def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Concorde'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 185000. # kg
vehicle.mass_properties.operating_empty = 78700. # kg
vehicle.mass_properties.takeoff = 183000. # kg, adjusted due to significant fuel burn on runway
vehicle.mass_properties.cargo = 1000. * Units.kilogram
vehicle.mass_properties.max_zero_fuel = 92000.
# envelope properties
vehicle.envelope.ultimate_load = 3.75
vehicle.envelope.limit_load = 2.5
# basic parameters
vehicle.reference_area = 358.25
vehicle.passengers = 100
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "sst"
vehicle.maximum_cross_sectional_area = 13.9
vehicle.total_length = 61.66
vehicle.design_mach_number = 2.02
vehicle.design_range = 4505 * Units.miles
vehicle.design_cruise_alt = 60000.0 * Units.ft
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 1.83
wing.sweeps.quarter_chord = 59.5 * Units.deg
wing.sweeps.leading_edge = 66.5 * Units.deg
wing.thickness_to_chord = 0.03
wing.taper = 0.
wing.spans.projected = 25.6
wing.chords.root = 33.8
wing.total_length = 33.8
wing.chords.tip = 1.1
wing.chords.mean_aerodynamic = 18.4
wing.areas.reference = 358.25
wing.areas.wetted = 601.
wing.areas.exposed = 326.5
wing.areas.affected = .6 * wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[14, 0, -.8]]
wing.aerodynamic_center = [35, 0, 0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.vortex_lift = True
wing.high_mach = True
wing.dynamic_pressure_ratio = 1.0
wing_airfoil = SUAVE.Components.Airfoils.Airfoil()
# This airfoil is not a true Concorde airfoil
wing_airfoil.coordinate_file = '../Vehicles/Airfoils/NACA65-203.txt'
wing.append_airfoil(wing_airfoil)
# set root sweep with inner section
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_1'
segment.percent_span_location = 0.
segment.twist = 0. * Units.deg
segment.root_chord_percent = 33.8 / 33.8
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 67. * Units.deg
segment.thickness_to_chord = 0.03
#segment.append_airfoil(wing_airfoil)
wing.Segments.append(segment)
# set section 2 start point
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_2'
segment.percent_span_location = 6.15 / (
25.6 / 2) + wing.Segments['section_1'].percent_span_location
segment.twist = 0. * Units.deg
segment.root_chord_percent = 13.8 / 33.8
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 48. * Units.deg
segment.thickness_to_chord = 0.03
segment.append_airfoil(wing_airfoil)
wing.Segments.append(segment)
# set section 3 start point
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_3'
segment.percent_span_location = 5.95 / (
25.6 / 2) + wing.Segments['section_2'].percent_span_location
segment.twist = 0. * Units.deg
segment.root_chord_percent = 4.4 / 33.8
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 71. * Units.deg
segment.thickness_to_chord = 0.03
segment.append_airfoil(wing_airfoil)
wing.Segments.append(segment)
# set tip
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'tip'
segment.percent_span_location = 1.
segment.twist = 0. * Units.deg
segment.root_chord_percent = 1.1 / 33.8
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.03
segment.append_airfoil(wing_airfoil)
wing.Segments.append(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# CG locations are approximate
# Masses from http://www.concordesst.com/fuelsys.html
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_9'
fuel_tank.mass_properties.center_of_gravity = np.array([[26.5, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 11096
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_10'
fuel_tank.mass_properties.center_of_gravity = np.array([[28.7, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 11943
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_1_and_4'
fuel_tank.mass_properties.center_of_gravity = np.array([[31.0, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 4198 + 4198
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_5_and_8'
fuel_tank.mass_properties.center_of_gravity = np.array([[32.9, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 7200 + 12838
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_6_and_7'
fuel_tank.mass_properties.center_of_gravity = np.array([[37.4, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 11587 + 7405
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_5A_and_7A'
fuel_tank.mass_properties.center_of_gravity = np.array([[40.2, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 2225 + 2225
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_2_and_3'
fuel_tank.mass_properties.center_of_gravity = np.array([[40.2, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 4570 + 4570
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
wing.Fuel_Tanks.append(fuel_tank)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 0.74 #
wing.sweeps.quarter_chord = 60 * Units.deg
wing.thickness_to_chord = 0.04
wing.taper = 0.14
wing.spans.projected = 6.0 #
wing.chords.root = 14.5
wing.total_length = 14.5
wing.chords.tip = 2.7
wing.chords.mean_aerodynamic = 8.66
wing.areas.reference = 33.91 #
wing.areas.wetted = 76.
wing.areas.exposed = 38.
wing.areas.affected = 33.91
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[42., 0, 1.]]
wing.aerodynamic_center = [50, 0, 0]
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.high_mach = True
wing.dynamic_pressure_ratio = 1.0
tail_airfoil = SUAVE.Components.Airfoils.Airfoil()
# This airfoil is not a true Concorde airfoil
tail_airfoil.coordinate_file = '../Vehicles/Airfoils/supersonic_tail.txt'
wing.append_airfoil(tail_airfoil)
# set root sweep with inner section
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_1'
segment.percent_span_location = 0.0
segment.twist = 0. * Units.deg
segment.root_chord_percent = 14.5 / 14.5
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 63. * Units.deg
segment.thickness_to_chord = 0.04
segment.append_airfoil(tail_airfoil)
wing.Segments.append(segment)
# set mid section start point
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_2'
segment.percent_span_location = 2.4 / (
6.0) + wing.Segments['section_1'].percent_span_location
segment.twist = 0. * Units.deg
segment.root_chord_percent = 7.5 / 14.5
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 40. * Units.deg
segment.thickness_to_chord = 0.04
segment.append_airfoil(tail_airfoil)
wing.Segments.append(segment)
# set tip
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'tip'
segment.percent_span_location = 1.
segment.twist = 0. * Units.deg
segment.root_chord_percent = 2.7 / 14.5
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = 0.04
segment.append_airfoil(tail_airfoil)
wing.Segments.append(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 4
fuselage.seat_pitch = 38. * Units.inches
fuselage.fineness.nose = 4.3
fuselage.fineness.tail = 6.4
fuselage.lengths.total = 61.66
fuselage.width = 2.88
fuselage.heights.maximum = 3.32 #
fuselage.heights.maximum = 3.32 #
fuselage.heights.at_quarter_length = 3.32 #
fuselage.heights.at_wing_root_quarter_chord = 3.32 #
fuselage.heights.at_three_quarters_length = 3.32 #
fuselage.areas.wetted = 442.
fuselage.areas.front_projected = 11.9
fuselage.effective_diameter = 3.1
fuselage.differential_pressure = 7.4e4 * Units.pascal # Maximum differential pressure
fuselage.OpenVSP_values = Data() # VSP uses degrees directly
fuselage.OpenVSP_values.nose = Data()
fuselage.OpenVSP_values.nose.top = Data()
fuselage.OpenVSP_values.nose.side = Data()
fuselage.OpenVSP_values.nose.top.angle = 20.0
fuselage.OpenVSP_values.nose.top.strength = 0.75
fuselage.OpenVSP_values.nose.side.angle = 20.0
fuselage.OpenVSP_values.nose.side.strength = 0.75
fuselage.OpenVSP_values.nose.TB_Sym = True
fuselage.OpenVSP_values.nose.z_pos = -.01
fuselage.OpenVSP_values.tail = Data()
fuselage.OpenVSP_values.tail.top = Data()
fuselage.OpenVSP_values.tail.side = Data()
fuselage.OpenVSP_values.tail.bottom = Data()
fuselage.OpenVSP_values.tail.top.angle = 0.0
fuselage.OpenVSP_values.tail.top.strength = 0.0
# CG locations are approximate
# Masses from http://www.concordesst.com/fuelsys.html
fuel_tank = SUAVE.Components.Energy.Storages.Fuel_Tanks.Fuel_Tank()
fuel_tank.tag = 'tank_11'
fuel_tank.mass_properties.center_of_gravity = np.array([[49.8, 0, 0]])
fuel_tank.mass_properties.fuel_mass_when_full = 10415
fuel_tank.fuel_type = SUAVE.Attributes.Propellants.Jet_A()
fuselage.Fuel_Tanks.append(fuel_tank)
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# the nacelle
# ------------------------------------------------------------------
nacelle = SUAVE.Components.Nacelles.Nacelle()
nacelle.diameter = 1.3
nacelle.tag = 'nacelle_L1'
nacelle.origin = [[36.56, 22, -1.9]]
nacelle.length = 12.0
nacelle.inlet_diameter = 1.1
nacelle.areas.wetted = 30.
