def main():
# This script could fail if either the design or analysis scripts fail,
# in case of failure check both. The design and analysis powers will
# differ because of karman-tsien compressibility corrections in the
# analysis scripts
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 50.0
prop_attributes.angular_velocity = 2000.*(2.*np.pi/60.0)
prop_attributes.tip_radius = 1.5
prop_attributes.hub_radius = 0.05
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 0.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 7000.
prop_attributes = propeller_design(prop_attributes)
# Find the operating conditions
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(prop_attributes.design_altitude)
V = prop_attributes.freestream_velocity
conditions = Data()
conditions.freestream = Data()
conditions.propulsion = Data()
conditions.freestream.update(atmosphere_conditions)
conditions.freestream.dynamic_viscosity = atmosphere_conditions.dynamic_viscosity
conditions.freestream.velocity = np.array([[V]])
conditions.propulsion.throttle = np.array([[1.0]])
# Create and attach this propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.inputs.omega = prop_attributes.angular_velocity
F, Q, P, Cplast = prop.spin(conditions)
# Truth values
F_truth = 166.41590262
Q_truth = 45.21732911
P_truth = 9470.2952633 # Over 9000!
Cplast_truth = 0.00085898
error = Data()
error.Thrust = np.max(np.abs(F-F_truth))
error.Power = np.max(np.abs(P-P_truth))
error.Torque = np.max(np.abs(Q-Q_truth))
error.Cp = np.max(np.abs(Cplast-Cplast_truth))
print 'Errors:'
print error
for k,v in error.items():
assert(np.abs(v)<0.001)
return
def basic_prop(Na=24, Nr=101):
# Design the Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_of_blades = 2
prop.freestream_velocity = 135. * Units['mph']
prop.angular_velocity = 1300. * Units.rpm
prop.tip_radius = 38. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_thrust = 1200.
prop.origin = [[0., 0., 0.]]
prop.number_azimuthal_stations = Na
prop.rotation = 1
prop.symmetry = True
prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
prop.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
prop.airfoil_polar_stations = list(np.zeros(Nr).astype(int))
prop = propeller_design(prop, Nr)
return prop
def ICE_CS(vehicle):
# Replace the C172 engine and propeller with a constant speed propeller
# Let's assume its an STC or 172RG
# build network
net = SUAVE.Components.Energy.Networks.Internal_Combustion_Propeller_Constant_Speed(
)
net.tag = 'internal_combustion'
net.number_of_engines = 1.
net.nacelle_diameter = 42 * Units.inches
net.engine_length = 0.01 * Units.inches
net.areas = Data()
net.rated_speed = 2700. * Units.rpm
net.rated_power = 180. * Units.hp
net.areas.wetted = 0.01
# Component 1 the engine
net.engine = SUAVE.Components.Energy.Converters.Internal_Combustion_Engine(
)
net.engine.sea_level_power = 180. * Units.horsepower
net.engine.flat_rate_altitude = 0.0
net.engine.rated_speed = 2700. * Units.rpm
net.engine.power_specific_fuel_consumption = 0.52
#
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_of_blades = 2.0
prop.freestream_velocity = 119. * Units.knots
prop.angular_velocity = 2650. * Units.rpm
prop.tip_radius = 76. / 2. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_power = .64 * 180. * Units.horsepower
prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
prop.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
prop = propeller_design(prop)
net.propeller = prop
# Replace the network
vehicle.propulsors.internal_combustion = net
return vehicle
def propeller():
'''
Parameters need to be added
'''
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 40.0 * Units['m/s'] # freestream
prop_attributes.angular_velocity = 150. * Units['rpm']
prop_attributes.tip_radius = 4.25 * Units.meters
prop_attributes.hub_radius = 0.05 * Units.meters
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 14.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 3500.0 * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
return prop
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Tecnam_P2006TElectric'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 1400 * Units.kilogram
vehicle.mass_properties.takeoff = 1400 * Units.kilogram
vehicle.mass_properties.operating_empty = 1000 * Units.kilogram
vehicle.mass_properties.max_zero_fuel = 1400 * Units.kilogram
vehicle.mass_properties.cargo = 80 * Units.kilogram
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
# basic parameters
vehicle.reference_area = 64.4 * Units['meters**2']
vehicle.passengers = 4
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Landing Gear
# ------------------------------------------------------------------
# used for noise calculations
#landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
#landing_gear.tag = "main_landing_gear"
#landing_gear.main_tire_diameter = 0.423 * Units.m
#landing_gear.nose_tire_diameter = 0.3625 * Units.m
#landing_gear.main_strut_length = 0.4833 * Units.m
#landing_gear.nose_strut_length = 0.3625 * Units.m
#landing_gear.main_units = 2 #number of main landing gear units
#landing_gear.nose_units = 1 #number of nose landing gear
#landing_gear.main_wheels = 1 #number of wheels on the main landing gear
#landing_gear.nose_wheels = 1 #number of wheels on the nose landing gear
#vehicle.landing_gear = landing_gear
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
aux = time.time()
fuselage.time = aux
fuselage.number_coach_seats = vehicle.passengers
fuselage.seats_abreast = 2
fuselage.seat_pitch = 0.995 * Units.meter
fuselage.fineness.nose = 1.27
fuselage.fineness.tail = 1 #3.31
fuselage.lengths.nose = 1.16 * Units.meter
fuselage.lengths.tail = 4.637 * Units.meter
fuselage.lengths.cabin = 2.653 * Units.meter
fuselage.lengths.total = 8.45 * Units.meter
fuselage.lengths.fore_space = 0.0 * Units.meter
fuselage.lengths.aft_space = 0.0 * Units.meter
fuselage.width = 1.1 * Units.meter #1.22
fuselage.heights.maximum = 1.41 * Units.meter
fuselage.effective_diameter = 2 * Units.meter
fuselage.areas.side_projected = 7.46 * Units['meters**2']
fuselage.areas.wetted = 25.0 * Units['meters**2']
fuselage.areas.front_projected = 1.54 * Units['meters**2']
fuselage.differential_pressure = 10**5 * Units.pascal # Maximum differential pressure
fuselage.heights.at_quarter_length = 1.077 * Units.meter
fuselage.heights.at_three_quarters_length = 0.5 * Units.meter #0.621 * Units.meter
fuselage.heights.at_wing_root_quarter_chord = 1.41 * Units.meter
## OpenVSP Design
fuselage.OpenVSP_values = Data() # VSP uses degrees directly
#MidFuselage1 Section
fuselage.OpenVSP_values.midfus1 = Data()
fuselage.OpenVSP_values.midfus1.z_pos = 0.03
#MidFuselage2 Section
fuselage.OpenVSP_values.midfus2 = Data()
fuselage.OpenVSP_values.midfus2.z_pos = 0.06
#MidFuselage3 Section
fuselage.OpenVSP_values.midfus3 = Data()
fuselage.OpenVSP_values.midfus3.z_pos = 0.04
#Tail Section
fuselage.OpenVSP_values.tail = Data()
fuselage.OpenVSP_values.tail.bottom = Data()
fuselage.OpenVSP_values.tail.z_pos = 0.039
fuselage.OpenVSP_values.tail.bottom.angle = -20.0
fuselage.OpenVSP_values.tail.bottom.strength = 1
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.thickness_to_chord = 0.15249979370200092
wing.taper = 0.9014171440917205
wing.spans.projected = 9.342124951930751 * Units.meter
wing.chords.root = 2.0273876759854024 * Units.meter
wing.chords.tip = wing.chords.root * wing.taper
wing.chords.mean_aerodynamic = (wing.chords.root * (2.0 / 3.0) *
((1.0 + wing.taper + wing.taper**2.0) /
(1.0 + wing.taper)))
wing.areas.reference = (wing.chords.root +
wing.chords.tip) * wing.spans.projected / 2
print wing.areas.reference
basearea = wing.areas.reference
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = wing.areas.wetted
wing.areas.affected = wing.areas.wetted
wing.twists.root = 0. * Units.degrees
wing.twists.tip = 0. * Units.degrees
wing.dihedral = 1. * Units.degrees
wing.origin = [2.986, 0, 1.077] # meters
wing.sweeps.leading_edge = 1.9 * Units.deg
wing.aspect_ratio = (wing.spans.projected *
wing.spans.projected) / wing.areas.reference
wing.span_efficiency = 0.95
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
## Wing Segments
# Root Segment
#segment = SUAVE.Components.Wings.Segment()
#segment.tag = 'root'
#segment.percent_span_location = 0.0
#segment.twist = 0. * Units.deg
#segment.root_chord_percent = 1.
#segment.dihedral_outboard = 1. * Units.degrees
#segment.sweeps.quarter_chord = 0. * Units.degrees
#segment.thickness_to_chord = 0.15
#airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
#airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
#segment.append_airfoil(airfoil)
#wing.Segments.append(segment)
# Tip Segment
#segment = SUAVE.Components.Wings.Segment()
#segment.tag = 'tip'
#segment.percent_span_location = 1.0
#segment.twist = 0. * Units.deg
#segment.root_chord_percent = 1.
#segment.dihedral_outboard = 1. * Units.degrees
#segment.sweeps.quarter_chord = 0. * Units.degrees
#segment.thickness_to_chord = 0.15
#airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
#airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
#segment.append_airfoil(airfoil)
#wing.Segments.append(segment)
# ------------------------------------------------------------------
# Flaps
# ------------------------------------------------------------------
wing.flaps.chord = wing.chords.root * 0.15
wing.flaps.span_start = 0.3 * wing.spans.projected
wing.flaps.span_end = 0.8 * wing.spans.projected
wing.flaps.area = wing.flaps.chord * (wing.flaps.span_end -
wing.flaps.span_start)
wing.flaps.type = 'single_slotted'
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 4.193
wing.areas.reference = 0.145 * basearea
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.97
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.chords.root
wing.chords.mean_aerodynamic = (wing.chords.root * (2.0 / 3.0) *
((1.0 + wing.taper + wing.taper**2.0) /
(1.0 + wing.taper)))
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = wing.areas.wetted
wing.areas.affected = wing.areas.wetted
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [7.789, 0.0, 0.3314] # meters
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.areas.reference = 0.099 * basearea
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = wing.areas.wetted
wing.areas.affected = wing.areas.wetted
wing.aspect_ratio = 1.407
wing.sweeps.quarter_chord = 38.75 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 0.4142
wing.span_efficiency = 0.97
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.chords.root = (2.0 * wing.areas.reference) / (wing.spans.projected *
(1 + wing.taper))
wing.chords.tip = wing.chords.root * wing.taper
wing.chords.mean_aerodynamic = (wing.chords.root * (2.0 / 3.0) *
((1.0 + wing.taper + wing.taper**2.0) /
(1.0 + wing.taper)))
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [7.25, 0, 0.497] # meters
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Propellers Powered By Batteries
# ------------------------------------------------------------------
# build network
net = SUAVE.Components.Energy.Networks.Lift_Forward_Propulsor()
net.nacelle_diameter_lift = 0.08 * Units.meters
net.nacelle_diameter_forward = 0.1732 * Units.meters
net.engine_length_lift = 0.47244 * Units.meters
net.engine_length_forward = 1.2 * Units.meters
net.number_of_engines_lift = 10
print 'Prop Optimization'
print 'Engines Lift Number'
print net.number_of_engines_lift
net.number_of_engines_forward = 2
net.thrust_angle_lift = 0.0 * Units.degrees
net.thrust_angle_forward = 0.0 * Units.degrees
net.voltage = 461. * Units['volt'] #461.
net.areas_forward = Data()
net.areas_forward.wetted = 1.1 * np.pi * net.nacelle_diameter_forward * net.engine_length_forward
net.areas_lift = Data()
net.areas_lift.wetted = 1.1 * np.pi * net.nacelle_diameter_forward * net.engine_length_lift
net.number_of_engines = 1
net.nacelle_diameter = net.nacelle_diameter_forward
net.areas = Data()
net.areas.wetted = net.areas_lift.wetted
net.engine_length = 1.
# Component 1 - Tip ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc_forward = esc
# Component 1 - High Lift ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc_lift = esc
# Component 2 - Tip Propeller
# Design the Propeller
#prop_attributes = Data()
#prop_attributes = propeller_design(prop_attributes)
prop_forward = SUAVE.Components.Energy.Converters.Propeller()
prop_forward.number_blades = 3.0
prop_forward.propulsive_efficiency = 0.85
prop_forward.freestream_velocity = 50.0 * Units['m/s'] # freestream m/s
prop_forward.angular_velocity = 27500. * Units['rpm'] # For 20x10 prop
prop_forward.tip_radius = 0.5 * Units.meter
prop_forward.hub_radius = 0.085 * Units.meter
prop_forward.design_Cl = 0.8
prop_forward.design_altitude = 14.0 * Units.km
prop_forward.design_thrust = 0.0 * Units.newton
prop_forward.design_power = 60000. * Units.watts
prop_forward = propeller_design(prop_forward)
net.propeller_forward = prop_forward
# Component 2 - High Lift Propeller
# Design the Propeller
#prop_attributes = Data()
prop_lift = SUAVE.Components.Energy.Converters.Propeller()
prop_lift.number_blades = 5.0
prop_lift.propulsive_efficiency = 0.85
prop_lift.freestream_velocity = 1. * Units['m/s'] # freestream m/s
prop_lift.angular_velocity = 2750 * Units['rpm'] # For 20x10 prop
prop_lift.tip_radius = 0.26 * Units.meter #0.2880360
prop_lift.hub_radius = 0.07772400 * Units.meter
prop_lift.design_Cl = 0.8
prop_lift.design_altitude = 0.0 * Units.meter
prop_lift.design_thrust = 0.0 * Units.newton
prop_lift.design_power = 10500. * Units.watts
prop_lift = propeller_design(prop_lift)
net.propeller_lift = prop_lift
# Component 3 - Tip Motor
motor = SUAVE.Components.Energy.Converters.Motor_Lo_Fid()
#motor.resistance = 1.
#motor.no_load_current = 7. * Units.ampere
#motor.speed_constant = 11.9999 * Units['rpm/volt'] # RPM/volt converted to (rad/s)/volt
#motor.propeller_radius = prop_forward.tip_radius
#motor.propeller_Cp = prop_forward.Cp[0]
#motor.gear_ratio = 12. # Gear ratio
#motor.gearbox_efficiency = .98 # Gear box efficiency
#motor.expected_current = 160. # Expected current
#motor.mass_properties.mass = 9.0 * Units.kg
#net.motor_forward = motor
motor.speed_constant = 180. * Units['rpm/volt'] # RPM/volt is standard
motor = size_from_kv(motor, motor.speed_constant)
motor.gear_ratio = 1. # Gear ratio, no gearbox
motor.gearbox_efficiency = .98 # Gear box efficiency, no gearbox
motor.motor_efficiency = 0.825
net.motor_forward = motor
# Component 3 - High Lift Motor
motor = SUAVE.Components.Energy.Converters.Motor_Lo_Fid()
#motor.resistance = 0.008
#motor.no_load_current = 4.5 * Units.ampere
#motor.speed_constant = 5800. * Units['rpm/volt'] # RPM/volt converted to (rad/s)/volt
#motor.propeller_radius = prop_lift.tip_radius
#motor.propeller_Cp = prop_lift.Cp[0]
#motor.gear_ratio = 12. # Gear ratio
#motor.gearbox_efficiency = .98 # Gear box efficiency
#motor.expected_current = 25. # Expected current
#motor.mass_properties.mass = 6.0 * Units.kg
#net.motor_lift = motor
motor.speed_constant = 90. * Units['rpm/volt'] # RPM/volt is standard
motor = size_from_kv(motor, motor.speed_constant)
motor.gear_ratio = 1. # Gear ratio, no gearbox
motor.gearbox_efficiency = .98 # Gear box efficiency, no gearbox
motor.motor_efficiency = 0.825
net.motor_lift = motor
# Component 4 - the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 50. * Units.watts
payload.mass_properties.mass = 5.0 * Units.kg
net.payload = payload
# Component 5 - the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 50. * Units.watts
net.avionics = avionics
# Component 6 - the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.mass_properties.mass = 386.0 * Units.kg
bat.specific_energy = 121.8 * Units.Wh / Units.kg #192.84
bat.specific_power = 0.312 * Units.kW / Units.kg #0.837
bat.resistance = 0.32
bat.max_voltage = 538. * Units['volt'] #10000.
initialize_from_mass(bat, bat.mass_properties.mass)
print bat.max_energy
net.battery = bat
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
# add the energy network to the vehicle
vehicle.append_component(net)
#now add weights objects
vehicle.landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
vehicle.control_systems = SUAVE.Components.Physical_Component()
vehicle.electrical_systems = SUAVE.Components.Physical_Component()
vehicle.avionics = SUAVE.Components.Energy.Peripherals.Avionics()
vehicle.passenger_weights = SUAVE.Components.Physical_Component()
#vehicle.furnishings = SUAVE.Components.Physical_Component()
#vehicle.apu = SUAVE.Components.Physical_Component()
#vehicle.hydraulics = SUAVE.Components.Physical_Component()
vehicle.optionals = SUAVE.Components.Physical_Component()
vehicle.wings[
'vertical_stabilizer'].rudder = SUAVE.Components.Physical_Component()
#print vehicle
return vehicle
def define_vehicle():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Solar'
vehicle.propulsors.propulsor = SUAVE.Components.Energy.Networks.Solar_Network()
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 200 * Units.kg
vehicle.mass_properties.operating_empty = 200 * Units.kg
vehicle.mass_properties.max_takeoff = 200 * Units.kg
# basic parameters
vehicle.reference_area = 80. # m^2
vehicle.envelope.ultimate_load = 2.0
vehicle.qm = 0.5*1.225*(25.**2.) #Max q
vehicle.Ltb = 10. # Tail boom length
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Main Wing'
wing.areas.reference = vehicle.reference_area #
wing.spans.projected = 40. #m
wing.aspect_ratio = (wing.spans.projected**2)/wing.areas.reference
wing.sweep = 0. * Units.deg #
wing.symmetric = True #
wing.thickness_to_chord = 0.12 #
wing.taper = 1. #
# size the wing planform
SUAVE.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.chords.mean_aerodynamic = wing.areas.reference/wing.spans.projected #
wing.areas.exposed = 0.8*wing.areas.wetted # might not be needed as input
wing.areas.affected = 0.6*wing.areas.wetted # part of high lift system
wing.span_efficiency = 0.97 #
wing.twists.root = 0.0*Units.degrees #
wing.twists.tip = 0.0*Units.degrees #
wing.highlift = False
wing.vertical = False
wing.eta = 1.0
wing.number_ribs = 26. # ?
wing.number_end_ribs = 2.
wing.transition_x_u = 0.
wing.transition_x_l = 0.
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Horizontal Stabilizer'
wing.areas.reference = vehicle.reference_area*.15 #m^2
wing.aspect_ratio = 20. #
wing.spans.projected = np.sqrt(wing.aspect_ratio*wing.areas.reference)
wing.sweep = 0 * Units.deg #
wing.symmetric = True
wing.thickness_to_chord = 0.12 #
wing.taper = 1. #
wing.twists.root = 0.0*Units.degrees #
wing.twists.tip = 0.0*Units.degrees #
# size the wing planform
SUAVE.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.chords.mean_aerodynamic = wing.areas.reference/wing.spans.projected #
wing.areas.exposed = 0.8*wing.areas.wetted # might not be needed as input
wing.areas.affected = 0.6*wing.areas.wetted # part of high lift system
wing.span_efficiency = 0.95 #
wing.twists.root = 0. #
wing.twists.tip = 0. #
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Vertical Stabilizer'
wing.areas.reference = vehicle.reference_area*.1 #m^2
wing.aspect_ratio = 20. #
wing.spans.projected = np.sqrt(wing.aspect_ratio*wing.areas.reference)
wing.sweep = 0 * Units.deg #
wing.symmetric = True
wing.thickness_to_chord = 0.12 #
wing.taper = 1. #
wing.twists.root = 0.0*Units.degrees #
wing.twists.tip = 0.0*Units.degrees #
# size the wing planform
SUAVE.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.chords.mean_aerodynamic = wing.areas.reference/wing.spans.projected #
wing.areas.exposed = 0.8*wing.areas.wetted # might not be needed as input
wing.areas.affected = 0.6*wing.areas.wetted # part of high lift system
wing.span_efficiency = 0.97 #
wing.twists.root = 0.0*Units.degrees #
wing.twists.tip = 0.0*Units.degrees #
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------
# Propulsor
#------------------------------------------------------------------
# build network
net = Solar_Network()
net.number_motors = 1.
net.nacelle_dia = 0.2
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = vehicle.reference_area
panel.efficiency = 0.2
panel.mass_properties.mass = panel.area*0.6
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 50.0 # freestream m/s
prop_attributes.angular_velocity = 300.*(2.*np.pi/60.0)
prop_attributes.tip_radius = 4.25
prop_attributes.hub_radius = 0.0508
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 23.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 10000.0
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.008
motor.no_load_current = 4.5
motor.speed_constant = 120.*(2.*np.pi/60.) # RPM/volt converted to rad/s
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 20. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 160. # Expected current
motor.mass_properties.mass = 2.0
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 100. #Watts
payload.mass_properties.mass = 25.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 0. #Watts
net.avionics = avionics
# Component 8 the Battery # I already assume 250 Wh/kg for batteries
bat = SUAVE.Components.Energy.Storages.Battery()
bat.mass_properties.mass = 50 * Units.kg
bat.type = 'Li-Ion'
bat.resistance = 0.0 #This needs updating
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 100.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# Calculate the vehicle mass
vehicle.mass_properties.breakdown = SUAVE.Methods.Weights.Correlations.Human_Powered.empty(vehicle)
# ------------------------------------------------------------------
# Simple Aerodynamics Model
# ------------------------------------------------------------------
aerodynamics = SUAVE.Attributes.Aerodynamics.Fidelity_Zero()
aerodynamics.initialize(vehicle)
vehicle.aerodynamics_model = aerodynamics
# ------------------------------------------------------------------
# Not so Simple Propulsion Model
# ------------------------------------------------------------------
vehicle.propulsion_model = net
# ------------------------------------------------------------------
# Define Configurations
# ------------------------------------------------------------------
# --- Takeoff Configuration ---
config = vehicle.new_configuration("takeoff")
# this configuration is derived from the baseline vehicle
# --- Cruise Configuration ---
config = vehicle.new_configuration("cruise")
# this configuration is derived from vehicle.Configs.takeoff
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Cessna_172_SP'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
vehicle.mass_properties.takeoff = 2550. * Units.pounds
vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
vehicle.mass_properties.cargo = 0.
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
cruise_speed = 124. * Units.kts
altitude = 8500. * Units.ft
atmo = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmo.compute_values(0.)
freestream0 = atmo.compute_values(altitude)
mach_number = (cruise_speed / freestream.speed_of_sound)[0][0]
vehicle.design_dynamic_pressure = (.5 * freestream0.density *
(cruise_speed * cruise_speed))[0][0]
vehicle.design_mach_number = mach_number
# basic parameters
vehicle.reference_area = 174. * Units.feet**2
vehicle.passengers = 4
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 174. * Units.feet**2
wing.spans.projected = 36. * Units.feet + 1. * Units.inches
wing.chords.root = 66. * Units.inches
wing.chords.tip = 45. * Units.inches
wing.chords.mean_aerodynamic = 58. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 3.0 * Units.degrees
wing.twists.tip = 1.5 * Units.degrees
wing.origin = [[80. * Units.inches, 0, 0]]
wing.aerodynamic_center = [22. * Units.inches, 0, 0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
# control surfaces -------------------------------------------
flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
flap.tag = 'flap'
flap.span_fraction_start = 0.15
flap.span_fraction_end = 0.324
flap.deflection = 1.0 * Units.deg
flap.chord_fraction = 0.19
wing.append_control_surface(flap)
slat = SUAVE.Components.Wings.Control_Surfaces.Slat()
slat.tag = 'slat'
slat.span_fraction_start = 0.324
slat.span_fraction_end = 0.963
slat.deflection = 1.0 * Units.deg
slat.chord_fraction = 0.1
wing.append_control_surface(slat)
SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 5800. * Units.inches**2
wing.spans.projected = 136. * Units.inches
wing.chords.root = 55. * Units.inches
wing.chords.tip = 30. * Units.inches
wing.chords.mean_aerodynamic = 43. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[246. * Units.inches, 0, 0]]
wing.aerodynamic_center = [20. * Units.inches, 0, 0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = False
wing.dynamic_pressure_ratio = 0.9
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.sweeps.quarter_chord = 25. * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 3500. * Units.inches**2
wing.spans.projected = 73. * Units.inches
wing.chords.root = 66. * Units.inches
wing.chords.tip = 27. * Units.inches
wing.chords.mean_aerodynamic = 48. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[237. * Units.inches, 0, 0]]
wing.aerodynamic_center = [20. * Units.inches, 0, 0]
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.number_coach_seats = 4.
fuselage.tag = 'fuselage'
fuselage.differential_pressure = 8 * Units.psi # Maximum differential pressure
fuselage.width = 42. * Units.inches # Width of the fuselage
fuselage.heights.maximum = 62. * Units.inches # Height of the fuselage
fuselage.lengths.total = 326. * Units.inches # Length of the fuselage
fuselage.lengths.empennage = 161. * Units.inches
fuselage.lengths.cabin = 105. * Units.inches
fuselage.lengths.structure = fuselage.lengths.total - fuselage.lengths.empennage
fuselage.mass_properties.volume = .4 * fuselage.lengths.total * (
np.pi / 4.) * (fuselage.heights.maximum**2.
) #try this as approximation
fuselage.mass_properties.internal_volume = .3 * fuselage.lengths.total * (
np.pi / 4.) * (fuselage.heights.maximum**2.)
fuselage.areas.wetted = 30000. * Units.inches**2.
fuselage.seats_abreast = 2.
fuselage.fineness.nose = 1.6
fuselage.fineness.tail = 2.
fuselage.lengths.nose = 60. * Units.inches
fuselage.heights.at_quarter_length = 62. * Units.inches
fuselage.heights.at_three_quarters_length = 62. * Units.inches
fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
fuselage.areas.front_projected = fuselage.width * fuselage.heights.maximum
fuselage.effective_diameter = 50. * Units.inches
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Landing gear
# ------------------------------------------------------------------
landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
main_gear = SUAVE.Components.Landing_Gear.Main_Landing_Gear()
nose_gear = SUAVE.Components.Landing_Gear.Nose_Landing_Gear()
main_gear.strut_length = 12. * Units.inches #guess based on picture
nose_gear.strut_length = 6. * Units.inches
landing_gear.main = main_gear
landing_gear.nose = nose_gear
#add to vehicle
vehicle.landing_gear = landing_gear
# ------------------------------------------------------------------
# Fuel
# ------------------------------------------------------------------
# define fuel weight needed to size fuel system
fuel = SUAVE.Attributes.Propellants.Aviation_Gasoline()
fuel.mass_properties = SUAVE.Components.Mass_Properties()
fuel.number_of_tanks = 1.
fuel.origin = wing.origin
fuel.internal_volume = fuel.mass_properties.mass / fuel.density #all of the fuel volume is internal
fuel.mass_properties.center_of_gravity = wing.mass_properties.center_of_gravity
fuel.mass_properties.mass = 319 * Units.lbs
vehicle.fuel = fuel
# ------------------------------------------------------------------
# Piston Propeller Network
# ------------------------------------------------------------------
# build network
net = SUAVE.Components.Energy.Networks.Internal_Combustion_Propeller()
net.tag = 'internal_combustion'
net.number_of_engines = 1.
net.identical_propellers = True
# the engine
engine = SUAVE.Components.Energy.Converters.Internal_Combustion_Engine()
engine.sea_level_power = 180. * Units.horsepower
engine.flat_rate_altitude = 0.0
engine.rated_speed = 2700. * Units.rpm
engine.power_specific_fuel_consumption = 0.52
net.engines.append(engine)
# the prop
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_of_blades = 2.0
prop.freestream_velocity = 119. * Units.knots
prop.angular_velocity = 2650. * Units.rpm
prop.tip_radius = 76. / 2. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_power = .64 * 180. * Units.horsepower
prop.variable_pitch = True
prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
prop.airfoil_polars = [[
'../Vehicles//Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
prop = propeller_design(prop)
net.propellers.append(prop)
# add the network to the vehicle
vehicle.append_component(net)
#find uninstalled avionics weight
Wuav = 2. * Units.lbs
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.mass_properties.uninstalled = Wuav
vehicle.avionics = avionics
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def main():
# This script could fail if either the design or analysis scripts fail,
# in case of failure check both. The design and analysis powers will
# differ because of karman-tsien compressibility corrections in the
# analysis scripts
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 50.0
prop_attributes.angular_velocity = 2000. * (2. * np.pi / 60.0)
prop_attributes.tip_radius = 1.5
prop_attributes.hub_radius = 0.05
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 0.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 7000.
prop_attributes = propeller_design(prop_attributes)
# Find the operating conditions
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(
prop_attributes.design_altitude)
V = prop_attributes.freestream_velocity
conditions = Data()
conditions.freestream = Data()
conditions.propulsion = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.freestream.update(atmosphere_conditions)
conditions.freestream.dynamic_viscosity = atmosphere_conditions.dynamic_viscosity
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions.propulsion.throttle = np.array([[1.0]])
conditions.frames.body.transform_to_inertial = np.array([np.eye(3)])
# Create and attach this propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.inputs.omega = np.array(prop_attributes.angular_velocity, ndmin=2)
F, Q, P, Cplast = prop.spin(conditions)
# Truth values
F_truth = 166.41590262
Q_truth = 45.21732911
P_truth = 9470.2952633 # Over 9000!
