def empty(vehicle,settings=None):
""" This is for an arbitrary aircraft configuration.
Assumptions:
N/A
Source:
N/A
Inputs:
N/A
Outputs:
output - a data dictionary with fields:
wt_payload - weight of the passengers plus baggage and paid cargo [kilograms]
wt_pax - weight of all the passengers [kilogram]
wt_bag - weight of all the baggage [kilogram]
wt_fuel - weight of the fuel carried [kilogram]
wt_empty - operating empty weight of the aircraft [kilograms]
Properties Used:
N/A
"""
# Unpack inputs
Nult = vehicle.envelope.ultimate_load
Nlim = vehicle.envelope.limit_load
TOW = vehicle.mass_properties.max_takeoff
wt_zf = vehicle.mass_properties.max_zero_fuel
wt_cargo = vehicle.mass_properties.cargo
num_pax = vehicle.passengers
ctrl_type = vehicle.systems.control
ac_type = vehicle.systems.accessories
S_gross_w = vehicle.reference_area
Wings = SUAVE.Components.Wings
Nets = SUAVE.Components.Energy.Networks
# Set the factors
if settings == None:
wt_factors = Data()
wt_factors.main_wing = 0.
wt_factors.empennage = 0.
wt_factors.fuselage = 0.
else:
wt_factors = settings.weight_reduction_factors
# Prime the totals
wt_main_wing = 0.0
wt_tail_horizontal = 0.0
wt_vtail_tot = 0.0
wt_propulsion = 0.0
wt_fuselage = 0.0
wt_rudder = 0.0
s_tail = 0.0
S_main_sum = 0.0
mac_main_w = 0.0
main_origin = np.array([0.,0.,0.])
# Work through the propulsions systems
for prop in vehicle.propulsors:
if isinstance(prop,Nets.Turbofan) or isinstance(prop,Nets.Turbojet_Super):
num_eng = prop.number_of_engines
thrust_sls = prop.sealevel_static_thrust
wt_engine_jet = Propulsion.engine_jet(thrust_sls)
wt_prop = Propulsion.integrated_propulsion(wt_engine_jet,num_eng)
if num_eng == 0.:
wt_prop = 0.
prop.mass_properties.mass = wt_prop
wt_propulsion += wt_prop
# Go through all the wings, except horizontal tail
for wing in vehicle.wings:
# Common Wing Parameters
S = wing.areas.reference
mac = wing.chords.mean_aerodynamic
b = wing.spans.projected
t_c = wing.thickness_to_chord
sweep = wing.sweeps.quarter_chord
# Main Wing
if isinstance(wing,Wings.Main_Wing):
# Unpack main wing specific parameters
lambda_w = wing.taper
area_fraction = S/S_gross_w
# Calculate the weights
rho = Aluminum().density
sigma = Aluminum().yield_tensile_strength
wt_wing = wing_main.wing_main(wing,Nult,TOW,wt_zf,rho,sigma,area_fraction)
# Apply weight factor
wt_wing = wt_wing*(1.-wt_factors.main_wing)
if np.isnan(wt_wing):
wt_wing = 0.
# Pack and sum
wing.mass_properties.mass = wt_wing
wt_main_wing += wt_wing
S_main_sum += S
mac_main_w += mac*S
main_origin += wing.origin[0][0]*S
# Vertical Tail
if isinstance(wing,Wings.Vertical_Tail):
# Unpack vertical tail specific parameters
t_tail = wing.t_tail
# Calculate the weights
tv = tail_vertical(S,Nult,b,TOW,t_c,sweep,S_gross_w,t_tail)
# Apply weight factor
wt_rud = tv.wt_rudder*(1.-wt_factors.empennage)
wt_vert_str = tv.wt_tail_vertical*(1.-wt_factors.empennage)
# See if there is a rudder attached, if not add one
if len(wing.control_surfaces)==0:
rudder = SUAVE.Components.Wings.Control_Surfaces.Rudder()
wing.append_control_surface(rudder)
else:
rudder = wing.control_surfaces.values()[0]
# Pack and sum
wing.mass_properties.mass = wt_vert_str
rudder.mass_properties.mass = wt_rudder
wt_vtail_tot += wt_vert_str + wt_rud
wt_rudder += wt_rud
s_tail += S
mac_main_w = mac_main_w/S_main_sum
main_origin = main_origin/S_main_sum
# Go through all the wings and find horizontal tails
for wing in vehicle.wings:
# Horizontal Tail
if isinstance(wing,Wings.Horizontal_Tail):
# Unpack horizontal tail specific parameters
h_tail_exposed = wing.areas.exposed / wing.areas.wetted
l_w2h = wing.origin[0][0] + wing.aerodynamic_center[0] - main_origin[0]
if np.isnan(mac_main_w):
mac_main_w = 0.
if np.isnan(l_w2h):
l_w2h = 0.
# Calculate the weights
wt_horiz = tail_horizontal(b,sweep,Nult,S,TOW,mac_main_w,mac,l_w2h,t_c, h_tail_exposed)
# Apply weight factor
wt_horiz = wt_horiz*(1.-wt_factors.empennage)
# Pack and sum
wing.mass_properties.mass = wt_horiz
wt_tail_horizontal += wt_horiz
s_tail += S
# Fuselages
for fuse in vehicle.fuselages:
# Unpack
S_fus = fuse.areas.wetted
diff_p_fus = fuse.differential_pressure
w_fus = fuse.width
h_fus = fuse.heights.maximum
l_fus = fuse.lengths.total
wing_c_r = vehicle.wings.main_wing.chords.root
if np.isnan(wing_c_r):
wing_c_r = 0.
