def empty(vehicle):
""" output = SUAVE.Methods.Weights.Correlations.Tube_Wing.empty(engine,wing,aircraft,fuselage,horizontal,vertical)
This is for a standard Tube and Wing aircraft configuration.
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]
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
"""
# 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 not vehicle.propulsors.has_key('Turbo Fan'):
wt_engine_jet = 0.0
wt_propulsion = 0.0
warnings.warn("There is no Turbo Fan Engine Weight being added to the Configuration", stacklevel=1)
else:
num_eng = vehicle.propulsors['Turbo Fan'].number_of_engines
thrust_sls = vehicle.propulsors['Turbo Fan'].thrust.design
wt_engine_jet = Propulsion.engine_jet(thrust_sls)
wt_propulsion = Propulsion.integrated_propulsion(wt_engine_jet,num_eng)
S_gross_w = vehicle.reference_area
#S_gross_w = vehicle.wings['Main Wing'].Areas.reference
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'].sweep
mac_w = vehicle.wings['Main Wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['Main Wing'].chords.root
wt_wing = 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'].sweep
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'].sweep
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
# process
# Calculating Empty Weight of Aircraft
wt_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(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.systems = output_2.wt_systems
output.wt_furnish = output_2.wt_furnish
output.horizontal_tail = wt_tail_horizontal
output.vertical_tail = output_3.wt_tail_vertical
output.rudder = output_3.wt_rudder
return output
def empty_custom_eng(vehicle, propulsor):
""" output = SUAVE.Methods.Weights.Correlations.Tube_Wing.empty(engine,wing,aircraft,fuselage,horizontal,vertical)
This is for a standard Tube and Wing aircraft configuration.
Inputs:
propulsor - a propulsor object
mass_properties.mass
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]
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
"""
# Unpack inputs
#thrust_sls = vehicle.propulsors['Turbo Fan'].thrust.design
S_gross_w = vehicle.reference_area
#S_gross_w = vehicle.wings['Main Wing'].Areas.reference
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'].sweep
mac_w = vehicle.wings['Main Wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['Main Wing'].chords.root
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
#l_w2h = vehicle.wings['Horizontal Stabilizer'].position[0] + vehicle.wings['Horizontal Stabilizer'].aerodynamic_center[0] - vehicle.wings['Main Wing'].position[0] - vehicle.wings['Main Wing'].aerodynamic_center[0] #Need to check this is the length of the horizontal tail moment arm
l_w2h = vehicle.w2h
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
S_h = vehicle.wings['Horizontal Stabilizer'].areas.reference
b_h = vehicle.wings['Horizontal Stabilizer'].spans.projected
sweep_h = vehicle.wings['Horizontal Stabilizer'].sweep
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
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'].sweep
t_tail = vehicle.wings['Vertical Stabilizer'].t_tail
# process
# Calculating Empty Weight of Aircraft
wt_wing = wing_main(S_gross_w, b, lambda_w, t_c_w, sweep_w, Nult, TOW,
wt_zf)
wt_landing_gear = landing_gear(TOW)
wt_propulsion = propulsor.mass_properties.mass
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)
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)
output_3 = tail_vertical(S_v, Nult, b_v, TOW, t_c_v, sweep_v, S_gross_w,
t_tail)
# 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.wings['Main Wing'].mass_properties.mass = wt_wing
vehicle.wings[
'Horizontal Stabilizer'].mass_properties.mass = wt_tail_horizontal
vehicle.wings[
'Vertical Stabilizer'].mass_properties.mass = output_3.wt_tail_vertical + output_3.wt_rudder
vehicle.fuselages.Fuselage.mass_properties.mass = wt_fuselage
#vehicle.propulsors['Turbo Fan'].mass_properties.mass = wt_engine_jet
# packup outputs
output = 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.systems = output_2.wt_systems
output.wt_furnish = output_2.wt_furnish
output.horizontal_tail = wt_tail_horizontal
output.vertical_tail = output_3.wt_tail_vertical
output.rudder = output_3.wt_rudder
return output
def empty(vehicle):
""" output = SUAVE.Methods.Weights.Correlations.Tube_Wing.empty(engine,wing,aircraft,fuselage,horizontal,vertical)
This is for a standard Tube and Wing aircraft configuration.