Cplast_truth = 0.00085898
error = Data()
error.Thrust = np.max(np.abs(F - F_truth))
error.Power = np.max(np.abs(P - P_truth))
error.Torque = np.max(np.abs(Q - Q_truth))
error.Cp = np.max(np.abs(Cplast - Cplast_truth))
print('Errors:')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
return
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'X57_Maxwell'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
vehicle.mass_properties.takeoff = 2550. * Units.pounds
vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
vehicle.mass_properties.cargo = 0.
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
# basic parameters
vehicle.reference_area = 15.45
vehicle.passengers = 4
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.span_efficiency = 0.9
wing.areas.reference = 15.45 * Units['meters**2']
wing.spans.projected = 11. * Units.meter
wing.chords.root = 1.67 * Units.meter
wing.chords.tip = 1.14 * Units.meter
wing.chords.mean_aerodynamic = 1.47 * Units.meter
wing.taper = wing.chords.root/wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 3.0 * Units.degrees
wing.twists.tip = 1.5 * Units.degrees
wing.origin = [2.032, 0., 0.]
wing.aerodynamic_center = [0.558, 0., 0.]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.span_efficiency = 0.95
wing.areas.reference = 3.74 * Units['meters**2']
wing.spans.projected = 3.454 * Units.meter
wing.sweeps.quarter_chord = 12.5 * Units.deg
wing.chords.root = 1.397 * Units.meter
wing.chords.tip = 0.762 * Units.meter
wing.chords.mean_aerodynamic = 1.09 * Units.meter
wing.taper = wing.chords.root/wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [6.248, 0., 0.784]
wing.aerodynamic_center = [0.508, 0., 0.]
wing.vertical = False
wing.symmetric = True
wing.high_lift = False
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.sweeps.quarter_chord = 25. * Units.deg
wing.thickness_to_chord = 0.12
wing.span_efficiency = 0.9
wing.areas.reference = 2.258 * Units['meters**2']
wing.spans.projected = 1.854 * Units.meter
wing.chords.root = 1.6764 * Units.meter
wing.chords.tip = 0.6858 * Units.meter
wing.chords.mean_aerodynamic = 1.21 * Units.meter
wing.taper = wing.chords.root/wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [6.01 ,0, 0.623]
wing.aerodynamic_center = [0.508 ,0,0]
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 2.
fuselage.fineness.nose = 1.6
fuselage.fineness.tail = 2.
fuselage.lengths.nose = 60. * Units.inches
fuselage.lengths.tail = 161. * Units.inches
fuselage.lengths.cabin = 105. * Units.inches
fuselage.lengths.total = 326. * Units.inches
fuselage.lengths.fore_space = 0.
fuselage.lengths.aft_space = 0.
fuselage.width = 42. * Units.inches
fuselage.heights.maximum = 62. * Units.inches
fuselage.heights.at_quarter_length = 62. * Units.inches
fuselage.heights.at_three_quarters_length = 62. * Units.inches
fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
fuselage.areas.side_projected = 8000. * Units.inches**2.
fuselage.areas.wetted = 30000. * Units.inches**2.
fuselage.areas.front_projected = 42.* 62. * Units.inches**2.
fuselage.effective_diameter = 50. * Units.inches
# add to vehicle
vehicle.append_component(fuselage)
#---------------------------------------------------------------------------------------------
# DEFINE PROPELLER
#---------------------------------------------------------------------------------------------
# build network
net = Battery_Propeller()
net.number_of_engines = 2.
net.nacelle_diameter = 42 * Units.inches
net.engine_length = 0.01 * Units.inches
net.areas = Data()
net.areas.wetted = 0.01*(2*np.pi*0.01/2)
# Component 1 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 2 the Propeller
# Design the Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_blades = 2.0
prop.freestream_velocity = 135.*Units['mph']
prop.angular_velocity = 1300. * Units.rpm # 2400
prop.tip_radius = 76./2. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_altitude = 12000. * Units.feet
prop.design_thrust = 800.
prop.origin = [[2.,2.5,0.784]] # prop influence
prop.symmetry = True
prop = propeller_design(prop)
net.propeller = prop
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
bat.mass_properties.mass = 500. * Units.kg
bat.specific_energy = 350. * Units.Wh/Units.kg
bat.resistance = 0.006
bat.max_voltage = 500.
initialize_from_mass(bat,bat.mass_properties.mass)
net.battery = bat
net.voltage = bat.max_voltage
# Component 9 Miscellaneous Systems
sys = SUAVE.Components.Systems.System()
sys.mass_properties.mass = 5 # kg
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller motor
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
etam = 0.95
v = bat.max_voltage *3/4
omeg = prop.angular_velocity
io = 4.0
start_kv = 1
end_kv = 25
# do optimization to find kv or just do a linspace then remove all negative values, take smallest one use 0.05 change
# essentially you are sizing the motor for a particular rpm which is sized for a design tip mach
# this reduces the bookkeeping errors
possible_kv_vals = np.linspace(start_kv,end_kv,(end_kv-start_kv)*20 +1 , endpoint = True) * Units.rpm
res_kv_vals = ((v-omeg/possible_kv_vals)*(1.-etam*v*possible_kv_vals/omeg))/io
positive_res_vals = np.extract(res_kv_vals > 0 ,res_kv_vals)
kv_idx = np.where(res_kv_vals == min(positive_res_vals))[0][0]
kv = possible_kv_vals[kv_idx]
res = min(positive_res_vals)
motor.mass_properties.mass = 10. * Units.kg
motor.origin = prop.origin
motor.propeller_radius = prop.tip_radius
motor.speed_constant = 0.35
motor.resistance = res
motor.no_load_current = io
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1. # Gear box efficiency
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 10. #Watts
payload.mass_properties.mass = 1.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 20. #Watts
net.avionics = avionics
# add the solar network to the vehicle
vehicle.append_component(net)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Cessna_172_SP'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
vehicle.mass_properties.takeoff = 2550. * Units.pounds
vehicle.mass_properties.max_zero_fuel = 2555. * Units.pounds
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
# basic parameters
vehicle.reference_area = 174. * Units.feet**2
vehicle.passengers = 4
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 174. * Units.feet**2
wing.spans.projected = 36. * Units.feet + 1. * Units.inches
wing.chords.root = 66. * Units.inches
wing.chords.tip = 45. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[80. * Units.inches, 0, 36.75 * Units.inches]]
wing.vertical = False
wing.symmetric = True
wing.high_lift = False
wing.dynamic_pressure_ratio = 1.0
# Wing Segments
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Root'
segment.percent_span_location = 0.0
segment.twist = 3. * Units.deg
segment.root_chord_percent = 1.
segment.thickness_to_chord = 0.12
segment.dihedral_outboard = 2.5 * Units.degrees
segment.sweeps.quarter_chord = 0
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Break'
segment.percent_span_location = 0.531177829
segment.twist = 2. * Units.deg
segment.root_chord_percent = 1.0
segment.thickness_to_chord = 0.12
segment.dihedral_outboard = 5 * Units.degrees
segment.sweeps.quarter_chord = -3. * Units.degrees
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Tip'
segment.percent_span_location = 1.
segment.twist = 1. * Units.degrees
segment.root_chord_percent = 0.67
segment.thickness_to_chord = 0.12
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = segment_properties(wing)
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Horizontal_Tail()
wing.tag = 'horizontal_stabilizer'
wing.sweeps.quarter_chord = 19.5 * Units.deg
wing.thickness_to_chord = 0.12
wing.spans.projected = 135. * Units.inches
wing.areas.reference = 5500 * Units.inches**2
wing.chords.root = 55. * Units.inches
wing.chords.tip = 28. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[253. * Units.inches, 0, 0]]
wing.vertical = False
wing.symmetric = True
wing.high_lift = False
wing.dynamic_pressure_ratio = 0.9
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.tag = 'vertical_stabilizer'
wing.sweeps.quarter_chord = 48. * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 3500. * Units.inches**2
wing.spans.projected = 56. * Units.inches
wing.chords.root = 64. * Units.inches
wing.chords.tip = 30. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[240. * Units.inches, 0, 0]]
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Strut
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'strut'
wing.sweeps.quarter_chord = 0. * Units.deg
wing.thickness_to_chord = 0.4
wing.areas.reference = 660. * Units.inches**2
wing.spans.projected = 200. * Units.inches
wing.chords.root = 6. * Units.inches
wing.chords.tip = 6. * Units.inches
wing.chords.mean_aerodynamic = 6. * Units.inches
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[
80. * Units.inches, 20. * Units.inches, -17. * Units.inches
]]
wing.dihedral = 30.0 * Units.degrees
wing.vertical = False
wing.symmetric = True
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.number_coach_seats = 4.
fuselage.tag = 'fuselage'
fuselage.differential_pressure = 0.
fuselage.width = 42. * Units.inches
fuselage.heights.maximum = 62. * Units.inches
fuselage.lengths.total = 326. * Units.inches
fuselage.lengths.empennage = 161. * Units.inches
fuselage.lengths.cabin = 105. * Units.inches
fuselage.lengths.structure = fuselage.lengths.total - fuselage.lengths.empennage
fuselage.mass_properties.volume = .4 * fuselage.lengths.total * (
np.pi / 4.) * (fuselage.heights.maximum**2.
) #try this as approximation
fuselage.mass_properties.internal_volume = .3 * fuselage.lengths.total * (
np.pi / 4.) * (fuselage.heights.maximum**2.)
fuselage.areas.wetted = 20.396
fuselage.seats_abreast = 2.
fuselage.fineness.nose = 1.6
fuselage.fineness.tail = 2.
fuselage.lengths.nose = 60. * Units.inches
fuselage.heights.at_quarter_length = 62. * Units.inches
fuselage.heights.at_three_quarters_length = 62. * Units.inches
fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
fuselage.areas.front_projected = fuselage.width * fuselage.heights.maximum
fuselage.effective_diameter = 50. * Units.inches
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.
segment.percent_z_location = 16.75 / 326.
segment.height = 0. * Units.inches
segment.width = 0. * Units.inches
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.070433529
segment.percent_z_location = 0.023187437
segment.height = 28.089241 * Units.inches
segment.width = 40. * Units.inches
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.16766091
segment.percent_z_location = 0.014375532
segment.height = 38.872257 * Units.inches
segment.width = 42. * Units.inches
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.243802224
segment.percent_z_location = 0.026573779
segment.height = 56.19732 * Units.inches
segment.width = 42. * Units.inches
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 0.443495521
segment.percent_z_location = 0.015
segment.height = 46 * Units.inches
segment.width = 42. * Units.inches
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 0.508342393
segment.percent_z_location = -0.000878305
segment.height = 34.067009 * Units.inches
segment.width = 42. * Units.inches
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_6'
segment.percent_x_location = 0.857801856
segment.percent_z_location = -0.022517885
segment.height = 15.538013 * Units.inches
segment.width = 10 * Units.inches
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Piston Propeller Network
# ------------------------------------------------------------------
# build network
net = SUAVE.Components.Energy.Networks.Internal_Combustion_Propeller()
net.tag = 'internal_combustion'
net.number_of_engines = 1.
net.identical_propellers = True
# the engine
engine = SUAVE.Components.Energy.Converters.Internal_Combustion_Engine()
engine.sea_level_power = 180. * Units.horsepower
engine.flat_rate_altitude = 0.0
engine.rated_speed = 2700. * Units.rpm
engine.power_specific_fuel_consumption = 0.52
net.engines.append(engine)
# the prop
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_of_blades = 2.0
prop.origin = [[0.0, 0.0, 16.75 * Units.inches]]
prop.freestream_velocity = 119. * Units.knots
prop.angular_velocity = 2650. * Units.rpm
prop.tip_radius = 76. / 2. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_power = .64 * 180. * Units.horsepower
prop.variable_pitch = False
prop.airfoil_geometry = ['./Airfoils/NACA_4412.txt']
prop.airfoil_polars = [[
'./Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
prop = propeller_design(prop)
net.propellers.append(prop)
# add the network to the vehicle
vehicle.append_component(net)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Lift_Cruise_CRM'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2450. * Units.lb
vehicle.mass_properties.operating_empty = 2250. * Units.lb # Approximate
vehicle.mass_properties.max_takeoff = 2450. * Units.lb # Approximate
vehicle.mass_properties.max_payload = 200. * Units.lb
vehicle.mass_properties.center_of_gravity = [[2.0144, 0.,
0.]] # Approximate
# basic parameters
vehicle.reference_area = 10.76
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
# ------------------------------------------------------------------
# WINGS
# ------------------------------------------------------------------
# WING PROPERTIES
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 10.76
wing.sweeps.quarter_chord = 0.0 * Units.degrees
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.spans.projected = 35.0 * Units.feet
wing.chords.root = 3.25 * Units.feet
wing.total_length = 3.25 * Units.feet
wing.chords.tip = 3.25 * Units.feet
wing.chords.mean_aerodynamic = 3.25 * Units.feet
wing.dihedral = 1.0 * Units.degrees
wing.areas.reference = 113.75 * Units.feet**2
wing.areas.wetted = 227.5 * Units.feet**2
wing.areas.exposed = 227.5 * Units.feet**2
wing.twists.root = 4.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[1.5, 0., 0.]]
wing.aerodynamic_center = [1.975, 0., 0.]
wing.winglet_fraction = 0.0
wing.symmetric = True
wing.vertical = False
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_1'
segment.percent_span_location = 0.
segment.twist = 0.
segment.root_chord_percent = 1.5
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 8.5 * Units.degrees
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_2'
segment.percent_span_location = 0.227
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0 * Units.degrees
segment.thickness_to_chord = 0.12
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_3'
segment.percent_span_location = 1.0
segment.twist = 0.
segment.root_chord_percent = 1.0
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0 * Units.degrees
segment.thickness_to_chord = 0.12
wing.Segments.append(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# WING PROPERTIES
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_tail'
wing.aspect_ratio = 4.0
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.spans.projected = 8.0 * Units.feet
wing.chords.root = 2.0 * Units.feet
wing.total_length = 2.0 * Units.feet
wing.chords.tip = 2.0 * Units.feet
wing.chords.mean_aerodynamic = 2.0 * Units.feet
wing.dihedral = 0. * Units.degrees
wing.areas.reference = 16.0 * Units.feet**2
wing.areas.wetted = 32.0 * Units.feet**2
wing.areas.exposed = 32.0 * Units.feet**2
wing.twists.root = 0. * Units.degrees
wing.twists.tip = 0. * Units.degrees
wing.origin = [[4.0, 0.0, 0.205]]
wing.aerodynamic_center = [4.2, 0., 0.]
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
# WING PROPERTIES
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_tail_1'
wing.aspect_ratio = 2.
wing.sweeps.quarter_chord = 20.0 * Units.degrees
wing.thickness_to_chord = 0.12
wing.taper = 0.5
wing.spans.projected = 3.0 * Units.feet
wing.chords.root = 2.0 * Units.feet
wing.total_length = 2.0 * Units.feet
wing.chords.tip = 1.0 * Units.feet
wing.chords.mean_aerodynamic = 1.5 * Units.feet
wing.areas.reference = 4.5 * Units.feet**2
wing.areas.wetted = 9.0 * Units.feet**2
wing.areas.exposed = 9.0 * Units.feet**2
wing.twists.root = 0. * Units.degrees
wing.twists.tip = 0. * Units.degrees
wing.origin = [[4.0, 4.0 * 0.3048, 0.205]]
wing.aerodynamic_center = [4.2, 0, 0]
wing.winglet_fraction = 0.0
wing.vertical = True
wing.symmetric = False
# add to vehicle
vehicle.append_component(wing)
# WING PROPERTIES
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_tail_2'
wing.aspect_ratio = 2.
wing.sweeps.quarter_chord = 20.0 * Units.degrees
wing.thickness_to_chord = 0.12
wing.taper = 0.5
wing.spans.projected = 3.0 * Units.feet
wing.chords.root = 2.0 * Units.feet
wing.total_length = 2.0 * Units.feet
wing.chords.tip = 1.0 * Units.feet
wing.chords.mean_aerodynamic = 1.5 * Units.feet
wing.areas.reference = 4.5 * Units.feet**2
wing.areas.wetted = 9.0 * Units.feet**2
wing.areas.exposed = 9.0 * Units.feet**2
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[4.0, -4.0 * 0.3048, 0.205]]
wing.aerodynamic_center = [4.2, 0, 0]
wing.winglet_fraction = 0.0
wing.vertical = True
wing.symmetric = False
# add to vehicle
vehicle.append_component(wing)
# ---------------------------------------------------------------
# FUSELAGE
# ---------------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.configuration = 'Tube_Wing'
fuselage.seats_abreast = 2.
fuselage.seat_pitch = 3.
fuselage.fineness.nose = 0.88
fuselage.fineness.tail = 1.13
fuselage.lengths.nose = 3.2 * Units.feet
fuselage.lengths.tail = 6.4 * Units.feet
fuselage.lengths.cabin = 6.4 * Units.feet
fuselage.lengths.total = 5.1
fuselage.width = 5.85 * Units.feet
fuselage.heights.maximum = 4.65 * Units.feet
fuselage.heights.at_quarter_length = 3.75 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 4.65 * Units.feet
fuselage.heights.at_three_quarters_length = 4.26 * Units.feet
fuselage.areas.wetted = 236. * Units.feet**2
fuselage.areas.front_projected = 0.14 * Units.feet**2
fuselage.effective_diameter = 5.85 * Units.feet
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.
segment.percent_z_location = -0.267 / 4.10534
segment.height = 0.1
segment.width = 0.1
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.2579 / 4.10534
segment.percent_z_location = -0.05881 / 1.372
segment.height = 0.5201
segment.width = 0.75
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.9939 / 4.10534
segment.percent_z_location = -0.0446 / 4.10534
segment.height = 1.18940
segment.width = 1.42045
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 1.95060 / 4.10534
segment.percent_z_location = 0
segment.height = 1.37248
segment.width = 1.35312
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 3.02797 / 4.10534
segment.percent_z_location = 0.25 / 4.10534
segment.height = 0.6
segment.width = 0.4
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 1.
segment.percent_z_location = 0.42522 / 4.10534
segment.height = 0.05
segment.width = 0.05
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#-------------------------------------------------------------------
# INNER BOOMS
#-------------------------------------------------------------------
long_boom = SUAVE.Components.Fuselages.Fuselage()
long_boom.tag = 'boom_1r'
long_boom.configuration = 'boom'
long_boom.origin = [[0.543, 1.630, -0.326]]
long_boom.seats_abreast = 0.
long_boom.seat_pitch = 0.0
long_boom.fineness.nose = 0.950
long_boom.fineness.tail = 1.029
long_boom.lengths.nose = 0.2
long_boom.lengths.tail = 0.2
long_boom.lengths.cabin = 2.5
long_boom.lengths.total = 3.5
long_boom.width = 0.15
long_boom.heights.maximum = 0.15
long_boom.heights.at_quarter_length = 0.15
long_boom.heights.at_three_quarters_length = 0.15
long_boom.heights.at_wing_root_quarter_chord = 0.15
long_boom.areas.wetted = 0.018
long_boom.areas.front_projected = 0.018
long_boom.effective_diameter = 0.15
long_boom.differential_pressure = 0.
long_boom.symmetric = True
long_boom.index = 1
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.
segment.percent_z_location = 0.0
segment.height = 0.05
segment.width = 0.05
long_boom.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.2 / 5.6
segment.percent_z_location = 0.
segment.height = 0.15
segment.width = 0.15
long_boom.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 5.4 / 5.6
segment.percent_z_location = 0.
segment.height = 0.15
segment.width = 0.15
long_boom.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 1.
segment.percent_z_location = 0.
segment.height = 0.05
segment.width = 0.05
long_boom.Segments.append(segment)
# add to vehicle
vehicle.append_component(long_boom)
# add left long boom
long_boom = deepcopy(vehicle.fuselages.boom_1r)
long_boom.origin[0][1] = -long_boom.origin[0][1]
long_boom.tag = 'Boom_1L'
long_boom.index = 1
vehicle.append_component(long_boom)
#-------------------------------------------------------------------
# OUTER BOOMS
#-------------------------------------------------------------------
short_boom = SUAVE.Components.Fuselages.Fuselage()
short_boom.tag = 'boom_2r'
short_boom.configuration = 'boom'
short_boom.origin = [[0.543, 2.826, -0.326]]
short_boom.seats_abreast = 0.
short_boom.seat_pitch = 0.0
short_boom.fineness.nose = 0.950
short_boom.fineness.tail = 1.029
short_boom.lengths.nose = 0.2
short_boom.lengths.tail = 0.2
short_boom.lengths.cabin = 2.0
short_boom.lengths.total = 3.3
short_boom.width = 0.15
short_boom.heights.maximum = 0.15
short_boom.heights.at_quarter_length = 0.15
short_boom.heights.at_three_quarters_length = 0.15
short_boom.heights.at_wing_root_quarter_chord = 0.15
short_boom.areas.wetted = 0.018
short_boom.areas.front_projected = 0.018
short_boom.effective_diameter = 0.15
short_boom.differential_pressure = 0.
short_boom.y_pitch_count = 2
short_boom.y_pitch = 1.196
short_boom.symmetric = True
short_boom.index = 1
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.
segment.percent_z_location = 0.0
segment.height = 0.05
segment.width = 0.05
short_boom.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.2 / 3.3
segment.percent_z_location = 0.
segment.height = 0.15
segment.width = 0.15
short_boom.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 3.1 / 3.3
segment.percent_z_location = 0.
segment.height = 0.15
segment.width = 0.15
short_boom.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 1.
segment.percent_z_location = 0.
segment.height = 0.05
segment.width = 0.05
short_boom.Segments.append(segment)
# add to vehicle
vehicle.append_component(short_boom)
# add outer right boom
short_boom = deepcopy(vehicle.fuselages.boom_2r)
short_boom.origin[0][1] = short_boom.y_pitch + short_boom.origin[0][1]
short_boom.tag = 'boom_3r'
short_boom.index = 1
vehicle.append_component(short_boom)
# add inner left boom
short_boom = deepcopy(vehicle.fuselages.boom_2r)
short_boom.origin[0][1] = -(short_boom.origin[0][1])
short_boom.tag = 'boom_2l'
short_boom.index = 1
vehicle.append_component(short_boom)
short_boom = deepcopy(vehicle.fuselages.boom_2r)
short_boom.origin[0][1] = -(short_boom.origin[0][1] + short_boom.y_pitch)
short_boom.tag = 'boom_3l'
short_boom.index = 1
vehicle.append_component(short_boom)
#------------------------------------------------------------------
# Nacelles
#------------------------------------------------------------------
rotor_nacelle = SUAVE.Components.Nacelles.Nacelle()
rotor_nacelle.tag = 'rotor_nacelle'
rotor_nacelle_origins = [[0.543, 1.63, -0.126], [0.543, -1.63, -0.126],
[3.843, 1.63, -0.126], [3.843, -1.63, -0.126],
[0.543, 2.826, -0.126], [0.543, -2.826, -0.126],
[3.843, 2.826, -0.126], [3.843, -2.826, -0.126],
[0.543, 4.022, -0.126], [0.543, -4.022, -0.126],
[3.843, 4.022, -0.126], [3.843, -4.022, -0.126]]
rotor_nacelle.length = 0.25
rotor_nacelle.diameter = 0.25
rotor_nacelle.orientation_euler_angles = [0, -90 * Units.degrees, 0.]
rotor_nacelle.flow_through = False
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_1'
nac_segment.percent_x_location = 0.0
nac_segment.height = 0.2
nac_segment.width = 0.2
rotor_nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_2'
nac_segment.percent_x_location = 0.25
nac_segment.height = 0.25
nac_segment.width = 0.25
rotor_nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_3'
nac_segment.percent_x_location = 0.5
nac_segment.height = 0.3
nac_segment.width = 0.3
rotor_nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_4'
nac_segment.percent_x_location = 0.75
nac_segment.height = 0.25
nac_segment.width = 0.25
rotor_nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_5'
nac_segment.percent_x_location = 1.0
nac_segment.height = 0.2
nac_segment.width = 0.2
rotor_nacelle.append_segment(nac_segment)
rotor_nacelle.areas.wetted = np.pi * rotor_nacelle.diameter * rotor_nacelle.length + 0.5 * np.pi * rotor_nacelle.diameter**2
for idx in range(12):
nacelle = deepcopy(rotor_nacelle)
nacelle.tag = 'nacelle_' + str(idx)
nacelle.origin = [rotor_nacelle_origins[idx]]
vehicle.append_component(nacelle)
#------------------------------------------------------------------
# network
#------------------------------------------------------------------
net = Lift_Cruise()
net.number_of_lift_rotor_engines = 12
net.number_of_propeller_engines = 1
net.nacelle_diameter = 0.6 * Units.feet
net.engine_length = 0.5 * Units.feet
net.areas = Data()
net.areas.wetted = np.pi * net.nacelle_diameter * net.engine_length + 0.5 * np.pi * net.nacelle_diameter**2
net.voltage = 500.
