def main():
#Parameters Required
#Using values for a Boeing 747-200
vehicle = SUAVE.Vehicle()
#print vehicle
vehicle.mass_properties.max_zero_fuel=238780*Units.kg
vehicle.mass_properties.max_takeoff =785000.*Units.lbs
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.chords.root = 42.9 * Units.feet #54.5ft
wing.sweeps.quarter_chord = 42.0 * Units.deg # Leading edge
wing.sweeps.leading_edge = 42.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = 14.7/42.9 #14.7/54.5
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([58.6,0,0]) * Units.feet
wing.aerodynamic_center = np.array([112.2*Units.feet,0.,0.])-wing.origin#16.16 * Units.meters,0.,0,])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.198])
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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.sweeps.quarter_chord = 44.0 * Units.deg # leading edge
wing.sweeps.leading_edge = 44.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
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.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)
vehicle.mass_properties.center_of_gravity=np.array([112.2,0,0]) * Units.feet
#configuration.mass_properties.zero_fuel_center_of_gravity=np.array([76.5,0,0])*Units.feet #just put a number here that got the expected value output; may want to change
fuel =SUAVE.Components.Physical_Component()
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
#now define configuration for calculation
configuration = Data()
configuration.mass_properties = Data()
configuration.mass_properties.center_of_gravity = vehicle.mass_properties.center_of_gravity
configuration.mass_properties.max_zero_fuel =vehicle.mass_properties.max_zero_fuel
configuration.fuel =fuel
configuration.mass_properties.zero_fuel_center_of_gravity=sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected =-1.56222373 #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()
vehicle.mass_properties.max_takeoff =4727*Units.kg #from Wikipedia
vehicle.mass_properties.empty =2515*Units.kg
vehicle.mass_properties.max_zero_fuel=vehicle.mass_properties.max_takeoff-vehicle.mass_properties.empty+15.*225*Units.lbs #15 passenger ac
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.chords.root = 7.9 * Units.feet
wing.sweeps.leading_edge = 4.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.sweeps.quarter_chord = 4.0 * Units.deg # Leading edge
wing.taper = 0.47
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([15.,0,0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0. , 0. ])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.152])
reference = SUAVE.Core.Container()
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights=Data()
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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 = 100.5 * Units.feet**2
wing.spans.projected = 22.5 * Units.feet
wing.sweeps.leading_edge = 21.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.sweeps.quarter_chord = 21.0 * Units.deg # leading edge
wing.taper = 3.1/6.17
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([36.3,0,0]) * Units.feet
wing.symmetric = True
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 = np.array([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 = 5.4 * Units.feet
fuselage.lengths.total = 44.0 * Units.feet
fuselage.width = 5.4 * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([17.2,0,0]) * Units.feet
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
#now define configuration for calculation
configuration = Data()
configuration.mass_properties = Data()
configuration.mass_properties.center_of_gravity = vehicle.mass_properties.center_of_gravity
configuration.mass_properties.max_zero_fuel = vehicle.mass_properties.max_zero_fuel
configuration.fuel =fuel
configuration.mass_properties.zero_fuel_center_of_gravity=sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
#Method Test
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = -2.48843437 #Should be -2.08
error.cm_a_beech_99 = (cm_a - expected)/expected
#Parameters Required
#Using values for an SIAI Marchetti S-211
vehicle = SUAVE.Vehicle()
vehicle.mass_properties.max_takeoff =2750*Units.kg #from Wikipedia
vehicle.mass_properties.empty =1850*Units.kg
vehicle.mass_properties.max_zero_fuel=vehicle.mass_properties.max_takeoff-vehicle.mass_properties.empty+2.*225*Units.lbs #2 passenger ac
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = 136.0 * Units.feet**2
wing.spans.projected = 26.3 * Units.feet
wing.chords.mean_aerodynamic = 5.4 * Units.feet
wing.chords.root = 7.03 * Units.feet
wing.chords.tip = 3.1 * Units.feet
wing.sweeps.quarter_chord = 19.5 * Units.deg # Leading edge
wing.sweeps.leading_edge = 19.5 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = 3.1/7.03
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([13.5,0,0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing),0.,0.])#16.6, 0. , 0. ]) * Units.feet - wing.origin
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.111])
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights=Data()
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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 = 36.46 * Units.feet**2
wing.spans.projected = 13.3 * Units.feet
wing.sweeps.quarter_chord= 18.5 * Units.deg # leading edge
wing.sweeps.leading_edge = 18.5 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = 1.6/3.88
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([26.07,0.,0.]) * Units.feet
wing.symmetric = True
wing.vertical = False
wing.dynamic_pressure_ratio = 0.9
wing.ep_alpha = 2.0*main_wing_CLa/np.pi/main_wing_ar
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0.0, 0.0])
wing.CL_alpha = datcom(wing,Mach)
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
vehicle.append_component(wing)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.x_root_quarter_chord = 12.67 * Units.feet
fuselage.lengths.total = 30.9 * Units.feet
fuselage.width = ((2.94+5.9)/2) * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([16.6,0,0]) * Units.feet
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
#now define configuration for calculation
configuration = Data()
configuration.mass_properties = Data()
configuration.mass_properties.center_of_gravity = vehicle.mass_properties.center_of_gravity
configuration.mass_properties.max_zero_fuel = vehicle.mass_properties.max_zero_fuel
configuration.fuel =fuel
configuration.mass_properties.zero_fuel_center_of_gravity=sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
#Method Test
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = -0.54071741 #Should be -0.6
error.cm_a_SIAI = (cm_a - expected)/expected
print error
for k,v in error.items():
assert(np.abs(v)<0.01)
return
def vehicle_setup():
vehicle = SUAVE.Vehicle()
vehicle.mass_properties.max_takeoff = 4727*Units.kg #from Wikipedia
vehicle.mass_properties.empty = 2515*Units.kg
vehicle.mass_properties.max_zero_fuel=vehicle.mass_properties.max_takeoff-vehicle.mass_properties.empty+15.*225*Units.lbs #15 passenger ac
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.chords.root = 7.9 * Units.feet
wing.sweeps.leading_edge = 4.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.sweeps.quarter_chord = 4.0 * Units.deg # Leading edge
wing.taper = 0.47
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([15.,0,0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0. , 0. ])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.152])
reference = SUAVE.Core.Container()
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights=Data()
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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 = 100.5 * Units.feet**2
wing.spans.projected = 22.5 * Units.feet
wing.sweeps.leading_edge = 21.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.sweeps.quarter_chord = 21.0 * Units.deg # leading edge
wing.taper = 3.1/6.17
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([36.3,0,0]) * Units.feet
wing.symmetric = True
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 = np.array([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 = 5.4 * Units.feet
fuselage.lengths.total = 44.0 * Units.feet
fuselage.width = 5.4 * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([17.2,0,0]) * Units.feet
fuel =SUAVE.Components.Physical_Component()
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
vehicle.fuel = fuel
vehicle.mass_properties.zero_fuel_center_of_gravity = sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
return vehicle
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.chords.root = 42.9 * Units.feet #54.5
wing.chords.tip = 14.7 * Units.feet
wing.chords.mean_aerodynamic = 27.3 * Units.feet
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.symmetric = True
wing.origin = np.array([58.6, 0., 3.6]) * Units.feet
reference = SUAVE.Core.Container()
vehicle.reference_area = wing.areas.reference
vehicle.append_component(wing)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
vertical = SUAVE.Components.Wings.Wing()
vertical.spans.exposed = 32.4 * Units.feet
vertical.chords.root = 38.7 * Units.feet # vertical.chords.fuselage_intersect
vertical.chords.tip = 13.4 * Units.feet
vertical.sweep = 50.0 * Units.deg # Leading Edge
vertical.x_root_LE1 = 180.0 * Units.feet
vertical.symmetric = False
vertical.exposed_root_chord_offset = 13.3 * Units.feet
ref_vertical = extend_to_ref_area(vertical)
wing.areas.reference = ref_vertical.areas.reference
wing.spans.projected = ref_vertical.spans.projected
wing.chords.root = ref_vertical.chords.root
dx_LE_vert = ref_vertical.root_LE_change
wing.chords.tip = vertical.chords.tip
wing.aspect_ratio = ref_vertical.aspect_ratio
wing.sweep = vertical.sweep
wing.taper = wing.chords.tip / wing.chords.root
wing.origin = np.array([vertical.x_root_LE1 + dx_LE_vert, 0., 0.])
