本文整理汇总了Python中SUAVE.Structure.Data.lift_curve_slope方法的典型用法代码示例。如果您正苦于以下问题：Python Data.lift_curve_slope方法的具体用法？Python Data.lift_curve_slope怎么用？Python Data.lift_curve_slope使用的例子？那么恭喜您, 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.Structure.Data的用法示例。
在下文中一共展示了Data.lift_curve_slope方法的2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.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import lift_curve_slope [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.chords.mean_aerodynamic = 27.3 * 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([58.6,0,0]) * Units.feet
wing.aerodynamic_center = np.array([112., 0. , 0. ]) * Units.feet- wing.origin
wing.eta = 1.0
wing.downwash_adj = 1.0
wing.ep_alpha = 1. - wing.downwash_adj
Mach = np.array([0.198])
reference = SUAVE.Structure.Container()
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
wing.CL_alpha = conditions.lift_curve_slope
vehicle.reference_area = wing.areas.reference
vehicle.append_component(wing)
lifting_surfaces = []
lifting_surfaces.append(wing)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Horizontal Stabilizer'
wing.areas.reference = 1490.55* Units.feet**2
wing.spans.projected = 71.6 * Units.feet
wing.sweep = 44.0 * Units.deg # leading edge
wing.taper = 7.5/32.6
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([187.0,0,0]) * Units.feet
wing.symmetric= True
wing.eta = 0.95
wing.downwash_adj = 1.0 - 2.0*vehicle.wings['Main Wing'].CL_alpha/np.pi/wing.aspect_ratio
wing.ep_alpha = 1. - wing.downwash_adj
wing.aerodynamic_center = [trapezoid_ac_x(wing), 0.0, 0.0] - wing.origin
wing.CL_alpha = datcom(wing,Mach)
vehicle.append_component(wing)
lifting_surfaces.append(wing)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'Fuselage'
fuselage.x_root_quarter_chord = 77.0 * Units.feet
fuselage.lengths.total = 229.7 * Units.feet
fuselage.width = 20.9 * 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]) * Units.feet
#Method Test
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = 0.93 # Should be -1.45
error = Data()
error.cm_a_747 = (cm_a - expected)/expected
#Parameters Required
#Using values for a Beech 99
vehicle = SUAVE.Vehicle()
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Main Wing'
wing.areas.reference = 280.0 * Units.feet**2
wing.spans.projected = 46.0 * Units.feet
wing.chords.mean_aerodynamic = 6.5 * Units.feet
wing.sweep = 3.0 * Units.deg # Leading edge
wing.taper = 0.47
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.origin = np.array([14.0,0,0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0. , 0. ]) - wing.origin
wing.eta = 1.0
wing.downwash_adj = 1.0
wing.ep_alpha = 1. - wing.downwash_adj
Mach = np.array([0.152])
reference = SUAVE.Structure.Container()
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
wing.CL_alpha = conditions.lift_curve_slope
vehicle.reference_area = wing.areas.reference
vehicle.append_component(wing)
lifting_surfaces = []
lifting_surfaces.append(wing)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Horizontal Stabilizer'
wing.areas.reference = 100.5 * Units.feet**2
wing.spans.projected = 22.5 * Units.feet
#.........这里部分代码省略.........
示例2: Data
# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import lift_curve_slope [as 别名]
wing.spans.projected = 196.0 * Units.feet
wing.chords.mean_aerodynamic = 27.3 * Units.feet
wing.chords.root = 44. * Units.feet #54.5ft
wing.sweep = 42.0 * Units.deg # Leading edge
wing.taper = 13.85/44. #14.7/54.5
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([59.,0,0]) * Units.feet
wing.aerodynamic_center = np.array([112.2*Units.feet,0.,0.])-wing.origin#16.16 * Units.meters,0.,0,])np.array([trapezoid_ac_x(wing),0., 0.])#
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
Mach = np.array([0.198])
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
wing.CL_alpha = conditions.lift_curve_slope
vehicle.reference_area = wing.areas.reference
vehicle.append_component(wing)
main_wing_CLa = wing.CL_alpha
main_wing_ar = wing.aspect_ratio
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Horizontal Stabilizer'
wing.areas.reference = 1490.55* Units.feet**2
wing.spans.projected = 71.6 * Units.feet
wing.sweep = 44.0 * Units.deg # leading edge
wing.taper = 7.5/32.6
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([187.0,0,0]) * Units.feet