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
lift_rotor_esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller(
)
lift_rotor_esc.efficiency = 0.95
net.lift_rotor_esc = lift_rotor_esc
propeller_esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller(
)
propeller_esc.efficiency = 0.95
net.propeller_esc = propeller_esc
#------------------------------------------------------------------
# Design Payload
#------------------------------------------------------------------
payload = SUAVE.Components.Energy.Peripherals.Avionics()
payload.power_draw = 0.
payload.mass_properties.mass = 200. * Units.kg
net.payload = payload
#------------------------------------------------------------------
# Design Avionics
#------------------------------------------------------------------
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 200. * Units.watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiNiMnCoO2_18650(
)
bat.mass_properties.mass = 500. * Units.kg
bat.max_voltage = net.voltage
initialize_from_mass(bat)
# Here we, are going to assume a battery pack module shape. This step is optional but
# required for thermal analysis of tge pack
number_of_modules = 10
bat.module_config.total = int(
np.ceil(bat.pack_config.total / number_of_modules))
bat.module_config.normal_count = int(
np.ceil(bat.module_config.total / bat.pack_config.series))
bat.module_config.parallel_count = int(
np.ceil(bat.module_config.total / bat.pack_config.parallel))
net.battery = bat
#------------------------------------------------------------------
# Design Rotors and Propellers
#------------------------------------------------------------------
# atmosphere and flight conditions for propeller/rotor design
g = 9.81 # gravitational acceleration
S = vehicle.reference_area # reference area
speed_of_sound = 340 # speed of sound
rho = 1.22 # reference density
fligth_CL = 0.75 # cruise target lift coefficient
AR = vehicle.wings.main_wing.aspect_ratio # aspect ratio
Cd0 = 0.06 # profile drag
Cdi = fligth_CL**2 / (np.pi * AR * 0.98) # induced drag
Cd = Cd0 + Cdi # total drag
V_inf = 110. * Units['mph'] # freestream velocity
Drag = S * (0.5 * rho * V_inf**2) * Cd # cruise drag
Hover_Load = vehicle.mass_properties.takeoff * g # hover load
net.identical_propellers = True
net.identical_lift_rotors = True
# Thrust Propeller
propeller = SUAVE.Components.Energy.Converters.Propeller()
propeller.number_of_blades = 3
propeller.freestream_velocity = V_inf
propeller.tip_radius = 1.0668
propeller.hub_radius = 0.21336
propeller.design_tip_mach = 0.5
propeller.angular_velocity = propeller.design_tip_mach * speed_of_sound / propeller.tip_radius
propeller.design_Cl = 0.7
propeller.design_altitude = 1000 * Units.feet
propeller.design_thrust = (Drag * 2.5) / net.number_of_propeller_engines
propeller.variable_pitch = True
propeller.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
propeller.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
propeller.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
propeller = propeller_design(propeller,
number_of_airfoil_section_points=50)
propeller.origin = [[16. * 0.3048, 0., 2.02 * 0.3048]]
net.propellers.append(propeller)
# Lift Rotors
lift_rotor = SUAVE.Components.Energy.Converters.Lift_Rotor()
lift_rotor.tip_radius = 2.8 * Units.feet
lift_rotor.hub_radius = 0.35 * Units.feet
lift_rotor.number_of_blades = 2
lift_rotor.design_tip_mach = 0.65
lift_rotor.disc_area = np.pi * (lift_rotor.tip_radius**2)
lift_rotor.freestream_velocity = 500. * Units['ft/min']
lift_rotor.angular_velocity = lift_rotor.design_tip_mach * speed_of_sound / lift_rotor.tip_radius
lift_rotor.design_Cl = 0.7
lift_rotor.design_altitude = 20 * Units.feet
lift_rotor.design_thrust = Hover_Load / (
net.number_of_lift_rotor_engines - 1
) # contingency for one-engine-inoperative condition
lift_rotor.variable_pitch = True
lift_rotor.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
lift_rotor.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
lift_rotor.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
lift_rotor = propeller_design(lift_rotor)
# Appending rotors with different origins
rotations = [1, -1, 1, -1, 1, -1, 1, -1, 1, -1, 1, -1]
origins = [[0.543, 1.63, -0.126], [0.543, -1.63, -0.126],
[3.843, 1.63, -0.126], [3.843, -1.63, -0.126],
[0.543, 2.826, -0.126], [0.543, -2.826, -0.126],
[3.843, 2.826, -0.126], [3.843, -2.826, -0.126],
[0.543, 4.022, -0.126], [0.543, -4.022, -0.126],
[3.843, 4.022, -0.126], [3.843, -4.022, -0.126]]
for ii in range(12):
lift_rotor = deepcopy(lift_rotor)
lift_rotor.tag = 'lift_rotor'
lift_rotor.rotation = rotations[ii]
lift_rotor.origin = [origins[ii]]
net.lift_rotors.append(lift_rotor)
net.number_of_lift_rotor_engines = 12
# append propellers to vehicle
net.lift_rotor = lift_rotor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
propeller_motor = SUAVE.Components.Energy.Converters.Motor()
propeller_motor.efficiency = 0.95
propeller_motor.nominal_voltage = bat.max_voltage
propeller_motor.mass_properties.mass = 2.0 * Units.kg
propeller_motor.origin = propeller.origin
propeller_motor.propeller_radius = propeller.tip_radius
propeller_motor.no_load_current = 2.0
propeller_motor = size_optimal_motor(propeller_motor, propeller)
net.propeller_motors.append(propeller_motor)
# Rotor (Lift) Motor
lift_rotor_motor = SUAVE.Components.Energy.Converters.Motor()
lift_rotor_motor.efficiency = 0.85
lift_rotor_motor.nominal_voltage = bat.max_voltage * 3 / 4
lift_rotor_motor.mass_properties.mass = 3. * Units.kg
lift_rotor_motor.origin = lift_rotor.origin
lift_rotor_motor.propeller_radius = lift_rotor.tip_radius
lift_rotor_motor.gearbox_efficiency = 1.0
lift_rotor_motor.no_load_current = 4.0
lift_rotor_motor = size_optimal_motor(lift_rotor_motor, lift_rotor)
# Appending motors with different origins
for _ in range(12):
lift_rotor_motor = deepcopy(lift_rotor_motor)
lift_rotor_motor.tag = 'motor'
net.lift_rotor_motors.append(lift_rotor_motor)
# append motor origin spanwise locations onto wing data structure
vehicle.append_component(net)
# Add extra drag sources from motors, props, and landing gear. All of these hand measured
motor_height = .25 * Units.feet
motor_width = 1.6 * Units.feet
propeller_width = 1. * Units.inches
propeller_height = propeller_width * .12
main_gear_width = 1.5 * Units.inches
main_gear_length = 2.5 * Units.feet
nose_gear_width = 2. * Units.inches
nose_gear_length = 2. * Units.feet
nose_tire_height = (0.7 + 0.4) * Units.feet
nose_tire_width = 0.4 * Units.feet
main_tire_height = (0.75 + 0.5) * Units.feet
main_tire_width = 4. * Units.inches
total_excrescence_area_spin = 12.*motor_height*motor_width + 2.*main_gear_length*main_gear_width \
+ nose_gear_width*nose_gear_length + 2*main_tire_height*main_tire_width\
+ nose_tire_height*nose_tire_width
total_excrescence_area_no_spin = total_excrescence_area_spin + 12 * propeller_height * propeller_width
vehicle.excrescence_area_no_spin = total_excrescence_area_no_spin
vehicle.excrescence_area_spin = total_excrescence_area_spin
vehicle.wings['main_wing'].motor_spanwise_locations = np.multiply(
2. / 36.25, [
-5.435, -5.435, -9.891, -9.891, -14.157, -14.157, 5.435, 5.435,
9.891, 9.891, 14.157, 14.157
])
vehicle.wings['main_wing'].winglet_fraction = 0.0
vehicle.wings['main_wing'].thickness_to_chord = 0.18
vehicle.wings['main_wing'].chords.mean_aerodynamic = 0.9644599977664836
vehicle.weight_breakdown = empty(vehicle)
compute_component_centers_of_gravity(vehicle)
vehicle.center_of_gravity()
return vehicle
def define_vehicle():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Solar'
vehicle.propulsors.propulsor = SUAVE.Components.Energy.Networks.Solar_Network(
)
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 200 * Units.kg
vehicle.mass_properties.operating_empty = 200 * Units.kg
vehicle.mass_properties.max_takeoff = 200 * Units.kg
# basic parameters
vehicle.reference_area = 80. # m^2
vehicle.envelope.ultimate_load = 2.0
vehicle.qm = 0.5 * 1.225 * (25.**2.) #Max q
vehicle.Ltb = 10. # Tail boom length
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Main Wing'
wing.areas.reference = vehicle.reference_area #
wing.spans.projected = 40. #m
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.sweep = 0. * Units.deg #
wing.symmetric = True #
wing.thickness_to_chord = 0.12 #
wing.taper = 1. #
# size the wing planform
SUAVE.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected #
wing.areas.exposed = 0.8 * wing.areas.wetted # might not be needed as input
wing.areas.affected = 0.6 * wing.areas.wetted # part of high lift system
wing.span_efficiency = 0.97 #
wing.twists.root = 0.0 * Units.degrees #
wing.twists.tip = 0.0 * Units.degrees #
wing.highlift = False
wing.vertical = False
wing.eta = 1.0
wing.number_ribs = 26. # ?
wing.number_end_ribs = 2.
wing.transition_x_u = 0.
wing.transition_x_l = 0.
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Horizontal Stabilizer'
wing.areas.reference = vehicle.reference_area * .15 #m^2
wing.aspect_ratio = 20. #
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.sweep = 0 * Units.deg #
wing.symmetric = True
wing.thickness_to_chord = 0.12 #
wing.taper = 1. #
wing.twists.root = 0.0 * Units.degrees #
wing.twists.tip = 0.0 * Units.degrees #
# size the wing planform
SUAVE.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected #
wing.areas.exposed = 0.8 * wing.areas.wetted # might not be needed as input
wing.areas.affected = 0.6 * wing.areas.wetted # part of high lift system
wing.span_efficiency = 0.95 #
wing.twists.root = 0. #
wing.twists.tip = 0. #
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Vertical Stabilizer'
wing.areas.reference = vehicle.reference_area * .1 #m^2
wing.aspect_ratio = 20. #
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.sweep = 0 * Units.deg #
wing.symmetric = True
wing.thickness_to_chord = 0.12 #
wing.taper = 1. #
wing.twists.root = 0.0 * Units.degrees #
wing.twists.tip = 0.0 * Units.degrees #
# size the wing planform
SUAVE.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected #
wing.areas.exposed = 0.8 * wing.areas.wetted # might not be needed as input
wing.areas.affected = 0.6 * wing.areas.wetted # part of high lift system
wing.span_efficiency = 0.97 #
wing.twists.root = 0.0 * Units.degrees #
wing.twists.tip = 0.0 * Units.degrees #
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------
# Propulsor
#------------------------------------------------------------------
# build network
net = Solar_Network()
net.number_motors = 1.
net.nacelle_dia = 0.2
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = vehicle.reference_area
panel.efficiency = 0.2
panel.mass_properties.mass = panel.area * 0.6
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 50.0 # freestream m/s
prop_attributes.angular_velocity = 300. * (2. * np.pi / 60.0)
prop_attributes.tip_radius = 4.25
prop_attributes.hub_radius = 0.0508
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 23.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 10000.0
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.008
motor.no_load_current = 4.5
motor.speed_constant = 120. * (2. * np.pi / 60.
) # RPM/volt converted to rad/s
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 20. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 160. # Expected current
motor.mass_properties.mass = 2.0
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 100. #Watts
payload.mass_properties.mass = 25.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 0. #Watts
net.avionics = avionics
# Component 8 the Battery # I already assume 250 Wh/kg for batteries
bat = SUAVE.Components.Energy.Storages.Battery()
bat.mass_properties.mass = 50 * Units.kg
bat.type = 'Li-Ion'
bat.resistance = 0.0 #This needs updating
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 100.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# Calculate the vehicle mass
vehicle.mass_properties.breakdown = SUAVE.Methods.Weights.Correlations.Human_Powered.empty(
vehicle)
# ------------------------------------------------------------------
# Simple Aerodynamics Model
# ------------------------------------------------------------------
aerodynamics = SUAVE.Attributes.Aerodynamics.Fidelity_Zero()
aerodynamics.initialize(vehicle)
vehicle.aerodynamics_model = aerodynamics
# ------------------------------------------------------------------
# Not so Simple Propulsion Model
# ------------------------------------------------------------------
vehicle.propulsion_model = net
# ------------------------------------------------------------------
# Define Configurations
# ------------------------------------------------------------------
# --- Takeoff Configuration ---
config = vehicle.new_configuration("takeoff")
# this configuration is derived from the baseline vehicle
# --- Cruise Configuration ---
config = vehicle.new_configuration("cruise")
# this configuration is derived from vehicle.Configs.takeoff
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def setup_vehicle():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
# Create a vehicle and set level properties
vehicle = SUAVE.Vehicle()
vehicle.tag = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2500. * Units.lb
vehicle.mass_properties.operating_empty = 2150. * Units.lb
vehicle.mass_properties.max_takeoff = 2500. * Units.lb
vehicle.mass_properties.max_payload = 100. * Units.lb
vehicle.mass_properties.center_of_gravity = [[2.0, 0.,
0.]] # I made this up
# basic parameters
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
# ------------------------------------------------------------------
# WINGS
# ------------------------------------------------------------------
# WING PROPERTIES
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.origin = [[1.5, 0., -0.5]]
wing.spans.projected = 35.0 * Units.feet
wing.chords.root = 3.25 * Units.feet
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Root'
segment.percent_span_location = 0.
segment.twist = 0.
segment.root_chord_percent = 1.5
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 8.5 * Units.degrees
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_2'
segment.percent_span_location = 0.227
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0 * Units.degrees
segment.thickness_to_chord = 0.12
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Tip'
segment.percent_span_location = 1.0
segment.twist = 0.
segment.root_chord_percent = 1.0
segment.dihedral_outboard = 0.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0 * Units.degrees
segment.thickness_to_chord = 0.12
wing.Segments.append(segment)
# Fill out more segment properties automatically
wing = segment_properties(wing)
wing = wing_segmented_planform(wing)
## ALSO SET THE VEHICLE REFERENCE AREA
vehicle.reference_area = wing.areas.reference
# add to vehicle
vehicle.append_component(wing)
# Add a horizontal tail
# WING PROPERTIES
wing = SUAVE.Components.Wings.Horizontal_Tail()
wing.tag = 'horizontal_tail'
wing.areas.reference = 2.0
wing.taper = 0.5
wing.sweeps.quarter_chord = 20. * Units.degrees
wing.aspect_ratio = 5.0
wing.thickness_to_chord = 0.12
wing.dihedral = 5. * Units.degrees
wing.origin = [[5.5, 0.0, 0.65]]
# Fill out more segment properties automatically
wing = wing_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# Add a vertical tail
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.tag = 'vertical_tail'
wing.areas.reference = 1.0
wing.taper = 0.5
wing.sweeps.quarter_chord = 30 * Units.degrees
wing.aspect_ratio = 2.5
wing.thickness_to_chord = 0.12
wing.origin = [[5.5, 0.0, 0.65]]
# Fill out more segment properties automatically
wing = wing_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# Add a fuseelage
# ---------------------------------------------------------------
# FUSELAGE
# ---------------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 2.
fuselage.fineness.nose = 0.88
fuselage.fineness.tail = 1.13
fuselage.lengths.nose = 3.2 * Units.feet
fuselage.lengths.tail = 6.4 * Units.feet
fuselage.lengths.cabin = 6.4 * Units.feet
fuselage.lengths.total = 6.0
fuselage.width = 5.85 * Units.feet
fuselage.heights.maximum = 4.65 * Units.feet
fuselage.heights.at_quarter_length = 3.75 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 4.65 * Units.feet
fuselage.heights.at_three_quarters_length = 4.26 * Units.feet
fuselage.areas.wetted = 236. * Units.feet**2
fuselage.areas.front_projected = 0.14 * Units.feet**2
fuselage.effective_diameter = 5.85 * Units.feet
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.
segment.percent_z_location = -0.05
segment.height = 0.1
segment.width = 0.1
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.06
segment.percent_z_location = -0.05
segment.height = 0.52
segment.width = 0.75
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.25
segment.percent_z_location = -.01
segment.height = 1.2
segment.width = 1.43
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.475
segment.percent_z_location = 0
segment.height = 1.4
segment.width = 1.4
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 0.75
segment.percent_z_location = 0.06
segment.height = 0.6
segment.width = 0.4
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 1.
segment.percent_z_location = 0.1
segment.height = 0.05
segment.width = 0.05
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#-------------------------------------------------------------------
# Booms
#-------------------------------------------------------------------
# Add booms for the motors
boom = SUAVE.Components.Fuselages.Fuselage()
boom.tag = 'boom_R'
boom.origin = [[0.525, 3.0, -0.35]]
boom.lengths.nose = 0.2
boom.lengths.tail = 0.2
boom.lengths.total = 4
boom.width = 0.15
boom.heights.maximum = 0.15
boom.heights.at_quarter_length = 0.15
boom.heights.at_three_quarters_length = 0.15
boom.heights.at_wing_root_quarter_chord = 0.15
boom.effective_diameter = 0.15
boom.areas.wetted = 2 * np.pi * (0.075) * 3.5
boom.areas.front_projected = np.pi * 0.15
boom.fineness.nose = 0.15 / 0.2
boom.fineness.tail = 0.15 / 0.2
vehicle.append_component(boom)
# Now attach the mirrored boom
other_boom = deepcopy(boom)
other_boom.origin[0][1] = -boom.origin[0][1]
other_boom.tag = 'boom_L'
vehicle.append_component(other_boom)
#------------------------------------------------------------------
# Network
#------------------------------------------------------------------
net = SUAVE.Components.Energy.Networks.Lift_Cruise()
net.number_of_lift_rotor_engines = 4
net.number_of_propeller_engines = 1
net.identical_propellers = True
net.identical_lift_rotors = True
net.voltage = 400.
#------------------------------------------------------------------
# Electronic Speed Controller
#------------------------------------------------------------------
lift_rotor_esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller(
)
lift_rotor_esc.efficiency = 0.95
net.lift_rotor_esc = lift_rotor_esc
propeller_esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller(
)
propeller_esc.efficiency = 0.95
net.propeller_esc = propeller_esc
#------------------------------------------------------------------
# Payload
#------------------------------------------------------------------
payload = SUAVE.Components.Energy.Peripherals.Avionics()
payload.power_draw = 0.
net.payload = payload
#------------------------------------------------------------------
# Avionics
#------------------------------------------------------------------
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 300. * Units.watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiNiMnCoO2_18650(
)
bat.mass_properties.mass = 1000. * Units.lb
bat.max_voltage = net.voltage
initialize_from_mass(bat)
net.battery = bat
#------------------------------------------------------------------
# Design Rotors and Propellers
#------------------------------------------------------------------
# The tractor propeller
propeller = SUAVE.Components.Energy.Converters.Propeller()
propeller.origin = [[0, 0, -0.325]]
propeller.number_of_blades = 3
propeller.tip_radius = 0.9
propeller.hub_radius = 0.1
propeller.angular_velocity = 2200 * Units.rpm
propeller.freestream_velocity = 100. * Units.knots
propeller.design_Cl = 0.7
propeller.design_altitude = 5000. * Units.feet
propeller.design_thrust = 500. * Units.lbf
propeller.airfoil_geometry = ['./Airfoils/NACA_4412.txt']
propeller.airfoil_polars = [[
'./Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
propeller.airfoil_polar_stations = np.zeros((20), dtype=np.int8).tolist()
propeller = propeller_design(propeller)
net.propellers.append(propeller)
# The lift rotors
lift_rotor = SUAVE.Components.Energy.Converters.Lift_Rotor()
lift_rotor.tip_radius = 1.5
lift_rotor.hub_radius = 0.15
lift_rotor.number_of_blades = 4
lift_rotor.design_tip_mach = 0.65
lift_rotor.freestream_velocity = 500. * Units['ft/min']
lift_rotor.angular_velocity = lift_rotor.design_tip_mach * Air(
).compute_speed_of_sound() / lift_rotor.tip_radius
lift_rotor.design_Cl = 0.7
lift_rotor.design_altitude = 3000. * Units.feet
lift_rotor.design_thrust = 2500 * Units.lbf / 4
lift_rotor.variable_pitch = False
lift_rotor.airfoil_geometry = ['./Airfoils/NACA_4412.txt']
lift_rotor.airfoil_polars = [[
'./Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'./Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
lift_rotor.airfoil_polar_stations = np.zeros((20), dtype=np.int8).tolist()
lift_rotor = propeller_design(lift_rotor)
# Appending rotors with different origins
rotations = [1, -1, -1, 1]
origins = [[0.6, 3., -0.125], [4.5, 3., -0.125], [0.6, -3., -0.125],
[4.5, -3., -0.125]]
for ii in range(4):
lift_rotor = deepcopy(lift_rotor)
lift_rotor.tag = 'lift_rotor'
lift_rotor.rotation = rotations[ii]
lift_rotor.origin = [origins[ii]]
net.lift_rotors.append(lift_rotor)
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
propeller_motor = SUAVE.Components.Energy.Converters.Motor()
propeller_motor.efficiency = 0.95
propeller_motor.nominal_voltage = bat.max_voltage
propeller_motor.mass_properties.mass = 2.0 * Units.kg
propeller_motor.origin = propeller.origin
propeller_motor.propeller_radius = propeller.tip_radius
propeller_motor.no_load_current = 2.0
propeller_motor = size_optimal_motor(propeller_motor, propeller)
net.propeller_motors.append(propeller_motor)
# Rotor (Lift) Motor
lift_rotor_motor = SUAVE.Components.Energy.Converters.Motor()
lift_rotor_motor.efficiency = 0.85
lift_rotor_motor.nominal_voltage = bat.max_voltage * 3 / 4
lift_rotor_motor.mass_properties.mass = 3. * Units.kg
lift_rotor_motor.origin = lift_rotor.origin
lift_rotor_motor.propeller_radius = lift_rotor.tip_radius
lift_rotor_motor.gearbox_efficiency = 1.0
lift_rotor_motor.no_load_current = 4.0
lift_rotor_motor = size_optimal_motor(lift_rotor_motor, lift_rotor)
for _ in range(4):
lift_rotor_motor = deepcopy(lift_rotor_motor)
lift_rotor_motor.tag = 'motor'
net.lift_rotor_motors.append(lift_rotor_motor)
vehicle.append_component(net)
# Now account for things that have been overlooked for now:
vehicle.excrescence_area = 0.1
return vehicle
def main():
# ------------------------------------------------------------------
# Propulsor
# ------------------------------------------------------------------
# build network
net = Solar_Network()
net.number_motors = 1.
net.nacelle_dia = 0.2
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = 100 * Units.m
panel.efficiency = 0.18
panel.mass_properties.mass = panel.area*.600
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
#Propeller design specs
design_altitude = 0.0 * Units.km
Velocity = 10.0 # freestream m/s
RPM = 5887
Blades = 2.0
Radius = .4064
Hub_Radius = 0.05
Design_Cl = 0.7
Thrust = 0.0 #Specify either thrust or power to design for
Power = 7500. #Specify either thrust or power to design for
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = Blades
prop_attributes.freestream_velocity = Velocity
prop_attributes.angular_velocity = RPM*(2.*np.pi/60.0)
prop_attributes.tip_radius = Radius
prop_attributes.hub_radius = Hub_Radius
prop_attributes.design_Cl = Design_Cl
prop_attributes.design_altitude = design_altitude
prop_attributes.design_thrust = Thrust
prop_attributes.design_power = Power
prop_attributes = propeller_design(prop_attributes)
# Create and attach this propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.01
motor.no_load_current = 8.0
motor.speed_constant = 140.*(2.*np.pi/60.) # RPM/volt converted to rad/s
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1.
motor.expected_current = 260.
motor.mass_properties.mass = 2.0
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 0. #Watts
payload.mass_properties.mass = 0. * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 0. #Watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Battery()
bat.mass_properties.mass = 50. #kg
bat.type = 'Li-Ion'
bat.resistance = 0.0
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 50.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# Setup the conditions to run the network
conditions = Data()
conditions.propulsion = Data()
conditions.freestream = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.frames.planet = Data()
numerics = Data()
# Calculate atmospheric properties
atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(design_altitude)
conditions.propulsion.throttle = np.array([[1.0],[1.0]])
conditions.freestream.velocity = np.array([[1.0],[1.0]])
conditions.freestream.density = np.array([rho,rho])
conditions.freestream.viscosity = np.array([mu, mu])
conditions.freestream.speed_of_sound = np.array([a, a])
conditions.freestream.altitude = np.array([[design_altitude], [design_altitude]])
conditions.propulsion.battery_energy = bat.max_energy()*np.ones_like(conditions.freestream.altitude)
conditions.frames.body.inertial_rotations = np.zeros([2,3])
conditions.frames.inertial.time = np.array([[0.0],[1.0]])
numerics.integrate_time = np.array([[0, 0],[0, 1]])
conditions.frames.planet.start_time = time.strptime("Sat, Jun 21 06:00:00 2014", "%a, %b %d %H:%M:%S %Y",)
conditions.frames.planet.latitude = np.array([[0.0],[0.0]])
conditions.frames.planet.longitude = np.array([[0.0],[0.0]])
conditions.freestream.temperature = np.array([T, T])
# Run the network and print the results
F, mdot, P = net(conditions,numerics)
# Truth results
truth_F = [[538.00449442], [538.00449442]]
truth_P = [[14272.1902522],[14272.1902522]]
truth_i = [[ 249.31622624],[ 249.31622624]]
truth_rpm = [[ 6668.4094191],[ 6668.4094191]]
truth_bat = [[45000000.] , [44987534.18868808]]
error = Data()
error.Thrust = np.max(np.abs(F-truth_F))
error.Propeller_Power = np.max(np.abs(P-truth_P))
error.RPM = np.max(np.abs(conditions.propulsion.rpm-truth_rpm))
error.Current = np.max(np.abs(conditions.propulsion.current-truth_i))
error.Battery = np.max(np.abs(bat.CurrentEnergy-truth_bat))
print error
for k,v in error.items():
assert(np.abs(v)<0.001)
return
def main():
# This script could fail if either the design or analysis scripts fail,
# in case of failure check both. The design and analysis powers will
# differ because of karman-tsien compressibility corrections in the
# analysis scripts
net = Battery_Propeller()
net.number_of_engines = 2
# Design Gearbox
gearbox = SUAVE.Components.Energy.Converters.Gearbox()
gearbox.gearwheel_radius1 = 1
gearbox.gearwheel_radius2 = 1
gearbox.efficiency = 0.95
gearbox.inputs.torque = 885.550158704757
gearbox.inputs.speed = 207.16160479940007
gearbox.inputs.power = 183451.9920076409
gearbox.compute()
# Design the Propeller with airfoil geometry defined
bad_prop = SUAVE.Components.Energy.Converters.Propeller()
bad_prop.tag = "Prop_W_Aifoil"
bad_prop.number_of_blades = 2
bad_prop.number_of_engines = 1
bad_prop.freestream_velocity = 1
bad_prop.tip_radius = 0.3
bad_prop.hub_radius = 0.21336
bad_prop.design_tip_mach = 0.1
bad_prop.angular_velocity = gearbox.inputs.speed
bad_prop.design_Cl = 0.7
bad_prop.design_altitude = 1. * Units.km
bad_prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
bad_prop.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
bad_prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
bad_prop.design_thrust = 100000
bad_prop = propeller_design(bad_prop)
prop_a = SUAVE.Components.Energy.Converters.Propeller()
prop_a.tag = "Prop_W_Aifoil"
prop_a.number_of_blades = 3
prop_a.number_of_engines = 1
prop_a.freestream_velocity = 49.1744
prop_a.tip_radius = 1.0668
prop_a.hub_radius = 0.21336
prop_a.design_tip_mach = 0.65
prop_a.angular_velocity = gearbox.inputs.speed # 207.16160479940007
prop_a.design_Cl = 0.7
prop_a.design_altitude = 1. * Units.km
prop_a.airfoil_geometry = [
'../Vehicles/Airfoils/NACA_4412.txt',
'../Vehicles/Airfoils/Clark_y.txt'
]
prop_a.airfoil_polars = [
[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
],
[
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_1000000.txt'
]
]
prop_a.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 1, 1, 1, 1, 1, 1, 1
]
prop_a.design_thrust = 3054.4809132125697
prop_a = propeller_design(prop_a)
# plot propeller
plot_propeller(prop_a)
# Design the Propeller with airfoil geometry defined
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.tag = "Prop_No_Aifoil"
prop.number_of_blades = 3
prop.number_of_engines = 1
prop.freestream_velocity = 49.1744
prop.tip_radius = 1.0668
prop.hub_radius = 0.21336
prop.design_tip_mach = 0.65
prop.angular_velocity = gearbox.inputs.speed
prop.design_Cl = 0.7
prop.design_altitude = 1. * Units.km
prop.origin = [[16. * 0.3048, 0., 2.02 * 0.3048]]
prop.design_power = gearbox.outputs.power
prop = propeller_design(prop)
# Design a Rotor with airfoil geometry defined
rot_a = SUAVE.Components.Energy.Converters.Rotor()
rot_a.tag = "Rot_W_Aifoil"
rot_a.tip_radius = 2.8 * Units.feet
rot_a.hub_radius = 0.35 * Units.feet
rot_a.number_of_blades = 2
rot_a.design_tip_mach = 0.65
rot_a.number_of_engines = 12
rot_a.disc_area = np.pi * (rot_a.tip_radius**2)
rot_a.induced_hover_velocity = 12.756071638899549
rot_a.freestream_velocity = 500. * Units['ft/min']
rot_a.angular_velocity = 258.9520059992501
rot_a.design_Cl = 0.7
rot_a.design_altitude = 20 * Units.feet
rot_a.design_thrust = 2271.2220451593753
rot_a.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
rot_a.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
rot_a.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
rot_a = propeller_design(rot_a)
# Design a Rotor without airfoil geometry defined
rot = SUAVE.Components.Energy.Converters.Rotor()
rot.tag = "Rot_No_Aifoil"
rot.tip_radius = 2.8 * Units.feet
rot.hub_radius = 0.35 * Units.feet
rot.number_of_blades = 2
rot.design_tip_mach = 0.65
rot.number_of_engines = 12
rot.disc_area = np.pi * (rot.tip_radius**2)
rot.induced_hover_velocity = 12.756071638899549
rot.freestream_velocity = 500. * Units['ft/min']
rot.angular_velocity = 258.9520059992501
rot.design_Cl = 0.7
rot.design_altitude = 20 * Units.feet
rot.design_thrust = 2271.2220451593753
rot = propeller_design(rot)
# Find the operating conditions
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(rot.design_altitude)
V = prop.freestream_velocity
Vr = rot.freestream_velocity
conditions = Data()
conditions.freestream = Data()
conditions.propulsion = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.freestream.update(atmosphere_conditions)
conditions.freestream.dynamic_viscosity = atmosphere_conditions.dynamic_viscosity
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions.propulsion.throttle = np.array([[1.0]])
conditions.frames.body.transform_to_inertial = np.array([np.eye(3)])
conditions_r = copy.deepcopy(conditions)
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions_r.frames.inertial.velocity_vector = np.array([[0, Vr, 0]])
# Create and attach this propeller
prop_a.inputs.omega = np.array(prop.angular_velocity, ndmin=2)
prop.inputs.omega = np.array(prop.angular_velocity, ndmin=2)
rot_a.inputs.omega = copy.copy(prop.inputs.omega)
rot.inputs.omega = copy.copy(prop.inputs.omega)
# propeller with airfoil results
F_a, Q_a, P_a, Cplast_a, output_a, etap_a = prop_a.spin(conditions)
plot_results(output_a, prop_a, 'blue', '-', 's')
# propeller without airfoil results
conditions.propulsion.pitch_command = np.array([[1.0]]) * Units.degree
F, Q, P, Cplast, output, etap = prop.spin(conditions)
plot_results(output, prop, 'red', '-', 'o')
# rotor with airfoil results
Fr_a, Qr_a, Pr_a, Cplastr_a, outputr_a, etapr = rot_a.spin(conditions_r)
plot_results(outputr_a, rot_a, 'green', '-', '^')
# rotor with out airfoil results
conditions_r.propulsion.pitch_command = np.array([[1.0]]) * Units.degree
Fr, Qr, Pr, Cplastr, outputr, etapr = rot.spin(conditions_r)
plot_results(outputr, rot, 'black', '-', 'P')
# Truth values for propeller with airfoil geometry defined
F_a_truth = 3390.61957425
Q_a_truth = 1001.363593
P_a_truth = 207444.08891448
Cplast_a_truth = 0.10692172
# Truth values for propeller without airfoil geometry defined
F_truth = 2705.62228566
Q_truth = 815.76685175
P_truth = 168995.57015083
Cplast_truth = 0.08710442
# Truth values for rotor with airfoil geometry defined
Fr_a_truth = 1447.00285504
Qr_a_truth = 191.05249889
Pr_a_truth = 39578.74227177
Cplastr_a_truth = 0.06225539
# Truth values for rotor without airfoil geometry defined
Fr_truth = 1290.26055703
Qr_truth = 179.49371868
Pr_truth = 37184.2068134
Cplastr_truth = 0.0584889
# Store errors
error = Data()
error.Thrust_a = np.max(np.abs(F_a - F_a_truth))
error.Torque_a = np.max(np.abs(Q_a - Q_a_truth))
error.Power_a = np.max(np.abs(P_a - P_a_truth))
error.Cp_a = np.max(np.abs(Cplast_a - Cplast_a_truth))
error.Thrust = np.max(np.abs(F - F_truth))
error.Torque = np.max(np.abs(Q - Q_truth))
error.Power = np.max(np.abs(P - P_truth))
error.Cp = np.max(np.abs(Cplast - Cplast_truth))
error.Thrustr_a = np.max(np.abs(Fr_a - Fr_a_truth))
error.Torquer_a = np.max(np.abs(Qr_a - Qr_a_truth))
error.Powerr_a = np.max(np.abs(Pr_a - Pr_a_truth))
error.Cpr_a = np.max(np.abs(Cplastr_a - Cplastr_a_truth))
error.Thrustr = np.max(np.abs(Fr - Fr_truth))
error.Torquer = np.max(np.abs(Qr - Qr_truth))
error.Powerr = np.max(np.abs(Pr - Pr_truth))
error.Cpr = np.max(np.abs(Cplastr - Cplastr_truth))
print('Errors:')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
return
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Solar'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 200. * Units.kg
vehicle.mass_properties.operating_empty = 200. * Units.kg
vehicle.mass_properties.max_takeoff = 200. * Units.kg
# basic parameters
vehicle.reference_area = 80.
