本文整理汇总了Python中SUAVE.Structure.Data.mass_properties方法的典型用法代码示例。如果您正苦于以下问题：Python Data.mass_properties方法的具体用法？Python Data.mass_properties怎么用？Python Data.mass_properties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.mass_properties方法的3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 mass_properties [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.........
示例2: main
# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import mass_properties [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3: datcom
# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import mass_properties [as 别名]
wing.vertical = False
wing.dynamic_pressure_ratio = 0.95
wing.ep_alpha = 2.0*main_wing_CLa/np.pi/main_wing_ar
wing.aerodynamic_center = [trapezoid_ac_x(wing), 0.0, 0.0]
wing.CL_alpha = datcom(wing,Mach)
vehicle.append_component(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([112.2,0,0]) * Units.feet
#Method Test
print '<<Test run of the taw_cmalpha() method>>'
print 'Boeing 747 at Mach {0}'.format(Mach[0])
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = -1.45
print 'Cm_alpha = {0:.4f}'.format(cm_a[0])
print 'Expected value = {}'.format(expected)
print 'Percent Error = {0:.2f}%'.format(100.0*(cm_a[0]-expected)/expected)
print 'Static Margin = {0:.4f}'.format(-cm_a[0]/vehicle.wings['Main Wing'].CL_alpha[0])
print ' '