ((lg_height + altitude) / wing_span) ** 1.5
/ (1. + 33. * ((lg_height + altitude) / wing_span) ** 1.5)
)
# Calculate v2 speed @ safety height for different alpha lift-off
alpha = np.linspace(0.0, min(ALPHA_LIMIT, alpha_v2), num=10)
vloff = np.zeros(np.size(alpha))
v2 = np.zeros(np.size(alpha))
atm_0 = Atmosphere(0.0)
for i in range(len(alpha)):
# Calculate lift coefficient
cl = cl0 + cl_alpha * alpha[i]
# Loop on estimated lift-off speed error induced by thrust estimation
rel_error = 0.1
vloff[i] = math.sqrt((mtow * g) / (0.5 * atm_0.density * wing_area * cl))
while rel_error > 0.05:
# Update thrust with vloff
flight_point = FlightPoint(
mach=vloff[i] / atm_0.speed_of_sound, altitude=0.0, engine_setting=EngineSetting.TAKEOFF,
thrust_rate=thrust_rate
)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
# Calculate vloff necessary to overcome weight
if thrust * math.sin(alpha[i]) > mtow * g:
break
else:
v = math.sqrt((mtow * g - thrust * math.sin(alpha[i])) / (0.5 * atm_0.density * wing_area * cl))
rel_error = abs(v - vloff[i]) / v
vloff[i] = v
# Perform climb with imposed rotational speed till reaching safety height
alpha_t = alpha[i]
gamma_t = 0.0
v_t = float(vloff[i])
altitude_t = 0.0
distance_t = 0.0
while altitude_t < SAFETY_HEIGHT:
# Estimation of thrust
atm = Atmosphere(altitude_t, altitude_in_feet=False)
flight_point = FlightPoint(
mach=v_t / atm.speed_of_sound, altitude=altitude_t, engine_setting=EngineSetting.TAKEOFF,
thrust_rate=thrust_rate
)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
# Calculate lift and drag
cl = cl0 + cl_alpha * alpha_t
lift = 0.5 * atm.density * wing_area * cl * v_t ** 2
cd = cd0 + k_ground(altitude_t) * coef_k * cl ** 2
drag = 0.5 * atm.density * wing_area * cd * v_t ** 2
# Calculate acceleration on x/z air axis
weight = mtow * g
acc_x = (thrust * math.cos(alpha_t) - weight * math.sin(gamma_t) - drag) / mtow
acc_z = (lift + thrust * math.sin(alpha_t) - weight * math.cos(gamma_t)) / mtow
# Calculate gamma change and new speed
delta_gamma = math.atan((acc_z * TIME_STEP) / (v_t + acc_x * TIME_STEP))
v_t_new = math.sqrt((acc_z * TIME_STEP) ** 2 + (v_t + acc_x * TIME_STEP) ** 2)
# Trapezoidal integration on distance/altitude
delta_altitude = (v_t_new * math.sin(gamma_t + delta_gamma) + v_t * math.sin(gamma_t)) / 2 * TIME_STEP
delta_distance = (v_t_new * math.cos(gamma_t + delta_gamma) + v_t * math.cos(gamma_t)) / 2 * TIME_STEP
# Update temporal values
alpha_t = min(alpha_v2, alpha_t + ALPHA_RATE * TIME_STEP)
gamma_t = gamma_t + delta_gamma
altitude_t = altitude_t + delta_altitude
distance_t = distance_t + delta_distance
v_t = v_t_new
# Save obtained v2
v2[i] = v_t
# If v2 target speed not reachable maximum lift-off speed chosen (alpha=0°)
if sum(v2 > v2_target) == 0:
alpha = 0.0
vloff = vloff[0] # FIXME: not reachable v2
warnings.warn("V2 @ 50ft requirement not reachable with max lift-off speed!")
else:
# If max alpha angle lead to v2 > v2 target take it
if v2[-1] > v2_target:
alpha = alpha[-1]
vloff = vloff[-1]
else:
alpha = np.interp(v2_target, v2, alpha)
vloff = np.interp(v2_target, v2, vloff)
outputs["vloff:speed"] = vloff
outputs["vloff:angle"] = alpha
class _v2(om.ExplicitComponent):
"""
Calculate V2 safety speed @ defined altitude considering a 30% safety margin on max lift capability (alpha imposed).