#
wt_fuse = tube(S_fus,diff_p_fus,w_fus,h_fus,l_fus,Nlim,wt_zf,wt_main_wing,wt_propulsion,wing_c_r)
wt_fuse = wt_fuse*(1.-wt_factors.fuselage)
fuse.mass_properties.mass = wt_fuse
wt_fuselage += wt_fuse
# Landing Gear
wt_landing_gear = landing_gear.landing_gear(TOW)
# Systems
output_2 = systems(num_pax, ctrl_type, s_tail, S_gross_w, ac_type)
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_main_wing + wt_fuselage + wt_landing_gear + wt_propulsion + output_2.wt_systems + \
wt_tail_horizontal + wt_vtail_tot)
# packup outputs
output = payload.payload(TOW, wt_empty, num_pax,wt_cargo)
output.wing = wt_main_wing
output.fuselage = wt_fuselage
output.propulsion = wt_propulsion
output.main_gear = wt_landing_gear * 0.9
output.nose_gear = wt_landing_gear * 0.1
output.horizontal_tail = wt_tail_horizontal
output.vertical_tail = wt_vtail_tot
output.rudder = wt_rudder
output.systems = output_2.wt_systems
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_2.wt_flt_ctrl
output.systems_breakdown.apu = output_2.wt_apu
output.systems_breakdown.hydraulics = output_2.wt_hyd_pnu
output.systems_breakdown.instruments = output_2.wt_instruments
output.systems_breakdown.avionics = output_2.wt_avionics
output.systems_breakdown.optionals = output_2.wt_opitems
output.systems_breakdown.electrical = output_2.wt_elec
output.systems_breakdown.air_conditioner = output_2.wt_ac
output.systems_breakdown.furnish = output_2.wt_furnish
control_systems = SUAVE.Components.Physical_Component()
control_systems.tag = 'control_systems'
electrical_systems = SUAVE.Components.Physical_Component()
electrical_systems.tag = 'electrical_systems'
passengers = SUAVE.Components.Physical_Component()
passengers.tag = 'passengers'
furnishings = SUAVE.Components.Physical_Component()
furnishings.tag = 'furnishings'
air_conditioner = SUAVE.Components.Physical_Component()
air_conditioner.tag = 'air_conditioner'
fuel = SUAVE.Components.Physical_Component()
fuel.tag = 'fuel'
apu = SUAVE.Components.Physical_Component()
apu.tag = 'apu'
hydraulics = SUAVE.Components.Physical_Component()
hydraulics.tag = 'hydraulics'
optionals = SUAVE.Components.Physical_Component()
optionals.tag = 'optionals'
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
main_gear = SUAVE.Components.Landing_Gear.Main_Landing_Gear()
nose_gear = SUAVE.Components.Landing_Gear.Nose_Landing_Gear()
#assign output weights to objects
main_gear.mass_properties.mass = output.main_gear
nose_gear.mass_properties.mass = output.nose_gear
control_systems.mass_properties.mass = output.systems_breakdown.control_systems
electrical_systems.mass_properties.mass = output.systems_breakdown.electrical
passengers.mass_properties.mass = output.pax + output.bag
furnishings.mass_properties.mass = output.systems_breakdown.furnish
avionics.mass_properties.mass = output.systems_breakdown.avionics \
+ output.systems_breakdown.instruments
air_conditioner.mass_properties.mass = output.systems_breakdown.air_conditioner
fuel.mass_properties.mass = output.fuel
apu.mass_properties.mass = output.systems_breakdown.apu
hydraulics.mass_properties.mass = output.systems_breakdown.hydraulics
optionals.mass_properties.mass = output.systems_breakdown.optionals
#assign components to vehicle
vehicle.systems.control_systems = control_systems
vehicle.systems.electrical_systems = electrical_systems
vehicle.systems.avionics = avionics
vehicle.systems.furnishings = furnishings
vehicle.systems.passengers = passengers
vehicle.systems.air_conditioner = air_conditioner
vehicle.systems.fuel = fuel
vehicle.systems.apu = apu
vehicle.systems.hydraulics = hydraulics
vehicle.systems.optionals = optionals
vehicle.landing_gear.nose_landing_gear = nose_gear
vehicle.landing_gear.main_landing_gear = main_gear
return output
def empty(vehicle):
""" This is for a BWB aircraft configuration.
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
Source:
N/A
Inputs:
engine - a data dictionary with the fields:
thrust_sls - sea level static thrust of a single engine [Newtons]
wing - a data dictionary with the fields:
gross_area - wing gross area [meters**2]
span - span of the wing [meters]
taper - taper ratio of the wing [dimensionless]
t_c - thickness-to-chord ratio of the wing [dimensionless]
sweep - sweep angle of the wing [radians]
mac - mean aerodynamic chord of the wing [meters]
r_c - wing root chord [meters]
aircraft - a data dictionary with the fields:
Nult - ultimate load of the aircraft [dimensionless]
Nlim - limit load factor at zero fuel weight of the aircraft [dimensionless]
TOW - maximum takeoff weight of the aircraft [kilograms]
zfw - maximum zero fuel weight of the aircraft [kilograms]
num_eng - number of engines on the aircraft [dimensionless]
num_pax - number of passengers on the aircraft [dimensionless]
wt_cargo - weight of the bulk cargo being carried on the aircraft [kilograms]
num_seats - number of seats installed on the aircraft [dimensionless]
ctrl - specifies if the control system is "fully powered", "partially powered", or not powered [dimensionless]
ac - determines type of instruments, electronics, and operating items based on types:
"short-range", "medium-range", "long-range", "business", "cargo", "commuter", "sst" [dimensionless]
fuselage - a data dictionary with the fields:
area - fuselage wetted area [meters**2]
diff_p - Maximum fuselage pressure differential [Pascal]
width - width of the fuselage [meters]
height - height of the fuselage [meters]
length - length of the fuselage [meters]
Outputs:
output - a data dictionary with fields:
wt_payload - weight of the passengers plus baggage and paid cargo [kilograms]
wt_pax - weight of all the passengers [kilogram]
wt_bag - weight of all the baggage [kilogram]
wt_fuel - weight of the fuel carried [kilogram]
wt_empty - operating empty weight of the aircraft [kilograms]
Properties Used:
N/A
"""
# Unpack inputs
Nult = vehicle.envelope.ultimate_load
Nlim = vehicle.envelope.limit_load
TOW = vehicle.mass_properties.max_takeoff
wt_zf = vehicle.mass_properties.max_zero_fuel
wt_cargo = vehicle.mass_properties.cargo
num_pax = vehicle.passengers
ctrl_type = vehicle.systems.control
ac_type = vehicle.systems.accessories
num_seats = vehicle.passengers
bwb_aft_centerbody_area = vehicle.fuselages[
'fuselage_bwb'].aft_centerbody_area
bwb_aft_centerbody_taper = vehicle.fuselages[
'fuselage_bwb'].aft_centerbody_taper
bwb_cabin_area = vehicle.fuselages['fuselage_bwb'].cabin_area
propulsor_name = list(vehicle.propulsors.keys())[0] #obtain the key f
#for the propulsor for assignment purposes
propulsors = vehicle.propulsors[propulsor_name]
num_eng = propulsors.number_of_engines
if propulsor_name == 'turbofan' or propulsor_name == 'Turbofan':
