本文整理汇总了Python中SUAVE.Core.Data.systems方法的典型用法代码示例。如果您正苦于以下问题：Python Data.systems方法的具体用法？Python Data.systems怎么用？Python Data.systems使用的例子？那么, 这里精选的方法代码示例或许可以为您提供帮助。您也可以进一步了解该方法所在类SUAVE.Core.Data的用法示例。
在下文中一共展示了Data.systems方法的4个代码示例，这些例子默认根据受欢迎程度排序。您可以为喜欢或者感觉有用的代码点赞，您的评价将有助于系统推荐出更棒的Python代码示例。
示例1: main
# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import systems [as 别名]
def main():
vehicle = vehicle_setup()
weight = Tube_Wing.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 12977.803363592691 #includes cargo #22177.6377131 #without cargo
actual.empty = 38688.08848390731
actual.wing = 6649.709658738429
actual.fuselage = 6642.061164271899
actual.propulsion = 6838.185174956626
actual.landing_gear = 3160.632
actual.systems = 13479.10479056802
actual.wt_furnish = 6431.80372889
actual.horizontal_tail = 1037.7414196819743
actual.vertical_tail = 629.0387683502595
actual.rudder = 251.61550734010382
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload)/actual.payload
error.pax = (actual.pax - weight.pax)/actual.pax
error.bag = (actual.bag - weight.bag)/actual.bag
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - weight.landing_gear)/actual.landing_gear
error.systems = (actual.systems - weight.systems)/actual.systems
error.wt_furnish = (actual.wt_furnish - weight.systems_breakdown.furnish)/actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder)/actual.rudder
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 618.485310343
actual.fuel = 144.69596603
actual.wing = 124.673093906
actual.fuselage = 119.522072873
actual.propulsion = 194.477769922 #includes power plant and propeller, does not include fuel system
actual.landing_gear = 44.8033840543+5.27975390045
actual.furnishing = 37.8341395817
actual.electrical = 36.7532226254
actual.control_systems = 14.8331955546
actual.fuel_systems = 15.6859717453
actual.systems = 108.096549345
error = Data()
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - (weight.landing_gear_main+weight.landing_gear_nose))/actual.landing_gear
error.furnishing = (actual.furnishing-weight.systems_breakdown.furnish)/actual.furnishing
error.electrical = (actual.electrical-weight.systems_breakdown.electrical)/actual.electrical
error.control_systems = (actual.control_systems-weight.systems_breakdown.control_systems)/actual.control_systems
error.fuel_systems = (actual.fuel_systems-weight.systems_breakdown.fuel_system)/actual.fuel_systems
error.systems = (actual.systems - weight.systems)/actual.systems
print('actual.systems=', actual.systems)
print('General Aviation Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1e-6)
# BWB WEIGHTS
vehicle = bwb_setup()
weight = BWB.empty(vehicle)
# regression values
actual = Data()
#.........这里部分代码省略.........
示例2: main
# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import systems [as 别名]
#.........这里部分代码省略.........
# setup
fan.polytropic_efficiency = 0.93
fan.pressure_ratio = 1.7
# add to network
turbofan.append(fan)
# ------------------------------------------------------------------
#Component 10 : thrust (to compute the thrust)
thrust = SUAVE.Components.Energy.Processes.Thrust()
thrust.tag ='compute_thrust'
#total design thrust (includes all the engines)
thrust.total_design = 2*24000. * Units.N #Newtons
#design sizing conditions
altitude = 35000.0*Units.ft
mach_number = 0.78
isa_deviation = 0.