vehicle.append_component(nacelle)
nacelle_2 = deepcopy(nacelle)
nacelle_2.tag = 'nacelle_2'
nacelle_2.origin = [[37., 5.3, -1.3]]
vehicle.append_component(nacelle_2)
nacelle_3 = deepcopy(nacelle)
nacelle_3.tag = 'nacelle_3'
nacelle_3.origin = [[37., -5.3, -1.3]]
vehicle.append_component(nacelle_3)
nacelle_4 = deepcopy(nacelle)
nacelle_4.tag = 'nacelle_4'
nacelle_4.origin = [[37., -6., -1.3]]
vehicle.append_component(nacelle_4)
# ------------------------------------------------------------------
# Turbojet Network
# ------------------------------------------------------------------
# instantiate the gas turbine network
turbojet = SUAVE.Components.Energy.Networks.Turbojet_Super()
turbojet.tag = 'turbojet'
# setup
turbojet.number_of_engines = 4.0
turbojet.engine_length = 12.0
turbojet.nacelle_diameter = 1.3
turbojet.inlet_diameter = 1.1
turbojet.areas = Data()
turbojet.areas.wetted = 120. / turbojet.number_of_engines
turbojet.origin = [[37., 6., -1.3], [37., 5.3, -1.3], [37., -5.3, -1.3],
[37., -6., -1.3]]
# working fluid
turbojet.working_fluid = SUAVE.Attributes.Gases.Air()
# ------------------------------------------------------------------
# Component 1 - Ram
# to convert freestream static to stagnation quantities
# instantiate
ram = SUAVE.Components.Energy.Converters.Ram()
ram.tag = 'ram'
# add to the network
turbojet.append(ram)
# ------------------------------------------------------------------
# Component 2 - Inlet Nozzle
# instantiate
inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
inlet_nozzle.tag = 'inlet_nozzle'
# setup
inlet_nozzle.polytropic_efficiency = 1.0
inlet_nozzle.pressure_ratio = 1.0
inlet_nozzle.pressure_recovery = 0.94
# add to network
turbojet.append(inlet_nozzle)
# ------------------------------------------------------------------
# Component 3 - Low Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'low_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.88
compressor.pressure_ratio = 3.1
# add to network
turbojet.append(compressor)
# ------------------------------------------------------------------
# Component 4 - High Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'high_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.88
compressor.pressure_ratio = 5.0
# add to network
turbojet.append(compressor)
# ------------------------------------------------------------------
# Component 5 - Low Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'low_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.89
# add to network
turbojet.append(turbine)
# ------------------------------------------------------------------
# Component 6 - High Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'high_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.87
# add to network
turbojet.append(turbine)
# ------------------------------------------------------------------
# Component 7 - Combustor
# instantiate
combustor = SUAVE.Components.Energy.Converters.Combustor()
combustor.tag = 'combustor'
# setup
combustor.efficiency = 0.94
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1440.
combustor.pressure_ratio = 0.92
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
# add to network
turbojet.append(combustor)
# ------------------------------------------------------------------
# Afterburner
# instantiate
afterburner = SUAVE.Components.Energy.Converters.Combustor()
afterburner.tag = 'afterburner'
# setup
afterburner.efficiency = 0.9
afterburner.alphac = 1.0
afterburner.turbine_inlet_temperature = 1500
afterburner.pressure_ratio = 1.0
afterburner.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
# add to network
turbojet.append(afterburner)
# ------------------------------------------------------------------
# Component 8 - Core Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Supersonic_Nozzle()
nozzle.tag = 'core_nozzle'
# setup
nozzle.pressure_recovery = 0.95
nozzle.pressure_ratio = 1.
# add to network
turbojet.append(nozzle)
# ------------------------------------------------------------------
#Component 10 : thrust (to compute the thrust)
thrust = SUAVE.Components.Energy.Processes.Thrust()
thrust.tag = 'compute_thrust'
#total design thrust (includes all the engines)
thrust.total_design = 40000. * Units.lbf
# Note: Sizing builds the network. It does not actually set the size of the turbojet
#design sizing conditions
altitude = 60000.0 * Units.ft
mach_number = 2.02
isa_deviation = 0.
# add to network
turbojet.thrust = thrust
#size the turbojet
turbojet_sizing(turbojet, mach_number, altitude)
# add gas turbine network gt_engine to the vehicle
vehicle.append_component(turbojet)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vsp_read_wing(wing_id, units_type='SI'):
"""This reads an OpenVSP wing vehicle geometry and writes it into a SUAVE wing format.
Assumptions:
1. OpenVSP wing is divided into segments ("XSecs" in VSP).
2. Written for OpenVSP 3.21.1
Source:
N/A
Inputs:
0. Pre-loaded VSP vehicle in memory, via vsp_read.
1. VSP 10-digit geom ID for wing.
2. units_type set to 'SI' (default) or 'Imperial'.
Outputs:
Writes SUAVE wing object, with these geometries, from VSP:
Wings.Wing. (* is all keys)
origin [m] in all three dimensions
spans.projected [m]
chords.root [m]
chords.tip [m]
aspect_ratio [-]
sweeps.quarter_chord [radians]
twists.root [radians]
twists.tip [radians]
thickness_to_chord [-]
dihedral [radians]
symmetric <boolean>
tag <string>
areas.exposed [m^2]
areas.reference [m^2]
areas.wetted [m^2]
Segments.
tag <string>
twist [radians]
percent_span_location [-] .1 is 10%
root_chord_percent [-] .1 is 10%
dihedral_outboard [radians]
sweeps.quarter_chord [radians]
thickness_to_chord [-]
airfoil <NACA 4-series, 6 series, or airfoil file>
Properties Used:
N/A
"""
# Check if this is vertical tail, this seems like a weird first step but it's necessary
# Get the initial rotation to get the dihedral angles
x_rot = vsp.GetParmVal(wing_id, 'X_Rotation', 'XForm')
if x_rot >= 70:
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.vertical = True
x_rot = (90 - x_rot) * Units.deg
else:
# Instantiate a wing
wing = SUAVE.Components.Wings.Wing()
# Set the units
if units_type == 'SI':
units_factor = Units.meter * 1.
else:
units_factor = Units.foot * 1.
# Apply a tag to the wing
if vsp.GetGeomName(wing_id):
tag = vsp.GetGeomName(wing_id)
tag = tag.translate(t_table)
wing.tag = tag
else:
wing.tag = 'winggeom'
# Top level wing parameters
# Wing origin
wing.origin[0][0] = vsp.GetParmVal(wing_id, 'X_Location',
'XForm') * units_factor
wing.origin[0][1] = vsp.GetParmVal(wing_id, 'Y_Location',
'XForm') * units_factor
wing.origin[0][2] = vsp.GetParmVal(wing_id, 'Z_Location',
'XForm') * units_factor
# Wing Symmetry
sym_planar = vsp.GetParmVal(wing_id, 'Sym_Planar_Flag', 'Sym')
sym_origin = vsp.GetParmVal(wing_id, 'Sym_Ancestor', 'Sym')
# Check for symmetry
if sym_planar == 2. and sym_origin == 1.: #origin at wing, not vehicle
wing.symmetric = True
else:
wing.symmetric = False
#More top level parameters
total_proj_span = vsp.GetParmVal(wing_id, 'TotalProjectedSpan',
'WingGeom') * units_factor
wing.aspect_ratio = vsp.GetParmVal(wing_id, 'TotalAR', 'WingGeom')
wing.areas.reference = vsp.GetParmVal(wing_id, 'TotalArea',
'WingGeom') * units_factor**2
wing.spans.projected = total_proj_span
# Check if this is a single segment wing
xsec_surf_id = vsp.GetXSecSurf(wing_id,
0) # This is how VSP stores surfaces.
x_sec_1 = vsp.GetXSec(xsec_surf_id, 1)
x_sec_1_span_parm = vsp.GetXSecParm(x_sec_1, 'Span')
x_sec_1_span = vsp.GetParmVal(x_sec_1_span_parm) * (
1 + wing.symmetric) * units_factor
if x_sec_1_span == wing.spans.projected:
single_seg = True
else:
single_seg = False
segment_num = vsp.GetNumXSec(
xsec_surf_id
) # Get number of wing segments (is one more than the VSP GUI shows).
