Python Data.empty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 empty [as 别名]
def Empty(vehicle):
    
    """
    Structural Weight correlation from all 415 samples of fixed-wing UAVs and sailplanes
    Equation 3.16 from 'Design of Solar Powered Airplanes for Continuous Flight' by Andre Noth
    Relatively valid for a wide variety of vehicles, may be optimistic
    Assumes a 'main wing' is attached
    
    """
    
    # Unpack
    S     = vehicle.reference_area
    AR    = vehicle.wings['main_wing'].aspect_ratio
    Earth = SUAVE.Attributes.Planets.Earth()
    g     = Earth.sea_level_gravity
    
    
    # Airframe weight
    Waf = (5.58*(S**1.59)*(AR**0.71))/g # All Samples
    #Waf = (0.44*(S**1.55)*(AR**1.30))/g  # Top 19
    
    # Pack
    weight = Data()
    weight.empty = Waf
    
    return weight

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import empty [as 别名]
def payload(TOW, empty, num_pax, wt_cargo, wt_passenger = 195*Units.lbs,wt_baggage = 30*Units.lbs):
    """ Calculate the weight of the payload and the resulting fuel mass
    
    Assumptions:
        based on FAA guidelines for weight of passengers
        
    Source: 
        N/A
        
    Inputs:
        TOW -                                                              [kilograms]
        wt_empty - Operating empty weight of the aircraft                  [kilograms]
        num_pax - number of passengers on the aircraft                     [dimensionless]
        wt_cargo - weight of cargo being carried on the aircraft           [kilogram]
        wt_passenger - weight of each passenger on the aircraft            [kilogram]
        wt_baggage - weight of the baggage for each passenger              [kilogram]
    
    Outputs:
        output - a data dictionary with fields:
            payload - weight of the passengers plus baggage and paid cargo [kilograms]
            pax - weight of all the passengers                             [kilogram]
            bag - weight of all the baggage                                [kilogram]
            fuel - weight of the fuel carried                              [kilogram]
            empty - operating empty weight of the aircraft                 [kilograms]
               
    Properties Used:
        N/A
    """ 
    
    # process
    wt_pax     = wt_passenger * num_pax
    wt_bag     = wt_baggage * num_pax 
    wt_payload = wt_pax + wt_bag + wt_cargo
    wt_fuel    = TOW - wt_payload - empty
    
    # packup outputs
    output = Data()
    output.payload = wt_payload
    output.pax     = wt_pax   
    output.bag     = wt_bag
    output.fuel    = wt_fuel
    output.empty   = empty
  
    return output

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import empty [as 别名]
def empty(vehicle):
    """ This comoutes the weigt of a UAV   
    
    Assumptions:
        Assumes a 'main wing' is attached

    Source:
        Structural Weight correlation from all 415 samples of fixed-wing UAVs and sailplanes
        Equation 3.16 from 'Design of Solar Powered Airplanes for Continuous Flight' by Andre Noth
        Relatively valid for a wide variety of vehicles, may be optimistic 

    Inputs:
        S                [meters**2]
        AR               [dimensionless]
        
    Outputs:
        weight           [kilograms]

    Properties Used:
        N/A
    """    
    
    # Unpack
    S     = vehicle.reference_area
    AR    = vehicle.wings['main_wing'].aspect_ratio
    Earth = SUAVE.Attributes.Planets.Earth()
    g     = Earth.sea_level_gravity
    
    
    # Airframe weight
    Waf = (5.58*(S**1.59)*(AR**0.71))/g # All Samples
    #Waf = (0.44*(S**1.55)*(AR**1.30))/g  # Top 19
    
    # Pack
    weight = Data()
    weight.empty = Waf
    
    return weight

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import empty [as 别名]
def empty(config,
          speed_of_sound                = 340.294,
          max_tip_mach                  = 0.65,
          disk_area_factor              = 1.15,
          max_thrust_to_weight_ratio    = 1.1,
          motor_efficiency              = 0.85 * 0.98):
    """weight = SUAVE.Methods.Weights.Buildups.Electric_Stopped_Rotor.empty(
            config,
            speed_of_sound              = 340.294,
            max_tip_mach                = 0.65,
            disk_area_factor            = 1.15,
            max_thrust_to_weight_ratio  = 1.1,
            motor_efficiency            = 0.85 * 0.98)