vehicle.envelope.ultimate_load = 2.0
vehicle.envelope.limit_load = 1.5
vehicle.envelope.maximum_dynamic_pressure = 0.5*1.225*(40.**2.) #Max q
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = vehicle.reference_area
wing.spans.projected = 40.0 * Units.meter
wing.aspect_ratio = (wing.spans.projected**2)/wing.areas.reference
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.symmetric = True
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.vertical = False
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
wing.chords.mean_aerodynamic = wing.areas.reference/wing.spans.projected
wing.chords.root = wing.areas.reference/wing.spans.projected
wing.chords.tip = wing.areas.reference/wing.spans.projected
wing.span_efficiency = 0.98
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.highlift = False
wing.vertical = False
wing.number_ribs = 26.
wing.number_end_ribs = 2.
wing.transition_x_upper = 0.6
wing.transition_x_lower = 1.0
wing.origin = [3.0,0.0,0.0] # meters
wing.aerodynamic_center = [3.0,0.0,0.0] # meters
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 20.
wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.95
wing.areas.reference = vehicle.reference_area * .15
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.spans.projected = np.sqrt(wing.aspect_ratio*wing.areas.reference)
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
wing.number_ribs = 5.0
wing.chords.root = wing.areas.reference/wing.spans.projected
wing.chords.tip = wing.areas.reference/wing.spans.projected
wing.chords.mean_aerodynamic = wing.areas.reference/wing.spans.projected
wing.origin = [10.,0.0,0.0] # meters
wing.aerodynamic_center = [0.5,0.0,0.0] # meters
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 20.
wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.97
wing.areas.reference = vehicle.reference_area * 0.1
wing.spans.projected = np.sqrt(wing.aspect_ratio*wing.areas.reference)
wing.chords.root = wing.areas.reference/wing.spans.projected
wing.chords.tip = wing.areas.reference/wing.spans.projected
wing.chords.mean_aerodynamic = wing.areas.reference/wing.spans.projected
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [10.,0.0,0.0] # meters
wing.aerodynamic_center = [0.5,0.0,0.0] # meters
wing.symmetric = True
wing.vertical = True
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------
# Propulsor
#------------------------------------------------------------------
# build network
net = Solar()
net.number_of_engines = 1.
net.nacelle_diameter = 0.2 * Units.meters
net.engine_length = 0.01 * Units.meters
net.areas = Data()
net.areas.wetted = 0.01*(2*np.pi*0.01/2.)
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = vehicle.reference_area * 0.9
panel.efficiency = 0.25
panel.mass_properties.mass = panel.area*(0.60 * Units.kg)
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 40.0 * Units['m/s']# freestream
prop_attributes.angular_velocity = 150. * Units['rpm']
prop_attributes.tip_radius = 4.25 * Units.meters
prop_attributes.hub_radius = 0.05 * Units.meters
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 14.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 3500.0 * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.008
motor.no_load_current = 4.5 * Units.ampere
motor.speed_constant = 120. * Units['rpm'] # RPM/volt converted to (rad/s)/volt
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 12. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 160. # Expected current
motor.mass_properties.mass = 2.0 * Units.kg
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 50. * Units.watts
payload.mass_properties.mass = 5.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 50. * Units.watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
bat.mass_properties.mass = 55.0 * Units.kg
bat.specific_energy = 450. * Units.Wh/Units.kg
bat.resistance = 0.05
initialize_from_mass(bat,bat.mass_properties.mass)
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 40.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# add the solar network to the vehicle
vehicle.append_component(net)
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'tail_sitter'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 0.82 * Units.kg
vehicle.mass_properties.operating_empty = 0.82 * Units.kg
vehicle.mass_properties.max_takeoff = 0.82 * Units.kg
# basic parameters
vehicle.reference_area = 0.1668
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.areas.reference = vehicle.reference_area
wing.spans.projected = 1.03 * Units.m
wing.aspect_ratio = (wing.spans.projected**2)/wing.areas.reference
wing.sweeps.quarter_chord = 5.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.dynamic_pressure_ratio = 1.0
wing.chords.mean_aerodynamic = 0.162 * Units.m
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.high_lift = False
wing.vertical = False
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------
# Propulsor
#------------------------------------------------------------------
# build network
net = SUAVE.Components.Energy.Networks.Battery_Propeller()
net.number_of_propeller_engines = 4.
net.voltage = 12.3
net.identical_propellers = True
# Component 1 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 2 the Propeller
# Design the Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_of_blades = 2.0
prop.freestream_velocity = 15.0 # freestream m/s
prop.angular_velocity = 7500. * Units['rpm']
prop.tip_radius = 4. * Units.inch
prop.hub_radius = 0.125 * Units.inch
prop.design_Cl = 0.7
prop.design_altitude = 0.1 * Units.km
prop.design_power = 200. * Units.watts
prop = propeller_design(prop)
origins = [[0., 0.15, -0.05], [0., -0.15, -0.05], [0., .35, 0.05], [0., 0.35, 0.05]]
for ii in range(4):
rotor = deepcopy(prop)
rotor.tag = 'propeller'
rotor.origin = [origins[ii]]
net.propellers.append(rotor)
# Component 3 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.speed_constant = 1500. * Units['rpm'] # RPM/volt converted to (rad/s)/volt
motor = size_from_kv(motor)
motor.gear_ratio = 1. # Gear ratio
motor.gearbox_efficiency = 1. # Gear box efficiency
motor.expected_current = 10. # Expected current
motor.propeller_radius = prop.tip_radius
for ii in range(4):
rotor_motor = deepcopy(motor)
rotor_motor.tag = 'motor'
rotor_motor.origin = [origins[ii]]
net.propeller_motors.append(rotor_motor)
# Component 4 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 0. #Watts
payload.mass_properties.mass = 0.0 * Units.kg
net.payload = payload
# Component 5 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 2. #Watts
net.avionics = avionics
# Component 6 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
bat.mass_properties.mass = 0.17 * Units.kg
bat.specific_energy = 175.*Units.Wh/Units.kg
bat.resistance = 0.003
bat.max_voltage = 11.1
initialize_from_mass(bat)
net.battery = bat
# add the solar network to the vehicle
vehicle.append_component(net)
return vehicle
def main():
# This script could fail if either the design or analysis scripts fail,
# in case of failure check both. The design and analysis powers will
# differ because of karman-tsien compressibility corrections in the
# analysis scripts
net = Battery_Propeller()
net.number_of_propeller_engines = 2
# Design Gearbox
gearbox = SUAVE.Components.Energy.Converters.Gearbox()
gearbox.gearwheel_radius1 = 1
gearbox.gearwheel_radius2 = 1
gearbox.efficiency = 0.95
gearbox.inputs.torque = 885.550158704757
gearbox.inputs.speed = 207.16160479940007
gearbox.inputs.power = 183451.9920076409
gearbox.compute()
# Design the Propeller with airfoil geometry defined
bad_prop = SUAVE.Components.Energy.Converters.Propeller()
bad_prop.tag = "Prop_W_Aifoil"
bad_prop.number_of_blades = 2
bad_prop.number_of_engines = 1
bad_prop.freestream_velocity = 1
bad_prop.tip_radius = 0.3
bad_prop.hub_radius = 0.21336
bad_prop.design_tip_mach = 0.1
bad_prop.angular_velocity = gearbox.inputs.speed
bad_prop.design_Cl = 0.7
bad_prop.design_altitude = 1. * Units.km
bad_prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
bad_prop.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
bad_prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
bad_prop.design_thrust = 100000
bad_prop = propeller_design(bad_prop)
prop_a = SUAVE.Components.Energy.Converters.Propeller()
prop_a.tag = "Prop_W_Aifoil"
prop_a.number_of_blades = 3
prop_a.number_of_engines = 1
prop_a.freestream_velocity = 49.1744
prop_a.tip_radius = 1.0668
prop_a.hub_radius = 0.21336
prop_a.design_tip_mach = 0.65
prop_a.angular_velocity = gearbox.inputs.speed # 207.16160479940007
prop_a.design_Cl = 0.7
prop_a.design_altitude = 1. * Units.km
prop_a.airfoil_geometry = [
'../Vehicles/Airfoils/NACA_4412.txt',
'../Vehicles/Airfoils/Clark_y.txt'
]
prop_a.airfoil_polars = [
[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
],
[
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/Clark_y_polar_Re_1000000.txt'
]
]
prop_a.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 1, 1, 1, 1, 1, 1, 1
]
prop_a.design_thrust = 3054.4809132125697
prop_a = propeller_design(prop_a)
# plot propeller
plot_propeller(prop_a)
# Design the Propeller with airfoil geometry defined
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.tag = "Prop_No_Aifoil"
prop.number_of_blades = 3
prop.number_of_engines = 1
prop.freestream_velocity = 49.1744
prop.tip_radius = 1.0668
prop.hub_radius = 0.21336
prop.design_tip_mach = 0.65
prop.angular_velocity = gearbox.inputs.speed
prop.design_Cl = 0.7
prop.design_altitude = 1. * Units.km
prop.origin = [[16. * 0.3048, 0., 2.02 * 0.3048]]
prop.design_power = gearbox.outputs.power
prop = propeller_design(prop)
# Design a Rotor with airfoil geometry defined
rot_a = SUAVE.Components.Energy.Converters.Rotor()
rot_a.tag = "Rot_W_Aifoil"
rot_a.tip_radius = 2.8 * Units.feet
rot_a.hub_radius = 0.35 * Units.feet
rot_a.number_of_blades = 2
rot_a.design_tip_mach = 0.65
rot_a.number_of_engines = 12
rot_a.disc_area = np.pi * (rot_a.tip_radius**2)
rot_a.freestream_velocity = 500. * Units['ft/min']
rot_a.angular_velocity = 258.9520059992501
rot_a.design_Cl = 0.7
rot_a.design_altitude = 20 * Units.feet
rot_a.design_thrust = 2271.2220451593753
rot_a.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
rot_a.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
rot_a.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
rot_a = propeller_design(rot_a)
# Design a Rotor without airfoil geometry defined
rot = SUAVE.Components.Energy.Converters.Rotor()
rot.tag = "Rot_No_Aifoil"
rot.tip_radius = 2.8 * Units.feet
rot.hub_radius = 0.35 * Units.feet
rot.number_of_blades = 2
rot.design_tip_mach = 0.65
rot.number_of_engines = 12
rot.disc_area = np.pi * (rot.tip_radius**2)
rot.freestream_velocity = 500. * Units['ft/min']
rot.angular_velocity = 258.9520059992501
rot.design_Cl = 0.7
rot.design_altitude = 20 * Units.feet
rot.design_thrust = 2271.2220451593753
rot = propeller_design(rot)
# Find the operating conditions
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(rot.design_altitude)
V = prop.freestream_velocity
Vr = rot.freestream_velocity
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
conditions._size = 1
conditions.freestream = Data()
conditions.propulsion = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.freestream.update(atmosphere_conditions)
conditions.freestream.dynamic_viscosity = atmosphere_conditions.dynamic_viscosity
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions.propulsion.throttle = np.array([[1.0]])
conditions.frames.body.transform_to_inertial = np.array([np.eye(3)])
conditions_r = copy.deepcopy(conditions)
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions_r.frames.inertial.velocity_vector = np.array([[0, Vr, 0]])
# Create and attach this propeller
prop_a.inputs.omega = np.array(prop.angular_velocity, ndmin=2)
prop.inputs.omega = np.array(prop.angular_velocity, ndmin=2)
rot_a.inputs.omega = copy.copy(prop.inputs.omega)
rot.inputs.omega = copy.copy(prop.inputs.omega)
# propeller with airfoil results
prop_a.inputs.pitch_command = 0.0 * Units.degree
F_a, Q_a, P_a, Cplast_a, output_a, etap_a = prop_a.spin(conditions)
plot_results(output_a, prop_a, 'blue', '-', 's')
# propeller without airfoil results
prop.inputs.pitch_command = 0.0 * Units.degree
F, Q, P, Cplast, output, etap = prop.spin(conditions)
plot_results(output, prop, 'red', '-', 'o')
# rotor with airfoil results
rot_a.inputs.pitch_command = 0.0 * Units.degree
Fr_a, Qr_a, Pr_a, Cplastr_a, outputr_a, etapr = rot_a.spin(conditions_r)
plot_results(outputr_a, rot_a, 'green', '-', '^')
# rotor with out airfoil results
rot.inputs.pitch_command = 0.0 * Units.degree
Fr, Qr, Pr, Cplastr, outputr, etapr = rot.spin(conditions_r)
plot_results(outputr, rot, 'black', '-', 'P')
# Truth values for propeller with airfoil geometry defined
F_a_truth = 3346.252387977257
Q_a_truth = 982.79237764
P_a_truth = 203596.8461368
Cplast_a_truth = 0.10493876
# Truth values for propeller without airfoil geometry defined
F_truth = 2629.0134410677906
Q_truth = 787.38493275
P_truth = 163115.92626346
Cplast_truth = 0.08407391
# Truth values for rotor with airfoil geometry defined
Fr_a_truth = 1500.9917429652628
Qr_a_truth = 141.39427382
Pr_a_truth = 29291.46467336
Cplastr_a_truth = 0.04607401
# Truth values for rotor without airfoil geometry defined
Fr_truth = 1250.1858330726345
Qr_truth = 121.95427719
Pr_truth = 25264.24377481
Cplastr_truth = 0.0397394
# Store errors
error = Data()
error.Thrust_a = np.max(np.abs(np.linalg.norm(F_a) - F_a_truth))
error.Torque_a = np.max(np.abs(Q_a - Q_a_truth))
error.Power_a = np.max(np.abs(P_a - P_a_truth))
error.Cp_a = np.max(np.abs(Cplast_a - Cplast_a_truth))
error.Thrust = np.max(np.abs(np.linalg.norm(F) - F_truth))
error.Torque = np.max(np.abs(Q - Q_truth))
error.Power = np.max(np.abs(P - P_truth))
error.Cp = np.max(np.abs(Cplast - Cplast_truth))
error.Thrustr_a = np.max(np.abs(np.linalg.norm(Fr_a) - Fr_a_truth))
error.Torquer_a = np.max(np.abs(Qr_a - Qr_a_truth))
error.Powerr_a = np.max(np.abs(Pr_a - Pr_a_truth))
error.Cpr_a = np.max(np.abs(Cplastr_a - Cplastr_a_truth))
error.Thrustr = np.max(np.abs(np.linalg.norm(Fr) - Fr_truth))
error.Torquer = np.max(np.abs(Qr - Qr_truth))
error.Powerr = np.max(np.abs(Pr - Pr_truth))
error.Cpr = np.max(np.abs(Cplastr - Cplastr_truth))
print('Errors:')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
return
def main():
# ------------------------------------------------------------------
# Propulsor
# ------------------------------------------------------------------
# build network
net = Solar()
net.number_of_engines = 1.
net.nacelle_dia = 0.2
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = 100 * Units.m
panel.efficiency = 0.18
panel.mass_properties.mass = panel.area*.600
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
# Propeller design specs
design_altitude = 0.0 * Units.km
Velocity = 10.0 # freestream m/s
RPM = 5887
Blades = 2.0
Radius = .4064
Hub_Radius = 0.05
Design_Cl = 0.7
Thrust = 0.0 #Specify either thrust or power to design for
Power = 7500. #Specify either thrust or power to design for
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = Blades
prop_attributes.freestream_velocity = Velocity
prop_attributes.angular_velocity = RPM*(2.*np.pi/60.0)
prop_attributes.tip_radius = Radius
prop_attributes.hub_radius = Hub_Radius
prop_attributes.design_Cl = Design_Cl
prop_attributes.design_altitude = design_altitude
prop_attributes.design_thrust = Thrust
prop_attributes.design_power = Power
prop_attributes = propeller_design(prop_attributes)
# Create and attach this propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.01
motor.no_load_current = 8.0
motor.speed_constant = 140.*(2.*np.pi/60.) # RPM/volt converted to rad/s
motor.propeller_radius = prop.prop_attributes.tip_radius
#motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1.
motor.expected_current = 260.
motor.mass_properties.mass = 2.0
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 0. #Watts
payload.mass_properties.mass = 0. * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 0. #Watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
batterymass = 50. #kg
bat.type = 'Li-Ion'
bat.resistance = 0.0
bat.energy_density = 250.
initialize_from_mass(bat,batterymass)
bat.current_energy = bat.max_energy
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 50.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# Setup the conditions to run the network
state = Data()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.numerics = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()
conditions = state.conditions
numerics = state.numerics
# Calculate atmospheric properties
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(prop_attributes.design_altitude)
rho = atmosphere_conditions.density[0,:]
a = atmosphere_conditions.speed_of_sound[0,:]
mu = atmosphere_conditions.dynamic_viscosity[0,:]
T = atmosphere_conditions.temperature[0,:]
conditions.propulsion.throttle = np.array([[1.0],[1.0]])
conditions.freestream.velocity = np.array([[1.0],[1.0]])
conditions.freestream.density = np.array([rho,rho])
conditions.freestream.dynamic_viscosity = np.array([mu, mu])
conditions.freestream.speed_of_sound = np.array([a, a])
conditions.freestream.altitude = np.array([[design_altitude], [design_altitude]])
conditions.propulsion.battery_energy = bat.max_energy*np.ones_like(conditions.freestream.altitude)
conditions.frames.body.inertial_rotations = np.zeros([2,3])
conditions.frames.inertial.time = np.array([[0.0],[1.0]])
numerics.time.integrate = np.array([[0, 0],[0, 1]])
numerics.time.differentiate = np.array([[0, 0],[0, 1]])
conditions.frames.planet.start_time = time.strptime("Sat, Jun 21 06:00:00 2014", "%a, %b %d %H:%M:%S %Y",)
conditions.frames.planet.latitude = np.array([[0.0],[0.0]])
conditions.frames.planet.longitude = np.array([[0.0],[0.0]])
conditions.freestream.temperature = np.array([T, T])
conditions.frames.body.transform_to_inertial = np.array([[[ 1., 0., 0.],
[ 0., 1., 0.],
[ 0., 0., 1.]],
[[ 1., 0., 0.],
[ 0., 1., 0.],
[ 0., 0., 1.]]])
conditions.propulsion.propeller_power_coefficient = np.array([[1.], [1.]]) * prop.prop_attributes.Cp
# Run the network and print the results
results = net(state)
F = results.thrust_force_vector
# Truth results
truth_F = [[ 545.35952329, 545.35952329]]
truth_i = [[ 249.31622624], [ 249.31622624]]
truth_rpm = [[ 6668.4094191], [ 6668.4094191]]
truth_bat = [[ 36000000. ], [ 35987534.18868808]]
error = Data()
error.Thrust = np.max(np.abs(F[:,0]-truth_F))
error.RPM = np.max(np.abs(conditions.propulsion.rpm-truth_rpm))
error.Current = np.max(np.abs(conditions.propulsion.current-truth_i))
error.Battery = np.max(np.abs(bat.current_energy-truth_bat))
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1e-6)
return
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'X57_Maxwell_Mod2'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
vehicle.mass_properties.takeoff = 2550. * Units.pounds
vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
vehicle.reference_area = 14.76
vehicle.passengers = 4
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "commuter"
cruise_speed = 135. * Units['mph']
altitude = 2500. * Units.ft
atmo = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmo.compute_values(0.)
freestream0 = atmo.compute_values(altitude)
mach_number = (cruise_speed / freestream.speed_of_sound)[0][0]
vehicle.design_dynamic_pressure = (.5 * freestream0.density *
(cruise_speed * cruise_speed))[0][0]
vehicle.design_mach_number = mach_number
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 14.76
wing.spans.projected = 11.4
wing.chords.root = 1.46
wing.chords.tip = 0.92
wing.chords.mean_aerodynamic = 1.19
wing.taper = wing.chords.root / wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 3.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[2.93, 0., 1.01]]
wing.aerodynamic_center = [3., 0., 1.01]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.winglet_fraction = 0.0
wing.dynamic_pressure_ratio = 1.0
airfoil = SUAVE.Components.Airfoils.Airfoil()
airfoil.coordinate_file = '../Vehicles/Airfoils/NACA_63_412.txt'
cg_x = wing.origin[0][0] + 0.25 * wing.chords.mean_aerodynamic
cg_z = wing.origin[0][2] - 0.2 * wing.chords.mean_aerodynamic
vehicle.mass_properties.center_of_gravity = [
[cg_x, 0., cg_z]
] # SOURCE: Design and aerodynamic analysis of a twin-engine commuter aircraft
# Wing Segments
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'inboard'
segment.percent_span_location = 0.0
segment.twist = 3. * Units.degrees
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.12
segment.append_airfoil(airfoil)
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'outboard'
segment.percent_span_location = 0.5438
segment.twist = 2. * Units.degrees
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.12
segment.append_airfoil(airfoil)
wing.append_segment(segment)
# Wing Segments
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'winglet'
segment.percent_span_location = 0.98
segment.twist = 1. * Units.degrees
segment.root_chord_percent = 0.630
segment.dihedral_outboard = 75. * Units.degrees
segment.sweeps.quarter_chord = 82. * Units.degrees
segment.thickness_to_chord = 0.12
segment.append_airfoil(airfoil)
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'tip'
segment.percent_span_location = 1.
segment.twist = 0. * Units.degrees
segment.root_chord_percent = 0.12
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.12
segment.append_airfoil(airfoil)
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = segment_properties(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 2.540
wing.spans.projected = 3.3 * Units.meter
wing.sweeps.quarter_chord = 0 * Units.deg
wing.chords.root = 0.769 * Units.meter
wing.chords.tip = 0.769 * Units.meter
wing.chords.mean_aerodynamic = 0.769 * Units.meter
wing.taper = 1.
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[7.7, 0., 0.25]]
wing.aerodynamic_center = [7.8, 0., 0.25]
wing.vertical = False
wing.winglet_fraction = 0.0
wing.symmetric = True
wing.high_lift = False
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.sweeps.quarter_chord = 25. * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 2.258 * Units['meters**2']
wing.spans.projected = 1.854 * Units.meter
wing.chords.root = 1.6764 * Units.meter
wing.chords.tip = 0.6858 * Units.meter
wing.chords.mean_aerodynamic = 1.21 * Units.meter
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[6.75, 0, 0.623]]
wing.aerodynamic_center = [0.508, 0, 0]
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.winglet_fraction = 0.0
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 2.
fuselage.fineness.nose = 1.6
fuselage.fineness.tail = 2.
fuselage.lengths.nose = 60. * Units.inches
fuselage.lengths.tail = 161. * Units.inches
fuselage.lengths.cabin = 105. * Units.inches
fuselage.lengths.total = 332.2 * Units.inches
fuselage.lengths.fore_space = 0.
fuselage.lengths.aft_space = 0.
fuselage.width = 42. * Units.inches
fuselage.heights.maximum = 62. * Units.inches
fuselage.heights.at_quarter_length = 62. * Units.inches
fuselage.heights.at_three_quarters_length = 62. * Units.inches
fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
fuselage.areas.side_projected = 8000. * Units.inches**2.
fuselage.areas.wetted = 30000. * Units.inches**2.
fuselage.areas.front_projected = 42. * 62. * Units.inches**2.
fuselage.effective_diameter = 50. * Units.inches
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0
segment.percent_z_location = 0
segment.height = 0.01
segment.width = 0.01
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.007279116466
segment.percent_z_location = 0.002502014453
segment.height = 0.1669064748
segment.width = 0.2780205877
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.01941097724
segment.percent_z_location = 0.001216095397
segment.height = 0.3129496403
segment.width = 0.4365777215
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.06308567604
segment.percent_z_location = 0.007395489231
segment.height = 0.5841726619
segment.width = 0.6735119903
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 0.1653761217
segment.percent_z_location = 0.02891281352
segment.height = 1.064028777
segment.width = 1.067200529
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 0.2426372155
segment.percent_z_location = 0.04214148761
segment.height = 1.293766653
segment.width = 1.183058255
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_6'
segment.percent_x_location = 0.2960174029
segment.percent_z_location = 0.04705241831
segment.height = 1.377026712
segment.width = 1.181540054
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_7'
segment.percent_x_location = 0.3809404284
segment.percent_z_location = 0.05313580461
segment.height = 1.439568345
segment.width = 1.178218989
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_8'
segment.percent_x_location = 0.5046854083
segment.percent_z_location = 0.04655492473
segment.height = 1.29352518
segment.width = 1.054390707
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_9'
segment.percent_x_location = 0.6454149933
segment.percent_z_location = 0.03741966266
segment.height = 0.8971223022
segment.width = 0.8501926505
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_10'
segment.percent_x_location = 0.985107095
segment.percent_z_location = 0.04540283436
segment.height = 0.2920863309
segment.width = 0.2012565415
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_11'
segment.percent_x_location = 1
segment.percent_z_location = 0.04787575562
segment.height = 0.1251798561
segment.width = 0.1206021048
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Nacelles
# ------------------------------------------------------------------
nacelle = SUAVE.Components.Nacelles.Nacelle()
nacelle.tag = 'nacelle_1'
nacelle.length = 2
nacelle.diameter = 42 * Units.inches
nacelle.areas.wetted = 0.01 * (2 * np.pi * 0.01 / 2)
nacelle.origin = [[2.5, 2.5, 1.0]]
nacelle.flow_through = False
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_1'
nac_segment.percent_x_location = 0.0
nac_segment.height = 0.0
nac_segment.width = 0.0
nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_2'
nac_segment.percent_x_location = 0.1
nac_segment.height = 0.5
nac_segment.width = 0.65
nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_3'
nac_segment.percent_x_location = 0.3
nac_segment.height = 0.52
nac_segment.width = 0.7
nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_4'
nac_segment.percent_x_location = 0.5
nac_segment.height = 0.5
nac_segment.width = 0.65
nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_5'
nac_segment.percent_x_location = 0.7
nac_segment.height = 0.4
nac_segment.width = 0.6
nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_6'
nac_segment.percent_x_location = 0.9
nac_segment.height = 0.3
nac_segment.width = 0.5
nacelle.append_segment(nac_segment)
nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
nac_segment.tag = 'segment_7'
nac_segment.percent_x_location = 1.0
nac_segment.height = 0.0
nac_segment.width = 0.0
nacelle.append_segment(nac_segment)
vehicle.append_component(nacelle)
nacelle_2 = deepcopy(nacelle)
nacelle_2.tag = 'nacelle_2'
nacelle_2.origin = [[2.5, -2.5, 1.0]]
vehicle.append_component(nacelle_2)
#---------------------------------------------------------------------------------------------
# DEFINE PROPELLER
#---------------------------------------------------------------------------------------------
# build network
net = Battery_Propeller()
net.number_of_propeller_engines = 2.
