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 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
def __call__(self, conditions):
""" process vehicle to setup geometry, condititon and configuration
Inputs:
conditions - DataDict() of aerodynamic conditions
results - DataDict() of
Outputs:
Assumptions:
"""
# unpack
configuration = self.configuration
geometry = self.geometry
stability_model = self.stability_model
q = conditions.freestream.dynamic_pressure
Sref = geometry.reference_area
mach = conditions.freestream.mach_number
velocity = conditions.freestream.velocity
density = conditions.freestream.density
Span = geometry.wings['main_wing'].spans.projected
mac = geometry.wings['main_wing'].chords.mean_aerodynamic
aero = conditions.aerodynamics
# set up data structures
static_stability = Data()
dynamic_stability = Data()
# Calculate CL_alpha
if not conditions.has_key('lift_curve_slope'):
conditions.lift_curve_slope = datcom(geometry.wings['main_wing'],
mach)
# Calculate change in downwash with respect to change in angle of attack
for surf in geometry.wings:
sref = surf.areas.reference
span = (surf.aspect_ratio * sref)**0.5
surf.CL_alpha = datcom(surf, mach)
surf.ep_alpha = Supporting_Functions.ep_alpha(
surf.CL_alpha, sref, span)
# Static Stability Methods
static_stability.cm_alpha = taw_cmalpha(geometry, mach, conditions,
configuration)
static_stability.cn_beta = taw_cnbeta(geometry, conditions,
configuration)
# Dynamic Stability
if np.count_nonzero(
configuration.mass_properties.moments_of_inertia.tensor) > 0:
# Dynamic Stability Approximation Methods - valid for non-zero I tensor
# Derivative of yawing moment with respect to the rate of yaw
cDw = aero.drag_breakdown.parasite[
'main_wing'].parasite_drag_coefficient # Might not be the correct value
l_v = geometry.wings['vertical_stabilizer'].origin[
0] + geometry.wings['vertical_stabilizer'].aerodynamic_center[
0] - geometry.wings['main_wing'].origin[
0] - geometry.wings['main_wing'].aerodynamic_center[0]
dynamic_stability.cn_r = Supporting_Functions.cn_r(
cDw, geometry.wings['vertical_stabilizer'].areas.reference,
Sref, l_v, span,
geometry.wings['vertical_stabilizer'].dynamic_pressure_ratio,
geometry.wings['vertical_stabilizer'].CL_alpha)
# Derivative of rolling moment with respect to roll rate
dynamic_stability.cl_p = 0 # Need to see if there is a low fidelity way to calculate cl_p
# Derivative of roll rate with respect to sideslip (dihedral effect)
dynamic_stability.cl_beta = 0 # Need to see if there is a low fidelity way to calculate cl_beta
# Derivative of pitching moment with respect to pitch rate
l_t = geometry.wings['horizontal_stabilizer'].origin[
0] + geometry.wings['horizontal_stabilizer'].aerodynamic_center[
0] - geometry.wings['main_wing'].origin[
0] - geometry.wings['main_wing'].aerodynamic_center[
0] #Need to check this is the length of the horizontal tail moment arm
dynamic_stability.cm_q = Supporting_Functions.cm_q(
conditions.lift_curve_slope, l_t,
mac) # Need to check Cm_i versus Cm_alpha
# Derivative of pitching rate with respect to d(alpha)/d(t)
dynamic_stability.cm_alpha_dot = Supporting_Functions.cm_alphadot(
static_stability.cm_alpha,
geometry.wings['horizontal_stabilizer'].ep_alpha, l_t,
mac) # Need to check Cm_i versus Cm_alpha
# Derivative of Z-axis force with respect to angle of attack
dynamic_stability.cz_alpha = Supporting_Functions.cz_alpha(
aero.drag_coefficient, conditions.lift_curve_slope)
stability_model.dutch_roll = Approximations.dutch_roll(
velocity, static_stability.cn_beta, Sref, density, Span,