        Calculates the empty fuselage mass for an electric stopped rotor including
        seats, avionics, servomotors, ballistic recovery system, rotor and hub
        assembly, transmission, and landing gear. Additionally incorporates
        results of the following common buildup scripts:

            fuselage,py
            prop.py
            wing.py
            wiring.py

        Originally written as part of an AA 290 project inteded for trade study
        of the Electric Stopped Rotor along with the following defined SUAVE config types:

            Electric Helicopter
            Electric Tiltrotor
            
        Sources:
        Project Vahana Conceptual Trade Study

        Inputs:

            config                          SUAVE Config Data Structure
            speed_of_sound                  Local Speed of Sound                [m/s]
            maximumTipMach                  Allowable Tip Mach Number           [Unitless]
            disk_area_factor                Inverse of Disk Area Efficiency     [Unitless]
            max_thrust_to_weight_ratio      Allowable Thrust to Weight Ratio    [Unitless]
            motor_efficiency                Motor Efficiency                    [Unitless]
        
        Outputs:

            output:                         Data Dictionary of Component Masses       [kg]

    """

    output = Data()

#-------------------------------------------------------------------------------
# Unpack Inputs
#-------------------------------------------------------------------------------

    rProp               = config.propulsors.network.propeller.prop_attributes.tip_radius
    mBattery            = config.propulsors.network.battery.mass_properties.mass
    mPayload            = config.propulsors.network.payload.mass_properties.mass
    MTOW                = config.mass_properties.max_takeoff
    nLiftProps          = config.propulsors.network.number_of_engines/2
    nThrustProps        = config.propulsors.network.number_of_engines/2
    nLiftBlades         = config.propulsors.network.propeller.prop_attributes.number_blades
    nThrustBlades       = config.propulsors.network.propeller.prop_attributes.number_blades
    fLength             = config.fuselages.fuselage.lengths.total
    fWidth              = config.fuselages.fuselage.width
    fHeight             = config.fuselages.fuselage.heights.maximum
    maxSpan             = config.wings['main_wing'].spans.projected
    
    sound               = speed_of_sound
    tipMach             = max_tip_mach
    k                   = disk_area_factor
    ToverW              = max_thrust_to_weight_ratio
    etaMotor            = motor_efficiency

    output.payload          = mPayload * Units.kg
    output.seats            = 30. *Units.kg
    output.avionics         = 15. *Units.kg
    output.motors           = config.propulsors.network.number_of_engines * 10. *Units.kg
    output.battery          = mBattery *Units.kg
    output.servos           = config.propulsors.network.number_of_engines * 0.65 *Units.kg
    output.brs              = 16. *Units.kg
    output.hubs             = config.propulsors.network.number_of_engines * 2. *Units.kg
    output.landing_gear     = MTOW * 0.02 *Units.kg

#-------------------------------------------------------------------------------
# Calculated Weights
#-------------------------------------------------------------------------------

    # Preparatory Calculations

    Vtip        = sound * tipMach                               # Prop Tip Velocity
    omega       = Vtip/0.8                                      # Prop Ang. Velocity
    maxLift     = config.mass_properties.max_takeoff * ToverW   # Maximum Thrust
    Ct          = maxLift/(1.225*np.pi*0.8**2*Vtip**2)          # Thrust Coefficient
    bladeSol    = 0.1                                           # Blade Solidity
    AvgCL       = 6 * Ct / bladeSol                             # Average Blade CL
    AvgCD       = 0.012                                         # Average Blade CD
   
    maxLiftPower    = 1.15*maxLift*(
                    k*np.sqrt(maxLift/(2*1.225*np.pi*0.8**2)) +
                    bladeSol*AvgCD/8*Vtip**3/(maxLift/(1.225*np.pi*0.8**2))
#.........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 empty [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 empty [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 empty [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 empty [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.........