x_sec = vsp.GetXSec(xsec_surf_id, 0)
chord_parm = vsp.GetXSecParm(x_sec, 'Root_Chord')
total_chord = vsp.GetParmVal(chord_parm)
span_sum = 0. # Non-projected.
proj_span_sum = 0. # Projected.
segment_spans = [None] * (segment_num) # Non-projected.
segment_dihedral = [None] * (segment_num)
segment_sweeps_quarter_chord = [None] * (segment_num)
# Check for wing segment *inside* fuselage, then skip XSec_0 to start at first exposed segment.
if total_chord == 1.:
start = 1
xsec_surf_id = vsp.GetXSecSurf(wing_id, 1)
x_sec = vsp.GetXSec(xsec_surf_id, 0)
chord_parm = vsp.GetXSecParm(x_sec, 'Tip_Chord')
root_chord = vsp.GetParmVal(chord_parm) * units_factor
else:
start = 0
root_chord = total_chord * units_factor
# -------------
# Wing segments
# -------------
if single_seg == False:
# Convert VSP XSecs to SUAVE segments. (Wing segments are defined by outboard sections in VSP, but inboard sections in SUAVE.)
for i in range(start, segment_num + 1):
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_' + str(i)
thick_cord = vsp.GetParmVal(wing_id, 'ThickChord',
'XSecCurve_' + str(i - 1))
segment.thickness_to_chord = thick_cord # Thick_cord stored for use in airfoil, below.
segment_root_chord = vsp.GetParmVal(
wing_id, 'Root_Chord', 'XSec_' + str(i)) * units_factor
segment.root_chord_percent = segment_root_chord / root_chord
segment.percent_span_location = proj_span_sum / (total_proj_span /
2)
segment.twist = vsp.GetParmVal(wing_id, 'Twist',
'XSec_' + str(i - 1)) * Units.deg
if i == start:
wing.thickness_to_chord = thick_cord
if i < segment_num: # This excludes the tip xsec, but we need a segment in SUAVE to store airfoil.
sweep = vsp.GetParmVal(wing_id, 'Sweep',
'XSec_' + str(i)) * Units.deg
sweep_loc = vsp.GetParmVal(wing_id, 'Sweep_Location',
'XSec_' + str(i))
AR = vsp.GetParmVal(wing_id, 'Aspect', 'XSec_' + str(i))
taper = vsp.GetParmVal(wing_id, 'Taper', 'XSec_' + str(i))
segment_sweeps_quarter_chord[i] = convert_sweep(
sweep, sweep_loc, 0.25, AR, taper)
segment.sweeps.quarter_chord = segment_sweeps_quarter_chord[
i] # Used again, below
# Used for dihedral computation, below.
segment_dihedral[i] = vsp.GetParmVal(
wing_id, 'Dihedral', 'XSec_' + str(i)) * Units.deg + x_rot
segment.dihedral_outboard = segment_dihedral[i]
segment_spans[i] = vsp.GetParmVal(
wing_id, 'Span', 'XSec_' + str(i)) * units_factor
proj_span_sum += segment_spans[i] * np.cos(segment_dihedral[i])
span_sum += segment_spans[i]
else:
segment.root_chord_percent = (vsp.GetParmVal(
wing_id, 'Tip_Chord',
'XSec_' + str(i - 1))) * units_factor / total_chord
# XSec airfoil
jj = i - 1 # Airfoil index i-1 because VSP airfoils and sections are one index off relative to SUAVE.
xsec_id = str(vsp.GetXSec(xsec_surf_id, jj))
airfoil = Airfoil()
if vsp.GetXSecShape(
xsec_id
) == vsp.XS_FOUR_SERIES: # XSec shape: NACA 4-series
camber = vsp.GetParmVal(wing_id, 'Camber',
'XSecCurve_' + str(jj))
if camber == 0.:
camber_loc = 0.
else:
camber_loc = vsp.GetParmVal(wing_id, 'CamberLoc',
'XSecCurve_' + str(jj))
airfoil.thickness_to_chord = thick_cord
camber_round = int(np.around(camber * 100))
camber_loc_round = int(np.around(camber_loc * 10))
thick_cord_round = int(np.around(thick_cord * 100))
airfoil.tag = 'NACA ' + str(camber_round) + str(
camber_loc_round) + str(thick_cord_round)
elif vsp.GetXSecShape(
xsec_id) == vsp.XS_SIX_SERIES: # XSec shape: NACA 6-series
thick_cord_round = int(np.around(thick_cord * 100))
a_value = vsp.GetParmVal(wing_id, 'A', 'XSecCurve_' + str(jj))
ideal_CL = int(
np.around(
vsp.GetParmVal(wing_id, 'IdealCl',
'XSecCurve_' + str(jj)) * 10))
series_vsp = int(
vsp.GetParmVal(wing_id, 'Series', 'XSecCurve_' + str(jj)))
series_dict = {
0: '63',
1: '64',
2: '65',
3: '66',
4: '67',
5: '63A',
6: '64A',
7: '65A'
} # VSP series values.
series = series_dict[series_vsp]
airfoil.tag = 'NACA ' + series + str(ideal_CL) + str(
thick_cord_round) + ' a=' + str(np.around(a_value, 1))
elif vsp.GetXSecShape(
xsec_id
) == vsp.XS_FILE_AIRFOIL: # XSec shape: 12 is type AF_FILE
airfoil.thickness_to_chord = thick_cord
airfoil.points = vsp.GetAirfoilCoordinates(
wing_id, float(jj / segment_num))
# VSP airfoil API calls get coordinates and write files with the final argument being the fraction of segment position, regardless of relative spans.
# (Write the root airfoil with final arg = 0. Write 4th airfoil of 5 segments with final arg = .8)
vsp.WriteSeligAirfoil(
str(wing.tag) + '_airfoil_XSec_' + str(jj) + '.dat',
wing_id, float(jj / segment_num))
airfoil.coordinate_file = 'str(wing.tag)' + '_airfoil_XSec_' + str(
jj) + '.dat'
airfoil.tag = 'AF_file'
segment.append_airfoil(airfoil)
wing.Segments.append(segment)
# Wing dihedral
proj_span_sum_alt = 0.
span_sum_alt = 0.
sweeps_sum = 0.
for ii in range(start, segment_num):
span_sum_alt += segment_spans[ii]
proj_span_sum_alt += segment_spans[ii] * np.cos(
segment_dihedral[ii]
) # Use projected span to find total wing dihedral.
sweeps_sum += segment_spans[ii] * np.tan(
segment_sweeps_quarter_chord[ii])
wing.dihedral = np.arccos(proj_span_sum_alt / span_sum_alt)
wing.sweeps.quarter_chord = -np.arctan(
sweeps_sum / span_sum_alt) # Minus sign makes it positive sweep.
# Add a tip segment, all values are zero except the tip chord
tc = vsp.GetParmVal(wing_id, 'Tip_Chord',
'XSec_' + str(segment_num - 1)) * units_factor
segment = SUAVE.Components.Wings.Segment()
segment.percent_span_location = 1.0
segment.root_chord_percent = tc / root_chord
# Chords
wing.chords.root = vsp.GetParmVal(wing_id, 'Tip_Chord',
'XSec_0') * units_factor
wing.chords.tip = tc
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Just double calculate and fix things:
wing = wing_segmented_planform(wing)
else:
# Single segment
# Get ID's
x_sec_1_dih_parm = vsp.GetXSecParm(x_sec_1, 'Dihedral')
x_sec_1_sweep_parm = vsp.GetXSecParm(x_sec_1, 'Sweep')
x_sec_1_sweep_loc_parm = vsp.GetXSecParm(x_sec_1, 'Sweep_Location')
x_sec_1_taper_parm = vsp.GetXSecParm(x_sec_1, 'Taper')
x_sec_1_rc_parm = vsp.GetXSecParm(x_sec_1, 'Root_Chord')
x_sec_1_tc_parm = vsp.GetXSecParm(x_sec_1, 'Tip_Chord')
# Calcs
sweep = vsp.GetParmVal(x_sec_1_sweep_parm) * Units.deg
sweep_loc = vsp.GetParmVal(x_sec_1_sweep_loc_parm)
taper = vsp.GetParmVal(x_sec_1_taper_parm)
c_4_sweep = convert_sweep(sweep, sweep_loc, 0.25, wing.aspect_ratio,
taper)
# Pull and pack
wing.sweeps.quarter_chord = c_4_sweep
wing.taper = taper
wing.dihedral = vsp.GetParmVal(x_sec_1_dih_parm) * Units.deg + x_rot
wing.chords.root = vsp.GetParmVal(x_sec_1_rc_parm) * units_factor
wing.chords.tip = vsp.GetParmVal(x_sec_1_tc_parm) * units_factor
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Just double calculate and fix things:
wing = wing_planform(wing)
# Twists
wing.twists.root = vsp.GetParmVal(wing_id, 'Twist', 'XSec_0') * Units.deg
wing.twists.tip = vsp.GetParmVal(
wing_id, 'Twist', 'XSec_' + str(segment_num - 1)) * Units.deg
return wing
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 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 read_vsp_wing(wing_id, units_type='SI',write_airfoil_file=True):
"""This reads an OpenVSP wing vehicle geometry and writes it into a SUAVE wing format.