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]
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
"""
# 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
V_descent = 80.
d_fus = vehicle.fuselages['fuselage'].effective_diameter
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)
# print wt_engine_jet
# print "Test2"
wt_propulsion = Propulsion.integrated_propulsion(wt_engine_jet, num_eng)
propulsors.mass_properties.mass = wt_propulsion
# print wt_propulsion
d_eng = propulsors.nacelle_diameter
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
# S_gross_w = vehicle.wings['main_wing'].Areas.reference
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'].sweep
mac_w = vehicle.wings['main_wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['main_wing'].chords.root
# print 'c_root =',wing_c_r
wt_wing = 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
A_h = vehicle.wings['horizontal_stabilizer'].aspect_ratio
sweep_h = vehicle.wings['horizontal_stabilizer'].sweep
mac_h = vehicle.wings['horizontal_stabilizer'].chords.mean_aerodynamic
c_r_h = vehicle.wings['horizontal_stabilizer'].chords.root
t_c_h = vehicle.wings['horizontal_stabilizer'].thickness_to_chord
# print 'c_r_h =',c_r_h
taper_h = vehicle.wings['horizontal_stabilizer'].taper
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)
Sel = 0.15 * S_h
wt_tail_horizontal = hor_tail(c_r_h, taper_h, l_w2h, Sel, S_h, A_h, b_h, sweep_h, TOW, Nult)
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
A_v = vehicle.wings['vertical_stabilizer'].aspect_ratio
t_c_v = vehicle.wings['vertical_stabilizer'].thickness_to_chord
sweep_v = vehicle.wings['vertical_stabilizer'].sweep
taper_v = vehicle.wings['vertical_stabilizer'].taper
c_r_v = vehicle.wings['vertical_stabilizer'].chords.root
mac_v = vehicle.wings['vertical_stabilizer'].chords.mean_aerodynamic
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_tail_vertical = vert_tail(S_v, b_v, c_r_v, taper_v, t_c_v, A_v, sweep_v, l_w2h, TOW)
vehicle.wings['vertical_stabilizer'].mass_properties.mass = wt_tail_vertical
# print 'b_v =',b_v
# print 'c_r_v =',c_r_v
# print 'c_t_v =',c_r_v*taper_v
# vehicle.wings['vertical_stabilizer'].mass_properties.mass = output_3.wt_tail_vertical + output_3.wt_rudder
# Calculating Empty Weight of Aircraft
wt_landing_gear, gear, h_gear = landing_gear(TOW, d_eng, h_fus, V_descent)
# print 'h_gear =', h_gear
l_nose = vehicle.fuselages['fuselage'].lengths.nose
l_center = vehicle.fuselages['fuselage'].lengths.cabin
l_tail = vehicle.fuselages['fuselage'].lengths.tail
S_cstot = 0.22 * mac_w * 0.6 * b * 1.5 + 0.3 * mac_v * 0.9 * b_v + 0.2 * mac_h * 0.9 * b_h # 1.5 factor is because of assumed spoilers on wings
Iy_SI = 1230140.7646609414 # kgm^2
Vmax = 240 # m/s
wt_fuselage = tube(S_fus, diff_p_fus, w_fus, h_fus, l_nose, l_center, l_tail, l_fus, Nlim, wt_zf, wt_wing,
wt_propulsion, wing_c_r, TOW, h_gear)
# output_2 = systems(num_seats, ctrl_type, S_h, S_v, S_gross_w, ac_type)
OEW = 109000 # kg #FIXME
V_fuel = 69000 / 820 # m^3 #FIXME
wt_engine_jet = 9000 # kg
output_sys = classIIsys(num_eng, wt_engine_jet, V_fuel, Iy_SI, S_cstot, l_fus, OEW, TOW, Vmax, b)
l_eng = 7 # m
output_prop = propgroup(num_eng, wt_engine_jet, d_eng, l_eng, output_sys.w_fuelsys, output_sys.w_starter, Nult,
Vmax)
wt_propulsion = output_prop.w_propulsion
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_wing + wt_fuselage + wt_landing_gear + wt_propulsion + output_sys.w_system +
wt_tail_horizontal + output_3.wt_tail_vertical + output_3.wt_rudder)
vehicle.fuselages['fuselage'].mass_properties.mass = wt_fuselage
# packup outputs
output = 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_sys.w_system
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_sys.w_flightcontrol
output.systems_breakdown.apu = output_sys.w_apu
output.systems_breakdown.hydraulics = output_sys.w_hydraullics
# output.systems_breakdown.instruments = output_sys.wt_instruments
output.systems_breakdown.avionics = output_sys.w_avionics
# output.systems_breakdown.optionals = output_sys.wt_opitems
output.systems_breakdown.electrical = output_sys.w_electrical
output.systems_breakdown.air_conditioner = output_sys.w_env
# 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):
""" output = SUAVE.Methods.Weights.Correlations.Tube_Wing.empty(engine,wing,aircraft,fuselage,horizontal,vertical)
This is for a standard Tube and Wing aircraft configuration.