# thrust_sls should be sea level static thrust. Using design thrust results in wrong propulsor
# weight estimation. Engine sizing should return this value.
# for now, using thrust_sls = design_thrust / 0.20, just for optimization evaluations
thrust_sls = propulsors.sealevel_static_thrust
wt_engine_jet = Propulsion.engine_jet(thrust_sls)
wt_propulsion = Propulsion.integrated_propulsion(
wt_engine_jet, num_eng)
propulsors.mass_properties.mass = wt_propulsion
else: #propulsor used is not a turbo_fan; assume mass_properties defined outside model
wt_propulsion = propulsors.mass_properties.mass
if wt_propulsion == 0:
warnings.warn(
"Propulsion mass= 0; there is no Engine Weight being added to the Configuration",
stacklevel=1)
S_gross_w = vehicle.reference_area
if 'main_wing' not in vehicle.wings:
wt_wing = 0.0
wing_c_r = 0.0
warnings.warn(
"There is no Wing Weight being added to the Configuration",
stacklevel=1)
else:
b = vehicle.wings['main_wing'].spans.projected
lambda_w = vehicle.wings['main_wing'].taper
t_c_w = vehicle.wings['main_wing'].thickness_to_chord
sweep_w = vehicle.wings['main_wing'].sweeps.quarter_chord
mac_w = vehicle.wings['main_wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['main_wing'].chords.root
S_h = vehicle.wings[
'main_wing'].areas.reference * 0.01 # control surface area on bwb
wt_wing = wing_main.wing_main(S_gross_w, b, lambda_w, t_c_w, sweep_w,
Nult, TOW, wt_zf)
vehicle.wings['main_wing'].mass_properties.mass = wt_wing
# Calculating Empty Weight of Aircraft
wt_landing_gear = landing_gear.landing_gear(TOW)
wt_cabin = cabin(bwb_cabin_area, TOW)
wt_aft_centerbody = aft_centerbody(num_eng, bwb_aft_centerbody_area,
bwb_aft_centerbody_taper, TOW)
output_2 = systems(num_seats, ctrl_type, S_h, S_gross_w, ac_type)
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_wing + wt_cabin + wt_aft_centerbody + wt_landing_gear +
wt_propulsion + output_2.wt_systems)
vehicle.fuselages['fuselage_bwb'].mass_properties.mass = wt_cabin
# packup outputs
output = payload.payload(TOW, wt_empty, num_pax, wt_cargo)
output.wing = wt_wing
output.fuselage = wt_cabin + wt_aft_centerbody
output.propulsion = wt_propulsion
output.landing_gear = wt_landing_gear
output.systems = output_2.wt_systems
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_2.wt_flt_ctrl
output.systems_breakdown.apu = output_2.wt_apu
output.systems_breakdown.hydraulics = output_2.wt_hyd_pnu
output.systems_breakdown.instruments = output_2.wt_instruments
output.systems_breakdown.avionics = output_2.wt_avionics
output.systems_breakdown.optionals = output_2.wt_opitems
output.systems_breakdown.electrical = output_2.wt_elec
output.systems_breakdown.air_conditioner = output_2.wt_ac
output.systems_breakdown.furnish = output_2.wt_furnish
#define weights components
try:
landing_gear_component = vehicle.landing_gear #landing gear previously defined
except AttributeError: # landing gear not defined
landing_gear_component = SUAVE.Components.Landing_Gear.Landing_Gear()
vehicle.landing_gear = landing_gear_component
control_systems = SUAVE.Components.Physical_Component()
electrical_systems = SUAVE.Components.Physical_Component()
passengers = SUAVE.Components.Physical_Component()
furnishings = SUAVE.Components.Physical_Component()
air_conditioner = SUAVE.Components.Physical_Component()
fuel = SUAVE.Components.Physical_Component()
apu = SUAVE.Components.Physical_Component()
hydraulics = SUAVE.Components.Physical_Component()
optionals = SUAVE.Components.Physical_Component()
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
#assign output weights to objects
landing_gear_component.mass_properties.mass = output.landing_gear
control_systems.mass_properties.mass = output.systems_breakdown.control_systems
electrical_systems.mass_properties.mass = output.systems_breakdown.electrical
passengers.mass_properties.mass = output.pax + output.bag
furnishings.mass_properties.mass = output.systems_breakdown.furnish
avionics.mass_properties.mass = output.systems_breakdown.avionics \
+ output.systems_breakdown.instruments
air_conditioner.mass_properties.mass = output.systems_breakdown.air_conditioner
fuel.mass_properties.mass = output.fuel
apu.mass_properties.mass = output.systems_breakdown.apu
hydraulics.mass_properties.mass = output.systems_breakdown.hydraulics
optionals.mass_properties.mass = output.systems_breakdown.optionals
#assign components to vehicle
vehicle.control_systems = control_systems
vehicle.electrical_systems = electrical_systems
vehicle.avionics = avionics
vehicle.furnishings = furnishings
vehicle.passenger_weights = passengers
vehicle.air_conditioner = air_conditioner
vehicle.fuel = fuel
vehicle.apu = apu
vehicle.hydraulics = hydraulics
vehicle.optionals = optionals
vehicle.landing_gear = landing_gear_component
return output
def empty(vehicle):
""" This is for a BWB aircraft configuration.