wing.effective_aspect_ratio = 2.2
wing.symmetric = False
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)
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 = 26.0 * Units.feet
fuselage.heights.at_three_quarters_length = 19.7 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 23.8 * 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([112.2, 0, 6.8
]) * Units.feet
#segment = SUAVE.Analyses.Mission.Segments.Base_Segment()
segment = SUAVE.Analyses.Mission.Segments.Segment()
segment.freestream = Data()
segment.freestream.mach_number = Mach
segment.atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
altitude = 0.0 * Units.feet
conditions = segment.atmosphere.compute_values(altitude / Units.km)
segment.a = conditions.speed_of_sound
segment.freestream.density = conditions.density
segment.freestream.dynamic_viscosity = conditions.dynamic_viscosity
segment.freestream.velocity = segment.freestream.mach_number * segment.a
#Method Test
cn_b = taw_cnbeta(vehicle, segment, configuration)
expected = 0.08122837 # 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
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.chords.root = 42.9 * Units.feet #54.5
wing.chords.tip = 14.7 * Units.feet
wing.chords.mean_aerodynamic = 27.3 * Units.feet
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.origin = np.array([58.6,0.,3.6]) * Units.feet
reference = SUAVE.Core.Container()
vehicle.reference_area = wing.areas.reference
vehicle.append_component(wing)
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
vertical = SUAVE.Components.Wings.Wing()
vertical.spans.exposed = 32.4 * Units.feet
vertical.chords.root = 38.7 * Units.feet # vertical.chords.fuselage_intersect
vertical.chords.tip = 13.4 * Units.feet
vertical.sweep = 50.0 * Units.deg # Leading Edge
vertical.x_root_LE1 = 180.0 * Units.feet
vertical.symmetric = False
vertical.exposed_root_chord_offset = 13.3 * Units.feet
ref_vertical = extend_to_ref_area(vertical)
wing.areas.reference = ref_vertical.areas.reference
wing.spans.projected = ref_vertical.spans.projected
wing.chords.root = ref_vertical.chords.root
dx_LE_vert = ref_vertical.root_LE_change
wing.chords.tip = vertical.chords.tip
wing.aspect_ratio = ref_vertical.aspect_ratio
wing.sweep = vertical.sweep
wing.taper = wing.chords.tip/wing.chords.root
wing.origin = np.array([vertical.x_root_LE1 + dx_LE_vert,0.,0.])
wing.effective_aspect_ratio = 2.2
wing.symmetric = False
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)
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 = 26.0 * Units.feet
fuselage.heights.at_three_quarters_length = 19.7 * Units.feet
fuselage.heights.at_wing_root_quarter_chord = 23.8 * 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([112.2,0,6.8]) * Units.feet
#segment = SUAVE.Analyses.Mission.Segments.Base_Segment()
segment = SUAVE.Analyses.Mission.Segments.Segment()
segment.freestream = Data()
segment.freestream.mach_number = Mach[0]
segment.atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
altitude = 0.0 * Units.feet
conditions = segment.atmosphere.compute_values(altitude / Units.km)
segment.a = conditions.speed_of_sound
segment.freestream.density = conditions.density
segment.freestream.dynamic_viscosity = conditions.dynamic_viscosity
segment.freestream.velocity = segment.freestream.mach_number * segment.a
#Method Test
cn_b = taw_cnbeta(vehicle,segment,configuration)
expected = 0.10045 # 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
reference.area = wing.area reference.mac = 27.3 * Units.feet reference.CL_alpha_wing = datcom(wing,Mach) wing.CL_alpha = reference.CL_alpha_wing horizontal = SUAVE.Components.Wings.Wing() horizontal.area = 1490.55* Units.feet**2 horizontal.span = 71.6 * Units.feet horizontal.sweep_le = 44.0 * Units.deg horizontal.taper = 7.5/32.6 horizontal.aspect_ratio = horizontal.span**2/horizontal.area horizontal.x_LE = 187.0 * Units.feet horizontal.symmetric= True horizontal.eta = 0.95 horizontal.downwash_adj = 1.0 - 2.0*reference.CL_alpha_wing/np.pi/wing.aspect_ratio horizontal.x_ac_LE = trapezoid_ac_x(horizontal) horizontal.CL_alpha = datcom(horizontal,Mach) Lifting_Surfaces = [] Lifting_Surfaces.append(wing) Lifting_Surfaces.append(horizontal) fuselage = SUAVE.Components.Fuselages.Fuselage() fuselage.x_root_quarter_chord = 77.0 * Units.feet fuselage.length = 229.7 * Units.feet fuselage.w_max = 20.9 * Units.feet aircraft = SUAVE.Vehicle() aircraft.reference = reference aircraft.Lifting_Surfaces = Lifting_Surfaces aircraft.fuselage = fuselage
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.chords.root = 42.9 * Units.feet #54.5ft
wing.sweep = 42.0 * Units.deg # Leading edge
wing.taper = 14.7 / 42.9 #14.7/54.5
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([58.6, 0, 0]) * Units.feet
wing.aerodynamic_center = np.array([
112.2 * Units.feet, 0., 0.