Assumptions:
1. OpenVSP wing is divided into segments ("XSecs" in VSP).
2. Written for OpenVSP 3.21.1
Source:
N/A
Inputs:
1. VSP 10-digit geom ID for wing.
2. units_type set to 'SI' (default) or 'Imperial'.
Outputs:
Writes SUAVE wing object, with these geometries, from VSP:
Wings.Wing. (* is all keys)
origin [m] in all three dimensions
spans.projected [m]
chords.root [m]
chords.tip [m]
aspect_ratio [-]
sweeps.quarter_chord [radians]
twists.root [radians]
twists.tip [radians]
thickness_to_chord [-]
dihedral [radians]
symmetric <boolean>
tag <string>
areas.reference [m^2]
areas.wetted [m^2]
Segments.
tag <string>
twist [radians]
percent_span_location [-] .1 is 10%
root_chord_percent [-] .1 is 10%
dihedral_outboard [radians]
sweeps.quarter_chord [radians]
thickness_to_chord [-]
airfoil <NACA 4-series, 6 series, or airfoil file>
Properties Used:
N/A
"""
# Check if this is vertical tail, this seems like a weird first step but it's necessary
# Get the initial rotation to get the dihedral angles
x_rot = vsp.GetParmVal( wing_id,'X_Rotation','XForm')
if x_rot >=70:
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.vertical = True
x_rot = (90-x_rot) * Units.deg
else:
# Instantiate a wing
wing = SUAVE.Components.Wings.Wing()
x_rot = x_rot * Units.deg
# Set the units
if units_type == 'SI':
units_factor = Units.meter * 1.
elif units_type == 'imperial':
units_factor = Units.foot * 1.
elif units_type == 'inches':
units_factor = Units.inch * 1.
# Apply a tag to the wing
if vsp.GetGeomName(wing_id):
tag = vsp.GetGeomName(wing_id)
tag = tag.translate(t_table)
wing.tag = tag
else:
wing.tag = 'winggeom'
scaling = vsp.GetParmVal(wing_id, 'Scale', 'XForm')
units_factor = units_factor*scaling
# Top level wing parameters
# Wing origin
wing.origin[0][0] = vsp.GetParmVal(wing_id, 'X_Location', 'XForm') * units_factor
wing.origin[0][1] = vsp.GetParmVal(wing_id, 'Y_Location', 'XForm') * units_factor
wing.origin[0][2] = vsp.GetParmVal(wing_id, 'Z_Location', 'XForm') * units_factor
# Wing Symmetry
sym_planar = vsp.GetParmVal(wing_id, 'Sym_Planar_Flag', 'Sym')
sym_origin = vsp.GetParmVal(wing_id, 'Sym_Ancestor_Origin_Flag', 'Sym')
# Check for symmetry
if sym_planar == 2. and sym_origin == 1.: #origin at wing, not vehicle
wing.symmetric = True
else:
wing.symmetric = False
#More top level parameters
total_proj_span = vsp.GetParmVal(wing_id, 'TotalProjectedSpan', 'WingGeom') * units_factor
wing.aspect_ratio = vsp.GetParmVal(wing_id, 'TotalAR', 'WingGeom')
wing.areas.reference = vsp.GetParmVal(wing_id, 'TotalArea', 'WingGeom') * units_factor**2
wing.spans.projected = total_proj_span
# Check if this is a single segment wing
xsec_surf_id = vsp.GetXSecSurf(wing_id, 0) # This is how VSP stores surfaces.
x_sec_1 = vsp.GetXSec(xsec_surf_id, 1)
if vsp.GetNumXSec(xsec_surf_id) == 2:
single_seg = True
else:
single_seg = False
segment_num = vsp.GetNumXSec(xsec_surf_id) # Get number of segments
span_sum = 0. # Non-projected.
proj_span_sum = 0. # Projected.
segment_spans = [None] * (segment_num) # Non-projected.
segment_dihedral = [None] * (segment_num)
segment_sweeps_quarter_chord = [None] * (segment_num)
# Necessary wing segment definitions start at XSec_1 (XSec_0 exists mainly to hold the root airfoil)
xsec_surf_id = vsp.GetXSecSurf(wing_id, 0)
x_sec = vsp.GetXSec(xsec_surf_id, 1)
chord_parm = vsp.GetXSecParm(x_sec,'Root_Chord')
root_chord = vsp.GetParmVal(chord_parm) * units_factor
# -------------
# Wing segments
# -------------
if single_seg == False:
# Convert VSP XSecs to SUAVE segments. (Wing segments are defined by outboard sections in VSP, but inboard sections in SUAVE.)
for i in range(1, segment_num+1):
# XSec airfoil
jj = i-1 # Airfoil index i-1 because VSP airfoils and sections are one index off relative to SUAVE.
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_' + str(i)
thick_cord = vsp.GetParmVal(wing_id, 'ThickChord', 'XSecCurve_' + str(jj))
segment.thickness_to_chord = thick_cord # Thick_cord stored for use in airfoil, below.
if i!=segment_num:
segment_root_chord = vsp.GetParmVal(wing_id, 'Root_Chord', 'XSec_' + str(i)) * units_factor
else:
segment_root_chord = 0.0
segment.root_chord_percent = segment_root_chord / root_chord
segment.percent_span_location = proj_span_sum / (total_proj_span/(1+wing.symmetric))
segment.twist = vsp.GetParmVal(wing_id, 'Twist', 'XSec_' + str(jj)) * Units.deg
if i==1:
wing.thickness_to_chord = thick_cord
if i < segment_num: # This excludes the tip xsec, but we need a segment in SUAVE to store airfoil.
sweep = vsp.GetParmVal(wing_id, 'Sweep', 'XSec_' + str(i)) * Units.deg
sweep_loc = vsp.GetParmVal(wing_id, 'Sweep_Location', 'XSec_' + str(i))
AR = 2*vsp.GetParmVal(wing_id, 'Aspect', 'XSec_' + str(i))
taper = vsp.GetParmVal(wing_id, 'Taper', 'XSec_' + str(i))
segment_sweeps_quarter_chord[i] = convert_sweep(sweep,sweep_loc,0.25,AR,taper)
segment.sweeps.quarter_chord = segment_sweeps_quarter_chord[i] # Used again, below
# Used for dihedral computation, below.
segment_dihedral[i] = vsp.GetParmVal(wing_id, 'Dihedral', 'XSec_' + str(i)) * Units.deg + x_rot
segment.dihedral_outboard = segment_dihedral[i]
segment_spans[i] = vsp.GetParmVal(wing_id, 'Span', 'XSec_' + str(i)) * units_factor
proj_span_sum += segment_spans[i] * np.cos(segment_dihedral[i])
span_sum += segment_spans[i]
else:
segment.root_chord_percent = (vsp.GetParmVal(wing_id, 'Tip_Chord', 'XSec_' + str(i-1))) * units_factor /root_chord
xsec_id = str(vsp.GetXSec(xsec_surf_id, jj))
airfoil = Airfoil()
if vsp.GetXSecShape(xsec_id) == vsp.XS_FOUR_SERIES: # XSec shape: NACA 4-series
camber = vsp.GetParmVal(wing_id, 'Camber', 'XSecCurve_' + str(jj))
if camber == 0.:
camber_loc = 0.
