Python Data.t_c方法代码示例

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import t_c [as 别名]
def main():

    # new units style
    a = 4 * Units.mm # convert into base units
    b = a / Units.mm # convert out of base units

    engine = Data()
    wing = Data()
    aircraft = Data()
    fuselage = Data()
    horizontal = Data()
    vertical = Data()

    # Parameters Required
    aircraft.Nult      = 1.5 * 2.5                       # Ultimate load
    aircraft.TOW       = 52300.  * Units.kilograms # Maximum takeoff weight in kilograms
    aircraft.zfw       = 42600. * Units.kilograms # Maximum zero fuel weight in kilograms
    aircraft.Nlim      = 2.5                       # Limit Load
    aircraft.num_eng   = 2.                        # Number of engines on the aircraft
    aircraft.num_pax   = 110.                      # Number of passengers
    aircraft.wt_cargo  = 0.  * Units.kilogram  # Mass of cargo
    aircraft.num_seats = 110.                      # Number of seats on aircraft
    aircraft.ctrl      = "partially powered"       # Specify fully powered, partially powered or anything else is fully aerodynamic
    aircraft.ac        = "medium-range"              # Specify what type of aircraft you have
    aircraft.w2h       = 16.     * Units.meters    # Length from the mean aerodynamic center of wing to mean aerodynamic center of the horizontal tail

    wing.gross_area    = 92.    * Units.meter**2  # Wing gross area in square meters
    wing.span          = 27.8     * Units.meter     # Span in meters
    wing.taper         = 0.28                       # Taper ratio
    wing.t_c           = 0.105                      # Thickness-to-chord ratio
    wing.sweep         = 23.5     * Units.deg       # sweep angle in degrees
    wing.c_r           = 5.4     * Units.meter     # Wing exposed root chord length
    wing.mac           = 12.     * Units.ft    # Length of the mean aerodynamic chord of the wing

    fuselage.area      = 320.      * Units.meter**2  # Fuselage wetted area 
    fuselage.diff_p    = 8.5     * Units.force_pound / Units.inches**2    # Maximum differential pressure
    fuselage.width     = 3.      * Units.meter     # Width of the fuselage
    fuselage.height    = 3.35    * Units.meter     # Height of the fuselage
    fuselage.length    = 36.24     * Units.meter     # Length of the fuselage

    engine.thrust_sls  = 18500.   * Units.force_pound    # Define Thrust in Newtons

    horizontal.area    = 26.     * Units.meters**2 # Area of the horizontal tail
    horizontal.span    = 12.08     * Units.meters    # Span of the horizontal tail
    horizontal.sweep   = 34.5     * Units.deg       # Sweep of the horizontal tail
    horizontal.mac     = 2.4      * Units.meters    # Length of the mean aerodynamic chord of the horizontal tail
    horizontal.t_c     = 0.11                      # Thickness-to-chord ratio of the horizontal tail
    horizontal.exposed = 0.9                         # Fraction of horizontal tail area exposed

    vertical.area      = 16.     * Units.meters**2 # Area of the vertical tail
    vertical.span      = 5.3     * Units.meters    # Span of the vertical tail
    vertical.t_c       = 0.12                      # Thickness-to-chord ratio of the vertical tail
    vertical.sweep     = 35.     * Units.deg       # Sweep of the vertical tail
    vertical.t_tail    = "no"                      # Set to "yes" for a T-tail

    aircraft.weight = Tube_Wing.empty(engine,wing,aircraft,fuselage,horizontal,vertical)

    outputWeight(aircraft,'weight_EMB190.dat')

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import t_c [as 别名]
    Data, Container, Data_Exception, Data_Warning,
)

import empty as empty

wing       = Data()
horizontal = Data()
vertical   = Data()
aircraft   = Data()

wing.sref   = 31.00
wing.span   = 34.10
wing.mac    = wing.sref/wing.span
wing.Nwr    = 2*41+22
wing.deltaw = (wing.span**2)/(wing.Nwr*wing.sref)
wing.t_c    = 0.128
wing.Nwer   = 10.

