Python Data.sweep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 sweep [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 sweep [as 别名]
def main():
    #Parameters Required
        #Using values for a Boeing 747-200
    vehicle = SUAVE.Vehicle()
    wing = SUAVE.Components.Wings.Wing()
    wing.tag = 'Main Wing'
    wing.areas.reference           = 5500.0 * Units.feet**2
    wing.spans.projected           = 196.0  * Units.feet
    wing.sweep       = 42.0   * Units.deg # Leading edge
    wing.taper          = 14.7/54.5
    wing.aspect_ratio   = wing.spans.projected**2/wing.areas.reference
    wing.symmetric      = True
    wing.origin           = np.array([0.0,0,3.6]) * Units.feet  
    
    reference               = SUAVE.Structure.Container()
    vehicle.reference_area   = wing.areas.reference
    vehicle.append_component(wing)
    
    lifting_surfaces    = []
    lifting_surfaces.append(wing)
    
    wing          = SUAVE.Components.Wings.Wing()
    wing.tag = 'Vertical Stabilizer'
    vertical = Data()
    vertical.span         = 32.4   * Units.feet
    vertical.root_chord   = 38.7   * Units.feet
    vertical.tip_chord    = 13.4   * Units.feet
    vertical.sweep     = 50.0   * Units.deg
    vertical.x_root_LE1   = 180.0  * Units.feet
    vertical.symmetric    = False
    dz_centerline         = 13.3   * Units.feet
    ref_vertical          = extend_to_ref_area(vertical,dz_centerline)    
    wing.areas.reference     = ref_vertical.ref_area
    wing.spans.projected     = ref_vertical.ref_span
    wing.sweep = 50.0   * Units.deg # leading edge
    wing.taper    = vertical.tip_chord/ref_vertical.ref_root_chord
    wing.aspect_ratio = ref_vertical.ref_aspect_ratio
    wing.origin     = np.array([vertical.x_root_LE1 + ref_vertical.root_LE_change,0.,0.]) * Units.feet
    wing.effective_aspect_ratio = 2.2
    wing.symmetric= True
    wing.aerodynamic_center = np.array([trapezoid_ac_x(wing),0.0,0.0])
    Mach = np.array([0.198])
    wing.CL_alpha = datcom(wing,Mach)   
    vehicle.append_component(wing)
    lifting_surfaces.append(wing)
    
    fuselage = SUAVE.Components.Fuselages.Fuselage()
    fuselage.tag = 'Fuselage'
    fuselage.areas.side_projected = 4696.16 * Units.feet**2
    fuselage.lengths.total = 229.7 * Units.feet
    fuselage.heights.maximum = 26.9 * Units.feet
    fuselage.width = 20.9    * Units.feet
    fuselage.heights.at_quarter_length = 26   * Units.feet
    fuselage.heights.at_three_quarters_length = 19.7 * Units.feet
    fuselage.heights.at_wing_root_quarter_chord = 15.8 * Units.feet
    vehicle.append_component(fuselage)
    
    configuration = Data()
    configuration.mass_properties = Data()
    configuration.mass_properties.center_of_gravity = Data()
    configuration.mass_properties.center_of_gravity = np.array([112.0,0,0]) * Units.feet    
    
    segment            = SUAVE.Attributes.Missions.Segments.Base_Segment()
    segment.freestream = Data()
    segment.freestream.mach_number          = 0.198
    segment.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
    altitude           = 0.0 * Units.feet
    segment.a          = segment.atmosphere.compute_values(altitude / Units.km, type="a")
    segment.freestream.density        = segment.atmosphere.compute_values(altitude / Units.km, type="rho")
    segment.freestream.viscosity        = segment.atmosphere.compute_values(altitude / Units.km, type="mew")
    segment.freestream.velocity      = segment.freestream.mach_number * segment.a    
    
    #Method Test   
    cn_b = taw_cnbeta(vehicle,segment,configuration)
    expected = -0.35 # Should be 0.184
    error = Data()
    error.cn_b_747 = (cn_b-expected)/expected  
    
    #Parameters Required
    #Using values for a Beechcraft Model 99
    #MODEL DOES NOT ACCOUNT FOR DESTABILIZING EFFECTS OF PROPELLERS!
    """wing               = SUAVE.Components.Wings.Wing()
    wing.area          = 280.0 * Units.feet**2
    wing.span          = 46.0  * Units.feet
    wing.sweep_le      = 3.0   * Units.deg
    wing.z_position    = 2.2   * Units.feet
    wing.taper         = 0.46
    wing.aspect_ratio  = wing.span**2/wing.area
    wing.symmetric     = True
    
    fuselage           = SUAVE.Components.Fuselages.Fuselage()
    fuselage.side_area = 185.36 * Units.feet**2
    fuselage.length    = 44.0   * Units.feet
    fuselage.h_max     = 6.0    * Units.feet
    fuselage.w_max     = 5.4    * Units.feet
    fuselage.height_at_vroot_quarter_chord   = 2.9 * Units.feet
    fuselage.height_at_quarter_length        = 4.8 * Units.feet
    fuselage.height_at_three_quarters_length = 4.3 * Units.feet
    
    nacelle           = SUAVE.Components.Fuselages.Fuselage()
#.........这里部分代码省略.........

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import sweep [as 别名]
    else:
        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)

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import sweep [as 别名]
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
horizontal.mac     = 5.      * Units.meters    # Length of the mean aerodynamic chord of the horizontal tail