net.identical_propellers = True
# Component 1 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 2 the Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.tag = 'propeller_1'
prop.number_of_blades = 2.0
prop.freestream_velocity = 135. * Units['mph']
prop.angular_velocity = 1300. * Units.rpm
prop.tip_radius = 76. / 2. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_altitude = 12000. * Units.feet
prop.design_thrust = 1200.
prop.origin = [[2., 2.5, 0.784]]
prop.rotation = -1
prop.symmetry = True
prop.variable_pitch = True
prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
prop.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
prop = propeller_design(prop)
prop_left = deepcopy(prop)
prop_left.tag = 'propeller_2'
prop_left.origin = [[2., -2.5, 0.784]]
prop_left.rotation = 1
net.propellers.append(prop)
net.propellers.append(prop_left)
# Component 3 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiNiMnCoO2_18650(
)
bat.mass_properties.mass = 500. * Units.kg
bat.max_voltage = 500.
initialize_from_mass(bat)
net.battery = bat
net.voltage = bat.max_voltage
# Component 4 Miscellaneous Systems
sys = SUAVE.Components.Systems.System()
sys.mass_properties.mass = 5 # kg
# Component 5 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.efficiency = 0.95
motor.gearbox_efficiency = 1.
motor.origin = [[2., 2.5, 0.784]]
motor.nominal_voltage = bat.max_voltage * 3 / 4
motor.propeller_radius = prop.tip_radius
motor.no_load_current = 4.0
motor = size_optimal_motor(motor, prop)
motor.mass_properties.mass = 10. * Units.kg
# append right motor
net.propeller_motors.append(motor)
# append left motor
motor_left = deepcopy(motor)
motor_left.origin = [[2., -2.5, 0.784]]
net.propeller_motors.append(motor_left)
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 10. # Watts
payload.mass_properties.mass = 1.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 20. # Watts
net.avionics = avionics
# add the solar network to the vehicle
vehicle.append_component(net)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def configs_setup(vehicle):
#---------------------------------------------------------------------------
# Initialize Configurations
#---------------------------------------------------------------------------
configs = SUAVE.Components.Configs.Config.Container()
base_config = SUAVE.Components.Configs.Config(vehicle)
base_config.tag = 'base'
configs.append(base_config)
#---------------------------------------------------------------------------
# Electric Helicopter Configuration
#---------------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'electric_helicopter'
config.propulsors.network.number_of_engines = 1
prop_attributes = Data()
prop_attributes.number_blades = 4.0
prop_attributes.freestream_velocity = 150. * Units['meter/second']
prop_attributes.angular_velocity = 3500. * Units['rpm']
prop_attributes.tip_radius = 3800. * Units.mm
prop_attributes.hub_radius = 500. * Units.mm
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 1. * Units.km
prop_attributes.design_thrust = 1600. * 9.81 * Units.newtons
prop_attributes.design_power = 0. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.origin = [0., 0., 0.]
config.propulsors.network.propeller = prop
configs.append(config)
#---------------------------------------------------------------------------
# Electric Stopped Rotor Configuration
#---------------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'electric_stopped_rotor'
config.propulsors.network.number_of_engines = 8
prop_attributes = Data()
prop_attributes.number_blades = 4.0
prop_attributes.freestream_velocity = 150. * Units['meter/second']
prop_attributes.angular_velocity = 3500. * Units['rpm']
prop_attributes.tip_radius = 800. * Units.mm #608
prop_attributes.hub_radius = 150. * Units.mm
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 1. * Units.km
prop_attributes.design_thrust = 200. * 9.81 * Units.newtons
prop_attributes.design_power = 0. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.origin = [0., 0., 0.]
config.propulsors.network.propeller = prop
thrust_prop_attributes = cp.deepcopy(prop_attributes)
thrust_prop_attributes.number_blades = 2.0
thrust_prop = SUAVE.Components.Energy.Converters.Propeller()
thrust_prop.prop_attributes = thrust_prop_attributes
config.propulsors.network.thrust_propeller = thrust_prop
configs.append(config)
#---------------------------------------------------------------------------
# Electric Tiltrotor Configuration
#---------------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'electric_tiltrotor'
config.propulsors.network.number_of_engines = 8
prop_attributes = Data()
prop_attributes.number_blades = 4.0
prop_attributes.freestream_velocity = 150. * Units['meter/second']
prop_attributes.angular_velocity = 3500. * Units['rpm']
prop_attributes.tip_radius = 800. * Units.mm #608
prop_attributes.hub_radius = 150. * Units.mm
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 1. * Units.km
prop_attributes.design_thrust = 200. * 9.81 * Units.newtons
prop_attributes.design_power = 0. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.origin = [0., 0., 0.]
config.propulsors.network.propeller = prop
configs.append(config)
return configs
def main():
# ------------------------------------------------------------------
# Propulsor
# ------------------------------------------------------------------
# build network
net = Solar_Network()
net.number_motors = 1.
net.nacelle_dia = 0.2
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = 100 * Units.m
panel.efficiency = 0.18
panel.mass_properties.mass = panel.area * .600
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
#Propeller design specs
design_altitude = 0.0 * Units.km
Velocity = 10.0 # freestream m/s
RPM = 5887
Blades = 2.0
Radius = .4064
Hub_Radius = 0.05
Design_Cl = 0.7
Thrust = 0.0 #Specify either thrust or power to design for
Power = 7500. #Specify either thrust or power to design for
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = Blades
prop_attributes.freestream_velocity = Velocity
prop_attributes.angular_velocity = RPM * (2. * np.pi / 60.0)
prop_attributes.tip_radius = Radius
prop_attributes.hub_radius = Hub_Radius
prop_attributes.design_Cl = Design_Cl
prop_attributes.design_altitude = design_altitude
prop_attributes.design_thrust = Thrust
prop_attributes.design_power = Power
prop_attributes = propeller_design(prop_attributes)
# Create and attach this propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.01
motor.no_load_current = 8.0
motor.speed_constant = 140. * (2. * np.pi / 60.
) # RPM/volt converted to rad/s
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1.
motor.expected_current = 260.
motor.mass_properties.mass = 2.0
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 0. #Watts
payload.mass_properties.mass = 0. * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 0. #Watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Battery()
bat.mass_properties.mass = 50. #kg
bat.type = 'Li-Ion'
bat.resistance = 0.0
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 50.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# Setup the conditions to run the network
conditions = Data()
conditions.propulsion = Data()
conditions.freestream = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.frames.planet = Data()
numerics = Data()
# Calculate atmospheric properties
atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(design_altitude)
conditions.propulsion.throttle = np.array([[1.0], [1.0]])
conditions.freestream.velocity = np.array([[1.0], [1.0]])
conditions.freestream.density = np.array([rho, rho])
conditions.freestream.viscosity = np.array([mu, mu])
conditions.freestream.speed_of_sound = np.array([a, a])
conditions.freestream.altitude = np.array([[design_altitude],
[design_altitude]])
conditions.propulsion.battery_energy = bat.max_energy() * np.ones_like(
conditions.freestream.altitude)
conditions.frames.body.inertial_rotations = np.zeros([2, 3])
conditions.frames.inertial.time = np.array([[0.0], [1.0]])
numerics.integrate_time = np.array([[0, 0], [0, 1]])
conditions.frames.planet.start_time = time.strptime(
"Sat, Jun 21 06:00:00 2014",
"%a, %b %d %H:%M:%S %Y",
)
conditions.frames.planet.latitude = np.array([[0.0], [0.0]])
conditions.frames.planet.longitude = np.array([[0.0], [0.0]])
conditions.freestream.temperature = np.array([T, T])
# Run the network and print the results
F, mdot, P = net(conditions, numerics)
# Truth results
truth_F = [[538.00449442], [538.00449442]]
truth_P = [[14272.1902522], [14272.1902522]]
truth_i = [[249.31622624], [249.31622624]]
truth_rpm = [[6668.4094191], [6668.4094191]]
truth_bat = [[45000000.], [44987534.18868808]]
error = Data()
error.Thrust = np.max(np.abs(F - truth_F))
error.Propeller_Power = np.max(np.abs(P - truth_P))
error.RPM = np.max(np.abs(conditions.propulsion.rpm - truth_rpm))
error.Current = np.max(np.abs(conditions.propulsion.current - truth_i))
error.Battery = np.max(np.abs(bat.CurrentEnergy - truth_bat))
print error
for k, v in error.items():
assert (np.abs(v) < 0.001)
return
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Lift_Cruise_CRM'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2450. * Units.lb
vehicle.mass_properties.operating_empty = 2250. * Units.lb # Approximate
vehicle.mass_properties.max_takeoff = 2450. * Units.lb # Approximate
vehicle.mass_properties.center_of_gravity = [2.0144, 0. , 0. ] # Approximate
# basic parameters
vehicle.reference_area = 10.76
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
# ------------------------------------------------------------------
# WINGS
# ------------------------------------------------------------------
# WING PROPERTIES
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 10.76
wing.sweeps.quarter_chord = 0.0 * Units.degrees
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.span_efficiency = 0.9
wing.spans.projected = 35.0 * Units.feet
wing.chords.root = 3.25 * Units.feet
wing.total_length = 3.25 * Units.feet
wing.chords.tip = 3.25 * Units.feet
wing.chords.mean_aerodynamic = 3.25 * Units.feet
wing.dihedral = 1.0 * Units.degrees
wing.areas.reference = 113.75 * Units.feet**2
wing.areas.wetted = 227.5 * Units.feet**2
wing.areas.exposed = 227.5 * Units.feet**2
wing.twists.root = 4.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [1.5, 0., 0. ]
wing.aerodynamic_center = [1.975 , 0., 0.]
wing.winglet_fraction = 0.0
wing.symmetric = True
wing.vertical = False
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_1'
segment.percent_span_location = 0.
segment.twist = 0.
segment.root_chord_percent = 1.5
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 8.5 * Units.degrees
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_2'
segment.percent_span_location = 0.227
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0 * Units.degrees
segment.thickness_to_chord = 0.12
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_3'
segment.percent_span_location = 1.0
segment.twist = 0.
segment.root_chord_percent = 1.0
segment.dihedral_outboard = 1.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0 * Units.degrees
segment.thickness_to_chord = 0.12
wing.Segments.append(segment)
# add to vehicle
vehicle.append_component(wing)
# WING PROPERTIES
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_tail'
wing.aspect_ratio = 4.0
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.9
wing.spans.projected = 8.0 * Units.feet
wing.chords.root = 2.0 * Units.feet
wing.total_length = 2.0 * Units.feet
wing.chords.tip = 2.0 * Units.feet
wing.chords.mean_aerodynamic = 2.0 * Units.feet
wing.dihedral = 0. * Units.degrees
wing.areas.reference = 16.0 * Units.feet**2
wing.areas.wetted = 32.0 * Units.feet**2
wing.areas.exposed = 32.0 * Units.feet**2
wing.twists.root = 0. * Units.degrees
wing.twists.tip = 0. * Units.degrees
wing.origin = [14.0*0.3048 , 0.0 , 0.205 ]
wing.aerodynamic_center = [15.0*0.3048 , 0., 0.]
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
# WING PROPERTIES
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_tail_1'
wing.aspect_ratio = 2.
wing.sweeps.quarter_chord = 20.0 * Units.degrees
wing.thickness_to_chord = 0.12
wing.taper = 0.5
wing.span_efficiency = 0.9
wing.spans.projected = 3.0 * Units.feet
wing.chords.root = 2.0 * Units.feet
wing.total_length = 2.0 * Units.feet
wing.chords.tip = 1.0 * Units.feet
wing.chords.mean_aerodynamic = 1.5 * Units.feet
wing.areas.reference = 4.5 * Units.feet**2
wing.areas.wetted = 9.0 * Units.feet**2
wing.areas.exposed = 9.0 * Units.feet**2
wing.twists.root = 0. * Units.degrees
wing.twists.tip = 0. * Units.degrees
wing.origin = [14.0*0.3048 , 4.0*0.3048 , 0.205 ]
wing.aerodynamic_center = 0.0
wing.winglet_fraction = 0.0
wing.vertical = True
wing.symmetric = False
# add to vehicle
vehicle.append_component(wing)
# WING PROPERTIES
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_tail_2'
wing.aspect_ratio = 2.
wing.sweeps.quarter_chord = 20.0 * Units.degrees
wing.thickness_to_chord = 0.12
wing.taper = 0.5
wing.span_efficiency = 0.9
wing.spans.projected = 3.0 * Units.feet
wing.chords.root = 2.0 * Units.feet
wing.total_length = 2.0 * Units.feet
wing.chords.tip = 1.0 * Units.feet
wing.chords.mean_aerodynamic = 1.5 * Units.feet
wing.areas.reference = 4.5 * Units.feet**2
wing.areas.wetted = 9.0 * Units.feet**2
wing.areas.exposed = 9.0 * Units.feet**2
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [14.0*0.3048 , -4.0*0.3048 , 0.205 ]
wing.aerodynamic_center = 0.0
wing.winglet_fraction = 0.0
wing.vertical = True
wing.symmetric = False
# add to vehicle
vehicle.append_component(wing)
# ---------------------------------------------------------------
# FUSELAGE
# ---------------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.configuration = 'Tube_Wing'
fuselage.origin = [[0. , 0., 0.]]
fuselage.seats_abreast = 2.
fuselage.seat_pitch = 3.
fuselage.fineness.nose = 0.88
fuselage.fineness.tail = 1.13
fuselage.lengths.nose = 3.2 * Units.feet
fuselage.lengths.tail = 6.4 * Units.feet
fuselage.lengths.cabin = 6.4 * Units.feet
fuselage.lengths.total = 16.0 * Units.feet
fuselage.width = 5.85 * Units.feet
fuselage.heights.maximum = 4.65 * Units.feet
fuselage.heights.at_quarter_length = 3.75 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 4.65 * Units.feet
fuselage.heights.at_three_quarters_length = 4.26 * Units.feet
fuselage.areas.wetted = 236. * Units.feet**2
fuselage.areas.front_projected = 0.14 * Units.feet**2
fuselage.effective_diameter = 5.85 * Units.feet
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_1'
segment.origin = [0., 0. ,0.]
segment.percent_x_location = 0.
segment.percent_z_location = 0.0
segment.height = 0.1 * Units.feet
segment.width = 0.1 * Units.feet
segment.length = 0.
segment.effective_diameter = 0.1 * Units.feet
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_2'
segment.origin = [4.*0.3048 , 0. ,0.1*0.3048 ]
segment.percent_x_location = 0.25
segment.percent_z_location = 0.05
segment.height = 3.75 * Units.feet
segment.width = 5.65 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 5.65 * Units.feet
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_3'
segment.origin = [8.*0.3048 , 0. ,0.34*0.3048 ]
segment.percent_x_location = 0.5
segment.percent_z_location = 0.071
segment.height = 4.65 * Units.feet
segment.width = 5.55 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 5.55 * Units.feet
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_4'
segment.origin = [12.*0.3048 , 0. ,0.77*0.3048 ]
segment.percent_x_location = 0.75
segment.percent_z_location = 0.089
segment.height = 4.73 * Units.feet
segment.width = 4.26 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 4.26 * Units.feet
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_5'
segment.origin = [16.*0.3048 , 0. ,2.02*0.3048 ]
segment.percent_x_location = 1.0
segment.percent_z_location = 0.158
segment.height = 0.67 * Units.feet
segment.width = 0.33 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 0.33 * Units.feet
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#-------------------------------------------------------------------
# BOOMS
#-------------------------------------------------------------------
boom = SUAVE.Components.Fuselages.Fuselage()
boom.tag = 'Boom_1R'
boom.configuration = 'Boom'
boom.origin = [[0.718,7.5*0.3048 , -0.15 ]]
boom.seats_abreast = 0.
boom.seat_pitch = 0.0
boom.fineness.nose = 0.950
boom.fineness.tail = 1.029
boom.lengths.nose = 0.5 * Units.feet
boom.lengths.tail = 0.5 * Units.feet
boom.lengths.cabin = 9.0 * Units.feet
boom.lengths.total = 10 * Units.feet
boom.width = 0.5 * Units.feet
boom.heights.maximum = 0.5 * Units.feet
boom.heights.at_quarter_length = 0.5 * Units.feet
boom.heights.at_three_quarters_length = 0.5 * Units.feet
boom.heights.at_wing_root_quarter_chord = 0.5 * Units.feet
boom.areas.wetted = 18 * Units.feet**2
boom.areas.front_projected = 0.26 * Units.feet**2
boom.effective_diameter = 0.5 * Units.feet
boom.differential_pressure = 0.
boom.y_pitch_count = 2
boom.y_pitch = (72/12)*0.3048
boom.symmetric = True
boom.boom_pitch = 6 * Units.feet
boom.index = 1
# add to vehicle
vehicle.append_component(boom)
# create pattern of booms on one side
original_boom_origin = boom.origin
if boom.y_pitch_count > 1 :
for n in range(boom.y_pitch_count):
if n == 0:
continue
else:
index = n+1
boom = deepcopy(vehicle.fuselages.boom_1r)
boom.origin[0][1] = n*boom.boom_pitch + original_boom_origin[0][1]
boom.tag = 'Boom_' + str(index) + 'R'
boom.index = n
vehicle.append_component(boom)
if boom.symmetric :
for n in range(boom.y_pitch_count):
index = n+1
boom = deepcopy(vehicle.fuselages.boom_1r)
boom.origin[0][1] = -n*boom.boom_pitch - original_boom_origin[0][1]
boom.tag = 'Boom_' + str(index) + 'L'
boom.index = n
vehicle.append_component(boom)
#------------------------------------------------------------------
# PROPULSOR
#------------------------------------------------------------------
net = Lift_Cruise()
net.number_of_engines_lift = 12
net.number_of_engines_forward = 1
net.thrust_angle_lift = 90. * Units.degrees
net.thrust_angle_forward = 0.
net.nacelle_diameter = 0.6 * Units.feet
net.engine_length = 0.5 * Units.feet
net.areas = Data()
net.areas.wetted = np.pi*net.nacelle_diameter*net.engine_length + 0.5*np.pi*net.nacelle_diameter**2
net.voltage = 500.
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
esc_lift = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc_lift.efficiency = 0.95
net.esc_lift = esc_lift
esc_thrust = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc_thrust.efficiency = 0.95
net.esc_forward = esc_thrust
#------------------------------------------------------------------
# Design Payload
#------------------------------------------------------------------
payload = SUAVE.Components.Energy.Peripherals.Avionics()
payload.power_draw = 0.
payload.mass_properties.mass = 200. * Units.kg
net.payload = payload
#------------------------------------------------------------------
# Design Avionics
#------------------------------------------------------------------
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 200. * Units.watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
bat.specific_energy = 300. * Units.Wh/Units.kg
bat.resistance = 0.005
bat.max_voltage = net.voltage
bat.mass_properties.mass = 300. * Units.kg
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
#------------------------------------------------------------------
# Design Rotors and Propellers
#------------------------------------------------------------------
# atmosphere and flight conditions for propeller/rotor design
g = 9.81 # gravitational acceleration
S = vehicle.reference_area # reference area
speed_of_sound = 340 # speed of sound
rho = 1.22 # reference density
fligth_CL = 0.75 # cruise target lift coefficient
AR = vehicle.wings.main_wing.aspect_ratio # aspect ratio
Cd0 = 0.06 # profile drag
Cdi = fligth_CL**2/(np.pi*AR*0.98) # induced drag
Cd = Cd0 + Cdi # total drag
V_inf = 110.* Units['mph'] # freestream velocity
Drag = S * (0.5*rho*V_inf**2 )*Cd # cruise drag
Hover_Load = vehicle.mass_properties.takeoff*g # hover load
# Thrust Propeller
propeller = SUAVE.Components.Energy.Converters.Propeller()
propeller.number_blades = 3
propeller.number_of_engines = net.number_of_engines_forward
propeller.freestream_velocity = V_inf
propeller.tip_radius = 1.0668
propeller.hub_radius = 0.21336
propeller.design_tip_mach = 0.65
propeller.angular_velocity = propeller.design_tip_mach *speed_of_sound /propeller.tip_radius
propeller.design_Cl = 0.7
propeller.design_altitude = 1. * Units.km
propeller.design_thrust = (Drag*2.5)/net.number_of_engines_forward
propeller = propeller_design(propeller)
propeller.origin = [[16.*0.3048 , 0. ,2.02*0.3048 ]]
net.propeller = propeller
# Lift Rotors
rotor = SUAVE.Components.Energy.Converters.Rotor()
rotor.tip_radius = 2.8 * Units.feet
rotor.hub_radius = 0.35 * Units.feet
rotor.number_blades = 2
rotor.design_tip_mach = 0.65
rotor.number_of_engines = net.number_of_engines_lift
rotor.disc_area = np.pi*(rotor.tip_radius**2)
rotor.induced_hover_velocity = np.sqrt(Hover_Load/(2*rho*rotor.disc_area*net.number_of_engines_lift))
rotor.freestream_velocity = 500. * Units['ft/min']
rotor.angular_velocity = rotor.design_tip_mach* speed_of_sound /rotor.tip_radius
rotor.design_Cl = 0.7
rotor.design_altitude = 20 * Units.feet
rotor.design_thrust = (Hover_Load * 2.5 )/net.number_of_engines_lift
rotor.x_pitch_count = 2
rotor.y_pitch_count = vehicle.fuselages['boom_1r'].y_pitch_count
rotor.y_pitch = vehicle.fuselages['boom_1r'].y_pitch
rotor = propeller_design(rotor)
rotor.origin = vehicle.fuselages['boom_1r'].origin
rotor.symmetric = True
# populating propellers on one side of wing
if rotor.y_pitch_count > 1 :
for n in range(rotor.y_pitch_count):
if n == 0:
continue
propeller_origin = [rotor.origin[0][0] , rotor.origin[0][1] + n*rotor.y_pitch ,rotor.origin[0][2]]
rotor.origin.append(propeller_origin)
# populating propellers on one side of the vehicle
if rotor.x_pitch_count > 1 :
relative_prop_origins = np.linspace(0,vehicle.fuselages['boom_1r'].lengths.total,rotor.x_pitch_count)
for n in range(len(rotor.origin)):
for m in range(len(relative_prop_origins)-1):
propeller_origin = [rotor.origin[n][0] + relative_prop_origins[m+1] , rotor.origin[n][1] ,rotor.origin[n][2] ]
rotor.origin.append(propeller_origin)
# propulating propellers on the other side of thevehicle
if rotor.symmetric :
for n in range(len(rotor.origin)):
propeller_origin = [rotor.origin[n][0] , -rotor.origin[n][1] ,rotor.origin[n][2] ]
rotor.origin.append(propeller_origin)
# re-compute number of lift rotors if changed
net.number_of_engines_lift = len(rotor.origin)
# append propellers to vehicle
net.rotor = rotor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
motor_forward = SUAVE.Components.Energy.Converters.Motor()
motor_forward.efficiency = 0.95
motor_forward.nominal_voltage = bat.max_voltage *3/4
motor_forward.mass_properties.mass = 2.0 * Units.kg
motor_forward.origin = propeller.origin
motor_forward.propeller_radius = propeller.tip_radius
motor_forward.gear_ratio = 1.
motor_forward.gearbox_efficiency = 1. # Gear box efficiency
motor_forward.no_load_current = 2.0
motor_forward = compute_optimal_motor_parameters(motor_forward,propeller)
net.motor_forward = motor_forward
# Rotor (Lift) Motor
motor_lift = SUAVE.Components.Energy.Converters.Motor()
motor_lift.efficiency = 0.95
motor_lift.nominal_voltage = bat.max_voltage
motor_lift.mass_properties.mass = 3. * Units.kg
motor_lift.origin = rotor.origin
motor_lift.propeller_radius = rotor.tip_radius
motor_lift.gear_ratio = 1.0
motor_lift.gearbox_efficiency = 1.0
motor_lift.no_load_current = 4.0
motor_lift = compute_optimal_motor_parameters(motor_lift,rotor)
net.motor_lift = motor_lift
# append motor origin spanwise locations onto wing data structure
vehicle.append_component(net)
# Add extra drag sources from motors, props, and landing gear. All of these hand measured
motor_height = .25 * Units.feet
motor_width = 1.6 * Units.feet
propeller_width = 1. * Units.inches
propeller_height = propeller_width *.12
main_gear_width = 1.5 * Units.inches
main_gear_length = 2.5 * Units.feet
nose_gear_width = 2. * Units.inches
nose_gear_length = 2. * Units.feet
nose_tire_height = (0.7 + 0.4) * Units.feet
nose_tire_width = 0.4 * Units.feet
main_tire_height = (0.75 + 0.5) * Units.feet
main_tire_width = 4. * Units.inches
total_excrescence_area_spin = 12.*motor_height*motor_width + 2.*main_gear_length*main_gear_width \
+ nose_gear_width*nose_gear_length + 2*main_tire_height*main_tire_width\
+ nose_tire_height*nose_tire_width
total_excrescence_area_no_spin = total_excrescence_area_spin + 12*propeller_height*propeller_width
vehicle.excrescence_area_no_spin = total_excrescence_area_no_spin
vehicle.excrescence_area_spin = total_excrescence_area_spin
vehicle.wings['main_wing'].motor_spanwise_locations = np.multiply(
2./36.25,
[-5.435, -5.435, -9.891, -9.891, -14.157, -14.157,
5.435, 5.435, 9.891, 9.891, 14.157, 14.157])
vehicle.wings['main_wing'].winglet_fraction = 0.0
vehicle.wings['main_wing'].thickness_to_chord = 0.18
vehicle.wings['main_wing'].chords.mean_aerodynamic = 0.9644599977664836
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Tiltwing'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2250. * Units.lb
vehicle.mass_properties.operating_empty = 2250. * Units.lb
vehicle.mass_properties.max_takeoff = 2250. * Units.lb
vehicle.mass_properties.center_of_gravity = [[ 2.0144, 0. , 0.]]
vehicle.passengers = 1
vehicle.reference_area = 10.58275476
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
# ------------------------------------------------------
# WINGS
# ------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'canard_wing'
wing.aspect_ratio = 11.37706641
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.spans.projected = 6.65
wing.chords.root = 0.95
wing.total_length = 0.95
wing.chords.tip = 0.95
wing.chords.mean_aerodynamic = 0.95
wing.dihedral = 0.0
wing.areas.reference = 6.31
wing.areas.wetted = 12.635
wing.areas.exposed = 12.635
wing.twists.root = 0.
wing.twists.tip = 0.
wing.origin = [[0.1, 0.0 , 0.0]]
wing.aerodynamic_center = [0.3, 0.0 , 0.0]
wing.winglet_fraction = 0.0
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 11.37706641
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.spans.projected = 6.65
wing.chords.root = 0.95
wing.total_length = 0.95
wing.chords.tip = 0.95
wing.chords.mean_aerodynamic = 0.95
wing.dihedral = 0.0
wing.areas.reference = 6.31
wing.areas.wetted = 12.635
wing.areas.exposed = 12.635
wing.twists.root = 0.
wing.twists.tip = 0.
wing.origin = [[ 5.138, 0.0 , 1.323 ]] # for images 1.54
wing.aerodynamic_center = [ 5.3, 0.0 , 1.323 ]
wing.winglet_fraction = 0.0
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------
# FUSELAGE
# ------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 0.
fuselage.seat_pitch = 1.
fuselage.fineness.nose = 1.5
fuselage.fineness.tail = 4.0
fuselage.lengths.nose = 1.7
fuselage.lengths.tail = 2.7
fuselage.lengths.cabin = 1.7
fuselage.lengths.total = 6.1
fuselage.width = 1.15
fuselage.heights.maximum = 1.7
fuselage.heights.at_quarter_length = 1.2
fuselage.heights.at_wing_root_quarter_chord = 1.7
fuselage.heights.at_three_quarters_length = 0.75
fuselage.areas.wetted = 12.97989862
fuselage.areas.front_projected = 1.365211404
fuselage.effective_diameter = 1.318423736
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.
segment.percent_z_location = 0.
segment.height = 0.09
segment.width = 0.23473
segment.length = 0.
segment.effective_diameter = 0.
