Python Data.vertical_tail方法代码示例

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import vertical_tail [as 别名]

#.........这里部分代码省略.........
            cts -      average tail surface chord                                      [meters]
            deltawts - average rib spacing to average chord ratio                      [dimensionless]
            Ntsr -     number of tail surface ribs (bts^2)/(deltats*Sts)               [dimensionless]
            t_cts -    tail airfoil thickness to chord ratio                           [dimensionless]
            
        vertical - a data dictionary with the fields:
            Sts -      tail surface area                                               [meters]
            bts -      tail surface span                                               [meters]
            cts -      average tail surface chord                                      [meters]
            deltawts - average rib spacing to average chord ratio                      [dimensionless]
            Ntsr -     number of tail surface ribs (bts**2)/(deltats*Sts)              [dimensionless]
            t_cts -    tail airfoil thickness to chord ratio                           [dimensionless]
            
        aircraft - a data dictionary with the fields:    
            nult -     ultimate load factor                                            [dimensionless]
            GW -       aircraft gross weight                                           [kilogram]
            qm -       dynamic pressure at maneuvering speed                           [Pascals]
            Ltb -      tailboom length                                                 [meters]
    
    Outputs:
        Wws -      weight of wing spar                                                 [kilogram]
        Wtss -     weight of tail surface spar                                         [kilogram]
        Wwr -      weight of wing ribs                                                 [kilogram]
        Wtsr -     weight of tail surface ribs                                         [kilogram]
        Wwer -     weight of wing end ribs                                             [kilogram]
        WwLE -     weight of wing leading edge                                         [kilogram]
        WtsLE -    weight of tail surface leading edge                                 [kilogram]
        WwTE -     weight of wing trailing edge                                        [kilogram]
        Wwc -      weight of wing covering                                             [kilogram]
        Wtsc -     weight of tail surface covering                                     [kilogram]
        Wtb -      tailboom weight                                                     [kilogram]
                
    Properties Used:
        N/A
    """ 
    
    #Unpack
    
    nult   = vehicle.envelope.ultimate_load
    gw     = vehicle.mass_properties.max_takeoff
    qm     = vehicle.envelope.maximum_dynamic_pressure
    
    # Wing weight
    if not vehicle.wings.has_key('main_wing'):
        wt_wing = 0.0
        warnings.warn("There is no Wing Weight being added to the Configuration", stacklevel=1)
    else:
        Sw      = vehicle.wings['main_wing'].areas.reference
        bw      = vehicle.wings['main_wing'].spans.projected
        cw      = vehicle.wings['main_wing'].chords.mean_aerodynamic
        Nwr     = vehicle.wings['main_wing'].number_ribs
        t_cw    = vehicle.wings['main_wing'].thickness_to_chord
        Nwer    = vehicle.wings['main_wing'].number_end_ribs
        wt_wing = wing.wing(Sw,bw,cw,Nwr,t_cw,Nwer,nult,gw)
        vehicle.wings['main_wing'].mass_properties.mass = wt_wing
    
    # Horizontal Tail weight
    if not vehicle.wings.has_key('horizontal_stabilizer'):
        wt_ht = 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
        chs    = vehicle.wings['horizontal_stabilizer'].chords.mean_aerodynamic
        Nhsr   = vehicle.wings['horizontal_stabilizer'].number_ribs
        t_ch   = vehicle.wings['horizontal_stabilizer'].thickness_to_chord
        wt_ht  = tail.tail(S_h,b_h,chs,Nhsr,t_ch,qm)
        vehicle.wings['horizontal_stabilizer'].mass_properties.mass = wt_ht
    
    # Vertical Tail weight
    if not vehicle.wings.has_key('vertical_stabilizer'):   
        wt_vt = 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
        cvs    = vehicle.wings['vertical_stabilizer'].chords.mean_aerodynamic
        Nvsr   = vehicle.wings['vertical_stabilizer'].number_ribs
        t_cv   = vehicle.wings['vertical_stabilizer'].thickness_to_chord
        wt_vt   = tail.tail(S_v,b_v,cvs,Nvsr,t_cv,qm)
        vehicle.wings['vertical_stabilizer'].mass_properties.mass = wt_vt

    # Fuselage weight
    if not vehicle.fuselages.has_key('fuselage'):   
        wt_tb = 0.0
        warnings.warn("There is no Fuselage Weight being added to the Configuration", stacklevel=1)    
    else: 
        Ltb     = vehicle.Ltb  
        wt_tb   = fuselage.fuselage(S_h,qm,Ltb)
        vehicle.Fuselages['fuselage'].mass_properties.mass = wt_tb
    
    weight                 = Data()
    weight.wing            = wt_wing
    weight.fuselage        = wt_tb
    weight.horizontal_tail = wt_ht
    weight.vertical_tail   = wt_vt
    
    weight.empty = wt_ht + wt_tb + wt_vt + wt_wing
    
    return weight

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import vertical_tail [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.........