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]
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
"""
# 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
#S_gross_w = vehicle.wings['main_wing'].Areas.reference
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'].sweep
mac_w = vehicle.wings['main_wing'].chords.mean_aerodynamic
wing_c_r = vehicle.wings['main_wing'].chords.root
wt_wing = 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'].sweep
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'].sweep
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(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(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
def empty(vehicle):
""" output = SUAVE.Methods.Weights.Correlations.Tube_Wing.empty(engine,wing,aircraft,fuselage,horizontal,vertical)
Computes the empty weight breakdown of a General Aviation type aircraft
Inputs:
engine - a data dictionary with the fields:
thrust_sls - sea level static thrust of a single engine [Newtons]
vehicle - a data dictionary with the fields:
reference_area [meters**2]
envelope - a data dictionary with the fields:
ultimate_load - ultimate load of the aircraft [dimensionless]
limit_load - limit load factor at zero fuel weight of the aircraft [dimensionless]
mass_properties - a data dictionary with the fields:
max_takeoff - max takeoff weight of the vehicle [kilograms]
max_zero_fuel - maximum zero fuel weight of the aircraft [kilograms]
cargo - cargo weight [kilograms]
passengers - number of passengers on the aircraft [dimensionless]
design_dynamic_pressure - dynamic pressure at cruise conditions [Pascal]
design_mach_number - mach number at cruise conditions [dimensionless]
propulsors - a data dictionary with the fields:
keys - identifier for the type of propulsor; different types have different fields
turbofan
thrust_sls - sealevel standard thrust [Newtons]
internal_combustion
rated_power - maximum rated power of the internal combustion engine [Watts]
number_engines - integer indicating the number of engines on the aircraft
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]
fuel - a data dictionary with the fields:
mass_properties - a data dictionary with the fields:
mass -mass of fuel [kilograms]
density - gravimetric density of fuel [kilograms/meter**3]
number_of_tanks - number of external fuel tanks [dimensionless]
internal_volume - internal fuel volume contained in the wing [meters**3]
wings - a data dictionary with the fields:
wing - a data dictionary with the fields:
span - span of the wing [meters]
taper - taper ratio of the wing [dimensionless]
thickness_to_chord - thickness-to-chord ratio of the wing [dimensionless]
chords - a data dictionary with the fields:
mean_aerodynamic - mean aerodynamic chord of the wing [meters]
root - root chord of the wing [meters]
sweeps - a data dictionary with the fields:
quarter_chord - quarter chord sweep angle of the wing [radians]
mac - mean aerodynamic chord of the wing [meters]
r_c - wing root chord [meters]
origin - location of the leading edge of the wing relative to the front of the fuselage [meters,meters,meters]
aerodynamic_center - location of the aerodynamic center of the horizontal_stabilizer relative to the leading edge of the wing [meters,meters,meters]
horizontal_stabilizer - a data dictionary with the fields:
areas - a data dictionary with the fields:
reference - reference area of the horizontal stabilizer [meters**2]
exposed - exposed area for the horizontal tail [meters**2]
taper - taper ratio of the horizontal stabilizer [dimensionless]
span - span of the horizontal tail [meters]
sweeps - a data dictionary with the fields:
quarter_chord - quarter chord sweep angle of the horizontal stabilizer [radians]
chords - a data dictionary with the fields:
mean_aerodynamic - mean aerodynamic chord of the horizontal stabilizer [meters]
root - root chord of the horizontal stabilizer
thickness_to_chord - thickness-to-chord ratio of the horizontal tail [dimensionless]