]) - wing.origin #16.16 * Units.meters,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
wing.symmetric = True
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
cm_a = taw_cmalpha(vehicle, Mach, conditions, configuration)
expected = -1.627 #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.chords.root = 7.9 * Units.feet
wing.sweep = 4.0 * Units.deg # Leading edge
wing.taper = 0.47
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([15., 0, 0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0., 0.])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
Mach = np.array([0.152])
reference = SUAVE.Core.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)
main_wing_CLa = wing.CL_alpha
main_wing_ar = wing.aspect_ratio
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.areas.reference = 100.5 * Units.feet**2
wing.spans.projected = 22.5 * Units.feet
wing.sweep = 21. * Units.deg # leading edge
wing.taper = 3.1 / 6.17
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.origin = np.array([36.3, 0, 0]) * Units.feet
wing.symmetric = True
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 = np.array([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 = 5.4 * Units.feet
fuselage.lengths.total = 44.0 * Units.feet
fuselage.width = 5.4 * 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([17.2, 0, 0
]) * Units.feet
#Method Test
cm_a = taw_cmalpha(vehicle, Mach, conditions, configuration)
expected = -2.521 #Should be -2.08
error.cm_a_beech_99 = (cm_a - expected) / expected
#Parameters Required
#Using values for an SIAI Marchetti S-211
vehicle = SUAVE.Vehicle()
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = 136.0 * Units.feet**2
wing.spans.projected = 26.3 * Units.feet
wing.chords.mean_aerodynamic = 5.4 * Units.feet
wing.chords.root = 7.03 * Units.feet
wing.chords.tip = 3.1 * Units.feet
wing.sweep = 19.5 * Units.deg # Leading edge
wing.taper = 3.1 / 7.03
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([13.5, 0, 0]) * Units.feet
wing.aerodynamic_center = np.array(
[trapezoid_ac_x(wing), 0.,
0.]) #16.6, 0. , 0. ]) * Units.feet - wing.origin
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
Mach = np.array([0.111])
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 = 36.46 * Units.feet**2
wing.spans.projected = 13.3 * Units.feet
wing.sweep = 18.5 * Units.deg # leading edge
wing.taper = 1.6 / 3.88
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.origin = np.array([26.07, 0., 0.]) * Units.feet
wing.symmetric = True
wing.vertical = False
wing.dynamic_pressure_ratio = 0.9
wing.ep_alpha = 2.0 * main_wing_CLa / np.pi / main_wing_ar
wing.aerodynamic_center = np.array([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 = 12.67 * Units.feet
fuselage.lengths.total = 30.9 * Units.feet
fuselage.width = ((2.94 + 5.9) / 2) * 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([16.6, 0, 0
]) * Units.feet
#Method Test
cm_a = taw_cmalpha(vehicle, Mach, conditions, configuration)
expected = -0.5409 #Should be -0.6
error.cm_a_SIAI = (cm_a - expected) / expected
for k, v in error.items():
assert (np.abs(v) < 0.01)
return
vertical = SUAVE.Components.Wings.Wing() vertical.span = 32.4 * Units.feet vertical.root_chord = 38.7 * Units.feet vertical.tip_chord = 13.4 * Units.feet vertical.sweep_le = 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) 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 = 2.2 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] = 112. * Units.feet segment = SUAVE.Attributes.Missions.Segments.Base_Segment() segment.M = 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.rho = segment.atmosphere.compute_values(altitude / Units.km, type="rho") segment.mew = segment.atmosphere.compute_values(altitude / Units.km, type="mew")
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.sweeps.leading_edge = 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.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
def vehicle_setup():
vehicle = SUAVE.Vehicle()
vehicle.mass_properties.max_takeoff = 4727*Units.kg #from Wikipedia
vehicle.mass_properties.empty = 2515*Units.kg
vehicle.mass_properties.max_zero_fuel =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.empty+15.*225*Units.lbs #15 passenger ac
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.chords.root = 7.9 * Units.feet
wing.sweeps.leading_edge = 4.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.sweeps.quarter_chord = 4.0 * Units.deg # Leading edge
wing.taper = 0.47
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = [[15.* Units.feet ,0,0]]
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0. , 0. ])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac = compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0][0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.152])
reference = SUAVE.Core.Container()
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights =Data()
conditions.weights.total_mass =np.array([[vehicle.mass_properties.max_takeoff]])
wing.CL_alpha = conditions.lift_curve_slope
vehicle.reference_area = wing.areas.reference
# control surfaces -------------------------------------------
flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
flap.tag = 'flap'
flap.span_fraction_start = 0.15 # not correct, only placeholder
flap.span_fraction_end = 0.324 # not correct, only placeholder
flap.deflection = 1.0 * Units.deg
flap.chord_fraction = 0.19 # not correct, only placeholder
wing.append_control_surface(flap)
slat = SUAVE.Components.Wings.Control_Surfaces.Slat()
slat.tag = 'slat'
slat.span_fraction_start = 0.324 # not correct, only placeholder
slat.span_fraction_end = 0.963 # not correct, only placeholder
slat.deflection = 1.0 * Units.deg
slat.chord_fraction = 0.1 # not correct, only placeholder
wing.append_control_surface(slat)
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 = 100.5 * Units.feet**2
wing.spans.projected = 22.5 * Units.feet
wing.sweeps.leading_edge = 21.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.sweeps.quarter_chord = 21.0 * Units.deg # leading edge
wing.taper = 3.1/6.17
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = [[36.3* Units.feet,0,0]]
wing.symmetric = True
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 = np.array([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 = 5.4 * Units.feet
fuselage.lengths.total = 44.0 * Units.feet
fuselage.width = 5.4 * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([[17.2,0,0]]) * Units.feet