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.97675/6.1
segment.percent_z_location = 0.21977/6.1
segment.height = 0.9027
segment.width = 1.01709
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 1.93556/6.1
segment.percent_z_location = 0.39371/6.1
segment.height = 1.30558
segment.width = 1.38871
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 3.44137/6.1
segment.percent_z_location = 0.57143/6.1
segment.height = 1.52588
segment.width = 1.47074
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 4.61031/6.1
segment.percent_z_location = 0.81577/6.1
segment.height = 1.14788
segment.width = 1.11463
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 1.
segment.percent_z_location = 1.19622/6.1
segment.height = 0.31818
segment.width = 0.23443
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#------------------------------------------------------------------
# network
#------------------------------------------------------------------
net = SUAVE.Components.Energy.Networks.Battery_Propeller()
net.number_of_propeller_engines = 8
net.thrust_angle = 0.0 * Units.degrees # conversion to radians,
net.nacelle_diameter = 0.2921 # https://www.magicall.biz/products/integrated-motor-controller-magidrive/
net.engine_length = 0.95
net.areas = Data()
net.areas.wetted = np.pi*net.nacelle_diameter*net.engine_length + 0.5*np.pi*net.nacelle_diameter**2
net.voltage = 400.
net.identical_propellers = True
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95
net.esc = esc
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 10. #Watts
payload.mass_properties.mass = 0.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 20. #Watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiNiMnCoO2_18650()
bat.mass_properties.mass = 200. * Units.kg
bat.max_voltage = net.voltage
initialize_from_mass(bat)
# Here we, are going to assume a battery pack module shape. This step is optional but
# required for thermal analysis of tge pack
number_of_modules = 10
bat.module_config.total = int(np.ceil(bat.pack_config.total/number_of_modules))
bat.module_config.normal_count = int(np.ceil(bat.module_config.total/bat.pack_config.series))
bat.module_config.parallel_count = int(np.ceil(bat.module_config.total/bat.pack_config.parallel))
net.battery = bat
# Component 9 Miscellaneous Systems
sys = SUAVE.Components.Systems.System()
sys.mass_properties.mass = 5 # kg
#------------------------------------------------------------------
# Nacelles
#------------------------------------------------------------------
nacelle = SUAVE.Components.Nacelles.Nacelle()
nacelle.diameter = 0.2921
nacelle.length = 0.95
nacelle_origins = [[-0.2, 1.347, 0.0], [-0.2, 3.2969999999999997, 0.0],
[-0.2, -1.347, 0.0], [-0.2, -3.2969999999999997, 0.0],
[4.938, 1.347, 1.54], [4.938, 3.2969999999999997, 1.54],
[4.938, -1.347, 1.54], [4.938, -3.2969999999999997, 1.54]]
nacelle.areas.wetted = np.pi*nacelle.diameter*nacelle.length + 0.5*np.pi*nacelle.diameter**2
for idx in range(8):
nacelle = deepcopy(nacelle)
nacelle.tag = 'nacelle_' + str(idx)
nacelle.origin = [nacelle_origins[idx]]
vehicle.append_component(nacelle)
#------------------------------------------------------------------
# Design Rotors
#------------------------------------------------------------------
# atmosphere conditions
speed_of_sound = 340
# Create propeller geometry
prop = SUAVE.Components.Energy.Converters.Propeller() # This is truly a prop because the default of the mission is pointing forward
prop.tip_radius = 0.8875
prop.hub_radius = 0.15
prop.disc_area = np.pi*(prop.tip_radius**2)
prop.design_tip_mach = 0.5
prop.number_of_blades = 3
prop.freestream_velocity = 10
prop.angular_velocity = prop.design_tip_mach*speed_of_sound/prop.tip_radius
prop.design_Cl = 0.7
prop.design_altitude = 500 * Units.feet
Hover_Load = vehicle.mass_properties.takeoff*9.81
prop.design_thrust = Hover_Load/(net.number_of_propeller_engines-1) # contingency for one-engine-inoperative condition
prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
prop.airfoil_polars = [['../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt' ,
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt' ,
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt' ,
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt' ,
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt' ]]
prop.airfoil_polar_stations = [0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0]
prop = propeller_design(prop)
prop.rotation = 1
# Front Rotors Locations
origins = [[-0.2, 1.347, 0.0], [-0.2, 3.2969999999999997, 0.0], [-0.2, -1.347, 0.0], [-0.2, -3.2969999999999997, 0.0],\
[4.938, 1.347, 1.54], [4.938, 3.2969999999999997, 1.54], [4.938, -1.347, 1.54], [4.938, -3.2969999999999997, 1.54]]
for ii in range(8):
rotor = deepcopy(prop)
rotor.tag = 'propeller'
rotor.origin = [origins[ii]]
net.propellers.append(rotor)
# Motor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.efficiency = 0.9
motor.nominal_voltage = bat.max_voltage *3/4
motor.propeller_radius = prop.tip_radius
motor.no_load_current = 2.0
motor = size_optimal_motor(motor,prop)
motor.mass_properties.mass = nasa_motor(motor.design_torque)
net.motor = motor
for ii in range(8):
rotor_motor = deepcopy(motor)
rotor_motor.tag = 'motor'
rotor_motor.origin = [origins[ii]]
net.propeller_motors.append(rotor_motor)
# Add extra drag sources from motors, props, and landing gear. All of these hand measured
motor_height = .25 * Units.feet
motor_width = 1.6 * Units.feet
propeller_width = 1. * Units.inches
propeller_height = propeller_width *.12
main_gear_width = 1.5 * Units.inches
main_gear_length = 2.5 * Units.feet
nose_gear_width = 2. * Units.inches
nose_gear_length = 2. * Units.feet
nose_tire_height = (0.7 + 0.4) * Units.feet
nose_tire_width = 0.4 * Units.feet
main_tire_height = (0.75 + 0.5) * Units.feet
main_tire_width = 4. * Units.inches
total_excrescence_area_spin = 12.*motor_height*motor_width + 2.* main_gear_length*main_gear_width \
+ nose_gear_width*nose_gear_length + 2 * main_tire_height*main_tire_width\
+ nose_tire_height*nose_tire_width
total_excrescence_area_no_spin = total_excrescence_area_spin + 12*propeller_height*propeller_width
vehicle.excrescence_area_no_spin = total_excrescence_area_no_spin
vehicle.excrescence_area_spin = total_excrescence_area_spin
# append motor origin spanwise locations onto wing data structure
motor_origins_front = np.array(origins[:4])
motor_origins_rear = np.array(origins[5:])
vehicle.wings['canard_wing'].motor_spanwise_locations = motor_origins_front[:,1]/ vehicle.wings['canard_wing'].spans.projected
vehicle.wings['canard_wing'].motor_spanwise_locations = motor_origins_front[:,1]/ vehicle.wings['canard_wing'].spans.projected
vehicle.wings['main_wing'].motor_spanwise_locations = motor_origins_rear[:,1]/ vehicle.wings['main_wing'].spans.projected
vehicle.append_component(net)
vehicle.weight_breakdown = empty(vehicle)
compute_component_centers_of_gravity(vehicle)
vehicle.center_of_gravity()
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Solar'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 200. * Units.kg
vehicle.mass_properties.operating_empty = 200. * Units.kg
vehicle.mass_properties.max_takeoff = 200. * Units.kg
# basic parameters
vehicle.reference_area = 80.
vehicle.envelope.ultimate_load = 2.0
vehicle.envelope.limit_load = 1.5
vehicle.envelope.maximum_dynamic_pressure = 0.5 * 1.225 * (40.**2.) #Max q
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = vehicle.reference_area
wing.spans.projected = 40.0 * Units.meter
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.symmetric = True
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.vertical = False
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.areas.reference / wing.spans.projected
wing.span_efficiency = 0.98
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.highlift = False
wing.vertical = False
wing.number_ribs = 26.
wing.number_end_ribs = 2.
wing.transition_x_upper = 0.6
wing.transition_x_lower = 1.0
wing.origin = [3.0, 0.0, 0.0] # meters
wing.aerodynamic_center = [3.0, 0.0, 0.0] # meters
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 20.
wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.95
wing.areas.reference = vehicle.reference_area * .15
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
wing.number_ribs = 5.0
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.areas.reference / wing.spans.projected
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected
wing.origin = [10., 0.0, 0.0] # meters
wing.aerodynamic_center = [0.5, 0.0, 0.0] # meters
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 20.
wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.97
wing.areas.reference = vehicle.reference_area * 0.1
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.areas.reference / wing.spans.projected
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [10., 0.0, 0.0] # meters
wing.aerodynamic_center = [0.5, 0.0, 0.0] # meters
wing.symmetric = True
wing.vertical = True
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------
# Propulsor
#------------------------------------------------------------------
# build network
net = Solar()
net.number_of_engines = 1.
net.nacelle_diameter = 0.2 * Units.meters
net.engine_length = 0.01 * Units.meters
net.areas = Data()
net.areas.wetted = 0.01 * (2 * np.pi * 0.01 / 2.)
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = vehicle.reference_area * 0.9
panel.efficiency = 0.25
panel.mass_properties.mass = panel.area * (0.60 * Units.kg)
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 40.0 * Units['m/s'] # freestream
prop_attributes.angular_velocity = 150. * Units['rpm']
prop_attributes.tip_radius = 4.25 * Units.meters
prop_attributes.hub_radius = 0.05 * Units.meters
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 14.0 * Units.km
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 3500.0 * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.008
motor.no_load_current = 4.5 * Units.ampere
motor.speed_constant = 120. * Units[
'rpm'] # RPM/volt converted to (rad/s)/volt
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 12. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 160. # Expected current
motor.mass_properties.mass = 2.0 * Units.kg
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 50. * Units.watts
payload.mass_properties.mass = 5.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 50. * Units.watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.mass_properties.mass = 55.0 * Units.kg
bat.specific_energy = 450. * Units.Wh / Units.kg
bat.resistance = 0.05
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 40.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# add the solar network to the vehicle
vehicle.append_component(net)
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'X57_Maxwell'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
vehicle.mass_properties.takeoff = 2550. * Units.pounds
vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
vehicle.mass_properties.cargo = 0.
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
# basic parameters
vehicle.reference_area = 15.45
vehicle.passengers = 4
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.sweeps.leading_edge = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 15.45 * Units['meters**2']
wing.spans.projected = 11. * Units.meter
wing.chords.root = 1.67 * Units.meter
wing.chords.tip = 1.14 * Units.meter
wing.chords.mean_aerodynamic = 1.47 * Units.meter
wing.taper = wing.chords.root / wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 3.0 * Units.degrees
wing.twists.tip = 1.5 * Units.degrees
wing.origin = [[2.032, 0., 0.784]]
wing.aerodynamic_center = [0.558, 0., 0.784]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 3.74 * Units['meters**2']
wing.spans.projected = 3.454 * Units.meter
wing.sweeps.quarter_chord = 12.5 * Units.deg
wing.chords.root = 1.397 * Units.meter
wing.chords.tip = 0.762 * Units.meter
wing.chords.mean_aerodynamic = 1.09 * Units.meter
wing.taper = wing.chords.root / wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[6.248, 0., 0]]
wing.aerodynamic_center = [0.508, 0., 0.]
wing.vertical = False
wing.symmetric = True
wing.high_lift = False
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.sweeps.quarter_chord = 25. * Units.deg
wing.thickness_to_chord = 0.12
wing.areas.reference = 2.258 * Units['meters**2']
wing.spans.projected = 1.854 * Units.meter
wing.chords.root = 1.6764 * Units.meter
wing.chords.tip = 0.6858 * Units.meter
wing.chords.mean_aerodynamic = 1.21 * Units.meter
wing.taper = wing.chords.root / wing.chords.tip
wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[6.01, 0, 0.623]]
wing.aerodynamic_center = [0.508, 0, 0]
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 2.
fuselage.fineness.nose = 1.6
fuselage.fineness.tail = 2.
fuselage.lengths.nose = 60. * Units.inches
fuselage.lengths.tail = 161. * Units.inches
fuselage.lengths.cabin = 105. * Units.inches
fuselage.lengths.total = 332.2 * Units.inches
fuselage.lengths.fore_space = 0.
fuselage.lengths.aft_space = 0.
fuselage.width = 42. * Units.inches
fuselage.heights.maximum = 62. * Units.inches
fuselage.heights.at_quarter_length = 62. * Units.inches
fuselage.heights.at_three_quarters_length = 62. * Units.inches
fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
fuselage.areas.side_projected = 8000. * Units.inches**2.
fuselage.areas.wetted = 30000. * Units.inches**2.
fuselage.areas.front_projected = 42. * 62. * Units.inches**2.
fuselage.effective_diameter = 50. * Units.inches
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0
segment.percent_z_location = 0
segment.height = 0.01
segment.width = 0.01
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.007279116466
segment.percent_z_location = 0.002502014453
segment.height = 0.1669064748
segment.width = 0.2780205877
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.01941097724
segment.percent_z_location = 0.001216095397
segment.height = 0.3129496403
segment.width = 0.4365777215
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.06308567604
segment.percent_z_location = 0.007395489231
segment.height = 0.5841726619
segment.width = 0.6735119903
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 0.1653761217
segment.percent_z_location = 0.02891281352
segment.height = 1.064028777
segment.width = 1.067200529
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 0.2426372155
segment.percent_z_location = 0.04214148761
segment.height = 1.293766653
segment.width = 1.183058255
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_6'
segment.percent_x_location = 0.2960174029
segment.percent_z_location = 0.04705241831
segment.height = 1.377026712
segment.width = 1.181540054
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_7'
segment.percent_x_location = 0.3809404284
segment.percent_z_location = 0.05313580461
segment.height = 1.439568345
segment.width = 1.178218989
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_8'
segment.percent_x_location = 0.5046854083
segment.percent_z_location = 0.04655492473
segment.height = 1.29352518
segment.width = 1.054390707
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_9'
segment.percent_x_location = 0.6454149933
segment.percent_z_location = 0.03741966266
segment.height = 0.8971223022
segment.width = 0.8501926505
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_10'
segment.percent_x_location = 0.985107095
segment.percent_z_location = 0.04540283436
segment.height = 0.2920863309
segment.width = 0.2012565415
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_11'
segment.percent_x_location = 1
segment.percent_z_location = 0.04787575562
segment.height = 0.1251798561
segment.width = 0.1206021048
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#---------------------------------------------------------------------------------------------
# DEFINE PROPELLER
#---------------------------------------------------------------------------------------------
# build network
net = Battery_Propeller()
net.number_of_engines = 2.
net.nacelle_diameter = 42 * Units.inches
net.engine_length = 0.01 * Units.inches
net.areas = Data()
net.areas.wetted = 0.01 * (2 * np.pi * 0.01 / 2)
# Component 1 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 2 the Propeller
# Design the Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_of_blades = 2.0
prop.freestream_velocity = 135. * Units['mph']
prop.angular_velocity = 1300. * Units.rpm
prop.tip_radius = 76. / 2. * Units.inches
prop.hub_radius = 8. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 12000. * Units.feet
prop.design_altitude = 12000. * Units.feet
prop.design_thrust = 1200.
prop.origin = [[2., 2.5, 0.784], [2., -2.5, 0.784]]
prop.rotation = [-1, 1]
prop.symmetry = True
prop.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
prop.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
prop.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
prop = propeller_design(prop)
net.propeller = prop
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.mass_properties.mass = 500. * Units.kg
bat.specific_energy = 350. * Units.Wh / Units.kg
bat.resistance = 0.006
bat.max_voltage = 500.
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
net.voltage = bat.max_voltage
# Component 9 Miscellaneous Systems
sys = SUAVE.Components.Systems.System()
sys.mass_properties.mass = 5 # kg
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller motor
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
etam = 0.95
v = bat.max_voltage * 3 / 4
omeg = prop.angular_velocity
io = 4.0
start_kv = 1
end_kv = 25
# do optimization to find kv or just do a linspace then remove all negative values, take smallest one use 0.05 change
# essentially you are sizing the motor for a particular rpm which is sized for a design tip mach
# this reduces the bookkeeping errors
possible_kv_vals = np.linspace(
start_kv, end_kv,
(end_kv - start_kv) * 20 + 1, endpoint=True) * Units.rpm
res_kv_vals = ((v - omeg / possible_kv_vals) *
(1. - etam * v * possible_kv_vals / omeg)) / io
positive_res_vals = np.extract(res_kv_vals > 0, res_kv_vals)
kv_idx = np.where(res_kv_vals == min(positive_res_vals))[0][0]
kv = possible_kv_vals[kv_idx]
res = min(positive_res_vals)
motor.mass_properties.mass = 10. * Units.kg
motor.origin = prop.origin
motor.propeller_radius = prop.tip_radius
motor.speed_constant = 0.35
motor.resistance = res
motor.no_load_current = io
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1. # Gear box efficiency
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 10. #Watts
payload.mass_properties.mass = 1.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 20. #Watts
net.avionics = avionics
# add the solar network to the vehicle
vehicle.append_component(net)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Vahana'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2250. * Units.lb
vehicle.mass_properties.operating_empty = 2250. * Units.lb
vehicle.mass_properties.max_takeoff = 2250. * Units.lb
vehicle.mass_properties.center_of_gravity = [2.0144, 0., 0.]
vehicle.reference_area = 10.58275476
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
# ------------------------------------------------------
# WINGS
# ------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'canard_wing'
wing.aspect_ratio = 11.37706641
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.span_efficiency = 0.9
wing.spans.projected = 6.65
wing.chords.root = 0.95
wing.total_length = 0.95
wing.chords.tip = 0.95
wing.chords.mean_aerodynamic = 0.95
wing.dihedral = 0.0
wing.areas.reference = 6.31
wing.areas.wetted = 12.635
wing.areas.exposed = 12.635
wing.twists.root = 0.
wing.twists.tip = 0.
wing.origin = [0.0, 0.0, 0.0]
wing.aerodynamic_center = [0., 0., 0.]
wing.winglet_fraction = 0.0
wing.symmetric = True
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_1'
segment.origin = [0., 0., 0.]
segment.percent_span_location = 0.
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_2'
segment.origin = [0., 0., 0.]
segment.percent_span_location = 1.
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# add to vehicle
vehicle.append_component(wing)
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 11.37706641
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.span_efficiency = 0.9
wing.spans.projected = 6.65
wing.chords.root = 0.95
wing.total_length = 0.95
wing.chords.tip = 0.95
wing.chords.mean_aerodynamic = 0.95
wing.dihedral = 0.0
wing.areas.reference = 6.31
wing.areas.wetted = 12.635
wing.areas.exposed = 12.635
wing.twists.root = 0.
wing.twists.tip = 0.
wing.origin = [5.138, 0.0, 1.24]
wing.aerodynamic_center = [0., 0., 0.]
wing.winglet_fraction = 0.0
wing.symmetric = True
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_1'
segment.origin = [0., 0., 0.]
segment.percent_span_location = 0.
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# Segment
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_2'
segment.origin = [0., 0., 0.]
segment.percent_span_location = 1.
segment.twist = 0.
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.18
wing.Segments.append(segment)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------
# FUSELAGE
# ------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.origin = [0., 0., 0.]
fuselage.seats_abreast = 0.
fuselage.seat_pitch = 1.
fuselage.fineness.nose = 1.5
fuselage.fineness.tail = 4.0
fuselage.lengths.nose = 1.7
fuselage.lengths.tail = 2.7
fuselage.lengths.cabin = 1.7
fuselage.lengths.total = 6.1
fuselage.width = 1.15
fuselage.heights.maximum = 1.7
fuselage.heights.at_quarter_length = 1.2
fuselage.heights.at_wing_root_quarter_chord = 1.7
fuselage.heights.at_three_quarters_length = 0.75
fuselage.areas.wetted = 12.97989862
fuselage.areas.front_projected = 1.365211404
fuselage.effective_diameter = 1.318423736
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_1'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.
segment.percent_z_location = 0.
segment.height = 0.
segment.width = 0.
segment.length = 0.
segment.effective_diameter = 0.
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_2'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.275
segment.percent_z_location = -0.009
segment.height = 0.309 * 2
segment.width = 0.28 * 2
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_3'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.768
segment.percent_z_location = 0.046
segment.height = 0.525 * 2
segment.width = 0.445 * 2
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_4'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.25 * 6.2
segment.percent_z_location = 0.209
segment.height = 0.7 * 2
segment.width = 0.55 * 2
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_5'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.5 * 6.2
segment.percent_z_location = 0.407
segment.height = 0.850 * 2
segment.width = 0.61 * 2
segment.effective_diameter = 0.
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_6'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.75
segment.percent_z_location = 0.771
segment.height = 0.63 * 2
segment.width = 0.442 * 2
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_7'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 1. * 6.2
segment.percent_z_location = 1.192
segment.height = 0.165 * 2
segment.width = 0.125 * 2
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#------------------------------------------------------------------
# PROPULSOR
#------------------------------------------------------------------
net = Vectored_Thrust()
net.number_of_engines = 8
net.thrust_angle = 90.0 * Units.degrees # conversion to radians,
net.nacelle_diameter = 0.2921 # https://www.magicall.biz/products/integrated-motor-controller-magidrive/
net.engine_length = 0.106
net.areas = Data()
net.areas.wetted = np.pi * net.nacelle_diameter * net.engine_length + 0.5 * np.pi * net.nacelle_diameter**2
net.voltage = 500.
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95
net.esc = esc
#------------------------------------------------------------------
# Design Payload
#------------------------------------------------------------------
payload = SUAVE.Components.Energy.Peripherals.Avionics()
payload.power_draw = 0.
payload.mass_properties.mass = 200. * Units.kg
net.payload = payload
#------------------------------------------------------------------
# Design Avionics
#------------------------------------------------------------------
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 200. * Units.watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.specific_energy = 300. * Units.Wh / Units.kg
bat.resistance = 0.005
bat.max_voltage = net.voltage
bat.mass_properties.mass = 350. * Units.kg
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
#------------------------------------------------------------------
# Design Rotors
#------------------------------------------------------------------
# atmosphere conditions
speed_of_sound = 340
rho = 1.22
fligth_CL = 0.75
AR = vehicle.wings.main_wing.aspect_ratio
Cd0 = 0.06
Cdi = fligth_CL**2 / (np.pi * AR * 0.98)
Cd = Cd0 + Cdi
# Create propeller geometry
rot = SUAVE.Components.Energy.Converters.Rotor()
rot.y_pitch = 1.850
rot.tip_radius = 0.8875
rot.hub_radius = 0.1
rot.disc_area = np.pi * (rot.tip_radius**2)
rot.design_tip_mach = 0.5
rot.number_blades = 3
rot.freestream_velocity = 85. * Units['ft/min'] # 110 mph
rot.angular_velocity = rot.design_tip_mach * speed_of_sound / rot.tip_radius
rot.design_Cl = 0.7
rot.design_altitude = 500 * Units.feet
Lift = vehicle.mass_properties.takeoff * 9.81
rot.design_thrust = (Lift * 1.5) / net.number_of_engines
rot.induced_hover_velocity = np.sqrt(
Lift / (2 * rho * rot.disc_area * net.number_of_engines))
rot = propeller_design(rot)
# Front Rotors Locations
rot_front = Data()
rot_front.origin = [[0.0, 1.347, 0.0]]
rot_front.symmetric = True
rot_front.x_pitch_count = 1
rot_front.y_pitch_count = 2
rot_front.y_pitch = 1.85
# populating rotors on one side of wing
if rot_front.y_pitch_count > 1:
for n in range(rot_front.y_pitch_count):
if n == 0:
continue
for i in range(len(rot_front.origin)):
propeller_origin = [
rot_front.origin[i][0],
rot_front.origin[i][1] + n * rot_front.y_pitch,
rot_front.origin[i][2]
]
rot_front.origin.append(propeller_origin)
# populating rotors on the other side of the vehicle
if rot_front.symmetric:
for n in range(len(rot_front.origin)):
propeller_origin = [
rot_front.origin[n][0], -rot_front.origin[n][1],
rot_front.origin[n][2]
]
rot_front.origin.append(propeller_origin)
# Rear Rotors Locations
rot_rear = Data()
rot_rear.origin = [[0.0, 1.347, 1.24]]
rot_rear.symmetric = True
rot_rear.x_pitch_count = 1
rot_rear.y_pitch_count = 2
rot_rear.y_pitch = 1.85
# populating rotors on one side of wing
if rot_rear.y_pitch_count > 1:
for n in range(rot_rear.y_pitch_count):
if n == 0:
continue
for i in range(len(rot_rear.origin)):
propeller_origin = [
rot_rear.origin[i][0],
rot_rear.origin[i][1] + n * rot_rear.y_pitch,
rot_rear.origin[i][2]
]
rot_rear.origin.append(propeller_origin)
# populating rotors on the other side of the vehicle
if rot_rear.symmetric:
for n in range(len(rot_rear.origin)):
propeller_origin = [
rot_rear.origin[n][0], -rot_rear.origin[n][1],
rot_rear.origin[n][2]
]
rot_rear.origin.append(propeller_origin)
# Assign all rotors (front and rear) to network
rot.origin = rot_front.origin + rot_rear.origin
# append rotors to vehicle
net.rotor = rot
# Motor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.mass_properties.mass = 9. * Units.kg
motor.origin = rot_front.origin + rot_rear.origin
motor.efficiency = 0.935
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1. # Gear box efficiency
motor.nominal_voltage = bat.max_voltage * 3 / 4
motor.propeller_radius = rot.tip_radius
motor.no_load_current = 2.0
motor = compute_optimal_motor_parameters(motor, rot)
net.motor = motor
vehicle.append_component(net)
# Add extra drag sources from motors, props, and landing gear. All of these hand measured
motor_height = .25 * Units.feet
motor_width = 1.6 * Units.feet
propeller_width = 1. * Units.inches
propeller_height = propeller_width * .12
main_gear_width = 1.5 * Units.inches
main_gear_length = 2.5 * Units.feet
nose_gear_width = 2. * Units.inches
nose_gear_length = 2. * Units.feet
nose_tire_height = (0.7 + 0.4) * Units.feet
nose_tire_width = 0.4 * Units.feet
main_tire_height = (0.75 + 0.5) * Units.feet
main_tire_width = 4. * Units.inches
total_excrescence_area_spin = 12.*motor_height*motor_width + 2.* main_gear_length*main_gear_width \
+ nose_gear_width*nose_gear_length + 2 * main_tire_height*main_tire_width\
+ nose_tire_height*nose_tire_width
total_excrescence_area_no_spin = total_excrescence_area_spin + 12 * propeller_height * propeller_width
vehicle.excrescence_area_no_spin = total_excrescence_area_no_spin
vehicle.excrescence_area_spin = total_excrescence_area_spin
# append motor origin spanwise locations onto wing data structure
motor_origins_front = np.array(rot_front.origin)
vehicle.wings['canard_wing'].motor_spanwise_locations = np.multiply(
0.19, motor_origins_front[:, 1])
motor_origins_rear = np.array(rot_rear.origin)
vehicle.wings['main_wing'].motor_spanwise_locations = np.multiply(
0.19, motor_origins_rear[:, 1])
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'multicopter'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2080. * Units.lb
vehicle.mass_properties.operating_empty = 1666. * Units.lb
vehicle.mass_properties.max_takeoff = 2080. * Units.lb
vehicle.mass_properties.center_of_gravity = [[2.6, 0., 0.]]