Find corresponding climb rate margin for imposed thrust rate.
Fuel burn is neglected : mass = MTOW.
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_alpha", np.nan, units="rad**-1")
self.add_input("data:aerodynamics:aircraft:low_speed:CD0", np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CD", np.nan)
self.add_input("data:aerodynamics:wing:low_speed:induced_drag_coefficient", np.nan)
self.add_input("data:geometry:wing:area", np.nan, units="m**2")
self.add_input("data:geometry:wing:span", np.nan, units="m")
self.add_input("data:geometry:landing_gear:height", np.nan, units="m")
self.add_input("data:weight:aircraft:MTOW", np.nan, units="kg")
self.add_output("v2:speed", units='m/s')
self.add_output("v2:angle", units='rad')
self.add_output("v2:climb_rate")
self.declare_partials("*", "*", method="fd")
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
cl_max_clean = inputs["data:aerodynamics:wing:low_speed:CL_max_clean"]
cl0 = inputs["data:aerodynamics:wing:low_speed:CL0_clean"] + inputs["data:aerodynamics:flaps:takeoff:CL"]
cl_alpha = inputs["data:aerodynamics:wing:low_speed:CL_alpha"]
cd0 = inputs["data:aerodynamics:aircraft:low_speed:CD0"] + inputs["data:aerodynamics:flaps:takeoff:CD"]
coef_k = inputs["data:aerodynamics:wing:low_speed:induced_drag_coefficient"]
wing_area = inputs["data:geometry:wing:area"]
wing_span = inputs["data:geometry:wing:span"]
lg_height = inputs["data:geometry:landing_gear:height"]
mtow = inputs["data:weight:aircraft:MTOW"]
# Define atmospheric condition for safety height
atm = Atmosphere(SAFETY_HEIGHT, altitude_in_feet=False)
# Define Cl considering 30% margin and estimate alpha
cl = cl_max_clean / 1.2 ** 2 # V2>=1.2*VS1
alpha_interp = np.linspace(0.0, 30.0, 31) * math.pi / 180.0
cl_interp = cl0 + alpha_interp * cl_alpha
alpha = np.interp(cl, cl_interp, alpha_interp)
# Calculate drag coefficient
k_ground = (
33. * ((lg_height + SAFETY_HEIGHT) / wing_span) ** 1.5
/ (1. + 33. * ((lg_height + SAFETY_HEIGHT) / wing_span) ** 1.5)
)
cd = cd0 + k_ground * coef_k * cl ** 2
# Find v2 safety speed for 0% climb rate
v2 = math.sqrt((mtow * g) / (0.5 * atm.density * wing_area * cl))
# Estimate climb rate considering alpha~0° and max thrust rate for CS23.65 (loop on error)
flight_point = FlightPoint(
mach=v2 / atm.speed_of_sound, altitude=SAFETY_HEIGHT, engine_setting=EngineSetting.TAKEOFF, thrust_rate=1.0
)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
gamma = math.asin(thrust / (mtow * g) - cd / cl)
rel_error = 0.1
while rel_error > 0.05:
new_gamma = math.asin(thrust / (mtow * g) - cd / cl * math.cos(gamma))
rel_error = abs((new_gamma - gamma) / new_gamma)
gamma = new_gamma
outputs["v2:speed"] = v2
outputs["v2:angle"] = alpha
outputs["v2:climb_rate"] = math.sin(gamma)
class _UpdateArea(om.ExplicitComponent):
"""
Computes area of horizontal tail plane (internal function)
"""
def __init__(self, **kwargs):
super().__init__(**kwargs)
self._engine_wrapper = None
def initialize(self):
self.options.declare("propulsion_id", default="", types=str)
def setup(self):
self._engine_wrapper = BundleLoader().instantiate_component(
self.options["propulsion_id"])
self._engine_wrapper.setup(self)
self.add_input("data:geometry:propulsion:count", val=np.nan)
self.add_input("settings:weight:aircraft:CG:range", val=0.3)
self.add_input("data:mission:sizing:takeoff:thrust_rate", val=np.nan)
self.add_input("data:geometry:wing:area", val=np.nan, units="m**2")
self.add_input("data:geometry:wing:MAC:at25percent:x",
val=np.nan,
units="m")
self.add_input(
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25",