fuel = SUAVE.Components.Physical_Component()
fuel.origin = wing.origin
fuel.mass_properties.center_of_gravity = wing.mass_properties.center_of_gravity
fuel.mass_properties.mass = vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg = vehicle.mass_properties.center_of_gravity
MTOW = vehicle.mass_properties.max_takeoff
fuel_cg = fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass = fuel.mass_properties.mass
sum_moments_less_fuel = (cg*MTOW-fuel_cg*fuel_mass)
vehicle.fuel = fuel
vehicle.mass_properties.zero_fuel_center_of_gravity = sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
return vehicle
wing.spans.projected = 196.0 * Units.feet
wing.chords.root = 54.5 * Units.feet
wing.chords.tip = 14.7 * Units.feet
wing.areas.reference = 5500.0 * Units.feet**2
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
wing.twists.root = 2.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [65.*Units.feet,0.,-3.6*Units.feet]
wing.aerodynamic_center = [trapezoid_ac_x(wing),0,0]
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Horizontal Stabilizer'
reference.area = wing.area
reference.mac = 27.3 * Units.feet
reference.CL_alpha_wing = datcom(wing, Mach)
wing.CL_alpha = reference.CL_alpha_wing
horizontal = SUAVE.Components.Wings.Wing()
horizontal.area = 1490.55 * Units.feet**2
horizontal.span = 71.6 * Units.feet
horizontal.sweep_le = 44.0 * Units.deg
horizontal.taper = 7.5 / 32.6
horizontal.aspect_ratio = horizontal.span**2 / horizontal.area
horizontal.x_LE = 187.0 * Units.feet
horizontal.symmetric = True
horizontal.eta = 0.95
horizontal.downwash_adj = 1.0 - 2.0 * reference.CL_alpha_wing / np.pi / wing.aspect_ratio
horizontal.x_ac_LE = trapezoid_ac_x(horizontal)
horizontal.CL_alpha = datcom(horizontal, Mach)
Lifting_Surfaces = []
Lifting_Surfaces.append(wing)
Lifting_Surfaces.append(horizontal)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.x_root_quarter_chord = 77.0 * Units.feet
fuselage.length = 229.7 * Units.feet
fuselage.w_max = 20.9 * Units.feet
aircraft = SUAVE.Vehicle()
aircraft.reference = reference
aircraft.Lifting_Surfaces = Lifting_Surfaces
aircraft.fuselage = fuselage
def main():
#Parameters Required
#Using values for a Boeing 747-200
vehicle = SUAVE.Vehicle()
#print vehicle
vehicle.mass_properties.max_zero_fuel=238780*Units.kg
vehicle.mass_properties.max_takeoff =785000.*Units.lbs
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.chords.root = 42.9 * Units.feet #54.5ft
wing.sweep = 42.0 * Units.deg # Leading edge
wing.taper = 14.7/42.9 #14.7/54.5
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([58.6,0,0]) * Units.feet
wing.aerodynamic_center = np.array([112.2*Units.feet,0.,0.])-wing.origin#16.16 * Units.meters,0.,0,])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweep)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.198])
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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
wing.symmetric = True
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)
vehicle.mass_properties.center_of_gravity=np.array([112.2,0,0]) * Units.feet
#configuration.mass_properties.zero_fuel_center_of_gravity=np.array([76.5,0,0])*Units.feet #just put a number here that got the expected value output; may want to change
fuel =SUAVE.Components.Physical_Component()
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
#now define configuration for calculation
configuration = Data()
configuration.mass_properties = Data()
configuration.mass_properties.center_of_gravity = vehicle.mass_properties.center_of_gravity
configuration.mass_properties.max_zero_fuel =vehicle.mass_properties.max_zero_fuel
configuration.fuel =fuel
configuration.mass_properties.zero_fuel_center_of_gravity=sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected =-1.56222373 #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()
vehicle.mass_properties.max_takeoff =4727*Units.kg #from Wikipedia
vehicle.mass_properties.empty =2515*Units.kg
vehicle.mass_properties.max_zero_fuel=vehicle.mass_properties.max_takeoff-vehicle.mass_properties.empty+15.*225*Units.lbs #15 passenger ac
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.chords.root = 7.9 * Units.feet
wing.sweep = 4.0 * Units.deg # Leading edge
wing.taper = 0.47
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([15.,0,0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0. , 0. ])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweep)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.152])
reference = SUAVE.Core.Container()
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights=Data()
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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 = 100.5 * Units.feet**2
wing.spans.projected = 22.5 * Units.feet
wing.sweep = 21. * Units.deg # leading edge
wing.taper = 3.1/6.17
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([36.3,0,0]) * Units.feet
wing.symmetric = True
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 = np.array([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 = 5.4 * Units.feet
fuselage.lengths.total = 44.0 * Units.feet
fuselage.width = 5.4 * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([17.2,0,0]) * Units.feet
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
#now define configuration for calculation
configuration = Data()
configuration.mass_properties = Data()
configuration.mass_properties.center_of_gravity = vehicle.mass_properties.center_of_gravity
configuration.mass_properties.max_zero_fuel = vehicle.mass_properties.max_zero_fuel
configuration.fuel =fuel
configuration.mass_properties.zero_fuel_center_of_gravity=sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
#Method Test
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = -2.48843437 #Should be -2.08
error.cm_a_beech_99 = (cm_a - expected)/expected
#Parameters Required
#Using values for an SIAI Marchetti S-211
vehicle = SUAVE.Vehicle()
vehicle.mass_properties.max_takeoff =2750*Units.kg #from Wikipedia
vehicle.mass_properties.empty =1850*Units.kg
vehicle.mass_properties.max_zero_fuel=vehicle.mass_properties.max_takeoff-vehicle.mass_properties.empty+2.*225*Units.lbs #2 passenger ac
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = 136.0 * Units.feet**2
wing.spans.projected = 26.3 * Units.feet
wing.chords.mean_aerodynamic = 5.4 * Units.feet
wing.chords.root = 7.03 * Units.feet
wing.chords.tip = 3.1 * Units.feet
wing.sweep = 19.5 * Units.deg # Leading edge