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
Source:
N/A
Inputs:
engine - a data dictionary with the fields:
thrust_sls - sea level static thrust of a single engine [Newtons]
wing - a data dictionary with the fields:
gross_area - wing gross area [meters**2]
span - span of the wing [meters]
taper - taper ratio of the wing [dimensionless]
t_c - thickness-to-chord ratio of the wing [dimensionless]
sweep - sweep angle of the wing [radians]
mac - mean aerodynamic chord of the wing [meters]
r_c - wing root chord [meters]
aircraft - a data dictionary with the fields:
Nult - ultimate load of the aircraft [dimensionless]
Nlim - limit load factor at zero fuel weight of the aircraft [dimensionless]
TOW - maximum takeoff weight of the aircraft [kilograms]
zfw - maximum zero fuel weight of the aircraft [kilograms]
num_eng - number of engines on the aircraft [dimensionless]
num_pax - number of passengers on the aircraft [dimensionless]
wt_cargo - weight of the bulk cargo being carried on the aircraft [kilograms]
num_seats - number of seats installed on the aircraft [dimensionless]
ctrl - specifies if the control system is "fully powered", "partially powered", or not powered [dimensionless]
ac - determines type of instruments, electronics, and operating items based on types:
"short-range", "medium-range", "long-range", "business", "cargo", "commuter", "sst" [dimensionless]
fuselage - a data dictionary with the fields:
area - fuselage wetted area [meters**2]
diff_p - Maximum fuselage pressure differential [Pascal]
width - width of the fuselage [meters]
height - height of the fuselage [meters]
length - length of the fuselage [meters]
Outputs:
output - a data dictionary with fields:
wt_payload - weight of the passengers plus baggage and paid cargo [kilograms]
wt_pax - weight of all the passengers [kilogram]
wt_bag - weight of all the baggage [kilogram]
wt_fuel - weight of the fuel carried [kilogram]
wt_empty - operating empty weight of the aircraft [kilograms]
Properties Used:
N/A
"""
# Unpack inputs
Nult = vehicle.envelope.ultimate_load
Nlim = vehicle.envelope.limit_load
TOW = vehicle.mass_properties.max_takeoff
wt_zf = vehicle.mass_properties.max_zero_fuel
wt_cargo = vehicle.mass_properties.cargo
num_pax = vehicle.passengers
ctrl_type = vehicle.systems.control
ac_type = vehicle.systems.accessories
num_seats = vehicle.passengers
bwb_aft_centerbody_area = vehicle.fuselages['fuselage_bwb'].aft_centerbody_area
bwb_aft_centerbody_taper = vehicle.fuselages['fuselage_bwb'].aft_centerbody_taper
bwb_cabin_area = vehicle.fuselages['fuselage_bwb'].cabin_area
propulsor_name = list(vehicle.propulsors.keys())[0] #obtain the key f
#for the propulsor for assignment purposes
propulsors = vehicle.propulsors[propulsor_name]
num_eng = propulsors.number_of_engines
if propulsor_name=='turbofan' or propulsor_name=='Turbofan':
# thrust_sls should be sea level static thrust. Using design thrust results in wrong propulsor
# weight estimation. Engine sizing should return this value.
# for now, using thrust_sls = design_thrust / 0.20, just for optimization evaluations
thrust_sls = propulsors.sealevel_static_thrust
wt_engine_jet = Propulsion.engine_jet(thrust_sls)
wt_propulsion = Propulsion.integrated_propulsion(wt_engine_jet,num_eng)
propulsors.mass_properties.mass = wt_propulsion
else: #propulsor used is not a turbo_fan; assume mass_properties defined outside model
wt_propulsion = propulsors.mass_properties.mass
if wt_propulsion==0:
warnings.warn("Propulsion mass= 0; there is no Engine Weight being added to the Configuration", stacklevel=1)
S_gross_w = vehicle.reference_area
if 'main_wing' not in vehicle.wings:
wt_wing = 0.0
wing_c_r = 0.0
warnings.warn("There is no Wing Weight being added to the Configuration", stacklevel=1)
else:
b = vehicle.wings['main_wing'].spans.projected
lambda_w = vehicle.wings['main_wing'].taper
t_c_w = vehicle.wings['main_wing'].thickness_to_chord
sweep_w = vehicle.wings['main_wing'].sweeps.quarter_chord
mac_w = vehicle.wings['main_wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['main_wing'].chords.root
S_h = vehicle.wings['main_wing'].areas.reference*0.01 # control surface area on bwb
wt_wing = wing_main.wing_main(S_gross_w,b,lambda_w,t_c_w,sweep_w,Nult,TOW,wt_zf)
vehicle.wings['main_wing'].mass_properties.mass = wt_wing
# Calculating Empty Weight of Aircraft
wt_landing_gear = landing_gear.landing_gear(TOW)
wt_cabin = cabin(bwb_cabin_area, TOW)
wt_aft_centerbody = aft_centerbody(num_eng, bwb_aft_centerbody_area, bwb_aft_centerbody_taper, TOW)
output_2 = systems(num_seats, ctrl_type, S_h , S_gross_w, ac_type)
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_wing + wt_cabin + wt_aft_centerbody + wt_landing_gear + wt_propulsion + output_2.wt_systems)
vehicle.fuselages['fuselage_bwb'].mass_properties.mass = wt_cabin
# packup outputs
output = payload.payload(TOW, wt_empty, num_pax,wt_cargo)
output.wing = wt_wing
output.fuselage = wt_cabin + wt_aft_centerbody
output.propulsion = wt_propulsion
output.landing_gear = wt_landing_gear
output.systems = output_2.wt_systems
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_2.wt_flt_ctrl
output.systems_breakdown.apu = output_2.wt_apu
output.systems_breakdown.hydraulics = output_2.wt_hyd_pnu
output.systems_breakdown.instruments = output_2.wt_instruments
output.systems_breakdown.avionics = output_2.wt_avionics
output.systems_breakdown.optionals = output_2.wt_opitems
output.systems_breakdown.electrical = output_2.wt_elec
output.systems_breakdown.air_conditioner = output_2.wt_ac
output.systems_breakdown.furnish = output_2.wt_furnish
#define weights components
try:
landing_gear_component=vehicle.landing_gear #landing gear previously defined
except AttributeError: # landing gear not defined
landing_gear_component=SUAVE.Components.Landing_Gear.Landing_Gear()
vehicle.landing_gear=landing_gear_component
control_systems = SUAVE.Components.Physical_Component()
electrical_systems = SUAVE.Components.Physical_Component()
passengers = SUAVE.Components.Physical_Component()
furnishings = SUAVE.Components.Physical_Component()
air_conditioner = SUAVE.Components.Physical_Component()
fuel = SUAVE.Components.Physical_Component()
apu = SUAVE.Components.Physical_Component()
hydraulics = SUAVE.Components.Physical_Component()
optionals = SUAVE.Components.Physical_Component()
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
#assign output weights to objects
landing_gear_component.mass_properties.mass = output.landing_gear
control_systems.mass_properties.mass = output.systems_breakdown.control_systems
electrical_systems.mass_properties.mass = output.systems_breakdown.electrical
passengers.mass_properties.mass = output.pax + output.bag
furnishings.mass_properties.mass = output.systems_breakdown.furnish
avionics.mass_properties.mass = output.systems_breakdown.avionics \
+ output.systems_breakdown.instruments
air_conditioner.mass_properties.mass = output.systems_breakdown.air_conditioner
fuel.mass_properties.mass = output.fuel
apu.mass_properties.mass = output.systems_breakdown.apu
hydraulics.mass_properties.mass = output.systems_breakdown.hydraulics
optionals.mass_properties.mass = output.systems_breakdown.optionals
#assign components to vehicle
vehicle.control_systems = control_systems
vehicle.electrical_systems = electrical_systems
vehicle.avionics = avionics
vehicle.furnishings = furnishings
vehicle.passenger_weights = passengers
vehicle.air_conditioner = air_conditioner
vehicle.fuel = fuel
vehicle.apu = apu
vehicle.hydraulics = hydraulics
vehicle.optionals = optionals
vehicle.landing_gear = landing_gear_component
return output
def empty(vehicle, settings=None):
""" This is for a standard Tube and Wing aircraft configuration.
