def weight(nexus):
vehicle = nexus.vehicle_configurations.base
# weight analysis
weights = nexus.analyses.base.weights.evaluate()
# print "nexus_rest"
compute_component_centers_of_gravity(vehicle)
nose_load_fraction = .06
compute_aircraft_center_of_gravity(vehicle, nose_load_fraction)
weights = nexus.analyses.cruise.weights.evaluate()
# print weights # FIXME
weights = nexus.analyses.landing.weights.evaluate()
# print weights
weights = nexus.analyses.takeoff.weights.evaluate()
# weights = nexus.analyses.short_field_takeoff.weights.evaluate()
empty_weight = vehicle.mass_properties.operating_empty
passenger_weight = 0
for config in nexus.vehicle_configurations:
# config.mass_properties.max_zero_fuel = empty_weight+passenger_weight
config.mass_properties.zero_fuel_center_of_gravity = vehicle.mass_properties.zero_fuel_center_of_gravity
config.fuel = vehicle.fuel
return nexus
def weight(nexus):
vehicle = nexus.vehicle_configurations.base
# weight analysis
weights = nexus.analyses.base.weights.evaluate()
# print "nexus_rest"
compute_component_centers_of_gravity(vehicle)
nose_load_fraction = .06
compute_aircraft_center_of_gravity(vehicle, nose_load_fraction)
weights = nexus.analyses.cruise.weights.evaluate()
# print weights # FIXME
weights = nexus.analyses.landing.weights.evaluate()
# print weights
weights = nexus.analyses.takeoff.weights.evaluate()
# weights = nexus.analyses.short_field_takeoff.weights.evaluate()
empty_weight = vehicle.mass_properties.operating_empty
passenger_weight = 0
for config in nexus.vehicle_configurations:
# config.mass_properties.max_zero_fuel = empty_weight+passenger_weight
config.mass_properties.zero_fuel_center_of_gravity = vehicle.mass_properties.zero_fuel_center_of_gravity
config.fuel = vehicle.fuel
return nexus
def simple_sizing(configs, analyses):
base = configs.base
base.pull_base()
# zero fuel weight
base.mass_properties.max_zero_fuel = 0.9 * base.mass_properties.max_takeoff
# wing areas
for wing in base.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# fuselage seats
base.fuselages['fuselage'].number_coach_seats = base.passengers
# weight analysis
#need to put here, otherwise it won't be updated
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
#compute centers of gravity
#need to put here, otherwise, results won't be stored
compute_component_centers_of_gravity(base,compute_propulsor_origin=True)
compute_aircraft_center_of_gravity(base)
# diff the new data
base.store_diff()
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = configs.landing
# make sure base data is current
landing.pull_base()
# landing weight
landing.mass_properties.landing = 0.85 * base.mass_properties.takeoff
# diff the new data
landing.store_diff()
# done!
return
def weight(nexus):
vehicle = nexus.vehicle_configurations.base
# weight analysis
weights = nexus.analyses.base.weights.evaluate()
nose_load_fraction = .06
compute_component_centers_of_gravity(vehicle, nose_load_fraction)
vehicle.center_of_gravity()
weights = nexus.analyses.cruise.weights.evaluate()
weights = nexus.analyses.landing.weights.evaluate()
weights = nexus.analyses.takeoff.weights.evaluate()
weights = nexus.analyses.short_field_takeoff.weights.evaluate()
empty_weight = vehicle.mass_properties.operating_empty
for config in nexus.vehicle_configurations:
config.mass_properties.zero_fuel_center_of_gravity = vehicle.mass_properties.zero_fuel_center_of_gravity
config.fuel = vehicle.fuel
return nexus
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 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 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 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 = '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
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.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_of_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.airfoil_geometry = ['../Vehicles/Airfoils/NACA_4412.txt']
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'
]]
rotor.airfoil_polar_stations = [
0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0, 0
]
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)
compute_component_centers_of_gravity(vehicle)
vehicle.center_of_gravity()
return vehicle