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import vertical_tail [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

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import vertical_tail [as 别名]

#.........这里部分代码省略.........
    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 = -13680.6265874
    actual.empty = 75346.5184349
    actual.wing = 27694.192985
    actual.fuselage = 11423.9380852
    actual.propulsion = 6855.68572746 
    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.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)<0.001)    
   
    
    return

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import vertical_tail [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.........

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import vertical_tail [as 别名]

#.........这里部分代码省略.........
                bts -      tail surface span (m)
                cts -      average tail surface chord (m)
                deltawts - average rib spacing to average chord ratio
                Ntsr -     number of tail surface ribs (bts**2)/(deltats*Sts)
                t_cts -    tail airfoil thickness to chord ratio
                
            aircraft - a data dictionary with the fields:    
                nult -     ultimate load factor
                GW -       aircraft gross weight
                qm -       dynamic pressure at maneuvering speed (N/m2)
                Ltb -      tailboom length (m)
        
            Outputs:
                Wws -      weight of wing spar (kg)
                Wtss -     weight of tail surface spar (kg)
                Wwr -      weight of wing ribs (kg)
                Wtsr -     weight of tail surface ribs (kg)
                Wwer -     weight of wing end ribs (kg)
                WwLE -     weight of wing leading edge (kg)
                WtsLE -    weight of tail surface leading edge (kg)
                WwTE -     weight of wing trailing edge (kg)
                Wwc -      weight of wing covering (kg)
                Wtsc -     weight of tail surface covering (kg)
                Wtb -      tailboom weight (kg)
                    
            Assumptions:
                All of this is from AIAA 89-2048, units are in kg. These weight estimates
                are from the MIT Daedalus and are valid for very lightweight
                carbon fiber composite structures. This may need to be solved iteratively since
                gross weight is an input.
                
        """
    
    #Unpack
    
    nult   = vehicle.envelope.ultimate_load
    gw     = vehicle.mass_properties.max_takeoff
    qm     = vehicle.qm
    
    # Wing weight
    if not vehicle.wings.has_key('main_wing'):
        wt_wing = 0.0
        warnings.warn("There is no Wing Weight being added to the Configuration", stacklevel=1)
    else:
        Sw      = vehicle.wings['main_wing'].areas.reference
        bw      = vehicle.wings['main_wing'].spans.projected
        cw      = vehicle.wings['main_wing'].chords.mean_aerodynamic
        Nwr     = vehicle.wings['main_wing'].number_ribs
        t_cw    = vehicle.wings['main_wing'].thickness_to_chord
        Nwer    = vehicle.wings['main_wing'].number_end_ribs
        wt_wing = wing.wing(Sw,bw,cw,Nwr,t_cw,Nwer,nult,gw)
        vehicle.wings['main_wing'].mass_properties.mass = wt_wing
    
    # Horizontal Tail weight
    if not vehicle.wings.has_key('horizontal_stabilizer'):
        wt_ht = 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
        chs    = vehicle.wings['horizontal_stabilizer'].chords.mean_aerodynamic
        Nhsr   = vehicle.wings['horizontal_stabilizer'].number_ribs
        t_ch   = vehicle.wings['horizontal_stabilizer'].thickness_to_chord
        wt_ht  = tail.tail(S_h,b_h,chs,Nhsr,t_ch,qm)
        vehicle.wings['horizontal_stabilizer'].mass_properties.mass = wt_ht
    
    # Vertical Tail weight
    if not vehicle.wings.has_key('vertical_stabilizer'):   
        wt_vt = 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
        cvs    = vehicle.wings['vertical_stabilizer'].chords.mean_aerodynamic
        Nvsr   = vehicle.wings['vertical_stabilizer'].number_ribs
        t_cv   = vehicle.wings['vertical_stabilizer'].thickness_to_chord
        wt_vt   = tail.tail(S_v,b_v,cvs,Nvsr,t_cv,qm)
        vehicle.wings['vertical_stabilizer'].mass_properties.mass = wt_vt

    ##Fuselage weight
    #Ltb     = vehicle.Ltb  
    #wt_tb   = fuselage.fuselage(S_h,qm,Ltb)
    #vehicle.Fuselages['fuselage'].mass_properties.mass = wt_tb
    
    weight                 = Data()
    weight.wing            = wt_wing
    #weight.fuselage        = wt_tb
    weight.horizontal_tail = wt_ht
    weight.vertical_tail   = wt_vt
    
    return weight