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
Source:
N/A
Inputs:
engine - a data dictionary with the fields:
thrust_sls - sea level static thrust of a single engine [Newtons]
wing - a data dictionary with the fields:
gross_area - wing gross area [meters**2]
span - span of the wing [meters]
taper - taper ratio of the wing [dimensionless]
t_c - thickness-to-chord ratio of the wing [dimensionless]
sweep - sweep angle of the wing [radians]
mac - mean aerodynamic chord of the wing [meters]
r_c - wing root chord [meters]
aircraft - a data dictionary with the fields:
Nult - ultimate load of the aircraft [dimensionless]
Nlim - limit load factor at zero fuel weight of the aircraft [dimensionless]
TOW - maximum takeoff weight of the aircraft [kilograms]
zfw - maximum zero fuel weight of the aircraft [kilograms]
num_eng - number of engines on the aircraft [dimensionless]
num_pax - number of passengers on the aircraft [dimensionless]
wt_cargo - weight of the bulk cargo being carried on the aircraft [kilograms]
num_seats - number of seats installed on the aircraft [dimensionless]
ctrl - specifies if the control system is "fully powered", "partially powered", or not powered [dimensionless]
ac - determines type of instruments, electronics, and operating items based on types:
"short-range", "medium-range", "long-range", "business", "cargo", "commuter", "sst" [dimensionless]
w2h - tail length (distance from the airplane c.g. to the horizontal tail aerodynamic center) [meters]
fuselage - a data dictionary with the fields:
area - fuselage wetted area [meters**2]
diff_p - Maximum fuselage pressure differential [Pascal]
width - width of the fuselage [meters]
height - height of the fuselage [meters]
length - length of the fuselage [meters]
horizontal
area - area of the horizontal tail [meters**2]
span - span of the horizontal tail [meters]
sweep - sweep of the horizontal tail [radians]
mac - mean aerodynamic chord of the horizontal tail [meters]
t_c - thickness-to-chord ratio of the horizontal tail [dimensionless]
exposed - exposed area ratio for the horizontal tail [dimensionless]
vertical
area - area of the vertical tail [meters**2]
span - sweight = weight * Units.lbpan of the vertical [meters]
t_c - thickness-to-chord ratio of the vertical tail [dimensionless]
sweep - sweep angle of the vertical tail [radians]
t_tail - factor to determine if aircraft has a t-tail, "yes" [dimensionless]
settings.weight_reduction_factors.
main_wing [dimensionless] (.1 is a 10% weight reduction)
empennage [dimensionless] (.1 is a 10% weight reduction)
fuselage [dimensionless] (.1 is a 10% weight reduction)
Outputs:
output - a data dictionary with fields:
wt_payload - weight of the passengers plus baggage and paid cargo [kilograms]
wt_pax - weight of all the passengers [kilogram]
wt_bag - weight of all the baggage [kilogram]
wt_fuel - weight of the fuel carried [kilogram]
wt_empty - operating empty weight of the aircraft [kilograms]
Properties Used:
N/A
"""
# Unpack inputs
Nult = vehicle.envelope.ultimate_load
Nlim = vehicle.envelope.limit_load
TOW = vehicle.mass_properties.max_takeoff
wt_zf = vehicle.mass_properties.max_zero_fuel
num_pax = vehicle.passengers
wt_cargo = vehicle.mass_properties.cargo
num_seats = vehicle.fuselages['fuselage'].number_coach_seats
ctrl_type = vehicle.systems.control
ac_type = vehicle.systems.accessories
if settings == None:
wt_factors = Data()
wt_factors.main_wing = 0.
wt_factors.empennage = 0.
wt_factors.fuselage = 0.
else:
wt_factors = settings.weight_reduction_factors
propulsor_name = list(vehicle.propulsors.keys())[
0] #obtain the key for the propulsor for assignment purposes
propulsors = vehicle.propulsors[propulsor_name]
num_eng = propulsors.number_of_engines
if propulsor_name == 'turbofan' or propulsor_name == 'Turbofan':