else:
camber_loc = vsp.GetParmVal(wing_id, 'CamberLoc', 'XSecCurve_' + str(jj))
airfoil.thickness_to_chord = thick_cord
camber_round = int(np.around(camber*100))
camber_loc_round = int(np.around(camber_loc*10))
thick_cord_round = int(np.around(thick_cord*100))
airfoil.tag = 'NACA ' + str(camber_round) + str(camber_loc_round) + str(thick_cord_round)
elif vsp.GetXSecShape(xsec_id) == vsp.XS_SIX_SERIES: # XSec shape: NACA 6-series
thick_cord_round = int(np.around(thick_cord*100))
a_value = vsp.GetParmVal(wing_id, 'A', 'XSecCurve_' + str(jj))
ideal_CL = int(np.around(vsp.GetParmVal(wing_id, 'IdealCl', 'XSecCurve_' + str(jj))*10))
series_vsp = int(vsp.GetParmVal(wing_id, 'Series', 'XSecCurve_' + str(jj)))
series_dict = {0:'63',1:'64',2:'65',3:'66',4:'67',5:'63A',6:'64A',7:'65A'} # VSP series values.
series = series_dict[series_vsp]
airfoil.tag = 'NACA ' + series + str(ideal_CL) + str(thick_cord_round) + ' a=' + str(np.around(a_value,1))
elif vsp.GetXSecShape(xsec_id) == vsp.XS_FILE_AIRFOIL: # XSec shape: 12 is type AF_FILE
airfoil.thickness_to_chord = thick_cord
# VSP airfoil API calls get coordinates and write files with the final argument being the fraction of segment position, regardless of relative spans.
# (Write the root airfoil with final arg = 0. Write 4th airfoil of 5 segments with final arg = .8)
if write_airfoil_file==True:
vsp.WriteSeligAirfoil(str(wing.tag) + '_airfoil_XSec_' + str(jj) +'.dat', wing_id, float(jj/segment_num))
airfoil.coordinate_file = str(wing.tag) + '_airfoil_XSec_' + str(jj) +'.dat'
airfoil.tag = 'airfoil'
segment.append_airfoil(airfoil)
wing.Segments.append(segment)
# Wing dihedral
proj_span_sum_alt = 0.
span_sum_alt = 0.
sweeps_sum = 0.
for ii in range(1, segment_num):
span_sum_alt += segment_spans[ii]
proj_span_sum_alt += segment_spans[ii] * np.cos(segment_dihedral[ii]) # Use projected span to find total wing dihedral.
sweeps_sum += segment_spans[ii] * np.tan(segment_sweeps_quarter_chord[ii])
wing.dihedral = np.arccos(proj_span_sum_alt / span_sum_alt)
wing.sweeps.quarter_chord = -np.arctan(sweeps_sum / span_sum_alt) # Minus sign makes it positive sweep.
# Add a tip segment, all values are zero except the tip chord
tc = vsp.GetParmVal(wing_id, 'Tip_Chord', 'XSec_' + str(segment_num-1)) * units_factor
# Chords
wing.chords.root = vsp.GetParmVal(wing_id, 'Tip_Chord', 'XSec_0') * units_factor
wing.chords.tip = tc
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Just double calculate and fix things:
wing = wing_segmented_planform(wing)
else:
# Single segment
# Get ID's
x_sec_1_dih_parm = vsp.GetXSecParm(x_sec_1,'Dihedral')
x_sec_1_sweep_parm = vsp.GetXSecParm(x_sec_1,'Sweep')
x_sec_1_sweep_loc_parm = vsp.GetXSecParm(x_sec_1,'Sweep_Location')
x_sec_1_taper_parm = vsp.GetXSecParm(x_sec_1,'Taper')
x_sec_1_rc_parm = vsp.GetXSecParm(x_sec_1,'Root_Chord')
x_sec_1_tc_parm = vsp.GetXSecParm(x_sec_1,'Tip_Chord')
x_sec_1_t_parm = vsp.GetXSecParm(x_sec_1,'ThickChord')
# Calcs
sweep = vsp.GetParmVal(x_sec_1_sweep_parm) * Units.deg
sweep_loc = vsp.GetParmVal(x_sec_1_sweep_loc_parm)
taper = vsp.GetParmVal(x_sec_1_taper_parm)
c_4_sweep = convert_sweep(sweep,sweep_loc,0.25,wing.aspect_ratio,taper)
# Pull and pack
wing.sweeps.quarter_chord = c_4_sweep
wing.taper = taper
wing.dihedral = vsp.GetParmVal(x_sec_1_dih_parm) * Units.deg + x_rot
wing.chords.root = vsp.GetParmVal(x_sec_1_rc_parm)* units_factor
wing.chords.tip = vsp.GetParmVal(x_sec_1_tc_parm) * units_factor
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
wing.thickness_to_chord = vsp.GetParmVal(x_sec_1_t_parm)
# Just double calculate and fix things:
wing = wing_planform(wing)
# Twists
wing.twists.root = vsp.GetParmVal(wing_id, 'Twist', 'XSec_0') * Units.deg
wing.twists.tip = vsp.GetParmVal(wing_id, 'Twist', 'XSec_' + str(segment_num-1)) * Units.deg
# check if control surface (sub surfaces) are defined
tags = []
LE_flags = []
span_fraction_starts = []
span_fraction_ends = []
chord_fractions = []
num_cs = vsp.GetNumSubSurf(wing_id)
# loop through wing and get all control surface parameters
for cs_idx in range(num_cs):
cs_id = vsp.GetSubSurf(wing_id,cs_idx)
param_names = vsp.GetSubSurfParmIDs(cs_id)
tags.append(vsp.GetSubSurfName(cs_id))
for p_idx in range(len(param_names)):
if 'LE_Flag' == vsp.GetParmName(param_names[p_idx]):
LE_flags.append(vsp.GetParmVal(param_names[p_idx]))
if 'UStart' == vsp.GetParmName(param_names[p_idx]):
span_fraction_starts.append(vsp.GetParmVal(param_names[p_idx]))
if 'UEnd' == vsp.GetParmName(param_names[p_idx]):
span_fraction_ends.append(vsp.GetParmVal(param_names[p_idx]))
if 'Length_C_Start' == vsp.GetParmName(param_names[p_idx]):
chord_fractions.append(vsp.GetParmVal(param_names[p_idx]))
# assign control surface parameters to wings. Outer most control surface on main/horizontal wing is assigned a aileron
for cs_idx in range(num_cs):
aileron_present = False
if num_cs > 1:
aileron_loc = np.argmax(np.array(span_fraction_starts))
if cs_idx == aileron_loc:
aileron_present = True
if LE_flags[cs_idx] == 1.0:
CS = SUAVE.Components.Wings.Control_Surfaces.Slat()
else:
if wing.vertical == True:
CS = SUAVE.Components.Wings.Control_Surfaces.Rudder()
else:
if aileron_present:
CS = SUAVE.Components.Wings.Control_Surfaces.Aileron()
else:
CS = SUAVE.Components.Wings.Control_Surfaces.Flap()
CS.tag = tags[cs_idx]
CS.span_fraction_start = span_fraction_starts[cs_idx]*3 - 1
CS.span_fraction_end = span_fraction_ends[cs_idx]*3 - 1
CS.chord_fraction = chord_fractions[cs_idx]
CS.span = (CS.span_fraction_end - CS.span_fraction_start)*wing.spans.projected
wing.append_control_surface(CS)
return wing
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Boeing_737800'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 79015.8 # kg
vehicle.mass_properties.takeoff = 79015.8 # kg
vehicle.mass_properties.operating_empty = 62746.4 # kg
vehicle.mass_properties.takeoff = 79015.8 # kg
vehicle.mass_properties.max_zero_fuel = 62732.0 # kg
vehicle.mass_properties.cargo = 10000. * Units.kilogram
vehicle.mass_properties.center_of_gravity = [[
15.30987849, 0., -0.48023939
]]
vehicle.mass_properties.moments_of_inertia.tensor = [[
3173074.17, 0, 28752.77565
], [0, 3019041.443, 0], [0, 0, 5730017.433]] # estimated, not correct
vehicle.design_mach_number = 0.78
vehicle.design_range = 3582 * Units.miles
vehicle.design_cruise_alt = 35000.0 * Units.ft
# envelope properties
vehicle.envelope.ultimate_load = 3.75
vehicle.envelope.limit_load = 1.5
# basic parameters
vehicle.reference_area = 124.862
vehicle.passengers = 170
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 10.18
wing.sweeps.quarter_chord = 25 * Units.deg
wing.thickness_to_chord = 0.1
wing.taper = 0.1
wing.spans.projected = 34.32
wing.chords.root = 7.760 * Units.meter
wing.chords.tip = 0.782 * Units.meter
wing.chords.mean_aerodynamic = 4.235 * Units.meter
wing.areas.reference = 124.862
wing.areas.wetted = 225.08
wing.twists.root = 4.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[13.61, 0, -0.93]]
wing.aerodynamic_center = [0, 0, 0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
# Wing Segments
root_airfoil = SUAVE.Components.Airfoils.Airfoil()
root_airfoil.coordinate_file = '../Vehicles/Airfoils/B737a.txt'
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Root'
segment.percent_span_location = 0.0
segment.twist = 4. * Units.deg
segment.root_chord_percent = 1.