val=np.nan,
units="m")
self.add_input("data:geometry:wing:MAC:length", val=np.nan, units="m")
self.add_input("data:geometry:propulsion:nacelle:height",
val=np.nan,
units="m")
self.add_input("data:weight:aircraft:MTOW", val=np.nan, units="kg")
self.add_input("data:weight:aircraft:MLW", val=np.nan, units="kg")
self.add_input("data:weight:aircraft:CG:aft:x", val=np.nan, units="m")
self.add_input("data:weight:airframe:landing_gear:main:CG:x",
val=np.nan,
units="m")
self.add_input("data:weight:aircraft_empty:CG:z",
val=np.nan,
units="m")
self.add_input("data:weight:propulsion:engine:CG:z",
val=np.nan,
units="m")
self.add_input("data:aerodynamics:wing:low_speed:CL0_clean",
val=np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CM0_clean",
val=np.nan)
self.add_input("data:aerodynamics:aircraft:landing:CL_max", val=np.nan)
self.add_input("data:aerodynamics:aircraft:takeoff:CL_max", val=np.nan)
self.add_input("data:aerodynamics:wing:low_speed:CL_max_clean",
val=np.nan)
self.add_input("data:aerodynamics:flaps:landing:CL", val=np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CL", val=np.nan)
self.add_input("data:aerodynamics:flaps:landing:CM", val=np.nan)
self.add_input("data:aerodynamics:flaps:takeoff:CM", val=np.nan)
self.add_input("data:aerodynamics:horizontal_tail:low_speed:CL_alpha",
val=np.nan,
units="rad**-1")
self.add_input("data:aerodynamics:horizontal_tail:efficiency",
val=np.nan)
self.add_input("landing:cl_htp", val=np.nan)
self.add_input("takeoff:cl_htp", val=np.nan)
self.add_input("low_speed:cl_alpha_htp_isolated", val=np.nan)
self.add_output("data:geometry:horizontal_tail:area",
val=4.0,
units="m**2")
self.declare_partials(
"*", "*",
method="fd") # FIXME: write partial avoiding discrete parameters
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
# Sizing constraints for the horizontal tail (methods from Torenbeek).
# Limiting cases: Rotating power at takeoff/landing, with the most
# forward CG position. Returns maximum area.
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs),
inputs["data:geometry:propulsion:count"])
n_engines = inputs["data:geometry:propulsion:count"]
cg_range = inputs["settings:weight:aircraft:CG:range"]
takeoff_t_rate = inputs["data:mission:sizing:takeoff:thrust_rate"]
wing_area = inputs["data:geometry:wing:area"]
x_wing_aero_center = inputs["data:geometry:wing:MAC:at25percent:x"]
lp_ht = inputs[
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25"]
wing_mac = inputs["data:geometry:wing:MAC:length"]
mtow = inputs["data:weight:aircraft:MTOW"]
mlw = inputs["data:weight:aircraft:MLW"]
x_cg_aft = inputs["data:weight:aircraft:CG:aft:x"]
z_cg_aircraft = inputs["data:weight:aircraft_empty:CG:z"]
z_cg_engine = inputs["data:weight:propulsion:engine:CG:z"]
x_lg = inputs["data:weight:airframe:landing_gear:main:CG:x"]
cl0_clean = inputs["data:aerodynamics:wing:low_speed:CL0_clean"]
cl_max_clean = inputs["data:aerodynamics:wing:low_speed:CL_max_clean"]
cl_max_landing = inputs["data:aerodynamics:aircraft:landing:CL_max"]
cl_max_takeoff = inputs["data:aerodynamics:aircraft:takeoff:CL_max"]
cl_flaps_landing = inputs["data:aerodynamics:flaps:landing:CL"]
cl_flaps_takeoff = inputs["data:aerodynamics:flaps:takeoff:CL"]
tail_efficiency_factor = inputs[
"data:aerodynamics:horizontal_tail:efficiency"]
cl_htp_landing = inputs["landing:cl_htp"]
cl_htp_takeoff = inputs["takeoff:cl_htp"]
cm_landing = inputs[
"data:aerodynamics:wing:low_speed:CM0_clean"] + inputs[
"data:aerodynamics:flaps:landing:CM"]
cm_takeoff = inputs[
"data:aerodynamics:wing:low_speed:CM0_clean"] + inputs[