# thrust_sls should be sea level static thrust. Using design thrust results in wrong propulsor
# weight estimation. Engine sizing should return this value.
# for now, using thrust_sls = design_thrust / 0.20, just for optimization evaluations
thrust_sls = propulsors.sealevel_static_thrust
wt_engine_jet = Propulsion.engine_jet(thrust_sls)
wt_propulsion = Propulsion.integrated_propulsion(
wt_engine_jet, num_eng)
propulsors.mass_properties.mass = wt_propulsion
else: #propulsor used is not a turbo_fan; assume mass_properties defined outside model
wt_propulsion = propulsors.mass_properties.mass
if wt_propulsion == 0:
warnings.warn(
"Propulsion mass= 0 ;e there is no Engine Weight being added to the Configuration",
stacklevel=1)
S_gross_w = vehicle.reference_area
if 'main_wing' not in vehicle.wings:
wt_wing = 0.0
wing_c_r = 0.0
warnings.warn(
"There is no Wing Weight being added to the Configuration",
stacklevel=1)
else:
b = vehicle.wings['main_wing'].spans.projected
lambda_w = vehicle.wings['main_wing'].taper
t_c_w = vehicle.wings['main_wing'].thickness_to_chord
sweep_w = vehicle.wings['main_wing'].sweeps.quarter_chord
mac_w = vehicle.wings['main_wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['main_wing'].chords.root
wt_wing = wing_main.wing_main(S_gross_w, b, lambda_w, t_c_w, sweep_w,
Nult, TOW, wt_zf)
wt_wing = wt_wing * (1. - wt_factors.main_wing)
vehicle.wings['main_wing'].mass_properties.mass = wt_wing
S_fus = vehicle.fuselages['fuselage'].areas.wetted
diff_p_fus = vehicle.fuselages['fuselage'].differential_pressure
w_fus = vehicle.fuselages['fuselage'].width
h_fus = vehicle.fuselages['fuselage'].heights.maximum
l_fus = vehicle.fuselages['fuselage'].lengths.total
if 'horizontal_stabilizer' not in vehicle.wings:
wt_tail_horizontal = 0.0
S_h = 0.0
warnings.warn(
"There is no Horizontal Tail Weight being added to the Configuration",
stacklevel=1)
else:
S_h = vehicle.wings['horizontal_stabilizer'].areas.reference
b_h = vehicle.wings['horizontal_stabilizer'].spans.projected
sweep_h = vehicle.wings['horizontal_stabilizer'].sweeps.quarter_chord
mac_h = vehicle.wings['horizontal_stabilizer'].chords.mean_aerodynamic
t_c_h = vehicle.wings['horizontal_stabilizer'].thickness_to_chord
h_tail_exposed = vehicle.wings[
'horizontal_stabilizer'].areas.exposed / vehicle.wings[
'horizontal_stabilizer'].areas.wetted
l_w2h = vehicle.wings['horizontal_stabilizer'].origin[
0] + vehicle.wings['horizontal_stabilizer'].aerodynamic_center[
0] - vehicle.wings['main_wing'].origin[0] - vehicle.wings[
'main_wing'].aerodynamic_center[
0] #Need to check this is the length of the horizontal tail moment arm
wt_tail_horizontal = tail_horizontal(b_h, sweep_h, Nult, S_h, TOW,
mac_w, mac_h, l_w2h, t_c_h,
h_tail_exposed)
wt_tail_horizontal = wt_tail_horizontal * (1. - wt_factors.empennage)
vehicle.wings[
'horizontal_stabilizer'].mass_properties.mass = wt_tail_horizontal
if 'vertical_stabilizer' not in vehicle.wings:
output_3 = Data()
output_3.wt_tail_vertical = 0.0
output_3.wt_rudder = 0.0
S_v = 0.0
warnings.warn(
"There is no Vertical Tail Weight being added to the Configuration",
stacklevel=1)
else:
S_v = vehicle.wings['vertical_stabilizer'].areas.reference
b_v = vehicle.wings['vertical_stabilizer'].spans.projected
t_c_v = vehicle.wings['vertical_stabilizer'].thickness_to_chord
sweep_v = vehicle.wings['vertical_stabilizer'].sweeps.quarter_chord
t_tail = vehicle.wings['vertical_stabilizer'].t_tail
output_3 = tail_vertical(S_v, Nult, b_v, TOW, t_c_v, sweep_v,
S_gross_w, t_tail)
wt_vtail_tot = output_3.wt_tail_vertical + output_3.wt_rudder
wt_vtail_tot = wt_vtail_tot * (1. - wt_factors.empennage)
vehicle.wings[
'vertical_stabilizer'].mass_properties.mass = wt_vtail_tot
# Calculating Empty Weight of Aircraft
wt_landing_gear = landing_gear.landing_gear(TOW)
wt_fuselage = tube(S_fus, diff_p_fus, w_fus, h_fus, l_fus, Nlim, wt_zf,
wt_wing, wt_propulsion, wing_c_r)
wt_fuselage = wt_fuselage * (1. - wt_factors.fuselage)
output_2 = systems(num_seats, ctrl_type, S_h, S_v, S_gross_w, ac_type)
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_wing + wt_fuselage + wt_landing_gear + wt_propulsion + output_2.wt_systems + \
wt_tail_horizontal + wt_vtail_tot)
vehicle.fuselages['fuselage'].mass_properties.mass = wt_fuselage
# packup outputs
output = payload.payload(TOW, wt_empty, num_pax, wt_cargo)
output.wing = wt_wing
output.fuselage = wt_fuselage
output.propulsion = wt_propulsion
output.landing_gear = wt_landing_gear
output.horizontal_tail = wt_tail_horizontal
output.vertical_tail = output_3.wt_tail_vertical * (1. -
wt_factors.empennage)
output.rudder = output_3.wt_rudder * (1. - wt_factors.empennage)
output.systems = output_2.wt_systems
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_2.wt_flt_ctrl
output.systems_breakdown.apu = output_2.wt_apu
output.systems_breakdown.hydraulics = output_2.wt_hyd_pnu
output.systems_breakdown.instruments = output_2.wt_instruments
output.systems_breakdown.avionics = output_2.wt_avionics
output.systems_breakdown.optionals = output_2.wt_opitems
output.systems_breakdown.electrical = output_2.wt_elec
output.systems_breakdown.air_conditioner = output_2.wt_ac
output.systems_breakdown.furnish = output_2.wt_furnish
#define weights components
try:
landing_gear_component = vehicle.landing_gear #landing gear previously defined
except AttributeError: # landing gear not defined
landing_gear_component = SUAVE.Components.Landing_Gear.Landing_Gear()
vehicle.landing_gear = landing_gear_component
control_systems = SUAVE.Components.Physical_Component()
electrical_systems = SUAVE.Components.Physical_Component()
passengers = SUAVE.Components.Physical_Component()
furnishings = SUAVE.Components.Physical_Component()
air_conditioner = SUAVE.Components.Physical_Component()
fuel = SUAVE.Components.Physical_Component()
apu = SUAVE.Components.Physical_Component()
hydraulics = SUAVE.Components.Physical_Component()
optionals = SUAVE.Components.Physical_Component()
rudder = SUAVE.Components.Physical_Component()
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
#assign output weights to objects
landing_gear_component.mass_properties.mass = output.landing_gear
control_systems.mass_properties.mass = output.systems_breakdown.control_systems
electrical_systems.mass_properties.mass = output.systems_breakdown.electrical
passengers.mass_properties.mass = output.pax + output.bag
furnishings.mass_properties.mass = output.systems_breakdown.furnish
avionics.mass_properties.mass = output.systems_breakdown.avionics \
+ output.systems_breakdown.instruments
air_conditioner.mass_properties.mass = output.systems_breakdown.air_conditioner
fuel.mass_properties.mass = output.fuel
apu.mass_properties.mass = output.systems_breakdown.apu
hydraulics.mass_properties.mass = output.systems_breakdown.hydraulics
optionals.mass_properties.mass = output.systems_breakdown.optionals
rudder.mass_properties.mass = output.rudder
#assign components to vehicle
vehicle.control_systems = control_systems
vehicle.electrical_systems = electrical_systems
vehicle.avionics = avionics
vehicle.furnishings = furnishings
vehicle.passenger_weights = passengers
vehicle.air_conditioner = air_conditioner
vehicle.fuel = fuel
vehicle.apu = apu
vehicle.hydraulics = hydraulics
vehicle.optionals = optionals
vehicle.landing_gear = landing_gear_component
vehicle.wings['vertical_stabilizer'].rudder = rudder
return output
def empty(vehicle):
""" This is for a standard Tube and Wing aircraft configuration.