segment.thickness_to_chord = 0.1
segment.dihedral_outboard = 2.5 * Units.degrees
segment.sweeps.quarter_chord = 28.225 * Units.degrees
segment.thickness_to_chord = .1
segment.append_airfoil(root_airfoil)
wing.append_segment(segment)
yehudi_airfoil = SUAVE.Components.Airfoils.Airfoil()
yehudi_airfoil.coordinate_file = '../Vehicles/Airfoils/B737b.txt'
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Yehudi'
segment.percent_span_location = 0.324
segment.twist = 0.047193 * Units.deg
segment.root_chord_percent = 0.5
segment.thickness_to_chord = 0.1
segment.dihedral_outboard = 5.5 * Units.degrees
segment.sweeps.quarter_chord = 25. * Units.degrees
segment.thickness_to_chord = .1
segment.append_airfoil(yehudi_airfoil)
wing.append_segment(segment)
mid_airfoil = SUAVE.Components.Airfoils.Airfoil()
mid_airfoil.coordinate_file = '../Vehicles/Airfoils/B737c.txt'
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_2'
segment.percent_span_location = 0.963
segment.twist = 0.00258 * Units.deg
segment.root_chord_percent = 0.220
segment.thickness_to_chord = 0.1
segment.dihedral_outboard = 5.5 * Units.degrees
segment.sweeps.quarter_chord = 56.75 * Units.degrees
segment.thickness_to_chord = .1
segment.append_airfoil(mid_airfoil)
wing.append_segment(segment)
tip_airfoil = SUAVE.Components.Airfoils.Airfoil()
tip_airfoil.coordinate_file = '../Vehicles/Airfoils/B737d.txt'
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Tip'
segment.percent_span_location = 1.
segment.twist = 0. * Units.degrees
segment.root_chord_percent = 0.10077
segment.thickness_to_chord = 0.1
segment.dihedral_outboard = 0.
segment.sweeps.quarter_chord = 0.
segment.thickness_to_chord = .1
segment.append_airfoil(tip_airfoil)
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# control surfaces -------------------------------------------
slat = SUAVE.Components.Wings.Control_Surfaces.Slat()
slat.tag = 'slat'
slat.span_fraction_start = 0.2
slat.span_fraction_end = 0.963
slat.deflection = 0.0 * Units.degrees
slat.chord_fraction = 0.075
wing.append_control_surface(slat)
flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
flap.tag = 'flap'
flap.span_fraction_start = 0.2
flap.span_fraction_end = 0.7
flap.deflection = 0.0 * Units.degrees
flap.configuration_type = 'double_slotted'
flap.chord_fraction = 0.30
wing.append_control_surface(flap)
aileron = SUAVE.Components.Wings.Control_Surfaces.Aileron()
aileron.tag = 'aileron'
aileron.span_fraction_start = 0.7
aileron.span_fraction_end = 0.963
aileron.deflection = 0.0 * Units.degrees
aileron.chord_fraction = 0.16
wing.append_control_surface(aileron)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Horizontal_Tail()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 4.99
wing.sweeps.quarter_chord = 28.2250 * Units.deg
wing.thickness_to_chord = 0.08
wing.taper = 0.3333
wing.spans.projected = 14.4
wing.chords.root = 4.2731
wing.chords.tip = 1.4243
wing.chords.mean_aerodynamic = 8.0
wing.areas.reference = 41.49
wing.areas.exposed = 59.354 # Exposed area of the horizontal tail
wing.areas.wetted = 71.81 # Wetted area of the horizontal tail
wing.twists.root = 3.0 * Units.degrees
wing.twists.tip = 3.0 * Units.degrees
wing.origin = [[33.02, 0, 1.466]]
wing.aerodynamic_center = [0, 0, 0]
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
# Wing Segments
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'root_segment'
segment.percent_span_location = 0.0
segment.twist = 0. * Units.deg
segment.root_chord_percent = 1.0
segment.dihedral_outboard = 8.63 * Units.degrees
segment.sweeps.quarter_chord = 28.2250 * Units.degrees
segment.thickness_to_chord = .1
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'tip_segment'
segment.percent_span_location = 1.
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.3333
segment.dihedral_outboard = 0 * Units.degrees
segment.sweeps.quarter_chord = 0 * Units.degrees
segment.thickness_to_chord = .1
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# control surfaces -------------------------------------------
elevator = SUAVE.Components.Wings.Control_Surfaces.Elevator()
elevator.tag = 'elevator'
elevator.span_fraction_start = 0.09
elevator.span_fraction_end = 0.92
elevator.deflection = 0.0 * Units.deg
elevator.chord_fraction = 0.3
wing.append_control_surface(elevator)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 1.98865
wing.sweeps.quarter_chord = 31.2 * Units.deg
wing.thickness_to_chord = 0.08
wing.taper = 0.1183
wing.spans.projected = 8.33
wing.total_length = wing.spans.projected
wing.chords.root = 10.1
wing.chords.tip = 1.20
wing.chords.mean_aerodynamic = 4.0
wing.areas.reference = 34.89
wing.areas.wetted = 57.25
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [[26.944, 0, 1.54]]
wing.aerodynamic_center = [0, 0, 0]
wing.vertical = True
wing.symmetric = False
wing.t_tail = 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 = 0. * Units.deg
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0 * Units.degrees
segment.sweeps.quarter_chord = 61.485 * Units.degrees
segment.thickness_to_chord = .1
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'segment_1'
segment.percent_span_location = 0.2962
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.45
segment.dihedral_outboard = 0. * Units.degrees
segment.sweeps.quarter_chord = 31.2 * Units.degrees
segment.thickness_to_chord = .1
wing.append_segment(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'segment_2'
segment.percent_span_location = 1.0
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.1183
segment.dihedral_outboard = 0.0 * Units.degrees
segment.sweeps.quarter_chord = 0.0
segment.thickness_to_chord = .1
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.number_coach_seats = vehicle.passengers
fuselage.seats_abreast = 6
fuselage.seat_pitch = 31. * Units.inches
fuselage.fineness.nose = 1.6
fuselage.fineness.tail = 2.
fuselage.lengths.nose = 6.4
fuselage.lengths.tail = 8.0
fuselage.lengths.cabin = 28.85
fuselage.lengths.total = 38.02
fuselage.lengths.fore_space = 6.
fuselage.lengths.aft_space = 5.