"data:aerodynamics:flaps:takeoff:CM"]
cl_alpha_htp_isolated = inputs["low_speed:cl_alpha_htp_isolated"]
z_eng = z_cg_aircraft - z_cg_engine
# Conditions for calculation
atm = Atmosphere(0.0)
rho = atm.density
# CASE1: TAKE-OFF ##############################################################################################
# method extracted from Torenbeek 1982 p325
# Calculation of take-off minimum speed
weight = mtow * g
vs0 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_takeoff))
vs1 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_clean))
# Rotation speed requirement from FAR 23.51 (depends on number of engines)
if n_engines == 1:
v_r = vs1 * 1.0
else:
v_r = vs1 * 1.1
# Definition of max forward gravity center position
x_cg = x_cg_aft - cg_range * wing_mac
# Definition of horizontal tail global position
x_ht = x_wing_aero_center + lp_ht
# Calculation of wheel factor
flight_point = FlightPoint(mach=v_r / atm.speed_of_sound,
altitude=0.0,
engine_setting=EngineSetting.TAKEOFF,
thrust_rate=takeoff_t_rate)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
fact_wheel = (
(x_lg - x_cg - z_eng * thrust / weight) / wing_mac * (vs0 / v_r)**2
) # FIXME: not clear if vs0 or vs1 should be used in formula
# Compute aerodynamic coefficients for takeoff @ 0° aircraft angle
cl0_takeoff = cl0_clean + cl_flaps_takeoff
# Calculation of correction coefficient n_h and n_q
n_h = (
x_ht - x_lg
) / lp_ht * tail_efficiency_factor # tail_efficiency_factor: dynamic pressure reduction at
# tail (typical value)
n_q = 1 + cl_alpha_htp_isolated / cl_htp_takeoff * _ANG_VEL * (
x_ht - x_lg) / v_r
# Calculation of volume coefficient based on Torenbeek formula
coef_vol = (cl_max_takeoff / (n_h * n_q * cl_htp_takeoff) *
(cm_takeoff / cl_max_takeoff - fact_wheel) +
cl0_takeoff / cl_htp_takeoff *
(x_lg - x_wing_aero_center) / wing_mac)
# Calculation of equivalent area
area_1 = coef_vol * wing_area * wing_mac / lp_ht
# CASE2: LANDING ###############################################################################################
# method extracted from Torenbeek 1982 p325
# Calculation of take-off minimum speed
weight = mlw * g
vs0 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_landing))
# Rotation speed requirement from FAR 23.73
v_r = vs0 * 1.3
# Calculation of wheel factor
flight_point = FlightPoint(
mach=v_r / atm.speed_of_sound,
altitude=0.0,
engine_setting=EngineSetting.IDLE,
thrust_rate=0.1
) # FIXME: fixed thrust rate (should depend on wished descent rate)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
fact_wheel = (
(x_lg - x_cg - z_eng * thrust / weight) / wing_mac * (vs0 / v_r)**2
) # FIXME: not clear if vs0 or vs1 should be used in formula
# Evaluate aircraft overall angle (aoa)
cl0_landing = cl0_clean + cl_flaps_landing
# Calculation of correction coefficient n_h and n_q
n_h = (
x_ht - x_lg
) / lp_ht * tail_efficiency_factor # tail_efficiency_factor: dynamic pressure reduction at
# tail (typical value)
n_q = 1 + cl_alpha_htp_isolated / cl_htp_landing * _ANG_VEL * (
x_ht - x_lg) / v_r
# Calculation of volume coefficient based on Torenbeek formula
coef_vol = (cl_max_landing / (n_h * n_q * cl_htp_landing) *
(cm_landing / cl_max_landing - fact_wheel) +
cl0_landing / cl_htp_landing *
(x_lg - x_wing_aero_center) / wing_mac)
# Calculation of equivalent area
area_2 = coef_vol * wing_area * wing_mac / lp_ht
if max(area_1, area_2) < 0.0:
print(
"Warning: HTP area estimated negative (in ComputeHTArea) forced to 1m²!"
)
outputs["data:geometry:horizontal_tail:area"] = 1.0
else:
outputs["data:geometry:horizontal_tail:area"] = max(area_1, area_2)