wing.taper = 3.1/7.03
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([13.5,0,0]) * Units.feet
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing),0.,0.])#16.6, 0. , 0. ]) * Units.feet - wing.origin
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweep)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.111])
conditions = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights=Data()
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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 = 36.46 * Units.feet**2
wing.spans.projected = 13.3 * Units.feet
wing.sweep = 18.5 * Units.deg # leading edge
wing.taper = 1.6/3.88
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([26.07,0.,0.]) * Units.feet
wing.symmetric = True
wing.vertical = False
wing.dynamic_pressure_ratio = 0.9
wing.ep_alpha = 2.0*main_wing_CLa/np.pi/main_wing_ar
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0.0, 0.0])
wing.CL_alpha = datcom(wing,Mach)
span_location_mac =compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset =.8*np.sin(wing.sweep)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0]=.3*wing.chords.mean_aerodynamic+mac_le_offset
vehicle.append_component(wing)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.x_root_quarter_chord = 12.67 * Units.feet
fuselage.lengths.total = 30.9 * Units.feet
fuselage.width = ((2.94+5.9)/2) * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([16.6,0,0]) * Units.feet
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
#now define configuration for calculation
configuration = Data()
configuration.mass_properties = Data()
configuration.mass_properties.center_of_gravity = vehicle.mass_properties.center_of_gravity
configuration.mass_properties.max_zero_fuel = vehicle.mass_properties.max_zero_fuel
configuration.fuel =fuel
configuration.mass_properties.zero_fuel_center_of_gravity=sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
#Method Test
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = -0.54071741 #Should be -0.6
error.cm_a_SIAI = (cm_a - expected)/expected
print error
for k,v in error.items():
assert(np.abs(v)<0.01)
return
reference.area = wing.area reference.mac = 27.3 * Units.feet reference.CL_alpha_wing = datcom(wing,Mach) wing.CL_alpha = reference.CL_alpha_wing horizontal = SUAVE.Components.Wings.Wing() horizontal.area = 1490.55* Units.feet**2 horizontal.span = 71.6 * Units.feet horizontal.sweep_le = 44.0 * Units.deg horizontal.taper = 7.5/32.6 horizontal.aspect_ratio = horizontal.span**2/horizontal.area horizontal.x_LE = 187.0 * Units.feet horizontal.symmetric= True horizontal.eta = 0.95 horizontal.downwash_adj = 1.0 - 2.0*reference.CL_alpha_wing/np.pi/wing.aspect_ratio horizontal.x_ac_LE = trapezoid_ac_x(horizontal) horizontal.CL_alpha = datcom(horizontal,Mach) Lifting_Surfaces = [] Lifting_Surfaces.append(wing) Lifting_Surfaces.append(horizontal) fuselage = SUAVE.Components.Fuselages.Fuselage() fuselage.x_root_quarter_chord = 77.0 * Units.feet fuselage.length = 229.7 * Units.feet fuselage.w_max = 20.9 * Units.feet aircraft = SUAVE.Vehicle() aircraft.reference = reference aircraft.Lifting_Surfaces = Lifting_Surfaces aircraft.fuselage = fuselage
def vehicle_setup():
vehicle = SUAVE.Vehicle()
#print vehicle
vehicle.mass_properties.max_zero_fuel = 238780 * Units.kg
vehicle.mass_properties.max_takeoff = 785000. * Units.lbs
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_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.chords.root = 42.9 * Units.feet #54.5ft
wing.chords.tip = 14.7 * Units.feet
wing.sweeps.quarter_chord = 42.0 * Units.deg # Leading edge
wing.sweeps.leading_edge = 42.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([58.6, 0, 3.6]) * Units.feet
wing.aerodynamic_center = np.array([
112.2 * Units.feet, 0., 0.
]) - wing.origin #16.16 * Units.meters,0.,0,])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac = compute_span_location_from_chord_length(
wing, wing.chords.mean_aerodynamic)
mac_le_offset = .8 * np.sin(
wing.sweeps.leading_edge
) * span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[
0] = .3 * wing.chords.mean_aerodynamic + mac_le_offset
Mach = np.array([0.198])
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = datcom(wing, Mach)
conditions.weights.total_mass = np.array(
[[vehicle.mass_properties.max_takeoff]])
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
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
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.sweeps.quarter_chord = 44.0 * Units.deg # leading edge
wing.sweeps.leading_edge = 44.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
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.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)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.spans.exposed = 32.4 * Units.feet
wing.chords.root = 38.7 * Units.feet # vertical.chords.fuselage_intersect
wing.chords.tip = 13.4 * Units.feet
wing.sweeps.quarter_chord = 50.0 * Units.deg # Leading Edge
wing.x_root_LE1 = 180.0 * Units.feet
wing.symmetric = False
wing.exposed_root_chord_offset = 13.3 * Units.feet
wing = extend_to_ref_area(wing)
wing.areas.reference = wing.extended.areas.reference
wing.spans.projected = wing.extended.spans.projected
wing.chords.root = 14.9612585185
dx_LE_vert = wing.extended.root_LE_change
wing.taper = 0.272993077083
wing.origin = np.array([wing.x_root_LE1 + dx_LE_vert, 0., 0.])
wing.aspect_ratio = (wing.spans.projected**2) / wing.areas.reference
wing.effective_aspect_ratio = 2.2
wing.symmetric = False
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0.0, 0.0])
wing.dynamic_pressure_ratio = .95
Mach = np.array([0.198])
wing.CL_alpha = 0.
wing.ep_alpha = 0.
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.lengths.total = 229.7 * Units.feet
fuselage.areas.side_projected = 4696.16 * Units.feet**2 #used for cnbeta
fuselage.heights.maximum = 26.9 * Units.feet #used for cnbeta
fuselage.heights.at_quarter_length = 26.0 * Units.feet #used for cnbeta
fuselage.heights.at_three_quarters_length = 19.7 * Units.feet #used for cnbeta
fuselage.heights.at_wing_root_quarter_chord = 23.8 * Units.feet #used for cnbeta
fuselage.x_root_quarter_chord = 77.0 * Units.feet #used for cmalpha
fuselage.lengths.total = 229.7 * Units.feet
fuselage.width = 20.9 * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([112.2, 0, 0
]) * Units.feet