fuselage.width = 3.74
fuselage.heights.maximum = 3.74
fuselage.heights.at_quarter_length = 3.74
fuselage.heights.at_three_quarters_length = 3.65
fuselage.heights.at_wing_root_quarter_chord = 3.74
fuselage.areas.side_projected = 142.1948
fuselage.areas.wetted = 385.51
fuselage.areas.front_projected = 12.57
fuselage.effective_diameter = 3.74
fuselage.differential_pressure = 5.0e4 * Units.pascal # Maximum differential pressure
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.0000
segment.percent_z_location = -0.00144
segment.height = 0.0100
segment.width = 0.0100
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.00576
segment.percent_z_location = -0.00144
segment.height = 0.7500
segment.width = 0.6500
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.02017
segment.percent_z_location = 0.00000
segment.height = 1.52783
segment.width = 1.20043
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.03170
segment.percent_z_location = 0.00000
segment.height = 1.96435
segment.width = 1.52783
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 0.04899
segment.percent_z_location = 0.00431
segment.height = 2.72826
segment.width = 1.96435
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 0.07781
segment.percent_z_location = 0.00861
segment.height = 3.49217
segment.width = 2.61913
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_6'
segment.percent_x_location = 0.10375
segment.percent_z_location = 0.01005
segment.height = 3.70130
segment.width = 3.05565
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_7'
segment.percent_x_location = 0.16427
segment.percent_z_location = 0.01148
segment.height = 3.92870
segment.width = 3.71043
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_8'
segment.percent_x_location = 0.22478
segment.percent_z_location = 0.01148
segment.height = 3.92870
segment.width = 3.92870
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_9'
segment.percent_x_location = 0.69164
segment.percent_z_location = 0.01292
segment.height = 3.81957
segment.width = 3.81957
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_10'
segment.percent_x_location = 0.71758
segment.percent_z_location = 0.01292
segment.height = 3.81957
segment.width = 3.81957
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_11'
segment.percent_x_location = 0.78098
segment.percent_z_location = 0.01722
segment.height = 3.49217
segment.width = 3.71043
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_12'
segment.percent_x_location = 0.85303
segment.percent_z_location = 0.02296
segment.height = 3.05565
segment.width = 3.16478
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_13'
segment.percent_x_location = 0.91931
segment.percent_z_location = 0.03157
segment.height = 2.40087
segment.width = 1.96435
fuselage.Segments.append(segment)
# Segment
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_14'
segment.percent_x_location = 1.00
segment.percent_z_location = 0.04593
segment.height = 1.09130
segment.width = 0.21826
fuselage.Segments.append(segment)
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Nacelles
# ------------------------------------------------------------------
nacelle = SUAVE.Components.Nacelles.Nacelle()
nacelle.tag = 'nacelle_1'
nacelle.length = 2.71
nacelle.inlet_diameter = 1.90
nacelle.diameter = 2.05
nacelle.areas.wetted = 1.1 * np.pi * nacelle.diameter * nacelle.length
nacelle.origin = [[13.72, -4.86, -1.9]]
nacelle.flow_through = True
nacelle.Airfoil.naca_4_series_airfoil = '2410'
nacelle_2 = deepcopy(nacelle)
nacelle_2.tag = 'nacelle_2'
nacelle_2.origin = [[13.72, 4.86, -1.9]]
vehicle.append_component(nacelle)
vehicle.append_component(nacelle_2)
# ------------------------------------------------------------------
# Turbofan Network
# ------------------------------------------------------------------
#instantiate the gas turbine network
turbofan = SUAVE.Components.Energy.Networks.Turbofan()
turbofan.tag = 'turbofan'
# setup
turbofan.number_of_engines = 2.0
turbofan.bypass_ratio = 5.4
turbofan.engine_length = 2.71
# This origin is overwritten by compute_component_centers_of_gravity(base,compute_propulsor_origin=True)
turbofan.origin = [[13.72, 4.86, -1.9], [13.72, -4.86, -1.9]]
# working fluid
turbofan.working_fluid = SUAVE.Attributes.Gases.Air()
# ------------------------------------------------------------------
# Component 1 - Ram
# to convert freestream static to stagnation quantities
# instantiate
ram = SUAVE.Components.Energy.Converters.Ram()
ram.tag = 'ram'
# add to the network
turbofan.append(ram)
# ------------------------------------------------------------------
# Component 2 - Inlet Nozzle
# instantiate
inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
inlet_nozzle.tag = 'inlet_nozzle'
# setup
inlet_nozzle.polytropic_efficiency = 0.98
inlet_nozzle.pressure_ratio = 0.98
# add to network
turbofan.append(inlet_nozzle)
# ------------------------------------------------------------------
# Component 3 - Low Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'low_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.91
compressor.pressure_ratio = 1.14
# add to network
turbofan.append(compressor)
# ------------------------------------------------------------------
# Component 4 - High Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'high_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.91
compressor.pressure_ratio = 13.415
# add to network
turbofan.append(compressor)
# ------------------------------------------------------------------
# Component 5 - Low Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'low_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.93
# add to network
turbofan.append(turbine)
# ------------------------------------------------------------------
# Component 6 - High Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'high_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.93
# add to network
turbofan.append(turbine)
# ------------------------------------------------------------------
# Component 7 - Combustor
# instantiate
combustor = SUAVE.Components.Energy.Converters.Combustor()
combustor.tag = 'combustor'
# setup
combustor.efficiency = 0.99
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1450
combustor.pressure_ratio = 0.95
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
# add to network
turbofan.append(combustor)
# ------------------------------------------------------------------
# Component 8 - Core Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
nozzle.tag = 'core_nozzle'
# setup
nozzle.polytropic_efficiency = 0.95
nozzle.pressure_ratio = 0.99
# add to network
turbofan.append(nozzle)
# ------------------------------------------------------------------
# Component 9 - Fan Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
nozzle.tag = 'fan_nozzle'
# setup
nozzle.polytropic_efficiency = 0.95
nozzle.pressure_ratio = 0.99
# add to network
turbofan.append(nozzle)
# ------------------------------------------------------------------
# Component 10 - Fan
# instantiate
fan = SUAVE.Components.Energy.Converters.Fan()
fan.tag = 'fan'
# setup
fan.polytropic_efficiency = 0.93
fan.pressure_ratio = 1.7
# add to network
turbofan.append(fan)
# ------------------------------------------------------------------
#Component 10 : thrust (to compute the thrust)
thrust = SUAVE.Components.Energy.Processes.Thrust()
thrust.tag = 'compute_thrust'
#total design thrust (includes all the engines)
thrust.total_design = 2 * 24000. * Units.N #Newtons
#design sizing conditions
altitude = 35000.0 * Units.ft
mach_number = 0.78
isa_deviation = 0.
#Engine setup for noise module
# add to network
turbofan.thrust = thrust
turbofan.core_nozzle_diameter = 0.92
turbofan.fan_nozzle_diameter = 1.659
turbofan.engine_height = 0.5 #Engine centerline heigh above the ground plane
turbofan.exa = 1 #distance from fan face to fan exit/ fan diameter)
turbofan.plug_diameter = 0.1 #dimater of the engine plug
turbofan.geometry_xe = 1. # Geometry information for the installation effects function
turbofan.geometry_ye = 1. # Geometry information for the installation effects function
turbofan.geometry_Ce = 2. # Geometry information for the installation effects function
#size the turbofan
turbofan_sizing(turbofan, mach_number, altitude)
# add gas turbine network turbofan to the vehicle
vehicle.append_component(turbofan)
# ------------------------------------------------------------------
# Fuel
# ------------------------------------------------------------------
fuel = SUAVE.Components.Physical_Component()
vehicle.fuel = fuel
fuel.mass_properties.mass = vehicle.mass_properties.max_takeoff - vehicle.mass_properties.max_fuel
fuel.origin = vehicle.wings.main_wing.mass_properties.center_of_gravity
fuel.mass_properties.center_of_gravity = vehicle.wings.main_wing.aerodynamic_center
# ------------------------------------------------------------------
# Landing Gear
# ------------------------------------------------------------------
landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
landing_gear.tag = "main_landing_gear"
landing_gear.main_tire_diameter = 1.12000 * Units.m
landing_gear.nose_tire_diameter = 0.6858 * Units.m
landing_gear.main_strut_length = 1.8 * Units.m
landing_gear.nose_strut_length = 1.3 * Units.m
landing_gear.main_units = 1 #number of nose landing gear
landing_gear.nose_units = 1 #number of nose landing gear
landing_gear.main_wheels = 2 #number of wheels on the main landing gear
landing_gear.nose_wheels = 2 #number of wheels on the nose landing gear
vehicle.landing_gear = landing_gear
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Boeing_BWB_450'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 823000. * Units.lb
vehicle.mass_properties.takeoff = 823000. * Units.lb
vehicle.mass_properties.max_zero_fuel = 0.9 * vehicle.mass_properties.max_takeoff
vehicle.mass_properties.cargo = 00. * Units.kilogram
# envelope properties
vehicle.envelope.ultimate_load = 2.5
vehicle.envelope.limit_load = 1.5
# basic parameters
vehicle.reference_area = 7840. * 2 * Units.feet**2
vehicle.passengers = 450.