# This needs updating
vehicle.passengers = 5
vehicle.reference_area = 73 * Units.feet**2
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 1.
wing.spans.projected = 0.1
wing.chords.root = 0.1
wing.chords.tip = 0.1
wing.origin = [[2.6, 0., 0.]]
wing.symbolic = True
vehicle.append_component(wing)
# ------------------------------------------------------
# FUSELAGE
# ------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.configuration = 'Tube_Wing'
fuselage.seats_abreast = 2.
fuselage.seat_pitch = 3.
fuselage.fineness.nose = 0.88
fuselage.fineness.tail = 1.13
fuselage.lengths.nose = 3.2 * Units.feet
fuselage.lengths.tail = 6.4 * Units.feet
fuselage.lengths.cabin = 6.4 * Units.feet
fuselage.lengths.total = 16.0 * Units.feet
fuselage.width = 5.85 * Units.feet
fuselage.heights.maximum = 4.65 * Units.feet
fuselage.heights.at_quarter_length = 3.75 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 4.65 * Units.feet
fuselage.heights.at_three_quarters_length = 4.26 * Units.feet
fuselage.areas.wetted = 236. * Units.feet**2
fuselage.areas.front_projected = 0.14 * Units.feet**2
fuselage.effective_diameter = 5.85 * Units.feet
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.
segment.percent_z_location = 0.0
segment.height = 0.1 * Units.feet
segment.width = 0.1 * Units.feet
segment.length = 0.
segment.effective_diameter = 0.1 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.origin = [4. * 0.3048, 0., 0.1 * 0.3048]
segment.percent_x_location = 0.25
segment.percent_z_location = 0.05
segment.height = 3.75 * Units.feet
segment.width = 5.65 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 5.65 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.origin = [8. * 0.3048, 0., 0.34 * 0.3048]
segment.percent_x_location = 0.5
segment.percent_z_location = 0.071
segment.height = 4.65 * Units.feet
segment.width = 5.55 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 5.55 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.origin = [12. * 0.3048, 0., 0.77 * 0.3048]
segment.percent_x_location = 0.75
segment.percent_z_location = 0.089
segment.height = 4.73 * Units.feet
segment.width = 4.26 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 4.26 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.origin = [16. * 0.3048, 0., 2.02 * 0.3048]
segment.percent_x_location = 1.0
segment.percent_z_location = 0.158
segment.height = 0.67 * Units.feet
segment.width = 0.33 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 0.33 * Units.feet
fuselage.append_segment(segment)
# add to vehicle
vehicle.append_component(fuselage)
# -----------------------------------------------------------------
# Design the Nacelle
# -----------------------------------------------------------------
nacelle = SUAVE.Components.Nacelles.Nacelle()
nacelle.diameter = 0.6 * Units.feet # need to check
nacelle.length = 0.5 * Units.feet
nacelle.tag = 'nacelle_1'
nacelle.areas.wetted = np.pi * nacelle.diameter * nacelle.length + 0.5 * np.pi * nacelle.diameter**2
nacelle.origin = [[0., 2., 1.4]]
vehicle.append_component(nacelle)
nacelle_2 = deepcopy(nacelle)
nacelle_2.tag = 'nacelle_2'
nacelle_2.origin = [[0.0, -2., 1.4]]
vehicle.append_component(nacelle_2)
nacelle_3 = deepcopy(nacelle)
nacelle_3.tag = 'nacelle_3'
nacelle_3.origin = [[2.5, 4., 1.4]]
vehicle.append_component(nacelle_3)
nacelle_4 = deepcopy(nacelle)
nacelle_4.tag = 'nacelle_4'
nacelle_4.origin = [[2.5, -4., 1.4]]
vehicle.append_component(nacelle_4)
nacelle_5 = deepcopy(nacelle)
nacelle_5.tag = 'nacelle_5'
nacelle_5.origin = [[5.0, 2., 1.4]]
vehicle.append_component(nacelle_5)
nacelle_6 = deepcopy(nacelle)
nacelle_6.tag = 'nacelle_6'
nacelle_6.origin = [[5.0, -2., 1.4]]
vehicle.append_component(nacelle_6)
#------------------------------------------------------------------
# Network
#------------------------------------------------------------------
net = SUAVE.Components.Energy.Networks.Battery_Propeller()
net.number_of_lift_rotor_engines = 6
net.nacelle_diameter = 0.6 * Units.feet # need to check
net.engine_length = 0.5 * Units.feet
net.areas = Data()
net.areas.wetted = np.pi * net.nacelle_diameter * net.engine_length + 0.5 * np.pi * net.nacelle_diameter**2
net.voltage = 500.
net.identical_lift_rotors = True
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95
net.esc = esc
#------------------------------------------------------------------
# Design Payload
#------------------------------------------------------------------
payload = SUAVE.Components.Energy.Peripherals.Avionics()
payload.power_draw = 0.
payload.mass_properties.mass = 200. * Units.kg
net.payload = payload
#------------------------------------------------------------------
# Design Avionics
#------------------------------------------------------------------
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 200. * Units.watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiNiMnCoO2_18650(
)
bat.mass_properties.mass = 300. * Units.kg
bat.max_voltage = net.voltage
initialize_from_mass(bat)
net.battery = bat
#------------------------------------------------------------------
# Design Rotors
#------------------------------------------------------------------
# atmosphere and flight conditions for propeller/lift_rotor design
g = 9.81 # gravitational acceleration
speed_of_sound = 340 # speed of sound
Hover_Load = vehicle.mass_properties.takeoff * g # hover load
design_tip_mach = 0.7 # design tip mach number
lift_rotor = SUAVE.Components.Energy.Converters.Lift_Rotor()
lift_rotor.tip_radius = 3.95 * Units.feet
lift_rotor.hub_radius = 0.6 * Units.feet
lift_rotor.disc_area = np.pi * (lift_rotor.tip_radius**2)
lift_rotor.number_of_blades = 3
lift_rotor.freestream_velocity = 10.0
lift_rotor.angular_velocity = (design_tip_mach *
speed_of_sound) / lift_rotor.tip_radius
lift_rotor.design_Cl = 0.7
lift_rotor.design_altitude = 1000 * Units.feet
lift_rotor.design_thrust = Hover_Load / (
net.number_of_lift_rotor_engines - 1
) # contingency for one-engine-inoperative condition
lift_rotor.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
lift_rotor.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
lift_rotor.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
lift_rotor = propeller_design(lift_rotor)
# Appending rotors with different origins
origins = [[0., 2., 1.4], [0.0, -2., 1.4], [2.5, 4., 1.4], [2.5, -4., 1.4],
[5.0, 2., 1.4], [5.0, -2., 1.4]]
for ii in range(6):
lift_rotor = deepcopy(lift_rotor)
lift_rotor.tag = 'lift_rotor'
lift_rotor.origin = [origins[ii]]
net.lift_rotors.append(lift_rotor)
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Motor
lift_motor = SUAVE.Components.Energy.Converters.Motor()
lift_motor.efficiency = 0.95
lift_motor.nominal_voltage = bat.max_voltage * 0.5
lift_motor.mass_properties.mass = 3. * Units.kg
lift_motor.origin = lift_rotor.origin
lift_motor.propeller_radius = lift_rotor.tip_radius
lift_motor.no_load_current = 2.0
lift_motor = size_optimal_motor(lift_motor, lift_rotor)
net.lift_motor = lift_motor
# Define motor sizing parameters
max_power = lift_rotor.design_power * 1.2
max_torque = lift_rotor.design_torque * 1.2
# test high temperature superconducting motor weight function
mass = hts_motor(max_power)
# test NDARC motor weight function
mass = nasa_motor(max_torque)
# test air cooled motor weight function
mass = air_cooled_motor(max_power)
lift_motor.mass_properties.mass = mass
# Appending motors with different origins
for ii in range(6):
lift_rotor_motor = deepcopy(lift_motor)
lift_rotor_motor.tag = 'motor'
net.lift_rotor_motors.append(lift_rotor_motor)
vehicle.append_component(net)
vehicle.weight_breakdown = empty(vehicle)
compute_component_centers_of_gravity(vehicle)
vehicle.center_of_gravity()
return vehicle
def main():
# ------------------------------------------------------------------
# Propulsor
# ------------------------------------------------------------------
# build network
net = Solar()
net.number_of_engines = 1.
net.nacelle_dia = 0.2
# Component 1 the Sun?
sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
net.solar_flux = sun
# Component 2 the solar panels
panel = SUAVE.Components.Energy.Converters.Solar_Panel()
panel.area = 100 * Units.m
panel.efficiency = 0.18
panel.mass_properties.mass = panel.area * .600
net.solar_panel = panel
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 5 the Propeller
# Propeller design specs
design_altitude = 0.0 * Units.km
Velocity = 10.0 # freestream m/s
RPM = 5887
Blades = 2.0
Radius = .4064
Hub_Radius = 0.05
Design_Cl = 0.7
Thrust = 0.0 #Specify either thrust or power to design for
Power = 7500. #Specify either thrust or power to design for
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = Blades
prop_attributes.freestream_velocity = Velocity
prop_attributes.angular_velocity = RPM * (2. * np.pi / 60.0)
prop_attributes.tip_radius = Radius
prop_attributes.hub_radius = Hub_Radius
prop_attributes.design_Cl = Design_Cl
prop_attributes.design_altitude = design_altitude
prop_attributes.design_thrust = Thrust
prop_attributes.design_power = Power
prop_attributes = propeller_design(prop_attributes)
# Create and attach this propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 4 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.01
motor.no_load_current = 8.0
motor.speed_constant = 140. * (2. * np.pi / 60.
) # RPM/volt converted to rad/s
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1.
motor.expected_current = 260.
motor.mass_properties.mass = 2.0
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 0. #Watts
payload.mass_properties.mass = 0. * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 0. #Watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
batterymass = 50. #kg
bat.type = 'Li-Ion'
bat.resistance = 0.0
bat.energy_density = 250.
initialize_from_mass(bat, batterymass)
bat.current_energy = bat.max_energy
net.battery = bat
#Component 9 the system logic controller and MPPT
logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
logic.system_voltage = 50.0
logic.MPPT_efficiency = 0.95
net.solar_logic = logic
# Setup the conditions to run the network
state = Data()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
)
state.numerics = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()
conditions = state.conditions
numerics = state.numerics
# Calculate atmospheric properties
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(
prop_attributes.design_altitude)
rho = atmosphere_conditions.density[0, :]
a = atmosphere_conditions.speed_of_sound[0, :]
mu = atmosphere_conditions.dynamic_viscosity[0, :]
T = atmosphere_conditions.temperature[0, :]
conditions.propulsion.throttle = np.array([[1.0], [1.0]])
conditions.freestream.velocity = np.array([[1.0], [1.0]])
conditions.freestream.density = np.array([rho, rho])
conditions.freestream.dynamic_viscosity = np.array([mu, mu])
conditions.freestream.speed_of_sound = np.array([a, a])
conditions.freestream.altitude = np.array([[design_altitude],
[design_altitude]])
conditions.propulsion.battery_energy = bat.max_energy * np.ones_like(
conditions.freestream.altitude)
conditions.frames.body.inertial_rotations = np.zeros([2, 3])
conditions.frames.inertial.time = np.array([[0.0], [1.0]])
numerics.time.integrate = np.array([[0, 0], [0, 1]])
numerics.time.differentiate = np.array([[0, 0], [0, 1]])
conditions.frames.planet.start_time = time.strptime(
"Sat, Jun 21 06:00:00 2014",
"%a, %b %d %H:%M:%S %Y",
)
conditions.frames.planet.latitude = np.array([[0.0], [0.0]])
conditions.frames.planet.longitude = np.array([[0.0], [0.0]])
conditions.freestream.temperature = np.array([T, T])
# Run the network and print the results
results = net(state)
F = results.thrust_force_vector
# Truth results
truth_F = [[522.40448791], [522.40448791]]
truth_i = [[314.90485916], [314.90485916]]
truth_rpm = [[6581.17653732], [6581.17653732]]
truth_bat = [[36000000.], [35984254.75704217]]
error = Data()
error.Thrust = np.max(np.abs(F[:, 0] - truth_F))
error.RPM = np.max(np.abs(conditions.propulsion.rpm - truth_rpm))
error.Current = np.max(np.abs(conditions.propulsion.current - truth_i))
error.Battery = np.max(np.abs(bat.current_energy - truth_bat))
print error
for k, v in error.items():
assert (np.abs(v) < 0.001)
return
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Pteryxyz_Electric'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 450. * Units.kg
vehicle.mass_properties.operating_empty = 450. * Units.kg
vehicle.mass_properties.max_takeoff = 450. * Units.kg
# basic parameters
vehicle.reference_area = 6.2
vehicle.envelope.ultimate_load = 2.0
vehicle.envelope.limit_load = 1.5
max_q = 56. #m/s
vehicle.envelope.maximum_dynamic_pressure = 0.5 * 1.225 * (max_q**2.
) #Max q
# ------------------------------------------------------------------
# Landing Gear
# ------------------------------------------------------------------
# used for noise calculations
landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
landing_gear.tag = "main_landing_gear"
landing_gear.main_tire_diameter = 0.423 * Units.m
landing_gear.nose_tire_diameter = 0.3625 * Units.m
landing_gear.main_strut_length = 0.4833 * Units.m
landing_gear.nose_strut_length = 0.3625 * Units.m
landing_gear.main_units = 2 # number of main landing gear units
landing_gear.nose_units = 1 # number of nose landing gear
landing_gear.main_wheels = 1 # number of wheels on the main landing gear
landing_gear.nose_wheels = 1 # number of wheels on the nose landing gear
vehicle.landing_gear = landing_gear
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = vehicle.reference_area
wing.spans.projected = 7.04 * Units.meter
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.symmetric = True
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.vertical = False
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.areas.reference / wing.spans.projected
wing.span_efficiency = 0.98
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.highlift = False
wing.vertical = False
wing.number_ribs = 26.
wing.number_end_ribs = 2.
wing.transition_x_upper = 0.6
wing.transition_x_lower = 1.0
wing.origin = [1.6, 0.0, 0.0] # meters
wing.aerodynamic_center = [1.9, 0.0, 0.0] # meters
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.spans.projected = 2.45 * Units.meter
wing.areas.reference = 1.20
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.95
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
wing.number_ribs = 5.0
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.areas.reference / wing.spans.projected
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected
wing.origin = [4.2, 0.0, 0.0] # meters
wing.aerodynamic_center = [0.5, 0.0, 0.0] # meters
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.spans.projected = 4.60 * Units.meter
wing.areas.reference = 0.5
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.97
wing.areas.reference = vehicle.reference_area * 0.1
wing.spans.projected = np.sqrt(wing.aspect_ratio * wing.areas.reference)
wing.chords.root = wing.areas.reference / wing.spans.projected
wing.chords.tip = wing.areas.reference / wing.spans.projected
wing.chords.mean_aerodynamic = wing.areas.reference / wing.spans.projected
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [3.7, 0.0, 0.0] # meters
wing.aerodynamic_center = [0.5, 0.0, 0.0] # meters
wing.symmetric = True
wing.vertical = True
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
wing.number_ribs = 5.
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
# fuselage.aerodynamic_center= [2.986,0,1.077]
fuselage.number_coach_seats = vehicle.passengers
fuselage.seats_abreast = 2
fuselage.seat_pitch = 0.995 * Units.meter
fuselage.fineness.nose = 1.27
fuselage.fineness.tail = 1 # 3.31
fuselage.lengths.nose = 1.16 * Units.meter
fuselage.lengths.tail = 4.637 * Units.meter
fuselage.lengths.cabin = 2.653 * Units.meter
fuselage.lengths.total = 8.45 * Units.meter
fuselage.lengths.fore_space = 0.0 * Units.meter
fuselage.lengths.aft_space = 0.0 * Units.meter
fuselage.width = 1.22 * Units.meter
fuselage.heights.maximum = 1.41 * Units.meter
fuselage.effective_diameter = 2 * Units.meter
fuselage.areas.side_projected = 7.46 * Units['meters**2']
fuselage.areas.wetted = 25.0 * Units['meters**2']
fuselage.areas.front_projected = 1.54 * Units['meters**2']
fuselage.differential_pressure = 0.0 * Units.pascal # Maximum differential pressure
fuselage.heights.at_quarter_length = 1.077 * Units.meter
fuselage.heights.at_three_quarters_length = 0.5 * Units.meter # 0.621 * Units.meter
fuselage.heights.at_wing_root_quarter_chord = 1.41 * Units.meter
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Propellers Powered By Batteries
# ------------------------------------------------------------------
# build network
net = Battery_Propeller()
net.number_of_engines = 1
net.nacelle_diameter = 0.58 * Units.meters
net.engine_length = 1.74 * Units.meters
net.thrust_angle = 0.0 * Units.degrees
net.voltage = 709.0 * Units.V
net.areas = Data()
net.areas.wetted = 1.1 * np.pi * net.nacelle_diameter * net.engine_length
max_power = 175 * Units.kW
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 4 - Propeller
# Design the Propeller
# Component 5 the Propeller front
# Design the Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_blades = 2.0
prop.freestream_velocity = 56.0 * Units['m/s'] # freestream
prop.angular_velocity = 10000. * Units['rpm']
prop.tip_radius = 1.40 * Units.meters
prop.hub_radius = 0.05 * Units.meters
prop.design_Cl = 0.7
prop.design_altitude = 5.0 * Units.km
prop.design_thrust = None
prop.design_power = 150000.0 * Units.watts
prop = propeller_design(prop)
net.propeller = prop
# Component 4 the Motor front
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.01
motor.no_load_current = 125. * Units.ampere
motor.speed_constant = 4.5 * Units[
'rpm/volt'] # RPM/volt converted to (rad/s)/volt
motor.propeller_radius = prop.tip_radius
motor.propeller_Cp = prop.power_coefficient
motor.gear_ratio = 1. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 212. # Expected current
motor.mass_properties.mass = 50.0 * Units.kg
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 50. * Units.watts
payload.mass_properties.mass = 5.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 50. * Units.watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.cell_energy = 9.36 #[Wh]
bat.cell_capacity = 2600 #[mAh]
bat.max_const_discharge = 35 #[A]
bat.v_max = 4.2 #[V]
bat.v_nom = 3.7 #[V]
bat.v_min = 2.5 #[V]
bat.cell_mass = 44.3 #[g]
bat.cell_current = 2.5 #[A]
bat.voltage = 709 #[V]
bat.power = 150 #[kW]
bat.boost_power = 175 #[kW]
bat.no_of_packs = 2
bat.current = bat.voltage * bat.power
bat.boost_current = bat.voltage * bat.boost_power
bat.current_per_pack = bat.current / bat.no_of_packs
bat.parallel_cells_per_pack = 4
bat.additional_cell = 1
bat.series_cells = bat.voltage / bat.v_max
bat.parallel_cells = bat.parallel_cells_per_pack * bat.additional_cell
bat.cells_per_pack = bat.series_cells * bat.parallel_cells
bat.cell_mass_ratio = 1.2
bat.mass_properties.mass = bat.cell_mass_ratio * bat.cells_per_pack * bat.cell_mass / 1000 * Units.kg
bat.pack_energy = bat.cells_per_pack * bat.cell_energy / 1000
bat.total_energy = bat.pack_energy * bat.no_of_packs
bat.specific_energy = 450. * Units.Wh / Units.kg
bat.resistance = 0.05
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
# add the energy network to the vehicle
vehicle.append_component(net)
# print vehicle
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Tecnam_P2006TElectric'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 1400 * Units.kilogram
vehicle.mass_properties.takeoff = 1400 * Units.kilogram
vehicle.mass_properties.operating_empty = 1000 * Units.kilogram
vehicle.mass_properties.max_zero_fuel = 1400 * Units.kilogram
vehicle.mass_properties.cargo = 80 * Units.kilogram
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
# basic parameters
vehicle.reference_area = 64.4 * Units['meters**2']
vehicle.passengers = 4
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Landing Gear
# ------------------------------------------------------------------
# used for noise calculations
landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
landing_gear.tag = "main_landing_gear"
landing_gear.main_tire_diameter = 0.423 * Units.m
landing_gear.nose_tire_diameter = 0.3625 * Units.m
landing_gear.main_strut_length = 0.4833 * Units.m
landing_gear.nose_strut_length = 0.3625 * Units.m
landing_gear.main_units = 2 #number of main landing gear units
landing_gear.nose_units = 1 #number of nose landing gear
landing_gear.main_wheels = 1 #number of wheels on the main landing gear
landing_gear.nose_wheels = 1 #number of wheels on the nose landing gear
vehicle.landing_gear = landing_gear
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.number_coach_seats = vehicle.passengers
fuselage.seats_abreast = 2
fuselage.seat_pitch = 0.995 * Units.meter
fuselage.fineness.nose = 1.27
fuselage.fineness.tail = 1 #3.31
fuselage.lengths.nose = 1.16 * Units.meter
fuselage.lengths.tail = 4.637 * Units.meter
fuselage.lengths.cabin = 2.653 * Units.meter
fuselage.lengths.total = 8.45 * Units.meter
fuselage.lengths.fore_space = 0.0 * Units.meter
fuselage.lengths.aft_space = 0.0 * Units.meter
fuselage.width = 1.22 * Units.meter
fuselage.heights.maximum = 1.41 * Units.meter
fuselage.effective_diameter = 2 * Units.meter
fuselage.areas.side_projected = 7.46 * Units['meters**2']
fuselage.areas.wetted = 25.0 * Units['meters**2']
fuselage.areas.front_projected = 1.54 * Units['meters**2']
fuselage.differential_pressure = 0.0 * Units.pascal # Maximum differential pressure
fuselage.heights.at_quarter_length = 1.077 * Units.meter
fuselage.heights.at_three_quarters_length = 0.5 * Units.meter #0.621 * Units.meter
fuselage.heights.at_wing_root_quarter_chord = 1.41 * Units.meter
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.thickness_to_chord = 0.2
wing.taper = 0.9
wing.spans.projected = 13.8 * Units.meter
wing.chords.root = 5.2 * Units.meter
wing.chords.tip = wing.chords.root * wing.taper
wing.chords.mean_aerodynamic = 0.6 * Units.meter
wing.areas.reference = (wing.chords.root +
wing.chords.tip) * wing.spans.projected / 2
wing.twists.root = 0. * Units.degrees
wing.twists.tip = 0. * Units.degrees
wing.dihedral = 1. * Units.degrees
wing.origin = [2.986, 0, 1.077] # meters
wing.sweeps.leading_edge = 1.9 * Units.deg
wing.aspect_ratio = (wing.spans.projected *
wing.spans.projected) / wing.areas.reference
wing.span_efficiency = 0.99 * (
1 - 0.0407 * (fuselage.width / wing.spans.projected) - 1.792 *
((fuselage.width / wing.spans.projected)**2))
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
## Wing Segments
# Root Segment
#segment = SUAVE.Components.Wings.Segment()
#segment.tag = 'root'
#segment.percent_span_location = 0.0
#segment.twist = 0. * Units.deg
#segment.root_chord_percent = 1.
#segment.dihedral_outboard = 1. * Units.degrees
#segment.sweeps.quarter_chord = 0. * Units.degrees
#segment.thickness_to_chord = 0.15
#airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
#airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
#segment.append_airfoil(airfoil)
#wing.Segments.append(segment)
# Tip Segment
#segment = SUAVE.Components.Wings.Segment()
#segment.tag = 'tip'
#segment.percent_span_location = 1.0
#segment.twist = 0. * Units.deg
#segment.root_chord_percent = 1.
#segment.dihedral_outboard = 1. * Units.degrees
#segment.sweeps.quarter_chord = 0. * Units.degrees
#segment.thickness_to_chord = 0.15
#airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
#airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
#segment.append_airfoil(airfoil)
#wing.Segments.append(segment)
# ------------------------------------------------------------------
# Flaps
# ------------------------------------------------------------------
wing.flaps.chord = 0.20
wing.flaps.span_start = 0.1053
wing.flaps.span_end = 0.6842
wing.flaps.type = 'single_slotted'
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 4.193
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.733
wing.spans.projected = 3.3 * Units.meter
wing.chords.root = 0.787 * Units.meter
wing.chords.tip = 0.787 * Units.meter
wing.chords.mean_aerodynamic = (wing.chords.root * (2.0 / 3.0) *
((1.0 + wing.taper + wing.taper**2.0) /
(1.0 + wing.taper))) * Units.meter
wing.areas.reference = 2.5971 * Units['meters**2']
wing.areas.exposed = 4.0 * Units['meters**2']
wing.areas.wetted = 4.0 * Units['meters**2']
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [7.789, 0.0, 0.3314] # meters
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 1.407
wing.sweeps.quarter_chord = 38.75 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = -0.107
wing.spans.projected = 1.574 * Units.meter
wing.chords.root = 1.2 * Units.meter
wing.chords.tip = 0.497 * Units.meter
wing.chords.mean_aerodynamic = (wing.chords.root * (2.0 / 3.0) *
((1.0 + wing.taper + wing.taper**2.0) /
(1.0 + wing.taper))) * Units.meter
wing.areas.reference = 1.761 * Units['meters**2']
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [7.25, 0, 0.497] # meters
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Propellers Powered By Batteries
# ------------------------------------------------------------------
# build network
net = SUAVE.Components.Energy.Networks.Lift_Forward_Propulsor_Network()
net.nacelle_diameter_lift = 0.08 * Units.meters
net.nacelle_diameter_forward = 0.1732 * Units.meters
net.engine_length_lift = 0.47244 * Units.meters
net.engine_length_forward = 1.2 * Units.meters
net.number_of_engines_lift = 12
net.number_of_engines_forward = 2
net.thrust_angle_lift = 0.0 * Units.degrees
net.thrust_angle_forward = 0.0 * Units.degrees
net.voltage = 461.
net.areas_forward = Data()
net.areas_forward.wetted = 1.1 * np.pi * net.nacelle_diameter_forward * net.engine_length_forward
net.areas_lift = Data()
net.areas_lift.wetted = 1.1 * np.pi * net.nacelle_diameter_forward * net.engine_length_lift
# Component 1 - Tip ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc_forward = esc
# Component 1 - High Lift ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc_lift = esc
# Component 2 - Tip Propeller
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 3.0
prop_attributes.freestream_velocity = 50.0 * Units['m/s'] # freestream m/s
prop_attributes.angular_velocity = 27500. * Units['rpm'] # For 20x10 prop
prop_attributes.tip_radius = 0.762 * Units.meter
prop_attributes.hub_radius = 0.085 * Units.meter
prop_attributes.design_Cl = 0.8
prop_attributes.design_altitude = 14.0 * Units.km
prop_attributes.design_thrust = 0.0 * Units.newton
prop_attributes.design_power = 250000. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop_forward = SUAVE.Components.Energy.Converters.Propeller()
prop_forward.prop_attributes = prop_attributes
net.propeller_forward = prop_forward
# Component 2 - High Lift Propeller
# Design the Propeller
prop_attributes = Data()
prop_attributes.number_blades = 5.0
prop_attributes.freestream_velocity = 1. * Units['m/s'] # freestream m/s
prop_attributes.angular_velocity = 2750 * Units['rpm'] # For 20x10 prop
prop_attributes.tip_radius = 0.2880360 * Units.meter
prop_attributes.hub_radius = 0.07772400 * Units.meter
prop_attributes.design_Cl = 0.8
prop_attributes.design_altitude = 0.0 * Units.meter
prop_attributes.design_thrust = 0.0 * Units.newton
prop_attributes.design_power = 10000. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop_lift = SUAVE.Components.Energy.Converters.Propeller()
prop_lift.prop_attributes = prop_attributes
net.propeller_lift = prop_lift
# Component 3 - Tip Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 1.
motor.no_load_current = 7. * Units.ampere
motor.speed_constant = 11.9999 * Units[
'rpm/volt'] # RPM/volt converted to (rad/s)/volt
motor.propeller_radius = prop_forward.prop_attributes.tip_radius
motor.propeller_Cp = prop_forward.prop_attributes.Cp[0]
motor.gear_ratio = 12. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 160. # Expected current
motor.mass_properties.mass = 9.0 * Units.kg
net.motor_forward = motor
# Component 3 - High Lift Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.008
motor.no_load_current = 4.5 * Units.ampere
motor.speed_constant = 5800. * Units[
'rpm/volt'] # RPM/volt converted to (rad/s)/volt
motor.propeller_radius = prop_lift.prop_attributes.tip_radius
motor.propeller_Cp = prop_lift.prop_attributes.Cp[0]
motor.gear_ratio = 12. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 25. # Expected current
motor.mass_properties.mass = 6.0 * Units.kg
net.motor_lift = motor
# Component 4 - the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 50. * Units.watts
payload.mass_properties.mass = 5.0 * Units.kg
net.payload = payload
# Component 5 - the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 50. * Units.watts
net.avionics = avionics
# Component 6 - the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.mass_properties.mass = 358.33 * Units.kg
bat.specific_energy = 192.84 * Units.Wh / Units.kg
bat.specific_power = 0.837 * Units.kW / Units.kg
bat.resistance = 0.0153
bat.max_voltage = 10000.
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
# add the energy network to the vehicle
vehicle.append_component(net)
#print vehicle
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Tiltwing'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2250. * Units.lb
vehicle.mass_properties.operating_empty = 2250. * Units.lb
vehicle.mass_properties.max_takeoff = 2250. * Units.lb
vehicle.mass_properties.center_of_gravity = [[2.0144, 0., 0.]]
vehicle.passengers = 1
vehicle.reference_area = 10.58275476
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
# ------------------------------------------------------
# WINGS
# ------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'canard_wing'
wing.aspect_ratio = 11.37706641
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.spans.projected = 6.65
wing.chords.root = 0.95
wing.total_length = 0.95
wing.chords.tip = 0.95
wing.chords.mean_aerodynamic = 0.95
wing.dihedral = 0.0
wing.areas.reference = 6.31
wing.areas.wetted = 12.635
wing.areas.exposed = 12.635
wing.twists.root = 0.
wing.twists.tip = 0.
wing.origin = [[0.1, 0.0, 0.0]]
wing.aerodynamic_center = [0., 0., 0.]
wing.winglet_fraction = 0.0
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 11.37706641
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.18
wing.taper = 1.
wing.spans.projected = 6.65
wing.chords.root = 0.95
wing.total_length = 0.95
wing.chords.tip = 0.95
wing.chords.mean_aerodynamic = 0.95
wing.dihedral = 0.0
wing.areas.reference = 6.31
wing.areas.wetted = 12.635
wing.areas.exposed = 12.635
wing.twists.root = 0.
wing.twists.tip = 0.
wing.origin = [[5.138, 0.0, 1.323]] # for images 1.54
wing.aerodynamic_center = [0., 0., 0.]
wing.winglet_fraction = 0.0
wing.symmetric = True
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------
# FUSELAGE
# ------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 0.
fuselage.seat_pitch = 1.
fuselage.fineness.nose = 1.5
fuselage.fineness.tail = 4.0
fuselage.lengths.nose = 1.7
fuselage.lengths.tail = 2.7
fuselage.lengths.cabin = 1.7
fuselage.lengths.total = 6.1
fuselage.width = 1.15
fuselage.heights.maximum = 1.7
fuselage.heights.at_quarter_length = 1.2
fuselage.heights.at_wing_root_quarter_chord = 1.7
fuselage.heights.at_three_quarters_length = 0.75
fuselage.areas.wetted = 12.97989862
fuselage.areas.front_projected = 1.365211404
fuselage.effective_diameter = 1.318423736
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.
segment.percent_z_location = 0.
segment.height = 0.09
segment.width = 0.23473
segment.length = 0.
segment.effective_diameter = 0.