#configuration.mass_properties.zero_fuel_center_of_gravity=np.array([76.5,0,0])*Units.feet #just put a number here that got the expected value output; may want to change
fuel = SUAVE.Components.Physical_Component()
fuel.origin = wing.origin
fuel.mass_properties.center_of_gravity = wing.mass_properties.center_of_gravity
fuel.mass_properties.mass = vehicle.mass_properties.max_takeoff - vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg = vehicle.mass_properties.center_of_gravity
MTOW = vehicle.mass_properties.max_takeoff
fuel_cg = fuel.origin + fuel.mass_properties.center_of_gravity
fuel_mass = fuel.mass_properties.mass
sum_moments_less_fuel = (cg * MTOW - fuel_cg * fuel_mass)
vehicle.fuel = fuel
vehicle.mass_properties.zero_fuel_center_of_gravity = sum_moments_less_fuel / vehicle.mass_properties.max_zero_fuel
return vehicle
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
wing.sweep = 21 * Units.deg # leading edge
wing.taper = 3.1/6.17
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([36.3,0,0]) * 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 = 5.4 * Units.feet
fuselage.lengths.total = 44.0 * Units.feet
fuselage.width = 5.4 * 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([17.2,0,0]) * Units.feet
#Method Test
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = 12.57 # Should be -2.08
error.cm_a_beech_99 = (cm_a - expected)/expected
#Parameters Required
#Using values for an SIAI Marchetti S-211
vehicle = SUAVE.Vehicle()
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'Main Wing'
wing.areas.reference = 136.0 * Units.feet**2
wing.spans.projected = 26.3 * Units.feet
wing.chords.mean_aerodynamic = 5.4 * Units.feet
wing.sweep = 195 * Units.deg # Leading edge
wing.taper = 3.1/7.03
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.origin = np.array([12.11,0,0]) * Units.feet
wing.aerodynamic_center = np.array([16.6, 0. , 0. ]) - wing.origin
wing.eta = 1.0
wing.downwash_adj = 1.0
wing.ep_alpha = 1. - wing.downwash_adj
Mach = np.array([0.111])
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 = 36.46 * Units.feet**2
wing.spans.projected = 13.3 * Units.feet
wing.sweep = 18.5 * Units.deg # leading edge
wing.taper = 1.6/3.88
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = np.array([26.07,0.,0.]) * Units.feet
wing.symmetric= True
wing.eta = 0.9
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 = 12.67 * Units.feet
fuselage.lengths.total = 30.9 * Units.feet
fuselage.width = ((2.94+5.9)/2) * 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([16.6,0,0]) * Units.feet
#Method Test
cm_a = taw_cmalpha(vehicle,Mach,conditions,configuration)
expected = -27.9 # should be -0.6
error.cm_a_SIAI = (cm_a - expected)/expected
for k,v in error.items():
assert(np.abs(v)<0.005)
return
def vehicle_setup():
vehicle = SUAVE.Vehicle()
#print vehicle
vehicle.mass_properties.max_zero_fuel = 238780*Units.kg
vehicle.mass_properties.max_takeoff = 833000.*Units.lbs
vehicle.design_mach_number = 0.92
vehicle.design_range = 6560 * Units.miles
vehicle.design_cruise_alt = 35000.0 * Units.ft
vehicle.envelope.limit_load = 2.5
vehicle.envelope.ultimate_load = 3.75
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "longe range"
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_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.chords.root = 42.9 * Units.feet #54.5ft
wing.chords.tip = 14.7 * Units.feet
wing.sweeps.quarter_chord = 42.0 * Units.deg # Leading edge
wing.sweeps.leading_edge = 42.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = [[58.6* Units.feet ,0,3.6* Units.feet ]]
wing.aerodynamic_center = np.array([112.2*Units.feet,0.,0.])-wing.origin[0]#16.16 * Units.meters,0.,0,])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac = compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset = .8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0] = .3*wing.chords.mean_aerodynamic+mac_le_offset
wing.thickness_to_chord = 0.14
Mach = np.array([0.198])
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights.total_mass = np.array([[vehicle.mass_properties.max_takeoff]])
wing.CL_alpha = conditions.lift_curve_slope
vehicle.reference_area = wing.areas.reference
# control surfaces -------------------------------------------
flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
flap.tag = 'flap'
flap.span_fraction_start = 0.15
flap.span_fraction_end = 0.324
flap.deflection = 1.0 * Units.deg
flap.chord_fraction = 0.19
wing.append_control_surface(flap)
slat = SUAVE.Components.Wings.Control_Surfaces.Slat()
slat.tag = 'slat'
slat.span_fraction_start = 0.324
slat.span_fraction_end = 0.963
slat.deflection = 1.0 * Units.deg
slat.chord_fraction = 0.1
wing.append_control_surface(slat)
vehicle.append_component(wing)
main_wing_CLa = wing.CL_alpha
main_wing_ar = wing.aspect_ratio
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Horizontal_Tail()
wing.tag = 'horizontal_stabilizer'
wing.areas.reference = 1490.55* Units.feet**2
wing.areas.wetted = 1490.55 * Units.feet ** 2
wing.areas.exposed = 1490.55 * Units.feet ** 2
wing.spans.projected = 71.6 * Units.feet
wing.sweeps.quarter_chord = 44.0 * Units.deg # leading edge
wing.sweeps.leading_edge = 44.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = 7.5/32.6
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.origin = [[187.0* Units.feet,0,0]]
wing.symmetric = True
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)
wing.thickness_to_chord = 0.1
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.tag = 'vertical_stabilizer'
wing.spans.exposed = 32.4 * Units.feet
wing.chords.root = 38.7 * Units.feet # vertical.chords.fuselage_intersect
wing.chords.tip = 13.4 * Units.feet
wing.sweeps.quarter_chord = 50.0 * Units.deg # Leading Edge
wing.x_root_LE1 = 180.0 * Units.feet
wing.symmetric = False
wing.exposed_root_chord_offset = 13.3 * Units.feet
wing = extend_to_ref_area(wing)
wing.areas.reference = wing.extended.areas.reference
wing.areas.wetted = wing.areas.reference
wing.areas.exposed = wing.areas.reference
wing.spans.projected = wing.extended.spans.projected
wing.chords.root = 14.9612585185
dx_LE_vert = wing.extended.root_LE_change
wing.taper = 0.272993077083
wing.origin = [[wing.x_root_LE1 + dx_LE_vert,0.,0.]]
wing.aspect_ratio = (wing.spans.projected**2)/wing.areas.reference
wing.effective_aspect_ratio = 2.2
wing.symmetric = False
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing),0.0,0.0])
wing.dynamic_pressure_ratio = .95
Mach = np.array([0.198])
wing.CL_alpha = 0.
wing.ep_alpha = 0.