horizontal.area   = 10. #FAKE NUMBER!!!!!
horizontal.span   = wing.sref*((1.+3./16.)/9.)
horizontal.mac    = horizontal.area/horizontal.span
horizontal.Nwr    = 16.
horizontal.deltah = (horizontal.span**2)/(horizontal.Nwr*horizontal.area)
horizontal.t_c    = 0.12 # I have no idea

vertical.area   = 10. #FAKE NUMBER!!!!!
vertical.span   = wing.sref*((11./16.)/9.)
vertical.mac    = horizontal.area/horizontal.span
vertical.Nwr    = 10.
vertical.deltah = (vertical.span**2)/(vertical.Nwr*vertical.area)
vertical.t_c    = 0.12 # I have no idea

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import t_c [as 别名]
        C     = 1.1
        beta2 = (1.0 -(M*cosL)**2)
        k1    = 2*C * cosL**2 * t_c / np.sqrt(beta2)
        k2    = C**2 * cosL**2 * t_c**2 * (1+5*cosL**2) / (2*beta2)
        ff    = 1 + k1 + k2
    
    return ff



if __name__ == '__main__':
    print 'RUNNING TEST'
    
    wing       = Data()
    wing.sweep = 25 * Units.deg
    wing.t_c   = 0.1
    
    Mach       = 0.75
    
    #Method Test
    print '<<Test run of the wing_form_factor() method>> \n'
    print 'Wing at Mach {}'.format(Mach)
    print 't/c = {0},  sweep = {1} degrees'.format(wing.t_c,wing.sweep/Units.deg)
    
    form_factor = wing_form_factor(wing, Mach)
    #Compare to AA241 Notes - http://adg.stanford.edu/aa241/AircraftDesign.html    
    expected    = 1.294
    
    print ' Form factor    = {0:.3f}m'.format(form_factor)
    print ' Expected value = {}m'.format(expected)
    print ' Percent Error  = {0:.2f}%'.format(100.0*(form_factor-expected)/expected)

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import t_c [as 别名]
aircraft.Nult      = 3.5                       # Ultimate load
aircraft.TOW       = 200000. * Units.kilograms # Maximum takeoff weight in kilograms
aircraft.zfw       = 150000. * Units.kilograms # Maximum zero fuel weight in kilograms
aircraft.Nlim      = 1.5                       # Limit Load
aircraft.num_eng   = 2.                        # Number of engines on the aircraft
aircraft.num_pax   = 125.                      # Number of passengers
aircraft.wt_cargo  = 10000.  * Units.kilogram  # Mass of cargo
aircraft.num_seats = 125.                      # Number of seats on aircraft
aircraft.ctrl      = "fully powered"           # Specify fully powered, partially powered or anything else is fully aerodynamic
aircraft.ac        = "long-range"              # Specify what type of aircraft you have
aircraft.w2h       = 20.     * Units.meters    # Length from the mean aerodynamic center of wing to mean aerodynamic center of the horizontal tail

wing.gross_area    = 500.    * Units.meter**2  # Wing gross area in square meters
wing.span          = 50.     * Units.meter     # Span in meters
wing.taper         = 0.2                       # Taper ratio
wing.t_c           = 0.08                      # Thickness-to-chord ratio
wing.sweep         = 35.     * Units.deg       # sweep angle in degrees
wing.c_r           = 15.     * Units.meter     # Wing root chord length
wing.mac           = 10.     * Units.meters    # Length of the mean aerodynamic chord of the wing

fuselage.area      = 10.     * Units.meter**2  # Fuselage cross-sectional area 
fuselage.diff_p    = 10**5   * Units.pascal    # Maximum differential pressure
fuselage.width     = 5.      * Units.meter     # Width of the fuselage
fuselage.height    = 4.5     * Units.meter     # Height of the fuselage
fuselage.length    = 60.     * Units.meter     # Length of the fuselage

engine.thrust_sls  = 1000.   * Units.newton    # Define Thrust in Newtons

horizontal.area    = 75.     * Units.meters**2 # Area of the horizontal tail
horizontal.span    = 15.     * Units.meters    # Span of the horizontal tail
horizontal.sweep   = 38.     * Units.deg       # Sweep of the horizontal tail