# add to network
turbofan.thrust = thrust
#size the turbofan
turbofan_sizing(turbofan,mach_number,altitude)
# add gas turbine network gt_engine to the vehicle
vehicle.append_component(turbofan)
vehicle.passengers = 170. # Number of passengers
vehicle.mass_properties.cargo = 0. * Units.kilogram # Mass of cargo
vehicle.systems.control = "fully powered" # Specify fully powered, partially powered or anything else is fully aerodynamic
vehicle.systems.accessories = "medium-range" # Specify what type of aircraft you have
vehicle.reference_area = 124.862 * Units.meter**2 # Wing gross area in square meters
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.spans.projected = 50. * Units.meter # Span in meters
wing.taper = 0.2 # Taper ratio
wing.thickness_to_chord = 0.08 # Thickness-to-chord ratio
wing.sweep = .4363323 * Units.rad # sweep angle in degrees
wing.chords.root = 15. * Units.meter # Wing root chord length
wing.chords.mean_aerodynamic = 10. * Units.meters # Length of the mean aerodynamic chord of the wing
wing.origin = [20,0,0] * Units.meters # Location of main wing from origin of the vehicle
wing.aerodynamic_center = [3,0,0] * Units.meters # Location of aerodynamic center from origin of the main wing
vehicle.append_component(wing)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.areas.wetted = 688.64 * Units.meter**2 # Fuselage wetted area
fuselage.differential_pressure = 55960.5 * Units.pascal # Maximum differential pressure
fuselage.width = 4. * Units.meter # Width of the fuselage
fuselage.heights.maximum = 4. * Units.meter # Height of the fuselage
fuselage.lengths.total = 58.4 * Units.meter # Length of the fuselage
fuselage.number_coach_seats = 200.
vehicle.append_component(fuselage)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.areas.reference = 75. * Units.meters**2 # Area of the horizontal tail
wing.spans.projected = 15. * Units.meters # Span of the horizontal tail
wing.sweep = 38. * Units.deg # Sweep of the horizontal tail
wing.chords.mean_aerodynamic = 5. * Units.meters # Length of the mean aerodynamic chord of the horizontal tail
wing.thickness_to_chord = 0.07 # Thickness-to-chord ratio of the horizontal tail示例3: main
# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import systems [as 别名]
def main():
vehicle = vehicle_setup()
weight = Tube_Wing.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 12990.957450008464 #includes cargo #22177.6377131 #without cargo
actual.empty = 38674.934397491539
actual.wing = 6649.7096587384294
actual.fuselage = 6642.0611642718986
actual.propulsion = 6838.1851749566231
actual.landing_gear = 3160.632
actual.systems = 13479.10479056802
actual.wt_furnish = 6431.80372889
actual.horizontal_tail = 1024.5873332662029
actual.vertical_tail = 629.03876835025949
actual.rudder = 251.61550734010382
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload)/actual.payload
error.pax = (actual.pax - weight.pax)/actual.pax
error.bag = (actual.bag - weight.bag)/actual.bag
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - weight.landing_gear)/actual.landing_gear
error.systems = (actual.systems - weight.systems)/actual.systems
error.wt_furnish = (actual.wt_furnish - weight.systems_breakdown.furnish)/actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder)/actual.rudder
print 'Results (kg)'
print weight
print 'Relative Errors'
print error
for k,v in error.items():
assert(np.abs(v)<1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 618.485310343
actual.fuel = 144.69596603
actual.wing = 124.673093906
actual.fuselage = 119.522072873
actual.propulsion = 194.477769922 #includes power plant and propeller, does not include fuel system
actual.landing_gear = 44.8033840543+5.27975390045
actual.furnishing = 37.8341395817
actual.electrical = 36.7532226254
actual.control_systems = 14.8331955546
actual.fuel_systems = 15.6859717453
actual.systems = 108.096549345
error = Data()
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - (weight.landing_gear_main+weight.landing_gear_nose))/actual.landing_gear
error.furnishing = (actual.furnishing-weight.systems_breakdown.furnish)/actual.furnishing
error.electrical = (actual.electrical-weight.systems_breakdown.electrical)/actual.electrical
error.control_systems = (actual.control_systems-weight.systems_breakdown.control_systems)/actual.control_systems
error.fuel_systems = (actual.fuel_systems-weight.systems_breakdown.fuel_system)/actual.fuel_systems
error.systems = (actual.systems - weight.systems)/actual.systems
print 'actual.systems=', actual.systems
print 'General Aviation Results (kg)'
print weight