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
Source:
N/A
Inputs:
engine - a data dictionary with the fields:
thrust_sls - sea level static thrust of a single engine [Newtons]
wing - a data dictionary with the fields:
gross_area - wing gross area [meters**2]
span - span of the wing [meters]
taper - taper ratio of the wing [dimensionless]
t_c - thickness-to-chord ratio of the wing [dimensionless]
sweep - sweep angle of the wing [radians]
mac - mean aerodynamic chord of the wing [meters]
r_c - wing root chord [meters]
aircraft - a data dictionary with the fields:
Nult - ultimate load of the aircraft [dimensionless]
Nlim - limit load factor at zero fuel weight of the aircraft [dimensionless]
TOW - maximum takeoff weight of the aircraft [kilograms]
zfw - maximum zero fuel weight of the aircraft [kilograms]
num_eng - number of engines on the aircraft [dimensionless]
num_pax - number of passengers on the aircraft [dimensionless]
wt_cargo - weight of the bulk cargo being carried on the aircraft [kilograms]
num_seats - number of seats installed on the aircraft [dimensionless]
ctrl - specifies if the control system is "fully powered", "partially powered", or not powered [dimensionless]
ac - determines type of instruments, electronics, and operating items based on types:
"short-range", "medium-range", "long-range", "business", "cargo", "commuter", "sst" [dimensionless]
w2h - tail length (distance from the airplane c.g. to the horizontal tail aerodynamic center) [meters]
fuselage - a data dictionary with the fields:
area - fuselage wetted area [meters**2]
diff_p - Maximum fuselage pressure differential [Pascal]
width - width of the fuselage [meters]
height - height of the fuselage [meters]
length - length of the fuselage [meters]
horizontal
area - area of the horizontal tail [meters**2]
span - span of the horizontal tail [meters]
sweep - sweep of the horizontal tail [radians]
mac - mean aerodynamic chord of the horizontal tail [meters]
t_c - thickness-to-chord ratio of the horizontal tail [dimensionless]
exposed - exposed area ratio for the horizontal tail [dimensionless]
vertical
area - area of the vertical tail [meters**2]
span - sweight = weight * Units.lbpan of the vertical [meters]
t_c - thickness-to-chord ratio of the vertical tail [dimensionless]
sweep - sweep angle of the vertical tail [radians]
t_tail - factor to determine if aircraft has a t-tail, "yes" [dimensionless]
Outputs:
output - a data dictionary with fields:
wt_payload - weight of the passengers plus baggage and paid cargo [kilograms]
wt_pax - weight of all the passengers [kilogram]
wt_bag - weight of all the baggage [kilogram]
wt_fuel - weight of the fuel carried [kilogram]
wt_empty - operating empty weight of the aircraft [kilograms]
Properties Used:
N/A
"""
# Unpack inputs
Nult = vehicle.envelope.ultimate_load
Nlim = vehicle.envelope.limit_load
TOW = vehicle.mass_properties.max_takeoff
wt_zf = vehicle.mass_properties.max_zero_fuel
num_pax = vehicle.passengers
wt_cargo = vehicle.mass_properties.cargo
num_seats = vehicle.fuselages['fuselage'].number_coach_seats
ctrl_type = vehicle.systems.control
ac_type = vehicle.systems.accessories
propulsor_name = vehicle.propulsors.keys()[0] #obtain the key for the propulsor for assignment purposes
propulsors = vehicle.propulsors[propulsor_name]
num_eng = propulsors.number_of_engines
if propulsor_name=='turbofan' or propulsor_name=='Turbofan':