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.97675 / 6.1
segment.percent_z_location = 0.21977 / 6.1
segment.height = 0.9027
segment.width = 1.01709
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 1.93556 / 6.1
segment.percent_z_location = 0.39371 / 6.1
segment.height = 1.30558
segment.width = 1.38871
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 3.44137 / 6.1
segment.percent_z_location = 0.57143 / 6.1
segment.height = 1.52588
segment.width = 1.47074
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 4.61031 / 6.1
segment.percent_z_location = 0.81577 / 6.1
segment.height = 1.14788
segment.width = 1.11463
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 1.
segment.percent_z_location = 1.19622 / 6.1
segment.height = 0.31818
segment.width = 0.23443
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
#------------------------------------------------------------------
# PROPULSOR
#------------------------------------------------------------------
net = Vectored_Thrust()
net.number_of_engines = 8
net.thrust_angle = 0.0 * Units.degrees # conversion to radians,
net.nacelle_diameter = 0.2921 # https://www.magicall.biz/products/integrated-motor-controller-magidrive/
net.engine_length = 0.95
net.areas = Data()
net.areas.wetted = np.pi * net.nacelle_diameter * net.engine_length + 0.5 * np.pi * net.nacelle_diameter**2
net.voltage = 400.
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95
net.esc = esc
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 10. #Watts
payload.mass_properties.mass = 0.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 20. #Watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.mass_properties.mass = 200. * Units.kg
bat.specific_energy = 200. * Units.Wh / Units.kg
bat.resistance = 0.006
bat.max_voltage = 400.
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
net.voltage = bat.max_voltage
# Component 9 Miscellaneous Systems
sys = SUAVE.Components.Systems.System()
sys.mass_properties.mass = 5 # kg
#------------------------------------------------------------------
# Design Rotors
#------------------------------------------------------------------
# atmosphere conditions
speed_of_sound = 340
rho = 1.22
fligth_CL = 0.75
AR = vehicle.wings.main_wing.aspect_ratio
Cd0 = 0.06
Cdi = fligth_CL**2 / (np.pi * AR * 0.98)
Cd = Cd0 + Cdi
# Create propeller geometry
rot = SUAVE.Components.Energy.Converters.Rotor()
rot.y_pitch = 1.850
rot.tip_radius = 0.8875
rot.hub_radius = 0.15
rot.disc_area = np.pi * (rot.tip_radius**2)
rot.design_tip_mach = 0.5
rot.number_of_blades = 3
rot.freestream_velocity = 10
rot.angular_velocity = rot.design_tip_mach * speed_of_sound / rot.tip_radius
rot.design_Cl = 0.7
rot.design_altitude = 500 * Units.feet
Lift = vehicle.mass_properties.takeoff * 9.81
rot.design_thrust = (Lift * 1.5) / net.number_of_engines
rot.induced_hover_velocity = np.sqrt(
Lift / (2 * rho * rot.disc_area * net.number_of_engines))
rot.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
rot.airfoil_polars = [[
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_50000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_100000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_200000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_500000.txt',
'../Vehicles/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt'
]]
rot.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
rot = propeller_design(rot)
rot.rotation = [1, 1, 1, 1, 1, 1, 1, 1]
# Front Rotors Locations
rot_front = Data()
rot_front.origin = [[-0.2, 1.347, 0.]]
rot_front.symmetric = True
rot_front.x_pitch_count = 1
rot_front.y_pitch_count = 2
rot_front.y_pitch = 1.95
# populating rotors on one side of wing
if rot_front.y_pitch_count > 1:
for n in range(rot_front.y_pitch_count):
if n == 0:
continue
for i in range(len(rot_front.origin)):
propeller_origin = [
rot_front.origin[i][0],
rot_front.origin[i][1] + n * rot_front.y_pitch,
rot_front.origin[i][2]
]
rot_front.origin.append(propeller_origin)
# populating rotors on the other side of the vehicle
if rot_front.symmetric:
for n in range(len(rot_front.origin)):
propeller_origin = [
rot_front.origin[n][0], -rot_front.origin[n][1],
rot_front.origin[n][2]
]
rot_front.origin.append(propeller_origin)
# Rear Rotors Locations
rot_rear = Data()
rot_rear.origin = [[5.138 - 0.2, 1.347, 1.54]]
rot_rear.symmetric = True
rot_rear.x_pitch_count = 1
rot_rear.y_pitch_count = 2
rot_rear.y_pitch = 1.95
# populating rotors on one side of wing
if rot_rear.y_pitch_count > 1:
for n in range(rot_rear.y_pitch_count):
if n == 0:
continue
for i in range(len(rot_rear.origin)):
propeller_origin = [
rot_rear.origin[i][0],
rot_rear.origin[i][1] + n * rot_rear.y_pitch,
rot_rear.origin[i][2]
]
rot_rear.origin.append(propeller_origin)
# populating rotors on the other side of the vehicle
if rot_rear.symmetric:
for n in range(len(rot_rear.origin)):
propeller_origin = [
rot_rear.origin[n][0], -rot_rear.origin[n][1],
rot_rear.origin[n][2]
]
rot_rear.origin.append(propeller_origin)
# Assign all rotors (front and rear) to network
rot.origin = rot_front.origin + rot_rear.origin
# append rotors to vehicle
net.rotor = rot
# Motor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.origin = rot_front.origin + rot_rear.origin
motor.efficiency = 0.935
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1. # Gear box efficiency
motor.nominal_voltage = bat.max_voltage * 3 / 4
motor.propeller_radius = rot.tip_radius
motor.no_load_current = 2.0
motor = size_optimal_motor(motor, rot)
motor.mass_properties.mass = nasa_motor(motor.design_torque)
net.motor = motor
vehicle.append_component(net)
# Add extra drag sources from motors, props, and landing gear. All of these hand measured
motor_height = .25 * Units.feet
motor_width = 1.6 * Units.feet
propeller_width = 1. * Units.inches
propeller_height = propeller_width * .12
main_gear_width = 1.5 * Units.inches
main_gear_length = 2.5 * Units.feet
nose_gear_width = 2. * Units.inches
nose_gear_length = 2. * Units.feet
nose_tire_height = (0.7 + 0.4) * Units.feet
nose_tire_width = 0.4 * Units.feet
main_tire_height = (0.75 + 0.5) * Units.feet
main_tire_width = 4. * Units.inches
total_excrescence_area_spin = 12.*motor_height*motor_width + 2.* main_gear_length*main_gear_width \
+ nose_gear_width*nose_gear_length + 2 * main_tire_height*main_tire_width\
+ nose_tire_height*nose_tire_width
total_excrescence_area_no_spin = total_excrescence_area_spin + 12 * propeller_height * propeller_width
vehicle.excrescence_area_no_spin = total_excrescence_area_no_spin
vehicle.excrescence_area_spin = total_excrescence_area_spin
# append motor origin spanwise locations onto wing data structure
motor_origins_front = np.array(rot_front.origin)
motor_origins_rear = np.array(rot_rear.origin)
vehicle.wings[
'canard_wing'].motor_spanwise_locations = motor_origins_front[:, 1] / vehicle.wings[
'canard_wing'].spans.projected
vehicle.wings[
'canard_wing'].motor_spanwise_locations = motor_origins_front[:, 1] / vehicle.wings[
'canard_wing'].spans.projected
vehicle.wings[
'main_wing'].motor_spanwise_locations = motor_origins_rear[:, 1] / vehicle.wings[
'main_wing'].spans.projected
net.origin = rot.origin
vehicle.weight_breakdown = empty(vehicle)
compute_component_centers_of_gravity(vehicle)
vehicle.center_of_gravity()
return vehicle
def main():
# This script could fail if either the design or analysis scripts fail,
# in case of failure check both. The design and analysis powers will
# differ because of karman-tsien compressibility corrections in the
# analysis scripts
net = Battery_Propeller()
net.number_of_engines = 2
# Design Gearbox
gearbox = SUAVE.Components.Energy.Converters.Gearbox()
gearbox.gearwheel_radius1 = 1
gearbox.gearwheel_radius2 = 1
gearbox.efficiency = 0.95
gearbox.inputs.torque = 800 # N-m
gearbox.inputs.speed = 209.43951023931953
gearbox.inputs.power = 7000
gearbox.compute()
# Design the Propeller with airfoil geometry defined
prop_a = SUAVE.Components.Energy.Converters.Propeller()
prop_a.number_blades = 2.0
prop_a.freestream_velocity = 50.0
prop_a.angular_velocity = gearbox.inputs.speed # 209.43951023931953
prop_a.tip_radius = 1.5
prop_a.hub_radius = 0.05
prop_a.design_Cl = 0.7
prop_a.design_altitude = 0.0 * Units.km
prop_a.airfoil_geometry = ['NACA_4412_geo.txt']
prop_a.airfoil_polars = ['NACA_4412_polar.txt']
prop_a.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
prop_a.design_power = gearbox.outputs.power # 7000
prop_a = propeller_design(prop_a)
# Design the Propeller with airfoil geometry defined
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_blades = 2.0
prop.freestream_velocity = 50.0
prop.angular_velocity = gearbox.inputs.speed # 209.43951023931953
prop.tip_radius = 1.5
prop.hub_radius = 0.05
prop.design_Cl = 0.7
prop.design_altitude = 0.0 * Units.km
prop.design_power = gearbox.outputs.power # 7000
prop = propeller_design(prop)
# Design a Rotor with airfoil geometry defined
rot_a = SUAVE.Components.Energy.Converters.Rotor()
rot_a.number_blades = 2.0
rot_a.freestream_velocity = 1 * Units.ft / Units.second
rot_a.angular_velocity = 2000. * (2. * np.pi / 60.0)
rot_a.tip_radius = 1.5
rot_a.hub_radius = 0.05
rot_a.design_Cl = 0.7
rot_a.design_altitude = 0.0 * Units.km
rot_a.design_thrust = 1000.0
rot_a.airfoil_geometry = ['NACA_4412_geo.txt']
rot_a.airfoil_polars = ['NACA_4412_polar.txt']
rot_a.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
rot_a.induced_hover_velocity = 13.5 #roughly equivalent to a Chinook at SL
rot_a = propeller_design(rot_a)
# Design a Rotor without airfoil geometry defined
rot = SUAVE.Components.Energy.Converters.Rotor()
rot.number_blades = 2.0
rot.freestream_velocity = 1 * Units.ft / Units.second
rot.angular_velocity = 2000. * (2. * np.pi / 60.0)
rot.tip_radius = 1.5
rot.hub_radius = 0.05
rot.design_Cl = 0.7
rot.design_altitude = 0.0 * Units.km
rot.design_thrust = 1000.0
rot.induced_hover_velocity = 13.5 #roughly equivalent to a Chinook at SL
rot = propeller_design(rot)
# Find the operating conditions
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(prop.design_altitude)
V = prop.freestream_velocity
Vr = rot.freestream_velocity
conditions = Data()
conditions.freestream = Data()
conditions.propulsion = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.freestream.update(atmosphere_conditions)
conditions.freestream.dynamic_viscosity = atmosphere_conditions.dynamic_viscosity
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions.propulsion.throttle = np.array([[1.0]])
conditions.frames.body.transform_to_inertial = np.array([np.eye(3)])
conditions_r = copy.deepcopy(conditions)
conditions.frames.inertial.velocity_vector = np.array([[V, 0, 0]])
conditions_r.frames.inertial.velocity_vector = np.array([[0, Vr, 0]])
# Create and attach this propeller
prop_a.inputs.omega = np.array(prop.angular_velocity, ndmin=2)
prop.inputs.omega = np.array(prop.angular_velocity, ndmin=2)
rot_a.inputs.omega = copy.copy(prop.inputs.omega)
rot.inputs.omega = copy.copy(prop.inputs.omega)
# propeller with airfoil results
F_a, Q_a, P_a, Cplast_a, output_a, etap_a = prop_a.spin(conditions)
# propeller without airfoil results
conditions.propulsion.pitch_command = np.array([[1.0]]) * Units.degree
F, Q, P, Cplast, output, etap = prop.spin_variable_pitch(conditions)
# rotor with airfoil results
Fr_a, Qr_a, Pr_a, Cplastr_a, outputr_a, etapr = rot_a.spin(conditions_r)
# rotor with out airfoil results
conditions_r.propulsion.pitch_command = np.array([[1.0]]) * Units.degree
Fr, Qr, Pr, Cplastr, outputr, etapr = rot.spin_variable_pitch(conditions_r)
# Truth values for propeller with airfoil geometry defined
F_a_truth = 97.95830293
Q_a_truth = 28.28338146
P_a_truth = 5923.6575610
Cplast_a_truth = 0.00053729
# Truth values for propeller without airfoil geometry defined
F_truth = 179.27020984
Q_truth = 48.06170453
P_truth = 10066.0198578
Cplast_truth = 0.00091302
# Truth values for rotor with airfoil geometry defined
Fr_a_truth = 893.16859917
Qr_a_truth = 77.57705597
Pr_a_truth = 16247.70060823
Cplastr_a_truth = 0.00147371
# Truth values for rotor without airfoil geometry defined
Fr_truth = 900.63698565
Qr_truth = 78.01972629
Pr_truth = 16340.41326374
Cplastr_truth = 0.00148212
# Store errors
error = Data()
error.Thrust_a = np.max(np.abs(F_a - F_a_truth))
error.Torque_a = np.max(np.abs(Q_a - Q_a_truth))
error.Power_a = np.max(np.abs(P_a - P_a_truth))
error.Cp_a = np.max(np.abs(Cplast_a - Cplast_a_truth))
error.Thrust = np.max(np.abs(F - F_truth))
error.Torque = np.max(np.abs(Q - Q_truth))
error.Power = np.max(np.abs(P - P_truth))
error.Cp = np.max(np.abs(Cplast - Cplast_truth))
error.Thrustr_a = np.max(np.abs(Fr_a - Fr_a_truth))
error.Torquer_a = np.max(np.abs(Qr_a - Qr_a_truth))
error.Powerr_a = np.max(np.abs(Pr_a - Pr_a_truth))
error.Cpr_a = np.max(np.abs(Cplastr_a - Cplastr_a_truth))
error.Thrustr = np.max(np.abs(Fr - Fr_truth))
error.Torquer = np.max(np.abs(Qr - Qr_truth))
error.Powerr = np.max(np.abs(Pr - Pr_truth))
error.Cpr = np.max(np.abs(Cplastr - Cplastr_truth))
print('Errors:')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
return
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'multicopter'
vehicle.configuration = 'eVTOL'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.takeoff = 2080. * Units.lb
vehicle.mass_properties.operating_empty = 1666. * Units.lb
vehicle.mass_properties.max_takeoff = 2080. * Units.lb
vehicle.mass_properties.center_of_gravity = [2.6, 0., 0.]
# This needs updating
vehicle.passengers = 5
vehicle.reference_area = 73 * Units.feet**2
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 1
wing.spans.projected = 0.01
vehicle.append_component(wing)
# ------------------------------------------------------
# FUSELAGE
# ------------------------------------------------------
# FUSELAGE PROPERTIES
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.configuration = 'Tube_Wing'
fuselage.seats_abreast = 2.
fuselage.seat_pitch = 3.
fuselage.fineness.nose = 0.88
fuselage.fineness.tail = 1.13
fuselage.lengths.nose = 3.2 * Units.feet
fuselage.lengths.tail = 6.4 * Units.feet
fuselage.lengths.cabin = 6.4 * Units.feet
fuselage.lengths.total = 16.0 * Units.feet
fuselage.width = 5.85 * Units.feet
fuselage.heights.maximum = 4.65 * Units.feet
fuselage.heights.at_quarter_length = 3.75 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 4.65 * Units.feet
fuselage.heights.at_three_quarters_length = 4.26 * Units.feet
fuselage.areas.wetted = 236. * Units.feet**2
fuselage.areas.front_projected = 0.14 * Units.feet**2
fuselage.effective_diameter = 5.85 * Units.feet
fuselage.differential_pressure = 0.
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_1'
segment.origin = [0., 0., 0.]
segment.percent_x_location = 0.
segment.percent_z_location = 0.0
segment.height = 0.1 * Units.feet
segment.width = 0.1 * Units.feet
segment.length = 0.
segment.effective_diameter = 0.1 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_2'
segment.origin = [4. * 0.3048, 0., 0.1 * 0.3048]
segment.percent_x_location = 0.25
segment.percent_z_location = 0.05
segment.height = 3.75 * Units.feet
segment.width = 5.65 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 5.65 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_3'
segment.origin = [8. * 0.3048, 0., 0.34 * 0.3048]
segment.percent_x_location = 0.5
segment.percent_z_location = 0.071
segment.height = 4.65 * Units.feet
segment.width = 5.55 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 5.55 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_4'
segment.origin = [12. * 0.3048, 0., 0.77 * 0.3048]
segment.percent_x_location = 0.75
segment.percent_z_location = 0.089
segment.height = 4.73 * Units.feet
segment.width = 4.26 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 4.26 * Units.feet
fuselage.append_segment(segment)
# Segment
segment = SUAVE.Components.Fuselages.Segment()
segment.tag = 'segment_5'
segment.origin = [16. * 0.3048, 0., 2.02 * 0.3048]
segment.percent_x_location = 1.0
segment.percent_z_location = 0.158
segment.height = 0.67 * Units.feet
segment.width = 0.33 * Units.feet
segment.length = 3.2 * Units.feet
segment.effective_diameter = 0.33 * Units.feet
fuselage.append_segment(segment)
# add to vehicle
vehicle.append_component(fuselage)
#------------------------------------------------------------------
# PROPULSOR
#------------------------------------------------------------------
net = Vectored_Thrust()
net.number_of_engines = 6
net.thrust_angle = 90. * Units.degrees
net.nacelle_diameter = 0.6 * Units.feet # need to check
net.engine_length = 0.5 * Units.feet
net.areas = Data()
net.areas.wetted = np.pi * net.nacelle_diameter * net.engine_length + 0.5 * np.pi * net.nacelle_diameter**2
net.voltage = 500.
#------------------------------------------------------------------
# Design Electronic Speed Controller
#------------------------------------------------------------------
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95
net.esc = esc
#------------------------------------------------------------------
# Design Payload
#------------------------------------------------------------------
payload = SUAVE.Components.Energy.Peripherals.Avionics()
payload.power_draw = 0.
payload.mass_properties.mass = 200. * Units.kg
net.payload = payload
#------------------------------------------------------------------
# Design Avionics
#------------------------------------------------------------------
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 200. * Units.watts
net.avionics = avionics
#------------------------------------------------------------------
# Design Battery
#------------------------------------------------------------------
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion(
)
bat.specific_energy = 350. * Units.Wh / Units.kg
bat.resistance = 0.005
bat.max_voltage = net.voltage
bat.mass_properties.mass = 300. * Units.kg
initialize_from_mass(bat, bat.mass_properties.mass)
net.battery = bat
#------------------------------------------------------------------
# Design Rotors
#------------------------------------------------------------------
# atmosphere and flight conditions for propeller/rotor design
g = 9.81 # gravitational acceleration
speed_of_sound = 340 # speed of sound
rho = 1.22 # reference density
Hover_Load = vehicle.mass_properties.takeoff * g # hover load
design_tip_mach = 0.7 # design tip mach number
rotor = SUAVE.Components.Energy.Converters.Rotor()
rotor.tip_radius = 3.95 * Units.feet
rotor.hub_radius = 0.6 * Units.feet
rotor.disc_area = np.pi * (rotor.tip_radius**2)
rotor.number_blades = 3
rotor.freestream_velocity = 500. * Units['ft/min']
rotor.angular_velocity = (design_tip_mach *
speed_of_sound) / rotor.tip_radius
rotor.design_Cl = 0.8
rotor.design_altitude = 1000 * Units.feet
rotor.design_thrust = (Hover_Load / net.number_of_engines) * 2.
rotor = propeller_design(rotor)
rotor.induced_hover_velocity = np.sqrt(
Hover_Load / (2 * rho * rotor.disc_area * net.number_of_engines))
# propulating propellers on the other side of the vehicle
rotor.origin = []
for fuselage in vehicle.fuselages:
if fuselage.tag == 'fuselage':
continue
else:
rotor.origin.append(fuselage.origin[0])
# append propellers to vehicle
net.rotor = rotor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.efficiency = 0.95
motor.nominal_voltage = bat.max_voltage
motor.mass_properties.mass = 3. * Units.kg
motor.origin = rotor.origin
motor.propeller_radius = rotor.tip_radius
motor.gear_ratio = 1.0
motor.gearbox_efficiency = 1.0
motor.no_load_current = 4.0
motor = size_optimal_motor(motor, rotor)
net.motor = motor
# Define motor sizing parameters
max_power = rotor.design_power * 1.2
max_torque = rotor.design_torque * 1.2
# test high temperature superconducting motor weight function
mass = hts_motor(max_power)
# test NDARC motor weight function
mass = nasa_motor(max_torque)
# test air cooled motor weight function
mass = air_cooled_motor(max_power)
motor.mass_properties.mass = mass
net.motor = motor
vehicle.append_component(net)
vehicle.weight_breakdown = empty(vehicle, None)
return vehicle
def main():
'''This script checks the functions in in Motor.py used to compute motor torques
and output voltage and currents'''
# Propeller
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.number_blades = 2.0
prop.freestream_velocity = 50.0
prop.angular_velocity = 209.43951023931953
prop.tip_radius = 1.5
prop.hub_radius = 0.05
prop.design_Cl = 0.7
prop.design_altitude = 0.0 * Units.km
prop.design_thrust = 2271.2220451593753
prop = propeller_design(prop)
# Motor
#------------------------------------------------------------------
# Design Motors
#------------------------------------------------------------------
# Propeller (Thrust) motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.mass_properties.mass = 9. * Units.kg
motor.efficiency = 0.935
motor.gear_ratio = 1.
motor.gearbox_efficiency = 1. # Gear box efficiency
motor.no_load_current = 2.0
motor.propeller_radius = prop.tip_radius
motor.nominal_voltage = 400
motor = size_optimal_motor(motor, prop)
# Propeller (Thrust) motor
motor_low_fid = SUAVE.Components.Energy.Converters.Motor_Lo_Fid()
motor_low_fid.motor_efficiency = 0.98
motor_low_fid.rated_power = 1000
motor_low_fid.rated_voltage = 200
motor_low_fid.mass_properties.mass = 9. * Units.kg
size_from_mass(motor_low_fid)
# Find the operating conditions
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere_conditions = atmosphere.compute_values(prop.design_altitude)
V = prop.freestream_velocity
conditions = Data()
conditions.freestream = Data()
conditions.propulsion = Data()
conditions.frames = Data()
conditions.frames.body = Data()
conditions.frames.inertial = Data()
conditions.freestream.update(atmosphere_conditions)
conditions.freestream.dynamic_viscosity = atmosphere_conditions.dynamic_viscosity
conditions.freestream.velocity = np.array([[V, 0, 0]])
conditions.propulsion.throttle = np.array([[1.0]])
conditions.frames.body.transform_to_inertial = np.array([np.eye(3)])
conditions.propulsion.propeller_power_coefficient = np.array([[0.02]])
#------------------------------------
# Motor Omega Function
#------------------------------------
# create copy of motor to test functions
motor_1 = motor
# Define function specific inputs
voltage_1 = 400
motor_1.inputs.voltage = np.array([[voltage_1]])
# Run Motor Omega Function
omega_1 = motor_1.omega(conditions)
torque_1 = motor_1.outputs.torque[0][0]
#------------------------------------
# Motor Current Function
#------------------------------------
# create copy of motor to test functions
motor_2 = motor
# Define function specific inputs
motor_2.inputs.voltage = np.array([[voltage_1]])
motor_2.outputs.omega = np.array([[prop.angular_velocity]])
# Run Motor Current Function
i, etam = motor_2.current(conditions)
current_2 = i[0][0]
#------------------------------------
# Motor Torque Function
#------------------------------------
# create copy of motor to test functions
motor_3 = motor
# Define function specific inputs
motor_3.inputs.voltage = np.array([[voltage_1]])
motor_3.inputs.omega = np.array([[prop.angular_velocity]])
# Run Motor Torque Function
motor_3.torque(conditions)
torque_3 = motor_3.outputs.torque[0][0]
#------------------------------------
# Motor Voltage-Current Function
#------------------------------------
# create copy of motor to test functions
motor_4 = motor
# Define function specific inputs
motor_4.inputs.torque = np.array([[torque_1]])
# Run Motor Voltage-Current Function
motor_4.voltage_current(conditions)
voltage_4 = motor_4.outputs.voltage[0][0]
current_4 = motor_4.outputs.current[0][0]
#------------------------------------
# Low Fidelity Motor
#------------------------------------
motor_low_fid.inputs.voltage = np.array([[voltage_1]])
p, i = motor_low_fid.power_lo(conditions)
power_out = p[0][0]
current = i[0][0]
# Truth values
omega_1_truth = 209.32373481
torque_1_truth = 1051.64631117327
current_2_truth = 332.1671891793585
torque_3_truth = 630.0492434009769
voltage_4_truth = 400.22093117369883
current_4_truth = 553.0983628781883
power_out_truth = 1960.0
error = Data()
error.omega_test = np.max(np.abs(omega_1_truth - omega_1[0]))
error.torque_test_1 = np.max(np.abs(torque_1_truth - torque_1))
error.current_test_1 = np.max(np.abs(current_2_truth - current_2))
error.torque_test_2 = np.max(np.abs(torque_3_truth - torque_3))
error.voltage_test = np.max(np.abs(voltage_4_truth - voltage_4))
error.current_test_2 = np.max(np.abs(current_4_truth - current_4))
error.power_out_test = np.max(np.abs(power_out_truth - power_out))
print('Errors:')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
return
def configs_setup(vehicle):
#---------------------------------------------------------------------------
# Initialize Configurations
#---------------------------------------------------------------------------
configs = SUAVE.Components.Configs.Config.Container()
base_config = SUAVE.Components.Configs.Config(vehicle)
base_config.tag = 'base'
configs.append(base_config)
#---------------------------------------------------------------------------
# Electric Helicopter Configuration
#---------------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'electric_helicopter'
config.propulsors.network.number_of_engines = 1
prop_attributes = Data()
prop_attributes.number_blades = 4.0
prop_attributes.freestream_velocity = 150. * Units['meter/second']
prop_attributes.angular_velocity = 3500. * Units['rpm']
prop_attributes.tip_radius = 3800. * Units.mm
prop_attributes.hub_radius = 500. * Units.mm
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 1. * Units.km
prop_attributes.design_thrust = 1600. * 9.81 * Units.newtons
prop_attributes.design_power = 0. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.origin = [0.,0.,0.]
config.propulsors.network.propeller = prop
configs.append(config)
#---------------------------------------------------------------------------
# Electric Stopped Rotor Configuration
#---------------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'electric_stopped_rotor'
config.propulsors.network.number_of_engines = 8
prop_attributes = Data()
prop_attributes.number_blades = 4.0
prop_attributes.freestream_velocity = 150. * Units['meter/second']
prop_attributes.angular_velocity = 3500. * Units['rpm']
prop_attributes.tip_radius = 800. * Units.mm #608
prop_attributes.hub_radius = 150. * Units.mm
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 1. * Units.km
prop_attributes.design_thrust = 200. * 9.81 * Units.newtons
prop_attributes.design_power = 0. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.origin = [0.,0.,0.]
config.propulsors.network.propeller = prop
thrust_prop_attributes = cp.deepcopy(prop_attributes)
thrust_prop_attributes.number_blades = 2.0
thrust_prop = SUAVE.Components.Energy.Converters.Propeller()
thrust_prop.prop_attributes = thrust_prop_attributes
config.propulsors.network.thrust_propeller = thrust_prop
configs.append(config)
#---------------------------------------------------------------------------
# Electric Tiltrotor Configuration
#---------------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'electric_tiltrotor'
config.propulsors.network.number_of_engines = 8
prop_attributes = Data()
prop_attributes.number_blades = 4.0
prop_attributes.freestream_velocity = 150. * Units['meter/second']
prop_attributes.angular_velocity = 3500. * Units['rpm']
prop_attributes.tip_radius = 800. * Units.mm #608
prop_attributes.hub_radius = 150. * Units.mm
prop_attributes.design_Cl = 0.7
prop_attributes.design_altitude = 1. * Units.km
prop_attributes.design_thrust = 200. * 9.81 * Units.newtons
prop_attributes.design_power = 0. * Units.watts
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller()
prop.prop_attributes = prop_attributes
prop.origin = [0.,0.,0.]
config.propulsors.network.propeller = prop
configs.append(config)
return configs