print 'Relative Errors'
print error
for k,v in error.items():
assert(np.abs(v)<1e-6)
return示例4: main
# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import systems [as 别名]
def main():
vehicle = SUAVE.Vehicle()# Create the vehicle for testing
# Parameters Required
vehicle.envelope.ultimate_load = 3.5 # Ultimate load
vehicle.mass_properties.max_takeoff = 79015.8 * Units.kilograms # Maximum takeoff weight in kilograms
vehicle.mass_properties.max_zero_fuel = 79015.8 * 0.9 * Units.kilograms # Maximum zero fuel weight in kilograms
vehicle.envelope.limit_load = 1.5 # Limit Load
turbofan = SUAVE.Components.Propulsors.TurboFanPASS()
turbofan.tag = 'turbo_fan'
turbofan.number_of_engines = 2. # Number of engines on the aircraft
turbofan.design_thrust = 200. * Units.newton # Define Thrust in Newtons
vehicle.append_component(turbofan)
vehicle.passengers = 170. # Number of passengers
vehicle.mass_properties.cargo = 0. * Units.kilogram # Mass of cargo
vehicle.systems.control = "fully powered" # Specify fully powered, partially powered or anything else is fully aerodynamic
vehicle.systems.accessories = "medium-range" # Specify what type of aircraft you have
vehicle.reference_area = 124.862 * Units.meter**2 # Wing gross area in square meters
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.spans.projected = 50. * Units.meter # Span in meters
wing.taper = 0.2 # Taper ratio
wing.thickness_to_chord = 0.08 # Thickness-to-chord ratio
wing.sweep = .4363323 * Units.rad # sweep angle in degrees
wing.chords.root = 15. * Units.meter # Wing root chord length
wing.chords.mean_aerodynamic = 10. * Units.meters # Length of the mean aerodynamic chord of the wing
wing.origin = [20,0,0] * Units.meters # Location of main wing from origin of the vehicle
wing.aerodynamic_center = [3,0,0] * Units.meters # Location of aerodynamic center from origin of the main wing
vehicle.append_component(wing)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.areas.wetted = 688.64 * Units.meter**2 # Fuselage wetted area
fuselage.differential_pressure = 55960.5 * Units.pascal # Maximum differential pressure
fuselage.width = 4. * Units.meter # Width of the fuselage
fuselage.heights.maximum = 4. * Units.meter # Height of the fuselage
fuselage.lengths.total = 58.4 * Units.meter # Length of the fuselage
fuselage.number_coach_seats = 200.
vehicle.append_component(fuselage)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.areas.reference = 75. * Units.meters**2 # Area of the horizontal tail
wing.spans.projected = 15. * Units.meters # Span of the horizontal tail
wing.sweep = 38. * Units.deg # Sweep of the horizontal tail
wing.chords.mean_aerodynamic = 5. * Units.meters # Length of the mean aerodynamic chord of the horizontal tail
wing.thickness_to_chord = 0.07 # Thickness-to-chord ratio of the horizontal tail
wing.areas.exposed = 199.7792 # Exposed area of the horizontal tail
wing.areas.wetted = 249.724 # Wetted area of the horizontal tail
wing.origin = [45,0,0] # Location of horizontal tail from origin of the vehicle
wing.aerodynamic_center = [3,0,0] # Location of aerodynamic center from origin of the horizontal tail
vehicle.append_component(wing)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.areas.reference = 60. * Units.meters**2 # Area of the vertical tail
wing.spans.projected = 15. * Units.meters # Span of the vertical tail
wing.thickness_to_chord = 0.07 # Thickness-to-chord ratio of the vertical tail
wing.sweep = 40. * Units.deg # Sweep of the vertical tail
wing.t_tail = "false" # Set to "yes" for a T-tail
vehicle.append_component(wing)
weight = Tube_Wing.empty(vehicle)
actual = Data()
actual.payload = 17349.9081525
actual.pax = 15036.5870655
actual.bag = 2313.321087
actual.fuel = -6993.89102491
actual.empty = 68659.7828724
actual.wing = 27694.192985
actual.fuselage = 11504.5186408
actual.propulsion = 88.3696093424
actual.landing_gear = 3160.632
actual.systems = 16655.7076511
actual.wt_furnish = 7466.1304102
actual.horizontal_tail = 2191.30720639
actual.vertical_tail = 5260.75341411
actual.rudder = 2104.30136565
error = Data()
error.payload = (actual.payload - weight.payload)/actual.payload
error.pax = (actual.pax - weight.pax)/actual.pax
error.bag = (actual.bag - weight.bag)/actual.bag
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - weight.landing_gear)/actual.landing_gear
error.systems = (actual.systems - weight.systems)/actual.systems
error.wt_furnish = (actual.wt_furnish - weight.wt_furnish)/actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder)/actual.rudder
for k,v in error.items():
#.........这里部分代码省略.........