wing.thickness_to_chord = 0.1
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.lengths.total = 229.7 * Units.feet
fuselage.areas.side_projected = 4696.16 * Units.feet**2 #used for cnbeta
fuselage.heights.maximum = 26.9 * Units.feet #used for cnbeta
fuselage.heights.at_quarter_length = 26.0 * Units.feet #used for cnbeta
fuselage.heights.at_three_quarters_length = 19.7 * Units.feet #used for cnbeta
fuselage.heights.at_wing_root_quarter_chord = 23.8 * Units.feet #used for cnbeta
fuselage.x_root_quarter_chord = 77.0 * Units.feet #used for cmalpha
fuselage.lengths.total = 229.7 * Units.feet
fuselage.width = 20.9 * Units.feet
fuselage.areas.wetted = 688.64 * Units.feet**2
fuselage.differential_pressure = 5.0e4 * Units.pascal
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity=np.array([[112.2,0,0]]) * Units.feet
# ------------------------------------------------------------------
# Turbofan Network
# ------------------------------------------------------------------
# instantiate the gas turbine network
turbofan = SUAVE.Components.Energy.Networks.Turbofan()
turbofan.tag = 'turbofan'
# setup
turbofan.number_of_engines = 4.0
turbofan.bypass_ratio = 4.8
turbofan.engine_length = 3.934
turbofan.nacelle_diameter = 2.428
turbofan.origin = [[36.56, 22, -1.9], [27, 12, -1.9],[36.56, -22, -1.9], [27, -12, -1.9]]
# compute engine areas
Awet = 1.1 * np.pi * turbofan.nacelle_diameter * turbofan.engine_length
# Assign engine areas
turbofan.areas.wetted = Awet
# working fluid
turbofan.working_fluid = SUAVE.Attributes.Gases.Air()
# ------------------------------------------------------------------
# Component 1 - Ram
# to convert freestream static to stagnation quantities
# instantiate
ram = SUAVE.Components.Energy.Converters.Ram()
ram.tag = 'ram'
# add to the network
turbofan.append(ram)
# ------------------------------------------------------------------
# Component 2 - Inlet Nozzle
# instantiate
inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
inlet_nozzle.tag = 'inlet_nozzle'
# setup
inlet_nozzle.polytropic_efficiency = 0.98
inlet_nozzle.pressure_ratio = 0.98
# add to network
turbofan.append(inlet_nozzle)
# ------------------------------------------------------------------
# Component 3 - Low Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'low_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.91
compressor.pressure_ratio = 1.14
# add to network
turbofan.append(compressor)
# ------------------------------------------------------------------
# Component 4 - High Pressure Compressor
# instantiate
compressor = SUAVE.Components.Energy.Converters.Compressor()
compressor.tag = 'high_pressure_compressor'
# setup
compressor.polytropic_efficiency = 0.91
compressor.pressure_ratio = 13.415
# add to network
turbofan.append(compressor)
# ------------------------------------------------------------------
# Component 5 - Low Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'low_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.93
# add to network
turbofan.append(turbine)
# ------------------------------------------------------------------
# Component 6 - High Pressure Turbine
# instantiate
turbine = SUAVE.Components.Energy.Converters.Turbine()
turbine.tag = 'high_pressure_turbine'
# setup
turbine.mechanical_efficiency = 0.99
turbine.polytropic_efficiency = 0.93
# add to network
turbofan.append(turbine)
# ------------------------------------------------------------------
# Component 7 - Combustor
# instantiate
combustor = SUAVE.Components.Energy.Converters.Combustor()
combustor.tag = 'combustor'
# setup
combustor.efficiency = 0.99
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1450
combustor.pressure_ratio = 0.95
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
# add to network
turbofan.append(combustor)
# ------------------------------------------------------------------
# Component 8 - Core Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
nozzle.tag = 'core_nozzle'
# setup
nozzle.polytropic_efficiency = 0.95
nozzle.pressure_ratio = 0.99
# add to network
turbofan.append(nozzle)
# ------------------------------------------------------------------
# Component 9 - Fan Nozzle
# instantiate
nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
nozzle.tag = 'fan_nozzle'
# setup
nozzle.polytropic_efficiency = 0.95
nozzle.pressure_ratio = 0.99
# add to network
turbofan.append(nozzle)
# ------------------------------------------------------------------
# Component 10 - Fan
# instantiate
fan = SUAVE.Components.Energy.Converters.Fan()
fan.tag = 'fan'
# setup
fan.polytropic_efficiency = 0.93
fan.pressure_ratio = 1.7
# add to network
turbofan.append(fan)
# ------------------------------------------------------------------
# Component 10 : thrust (to compute the thrust)
thrust = SUAVE.Components.Energy.Processes.Thrust()
thrust.tag = 'compute_thrust'
# total design thrust (includes all the engines)
# picked lower range of 747-400 at https://en.wikipedia.org/wiki/Boeing_747
thrust.total_design = 4 * 276000. * Units.N # Newtons
# design sizing conditions
altitude = 35000.0 * Units.ft
mach_number = 0.92
isa_deviation = 0.
# Engine setup for noise module
# add to network
turbofan.thrust = thrust
# size the turbofan
turbofan_sizing(turbofan, mach_number, altitude)
# add gas turbine network turbofan to the vehicle
vehicle.append_component(turbofan)
#configuration.mass_properties.zero_fuel_center_of_gravity=np.array([76.5,0,0])*Units.feet #just put a number here that got the expected value output; may want to change
fuel =SUAVE.Components.Physical_Component()
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
vehicle.fuel = fuel
vehicle.mass_properties.zero_fuel_center_of_gravity = sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
return vehicle
vertical.root_chord = 38.7 * Units.feet vertical.tip_chord = 13.4 * Units.feet vertical.sweep_le = 50.0 * Units.deg vertical.x_root_LE1 = 181.0 * Units.feet dz_centerline = 13.5 * 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 = 2.25 vertical_symm = copy.deepcopy(vertical) vertical_symm.span = 2.0*vertical.span vertical_symm.area = 2.0*vertical.area vertical.x_ac_LE = trapezoid_ac_x(vertical_symm) aircraft = SUAVE.Vehicle() aircraft.wing = wing aircraft.fuselage = fuselage aircraft.vertical = vertical aircraft.Mass_Props.pos_cg[0] = 112.2 * Units.feet segment = SUAVE.Attributes.Missions.Segments.Segment() segment.M = 0.198 segment.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976() altitude = 0.0 * Units.feet segment.compute_atmosphere(altitude / Units.km) segment.v_inf = segment.M * segment.a #Method Test
def vehicle_setup():
vehicle = SUAVE.Vehicle()
vehicle.mass_properties.max_takeoff = 2750 * Units.kg #from Wikipedia
vehicle.mass_properties.empty = 1850 * Units.kg
vehicle.mass_properties.max_zero_fuel = vehicle.mass_properties.max_takeoff - vehicle.mass_properties.empty + 2. * 225 * Units.lbs #2 passenger ac
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'main_wing'
wing.areas.reference = 136.0 * Units.feet**2
wing.spans.projected = 26.3 * Units.feet
wing.chords.mean_aerodynamic = 5.4 * Units.feet