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 289.**2 / (7840. * 2)
wing.thickness_to_chord = 0.15
wing.taper = 0.0138
wing.spans.projected = 289.0 * Units.feet
wing.chords.root = 145.0 * Units.feet
wing.chords.tip = 3.5 * Units.feet
wing.chords.mean_aerodynamic = 80. * Units.feet
wing.areas.reference = 7840. * 2 * Units.feet**2
wing.sweeps.quarter_chord = 33. * Units.degrees
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.dihedral = 2.5 * Units.degrees
wing.origin = [[0., 0., 0]]
wing.aerodynamic_center = [0, 0, 0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_1'
segment.percent_span_location = 0.0
segment.twist = 0. * Units.deg
segment.root_chord_percent = 1.
segment.dihedral_outboard = 0. * Units.degrees
segment.sweeps.quarter_chord = 40.0 * Units.degrees
segment.thickness_to_chord = 0.165
segment.vsp_mesh = Data()
segment.vsp_mesh.inner_radius = 4.
segment.vsp_mesh.outer_radius = 4.
segment.vsp_mesh.inner_length = .14
segment.vsp_mesh.outer_length = .14
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_2'
segment.percent_span_location = 0.052
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.921
segment.dihedral_outboard = 0. * Units.degrees
segment.sweeps.quarter_chord = 52.5 * Units.degrees
segment.thickness_to_chord = 0.167
segment.vsp_mesh = Data()
segment.vsp_mesh.inner_radius = 4.
segment.vsp_mesh.outer_radius = 4.
segment.vsp_mesh.inner_length = .14
segment.vsp_mesh.outer_length = .14
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_3'
segment.percent_span_location = 0.138
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.76
segment.dihedral_outboard = 1.85 * Units.degrees
segment.sweeps.quarter_chord = 36.9 * Units.degrees
segment.thickness_to_chord = 0.171
segment.vsp_mesh = Data()
segment.vsp_mesh.inner_radius = 4.
segment.vsp_mesh.outer_radius = 4.
segment.vsp_mesh.inner_length = .14
segment.vsp_mesh.outer_length = .14
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_4'
segment.percent_span_location = 0.221
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.624
segment.dihedral_outboard = 1.85 * Units.degrees
segment.sweeps.quarter_chord = 30.4 * Units.degrees
segment.thickness_to_chord = 0.175
segment.vsp_mesh = Data()
segment.vsp_mesh.inner_radius = 4.
segment.vsp_mesh.outer_radius = 2.8
segment.vsp_mesh.inner_length = .14
segment.vsp_mesh.outer_length = .14
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_5'
segment.percent_span_location = 0.457
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.313
segment.dihedral_outboard = 1.85 * Units.degrees
segment.sweeps.quarter_chord = 30.85 * Units.degrees
segment.thickness_to_chord = 0.118
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_6'
segment.percent_span_location = 0.568
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.197
segment.dihedral_outboard = 1.85 * Units.degrees
segment.sweeps.quarter_chord = 34.3 * Units.degrees
segment.thickness_to_chord = 0.10
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'section_7'
segment.percent_span_location = 0.97
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.086
segment.dihedral_outboard = 73. * Units.degrees
segment.sweeps.quarter_chord = 55. * Units.degrees
segment.thickness_to_chord = 0.10
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'tip'
segment.percent_span_location = 1
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.0241
segment.dihedral_outboard = 0. * Units.degrees
segment.sweeps.quarter_chord = 0. * Units.degrees
segment.thickness_to_chord = 0.10
wing.Segments.append(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Nacelle
# ------------------------------------------------------------------
nacelle = SUAVE.Components.Nacelles.Nacelle()
nacelle.diameter = 3.96 * Units.meters
nacelle.length = 289. * Units.inches
nacelle.tag = 'nacelle'
nacelle.origin = [[
133.0 * Units.feet, 25.0 * Units.feet, 6.5 * Units.feet
]]
nacelle.areas.wetted = nacelle.length * (2 * np.pi * nacelle.diameter / 2.)
nacelle_2 = deepcopy(nacelle)
nacelle_2.tag = 'nacelle_2'
nacelle_2.origin = [[
145.0 * Units.feet, 0.0 * Units.feet, 6.5 * Units.feet
]]
nacelle_3 = deepcopy(nacelle)
nacelle_3.tag = 'nacelle_3'
nacelle_3.origin = [[
133.0 * Units.feet, -25.0 * Units.feet, 6.5 * Units.feet
]]
vehicle.append_component(nacelle)
vehicle.append_component(nacelle_2)
vehicle.append_component(nacelle_3)
# ------------------------------------------------------------------
# Turbofan Network
# ------------------------------------------------------------------
#instantiate the gas turbine network
turbofan = SUAVE.Components.Energy.Networks.Turbofan()
turbofan.tag = 'turbofan1'
# setup
turbofan.number_of_engines = 3.0
turbofan.bypass_ratio = 8.1
turbofan.engine_length = 289. * Units.inches
turbofan.origin = [
[133.0 * Units.feet, 25.0 * Units.feet, 6.5 * Units.feet],
[145.0 * Units.feet, 0.0 * Units.feet, 6.5 * Units.feet],
[133.0 * Units.feet, -25.0 * Units.feet, 6.5 * Units.feet]
]
# working fluid
turbofan.working_fluid = SUAVE.Attributes.Gases.Air()
# ------------------------------------------------------------------
# Component 1 - Ram
# to convert freestream static to stagnation quantities
# instantiate
ram = SUAVE.Components.Energy.Converters.Ram()
ram.tag = 'ram'
# add to the network
turbofan.append(ram)
# ------------------------------------------------------------------
# Component 2 - Inlet Nozzle
# instantiate
inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
inlet_nozzle.tag = 'inlet_nozzle'
# setup
inlet_nozzle.polytropic_efficiency = 1.0
inlet_nozzle.pressure_ratio = 1.0
# add to network
turbofan.append(inlet_nozzle)
# ------------------------------------------------------------------
# Component 3 - Low Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'low_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.91
compressor.pressure_ratio = 1.1
# add to network
turbofan.append(compressor)
# ------------------------------------------------------------------
# Component 4 - High Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'high_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.91
compressor.pressure_ratio = 23.0
#compressor.hub_to_tip_ratio = 0.325
# add to network
turbofan.append(compressor)
# ------------------------------------------------------------------
# Component 5 - Low Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'low_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.93
# add to network
turbofan.append(turbine)
# ------------------------------------------------------------------
# Component 6 - High Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'high_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.93
# add to network
turbofan.append(turbine)
# ------------------------------------------------------------------
# Component 7 - Combustor
# instantiate
combustor = SUAVE.Components.Energy.Converters.Combustor()
combustor.tag = 'combustor'
# setup
combustor.efficiency = 1.0
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1592. * Units.kelvin
combustor.pressure_ratio = 0.95
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
# add to network
turbofan.append(combustor)
# ------------------------------------------------------------------
# Component 8 - Core Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
nozzle.tag = 'core_nozzle'
# setup
nozzle.polytropic_efficiency = 0.95
nozzle.pressure_ratio = 0.99
# add to network
turbofan.append(nozzle)
# ------------------------------------------------------------------
# Component 9 - Fan Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
nozzle.tag = 'fan_nozzle'
# setup
nozzle.polytropic_efficiency = 0.95
nozzle.pressure_ratio = 0.99
# add to network
turbofan.append(nozzle)
# ------------------------------------------------------------------
# Component 10 - Fan
# instantiate
fan = SUAVE.Components.Energy.Converters.Fan()
fan.tag = 'fan'
# setup
fan.polytropic_efficiency = 0.93
fan.pressure_ratio = 1.58
# add to network
turbofan.append(fan)
# ------------------------------------------------------------------
#Component 10 : thrust (to compute the thrust)
thrust = SUAVE.Components.Energy.Processes.Thrust()
thrust.tag = 'compute_thrust'
#total design thrust (includes all the engines)
thrust.total_design = 2.0 * 512000 * Units.N
thrust.bypass_ratio = 8.4
#design sizing conditions
altitude = 0. * Units.km
mach_number = 0.01
isa_deviation = 0.
# add to network
turbofan.thrust = thrust
#size the turbofan
turbofan_sizing(turbofan, mach_number, altitude)
#turbofan.size(mach_number,altitude)
#computing the engine length and diameter
for nac in vehicle.nacelles:
compute_turbofan_geometry(turbofan, nac)
vehicle.append_component(turbofan)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
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 = wing_segmented_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 = 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 = 3.0
prop.freestream_velocity = 150. * Units.knots
prop.angular_velocity = 2400. * Units.rpm
prop.tip_radius = 1.72 / 2
prop.hub_radius = 10. * Units.inches
prop.design_Cl = 0.8
prop.design_altitude = 9000. * Units.feet
prop.design_power = 98 * 0.65 * Units.hp # assume 65 BHP at cruise
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, number_of_airfoil_section_points=102)
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 = 400.
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, 0.95]]
motor.nominal_voltage = bat.max_voltage * 0.8
motor.propeller_radius = prop.tip_radius
motor.no_load_current = 0.1
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, 0.95]]
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