mac - mean aerodynamic chord of the horizontal tail [meters]
origin - location of the leading of the horizontal tail relative to the front of the fuselage [meters,meters,meters]
aerodynamic_center - location of the aerodynamic center of the horizontal_stabilizer relative to the leading edge of the horizontal stabilizer [meters,meters,meters]
vertical - a data dictionary with the fields:
areas - a data dictionary with the fields:
reference - reference area of the vertical stabilizer [meters**2]
span - span of the vertical tail [meters]
taper - taper ratio of the horizontal stabilizer [dimensionless]
t_c - thickness-to-chord ratio of the vertical tail [dimensionless]
sweeps - a data dictionary with the fields:
quarter_chord - quarter chord sweep angle of the vertical stabilizer [radians]
t_tail - flag to determine if aircraft has a t-tail, "yes" [dimensionless]
fuselages - a data dictionary with the fields:
fuselage - a data dictionary with the fields:
areas - a data dictionary with the fields:
wetted - wetted area of the fuselage [meters**2]
differential_pressure - Maximum fuselage pressure differential [Pascal]
width - width of the fuselage [meters]
heights - a data dictionary with the fields:
maximum - height of the fuselage [meters]
lengths- a data dictionary with the fields:
structure - structural length of the fuselage [meters]
mass_properties - a data dictionary with the fields:
volume - total volume of the fuselage [meters**3]
internal_volume - internal volume of the fuselage [meters**3]
number_coach_sets - number of seats on the aircraft [dimensionless]
landing_gear - a data dictionary with the fields:
main - a data dictionary with the fields:
strut_length - strut length of the main gear [meters]
nose - a data dictionary with the fields:
strut_length - strut length of the nose gear [meters]
avionics - a data dictionary, used to determine if avionics weight is calculated, don't include if vehicle has none
air_conditioner - a data dictionary, used to determine if air conditioner weight is calculated, don't include if vehicle has none
Outputs:
output - a data dictionary with fields:
wing - wing weight [kilograms]
fuselage - fuselage weight [kilograms]
propulsion - propulsion [kilograms]
landing_gear_main - main gear weight [kilograms]
landing_gear_nose - nose gear weight [kilograms]
horizonal_tail - horizontal stabilizer weight [kilograms]
vertical_tail - vertical stabilizer weight [kilograms]
systems - total systems weight [kilograms]
systems_breakdown - a data dictionary with fields:
control_systems - control systems weight [kilograms]
hydraulics - hydraulics weight [kilograms]
avionics - avionics weight [kilograms]
electric - electrical systems weight [kilograms]
air_conditioner - air conditioner weight [kilograms]
furnish - furnishing weight [kilograms]
fuel_system - fuel system weight [ kilograms]
Wing, empannage, fuselage, propulsion and individual systems masses updated with their calculated values
Assumptions:
calculated aircraft weight from correlations created per component of historical aircraft
"""
# Unpack inputs
S_gross_w = vehicle.reference_area
fuel = vehicle.fuel
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
q_c = vehicle.design_dynamic_pressure
mach_number = vehicle.design_mach_number
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':
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
elif propulsor_name == 'internal_combustion':
rated_power = propulsors.rated_power/num_eng
wt_engine_piston = Propulsion.engine_piston(rated_power)
wt_propulsion = Propulsion.integrated_propulsion_general_aviation(wt_engine_piston,num_eng)
propulsors.mass_properties.mass = wt_propulsion
else: #propulsor used is not an IC Engine or Turbofan; 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)
#find fuel volume
if not vehicle.has_key('fuel'):
warnings.warn("fuel mass= 0 ; fuel system volume is calculated incorrectly ", stacklevel=1)
N_tank = 0
V_fuel = 0.
V_fuel_int = 0.