# thrust_sls should be sea level static thrust. Using design thrust results in wrong propulsor
# weight estimation. Engine sizing should return this value.
# for now, using thrust_sls = design_thrust / 0.20, just for optimization evaluations
thrust_sls = propulsors.sealevel_static_thrust
wt_engine_jet = Propulsion.engine_jet(thrust_sls)
wt_propulsion = Propulsion.integrated_propulsion(wt_engine_jet,num_eng)
propulsors.mass_properties.mass = wt_propulsion
else: #propulsor used is not a turbo_fan; assume mass_properties defined outside model
wt_propulsion = propulsors.mass_properties.mass
if wt_propulsion==0:
warnings.warn("Propulsion mass= 0 ;e there is no Engine Weight being added to the Configuration", stacklevel=1)
S_gross_w = vehicle.reference_area
if not vehicle.wings.has_key('main_wing'):
wt_wing = 0.0
wing_c_r = 0.0
warnings.warn("There is no Wing Weight being added to the Configuration", stacklevel=1)
else:
b = vehicle.wings['main_wing'].spans.projected
lambda_w = vehicle.wings['main_wing'].taper
t_c_w = vehicle.wings['main_wing'].thickness_to_chord
sweep_w = vehicle.wings['main_wing'].sweeps.quarter_chord
mac_w = vehicle.wings['main_wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['main_wing'].chords.root
wt_wing = wing_main.wing_main(S_gross_w,b,lambda_w,t_c_w,sweep_w,Nult,TOW,wt_zf)
vehicle.wings['main_wing'].mass_properties.mass = wt_wing
S_fus = vehicle.fuselages['fuselage'].areas.wetted
diff_p_fus = vehicle.fuselages['fuselage'].differential_pressure
w_fus = vehicle.fuselages['fuselage'].width
h_fus = vehicle.fuselages['fuselage'].heights.maximum
l_fus = vehicle.fuselages['fuselage'].lengths.total
if not vehicle.wings.has_key('horizontal_stabilizer'):
wt_tail_horizontal = 0.0
S_h = 0.0
warnings.warn("There is no Horizontal Tail Weight being added to the Configuration", stacklevel=1)
else:
S_h = vehicle.wings['horizontal_stabilizer'].areas.reference
b_h = vehicle.wings['horizontal_stabilizer'].spans.projected
sweep_h = vehicle.wings['horizontal_stabilizer'].sweeps.quarter_chord
mac_h = vehicle.wings['horizontal_stabilizer'].chords.mean_aerodynamic
t_c_h = vehicle.wings['horizontal_stabilizer'].thickness_to_chord
h_tail_exposed = vehicle.wings['horizontal_stabilizer'].areas.exposed / vehicle.wings['horizontal_stabilizer'].areas.wetted
l_w2h = vehicle.wings['horizontal_stabilizer'].origin[0] + vehicle.wings['horizontal_stabilizer'].aerodynamic_center[0] - vehicle.wings['main_wing'].origin[0] - vehicle.wings['main_wing'].aerodynamic_center[0] #Need to check this is the length of the horizontal tail moment arm
wt_tail_horizontal = tail_horizontal(b_h,sweep_h,Nult,S_h,TOW,mac_w,mac_h,l_w2h,t_c_h, h_tail_exposed)
vehicle.wings['horizontal_stabilizer'].mass_properties.mass = wt_tail_horizontal
if not vehicle.wings.has_key('vertical_stabilizer'):
output_3 = Data()
output_3.wt_tail_vertical = 0.0
output_3.wt_rudder = 0.0
S_v = 0.0
warnings.warn("There is no Vertical Tail Weight being added to the Configuration", stacklevel=1)
else:
S_v = vehicle.wings['vertical_stabilizer'].areas.reference
b_v = vehicle.wings['vertical_stabilizer'].spans.projected
t_c_v = vehicle.wings['vertical_stabilizer'].thickness_to_chord
sweep_v = vehicle.wings['vertical_stabilizer'].sweeps.quarter_chord
t_tail = vehicle.wings['vertical_stabilizer'].t_tail
output_3 = tail_vertical(S_v,Nult,b_v,TOW,t_c_v,sweep_v,S_gross_w,t_tail)
vehicle.wings['vertical_stabilizer'].mass_properties.mass = output_3.wt_tail_vertical + output_3.wt_rudder
# Calculating Empty Weight of Aircraft
wt_landing_gear = landing_gear.landing_gear(TOW)
wt_fuselage = tube(S_fus, diff_p_fus,w_fus,h_fus,l_fus,Nlim,wt_zf,wt_wing,wt_propulsion, wing_c_r)
output_2 = systems(num_seats, ctrl_type, S_h, S_v, S_gross_w, ac_type)
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_wing + wt_fuselage + wt_landing_gear + wt_propulsion + output_2.wt_systems + \
wt_tail_horizontal + output_3.wt_tail_vertical + output_3.wt_rudder)
vehicle.fuselages['fuselage'].mass_properties.mass = wt_fuselage
# packup outputs
output = payload.payload(TOW, wt_empty, num_pax,wt_cargo)
output.wing = wt_wing
output.fuselage = wt_fuselage
output.propulsion = wt_propulsion
output.landing_gear = wt_landing_gear
output.horizontal_tail = wt_tail_horizontal
output.vertical_tail = output_3.wt_tail_vertical
output.rudder = output_3.wt_rudder
output.systems = output_2.wt_systems
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_2.wt_flt_ctrl
output.systems_breakdown.apu = output_2.wt_apu
output.systems_breakdown.hydraulics = output_2.wt_hyd_pnu
output.systems_breakdown.instruments = output_2.wt_instruments
output.systems_breakdown.avionics = output_2.wt_avionics
output.systems_breakdown.optionals = output_2.wt_opitems
output.systems_breakdown.electrical = output_2.wt_elec
output.systems_breakdown.air_conditioner = output_2.wt_ac
output.systems_breakdown.furnish = output_2.wt_furnish
#define weights components
try:
landing_gear_component=vehicle.landing_gear #landing gear previously defined
except AttributeError: # landing gear not defined
landing_gear_component=SUAVE.Components.Landing_Gear.Landing_Gear()
vehicle.landing_gear=landing_gear_component
control_systems = SUAVE.Components.Physical_Component()
electrical_systems= SUAVE.Components.Physical_Component()
passengers = SUAVE.Components.Physical_Component()
furnishings = SUAVE.Components.Physical_Component()
air_conditioner = SUAVE.Components.Physical_Component()
fuel = SUAVE.Components.Physical_Component()
apu = SUAVE.Components.Physical_Component()
hydraulics = SUAVE.Components.Physical_Component()
optionals = SUAVE.Components.Physical_Component()
rudder = SUAVE.Components.Physical_Component()
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
#assign output weights to objects
landing_gear_component.mass_properties.mass = output.landing_gear
control_systems.mass_properties.mass = output.systems_breakdown.control_systems
electrical_systems.mass_properties.mass = output.systems_breakdown.electrical
passengers.mass_properties.mass = output.pax + output.bag
furnishings.mass_properties.mass = output.systems_breakdown.furnish
avionics.mass_properties.mass = output.systems_breakdown.avionics \
+ output.systems_breakdown.instruments
air_conditioner.mass_properties.mass = output.systems_breakdown.air_conditioner
fuel.mass_properties.mass = output.fuel
apu.mass_properties.mass = output.systems_breakdown.apu
hydraulics.mass_properties.mass = output.systems_breakdown.hydraulics
optionals.mass_properties.mass = output.systems_breakdown.optionals
rudder.mass_properties.mass = output.rudder
#assign components to vehicle
vehicle.control_systems = control_systems
vehicle.electrical_systems = electrical_systems
vehicle.avionics = avionics
vehicle.furnishings = furnishings
vehicle.passenger_weights = passengers
vehicle.air_conditioner = air_conditioner
vehicle.fuel = fuel
vehicle.apu = apu
vehicle.hydraulics = hydraulics
vehicle.optionals = optionals
vehicle.landing_gear = landing_gear_component
vehicle.wings['vertical_stabilizer'].rudder = rudder
return output