wing.chords.root = 7.03 * Units.feet
wing.chords.tip = 3.1 * Units.feet
wing.sweeps.quarter_chord = 19.5 * Units.deg # Leading edge
wing.sweeps.leading_edge = 19.5 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = 3.1 / 7.03
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([13.5, 0, 0]) * Units.feet
wing.aerodynamic_center = np.array(
[trapezoid_ac_x(wing), 0.,
0.]) #16.6, 0. , 0. ]) * Units.feet - wing.origin
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac = compute_span_location_from_chord_length(
wing, wing.chords.mean_aerodynamic)
mac_le_offset = .8 * np.sin(
wing.sweeps.leading_edge
) * span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[
0] = .3 * wing.chords.mean_aerodynamic + mac_le_offset
Mach = np.array([0.111])
conditions = Data()
conditions.lift_curve_slope = datcom(wing, Mach)
conditions.weights = Data()
conditions.weights.total_mass = np.array(
[[vehicle.mass_properties.max_takeoff]])
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 = 36.46 * Units.feet**2
wing.spans.projected = 13.3 * Units.feet
wing.sweeps.quarter_chord = 18.5 * Units.deg # leading edge
wing.sweeps.leading_edge = 18.5 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = 1.6 / 3.88
wing.aspect_ratio = wing.spans.projected**2 / wing.areas.reference
wing.origin = np.array([26.07, 0., 0.]) * Units.feet
wing.symmetric = True
wing.vertical = False
wing.dynamic_pressure_ratio = 0.9
wing.ep_alpha = 2.0 * main_wing_CLa / np.pi / main_wing_ar
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing), 0.0, 0.0])
wing.CL_alpha = datcom(wing, Mach)
span_location_mac = compute_span_location_from_chord_length(
wing, wing.chords.mean_aerodynamic)
mac_le_offset = .8 * np.sin(
wing.sweeps.leading_edge
) * span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[
0] = .3 * wing.chords.mean_aerodynamic + mac_le_offset
vehicle.append_component(wing)
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.x_root_quarter_chord = 12.67 * Units.feet
fuselage.lengths.total = 30.9 * Units.feet
fuselage.width = ((2.94 + 5.9) / 2) * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity = np.array([16.6, 0, 0
]) * Units.feet
fuel = SUAVE.Components.Physical_Component()
fuel.origin = wing.origin
fuel.mass_properties.center_of_gravity = wing.mass_properties.center_of_gravity
fuel.mass_properties.mass = vehicle.mass_properties.max_takeoff - vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg = vehicle.mass_properties.center_of_gravity
MTOW = vehicle.mass_properties.max_takeoff
fuel_cg = fuel.origin + fuel.mass_properties.center_of_gravity
fuel_mass = fuel.mass_properties.mass
sum_moments_less_fuel = (cg * MTOW - fuel_cg * fuel_mass)
vehicle.fuel = fuel
vehicle.mass_properties.zero_fuel_center_of_gravity = sum_moments_less_fuel / vehicle.mass_properties.max_zero_fuel
return vehicle
def vehicle_setup():
vehicle = SUAVE.Vehicle()
#print vehicle
vehicle.mass_properties.max_zero_fuel=238780*Units.kg
vehicle.mass_properties.max_takeoff =785000.*Units.lbs
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_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.chords.root = 42.9 * Units.feet #54.5ft
wing.chords.tip = 14.7 * Units.feet
wing.sweeps.quarter_chord = 42.0 * Units.deg # Leading edge
wing.sweeps.leading_edge = 42.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
wing.taper = wing.chords.tip / wing.chords.root
wing.aspect_ratio = wing.spans.projected**2/wing.areas.reference
wing.symmetric = True
wing.vertical = False
wing.origin = np.array([58.6,0,3.6]) * Units.feet
wing.aerodynamic_center = np.array([112.2*Units.feet,0.,0.])-wing.origin#16.16 * Units.meters,0.,0,])
wing.dynamic_pressure_ratio = 1.0
wing.ep_alpha = 0.0
span_location_mac = compute_span_location_from_chord_length(wing, wing.chords.mean_aerodynamic)
mac_le_offset = .8*np.sin(wing.sweeps.leading_edge)*span_location_mac #assume that 80% of the chord difference is from leading edge sweep
wing.mass_properties.center_of_gravity[0] = .3*wing.chords.mean_aerodynamic+mac_le_offset
Mach = np.array([0.198])
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = datcom(wing,Mach)
conditions.weights.total_mass=np.array([[vehicle.mass_properties.max_takeoff]])
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
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
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.sweeps.quarter_chord = 44.0 * Units.deg # leading edge
wing.sweeps.leading_edge = 44.0 * Units.deg # Same as the quarter chord sweep (ignore why EMB)
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.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)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.spans.exposed = 32.4 * Units.feet
wing.chords.root = 38.7 * Units.feet # vertical.chords.fuselage_intersect
wing.chords.tip = 13.4 * Units.feet
wing.sweeps.quarter_chord = 50.0 * Units.deg # Leading Edge
wing.x_root_LE1 = 180.0 * Units.feet
wing.symmetric = False
wing.exposed_root_chord_offset = 13.3 * Units.feet
wing = extend_to_ref_area(wing)
wing.areas.reference = wing.extended.areas.reference
wing.spans.projected = wing.extended.spans.projected
wing.chords.root = 14.9612585185
dx_LE_vert = wing.extended.root_LE_change
wing.taper = 0.272993077083
wing.origin = np.array([wing.x_root_LE1 + dx_LE_vert,0.,0.])
wing.aspect_ratio = (wing.spans.projected**2)/wing.areas.reference
wing.effective_aspect_ratio = 2.2
wing.symmetric = False
wing.aerodynamic_center = np.array([trapezoid_ac_x(wing),0.0,0.0])
wing.dynamic_pressure_ratio = .95
Mach = np.array([0.198])
wing.CL_alpha = 0.
wing.ep_alpha = 0.
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.lengths.total = 229.7 * Units.feet
fuselage.areas.side_projected = 4696.16 * Units.feet**2 #used for cnbeta
fuselage.heights.maximum = 26.9 * Units.feet #used for cnbeta
fuselage.heights.at_quarter_length = 26.0 * Units.feet #used for cnbeta
fuselage.heights.at_three_quarters_length = 19.7 * Units.feet #used for cnbeta
fuselage.heights.at_wing_root_quarter_chord = 23.8 * Units.feet #used for cnbeta
fuselage.x_root_quarter_chord = 77.0 * Units.feet #used for cmalpha
fuselage.lengths.total = 229.7 * Units.feet
fuselage.width = 20.9 * Units.feet
vehicle.append_component(fuselage)
vehicle.mass_properties.center_of_gravity=np.array([112.2,0,0]) * Units.feet
#configuration.mass_properties.zero_fuel_center_of_gravity=np.array([76.5,0,0])*Units.feet #just put a number here that got the expected value output; may want to change
fuel =SUAVE.Components.Physical_Component()
fuel.origin =wing.origin
fuel.mass_properties.center_of_gravity =wing.mass_properties.center_of_gravity
fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_zero_fuel
#find zero_fuel_center_of_gravity
cg =vehicle.mass_properties.center_of_gravity
MTOW =vehicle.mass_properties.max_takeoff
fuel_cg =fuel.origin+fuel.mass_properties.center_of_gravity
fuel_mass =fuel.mass_properties.mass
sum_moments_less_fuel=(cg*MTOW-fuel_cg*fuel_mass)
vehicle.fuel = fuel
vehicle.mass_properties.zero_fuel_center_of_gravity = sum_moments_less_fuel/vehicle.mass_properties.max_zero_fuel
return vehicle
