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')
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.Core.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.Analyses.Mission.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.dynamic_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()
nacelle.side_area = 34.45 * Units.feet**2
nacelle.x_front = 7.33 * Units.feet
nacelle.length = 14.13 * Units.feet
nacelle.h_max = 3.68 * Units.feet
nacelle.w_max = 2.39 * Units.feet
nacelle.height_at_quarter_length = 3.08 * Units.feet
nacelle.height_at_three_quarters_length = 2.12 * Units.feet
other_bodies = [nacelle,nacelle]
vertical = SUAVE.Components.Wings.Wing()
vertical.span = 6.6 * Units.feet
vertical.root_chord = 8.2 * Units.feet
vertical.tip_chord = 3.6 * Units.feet
vertical.sweep_le = 47.0 * Units.deg
vertical.x_root_LE1 = 34.8 * Units.feet
vertical.symmetric = False
dz_centerline = 2.0 * Units.feet
ref_vertical = extend_to_ref_area(vertical,dz_centerline)
vertical.span = ref_vertical.ref_span
vertical.area = ref_vertical.ref_area
vertical.aspect_ratio = ref_vertical.ref_aspect_ratio
vertical.x_root_LE = vertical.x_root_LE1 + ref_vertical.root_LE_change
vertical.taper = vertical.tip_chord/ref_vertical.ref_root_chord
vertical.effective_aspect_ratio = 1.57
vertical.x_ac_LE = trapezoid_ac_x(vertical)
aircraft = SUAVE.Vehicle()
aircraft.wing = wing
aircraft.fuselage = fuselage
aircraft.other_bodies = other_bodies
aircraft.vertical = vertical
aircraft.Mass_Props.pos_cg[0] = 17.2 * Units.feet
segment = SUAVE.Analyses.Mission.Segments.Base_Segment()
segment.M = 0.152
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.rho = segment.atmosphere.compute_values(altitude / Units.km, type="rho")
segment.mew = segment.atmosphere.compute_values(altitude / Units.km, type="mew")
segment.v_inf = segment.M * segment.a
#Method Test
expected = 0.12
print 'Beech 99 at M = {0} and h = {1} meters'.format(segment.M, altitude)
cn_b = taw_cnbeta(aircraft,segment)
print 'Cn_beta = {0:.4f}'.format(cn_b)
print 'Expected value = {}'.format(expected)
print 'Percent Error = {0:.2f}%'.format(100.0*(cn_b-expected)/expected)
print ' '
#Parameters Required
#Using values for an SIAI Marchetti S-211
wing = SUAVE.Components.Wings.Wing()
wing.area = 136.0 * Units.feet**2
wing.span = 26.3 * Units.feet
wing.sweep_le = 19.5 * Units.deg
wing.z_position = -1.1 * Units.feet
wing.taper = 3.1/7.03
wing.aspect_ratio = wing.span**2/wing.area
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.side_area = 116.009 * Units.feet**2
fuselage.length = 30.9 * Units.feet
fuselage.h_max = 5.1 * Units.feet
fuselage.w_max = 5.9 * Units.feet
fuselage.height_at_vroot_quarter_chord = 4.1 * Units.feet
fuselage.height_at_quarter_length = 4.5 * Units.feet
fuselage.height_at_three_quarters_length = 4.3 * Units.feet
other_bodies = []
vertical = SUAVE.Components.Wings.Wing()
vertical.span = 5.8 * Units.feet
vertical.root_chord = 5.7 * Units.feet
vertical.tip_chord = 2.0 * Units.feet
vertical.sweep_le = 40.2 * Units.deg
vertical.x_root_LE1 = 22.62 * Units.feet
vertical.symmetric = False
dz_centerline = 2.9 * Units.feet
ref_vertical = extend_to_ref_area(vertical,dz_centerline)
vertical.span = ref_vertical.ref_span
vertical.area = ref_vertical.ref_area
vertical.aspect_ratio = ref_vertical.ref_aspect_ratio
vertical.x_root_LE = vertical.x_root_LE1 + ref_vertical.root_LE_change
vertical.taper = vertical.tip_chord/ref_vertical.ref_root_chord
vertical.effective_aspect_ratio = 2.65
vertical.x_ac_LE = trapezoid_ac_x(vertical)
aircraft = SUAVE.Vehicle()
aircraft.wing = wing
aircraft.fuselage = fuselage
aircraft.other_bodies = other_bodies
aircraft.vertical = vertical
aircraft.Mass_Props.pos_cg[0] = 16.6 * Units.feet
segment = SUAVE.Analyses.Mission.Segments.Base_Segment()
segment.M = 0.111
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.rho = segment.atmosphere.compute_values(altitude / Units.km, type="rho")
segment.mew = segment.atmosphere.compute_values(altitude / Units.km, type="mew")
segment.v_inf = segment.M * segment.a
#Method Test
print 'SIAI Marchetti S-211 at M = {0} and h = {1} meters'.format(segment.M, altitude)
cn_b = taw_cnbeta(aircraft,segment)
expected = 0.160
print 'Cn_beta = {0:.4f}'.format(cn_b)
print 'Expected value = {}'.format(expected)
print 'Percent Error = {0:.2f}%'.format(100.0*(cn_b-expected)/expected)
print ' '"""
for k,v in error.items():
assert(np.abs(v)<0.1)
return