else:
m_fuel = fuel.mass_properties.mass
landing_weight = TOW-m_fuel #just assume this for now
N_tank = fuel.number_of_tanks
V_fuel_int = fuel.internal_volume #fuel in internal (as opposed to external tanks)
V_fuel = m_fuel/fuel.density #total fuel
fuel.mass_properties.volume = V_fuel
#main wing
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
AR_w = (b**2.)/S_gross_w
taper_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
#now run weight script for the wing
wt_wing = wing_main(S_gross_w, m_fuel, AR_w, sweep_w, q_c, taper_w, t_c_w,Nult,TOW)
vehicle.wings['main_wing'].mass_properties.mass = wt_wing
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
AR_h = (b_h**2.)/S_h
taper_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
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] #used for fuselage weight
wt_tail_horizontal = tail_horizontal(S_h, AR_h, sweep_h, q_c, taper_h, t_c_h,Nult,TOW)
vehicle.wings['horizontal_stabilizer'].mass_properties.mass = wt_tail_horizontal
#vertical stabilizer
if not vehicle.wings.has_key('vertical_stabilizer'):
output_3 = Data()
output_3.wt_tail_vertical = 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
AR_v = (b_v**2.)/S_v
taper_v = vehicle.wings['vertical_stabilizer'].taper
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, AR_v, sweep_v, q_c, taper_v, t_c_v, Nult,TOW,t_tail)
vehicle.wings['vertical_stabilizer'].mass_properties.mass = output_3.wt_tail_vertical
if not vehicle.has_key('fuselages.fuselage'):
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.structure
V_fuse = vehicle.fuselages['fuselage'].mass_properties.volume
V_int = vehicle.fuselages['fuselage'].mass_properties.internal_volume
num_seats = vehicle.fuselages['fuselage'].number_coach_seats
#calculate fuselage weight
wt_fuselage = fuselage(S_fus, Nult, TOW, w_fus, h_fus, l_fus, l_w2h, q_c, V_fuse, diff_p_fus)
else:
print 'got here'
warnings.warn('There is no Fuselage weight being added to the vehicle', stacklevel=1)
#landing gear
if not vehicle.has_key('landing_gear'):
warnings.warn('There is no Landing Gear weight being added to the vehicle', stacklevel=1)
wt_landing_gear = Data()
wt_landing_gear.main = 0.0
wt_landing_gear.nose = 0.0
else:
landing_gear_component = vehicle.landing_gear #landing gear previously defined
strut_length_main = landing_gear_component.main.strut_length
strut_length_nose = landing_gear_component.nose.strut_length
wt_landing_gear = landing_gear(landing_weight, Nult, strut_length_main, strut_length_nose)
landing_gear_component.main.mass_properties.mass = wt_landing_gear.main
landing_gear_component.nose.mass_properties.mass = wt_landing_gear.nose
if not vehicle.has_key('avionics'):
warnings.warn('There is no avionics weight being added to the vehicle; many weight correlations are dependant on this', stacklevel=1)
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
W_uav = 0.
else:
avionics = vehicle.avionics
W_uav = avionics.mass_properties.uninstalled
has_air_conditioner = 0
if vehicle.has_key('air_conditioner'):
has_air_conditioner = 1
# Calculating Empty Weight of Aircraft
output_2 = systems(W_uav,V_fuel, V_fuel_int, N_tank, num_eng, l_fus, b, TOW, Nult, num_seats, mach_number, has_air_conditioner)
# Calculate the equipment empty weight of the aircraft
wt_empty = (wt_wing + wt_fuselage + wt_landing_gear.main+wt_landing_gear.nose + wt_propulsion + output_2.wt_systems + \
wt_tail_horizontal + output_3.wt_tail_vertical)
vehicle.fuselages['fuselage'].mass_properties.mass = wt_fuselage
# packup outputs
output = payload(TOW, wt_empty, num_pax,wt_cargo)
output.wing = wt_wing
output.fuselage = wt_fuselage
output.propulsion = wt_propulsion
output.landing_gear_main = wt_landing_gear.main
output.landing_gear_nose = wt_landing_gear.nose
output.horizontal_tail = wt_tail_horizontal
output.vertical_tail = output_3.wt_tail_vertical
output.systems = output_2.wt_systems
output.systems_breakdown = Data()
output.systems_breakdown.control_systems = output_2.wt_flt_ctrl
output.systems_breakdown.hydraulics = output_2.wt_hyd_pnu
output.systems_breakdown.avionics = output_2.wt_avionics
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
output.systems_breakdown.fuel_system = output_2.wt_fuel_sys
#define weights components
control_systems = SUAVE.Components.Physical_Component()
electrical_systems= SUAVE.Components.Physical_Component()
passengers = SUAVE.Components.Physical_Component()
furnishings = SUAVE.Components.Physical_Component()
apu = SUAVE.Components.Physical_Component()
hydraulics = SUAVE.Components.Physical_Component()
#assign output weights to objects
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
hydraulics.mass_properties.mass = output.systems_breakdown.hydraulics
if has_air_conditioner:
vehicle.air_conditioner.mass_properties.mass = output.systems_breakdown.air_conditioner
#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.apu = apu
vehicle.hydraulics = hydraulics
vehicle.landing_gear = landing_gear_component
#note; air conditioner optional, and weight is added to the air_conditioner object directly
return output