configuration.mass_properties.moments_of_inertia.tensor[2][2],
dynamic_stability.cn_r)
if dynamic_stability.cl_p != 0:
stability_model.roll_tau = Approximations.roll(
configuration.mass_properties.momen[2][2], Sref, density,
velocity, Span, dynamic_stability.cl_p)
dynamic_stability.cy_phi = Supporting_Functions.cy_phi(
aero.lift_coefficient)
dynamic_stability.cl_r = Supporting_Functions.cl_r(
aero.lift_coefficient) # Will need to be changed
stability_model.spiral_tau = Approximations.spiral(
conditions.weights.total_mass, velocity, density, Sref,
dynamic_stability.cl_p, static_stability.cn_beta,
dynamic_stability.cy_phi, dynamic_stability.cl_beta,
dynamic_stability.cn_r, dynamic_stability.cl_r)
stability_model.short_period = Approximations.short_period(
velocity, density, Sref, mac, dynamic_stability.cm_q,
dynamic_stability.cz_alpha, conditions.weights.total_mass,
static_stability.cm_alpha,
configuration.mass_properties.moments_of_inertia.tensor[1][1],
dynamic_stability.cm_alpha_dot)
stability_model.phugoid = Approximations.phugoid(
conditions.freestream.gravity, conditions.freestream.velocity,
aero.drag_coefficient, aero.lift_coefficient)
# Dynamic Stability Full Linearized Methods
if aero.has_key(
'cy_beta'
) and dynamic_stability.cl_p != 0 and dynamic_stability.cl_beta != 0:
theta = conditions.frames.wind.body_rotations[:, 1]
dynamic_stability.cy_beta = aero.cy_beta
dynamic_stability.cl_psi = Supporting_Functions.cy_psi(
aero.lift_coefficient, theta)
dynamic_stability.cL_u = 0
dynamic_stability.cz_u = Supporting_Functions.cz_u(
aero.lift_coefficient, velocity, dynamic_stability.cL_u)
dynamic_stability.cz_alpha_dot = Supporting_Functions.cz_alphadot(
static_stability.cm_alpha,
geometry.wings['horizontal_stabilizer'].ep_alpha)
dynamic_stability.cz_q = Supporting_Functions.cz_q(
static_stability.cm_alpha)
dynamic_stability.cx_u = Supporting_Functions.cx_u(
aero.drag_coefficient)
dynamic_stability.cx_alpha = Supporting_Functions.cx_alpha(
aero.lift_coefficient, conditions.lift_curve_slope)
lateral_directional = Full_Linearized_Equations.lateral_directional(
velocity, static_stability.cn_beta, Sref, density, Span,
configuration.mass_properties.moments_of_inertia.tensor[2]
[2], dynamic_stability.cn_r,
configuration.mass_properties.Moments_Of_Inertia.tensor[0]
[0], dynamic_stability.cl_p,
configuration.mass_properties.moments_of_inertia.tensor[0]
[2], dynamic_stability.cl_r, dynamic_stability.cl_beta,
dynamic_stability.cn_p, dynamic_stability.cy_phi,
dynamic_stability.cy_psi, dynamic_stability.cy_beta,
conditions.weights.total_mass)
longitudinal = Full_Linearized_Equations.longitudinal(
velocity, density, Sref, mac, dynamic_stability.cm_q,
dynamic_stability.cz_alpha, conditions.weights.total_mass,
static_stability.cm_alpha,
configuration.mass_properties.moments_of_inertia.tensor[1]
[1], dynamic_stability.cm_alpha_dot,
dynamic_stability.cz_u, dynamic_stability.cz_alpha_dot,
dynamic_stability.cz_q, -aero.lift_coefficient, theta,
dynamic_stability.cx_u, dynamic_stability.cx_alpha)
stability_model.dutch_roll.natural_frequency = lateral_directional.dutch_natural_frequency
stability_model.dutch_roll.damping_ratio = lateral_directional.dutch_damping_ratio
stability_model.spiral_tau = lateral_directional.spiral_tau
stability_model.roll_tau = lateral_directional.roll_tau
stability_model.short_period.natural_frequency = longitudinal.short_natural_frequency
stability_model.short_period.damping_ratio = longitudinal.short_damping_ratio
stability_model.phugoid.natural_frequency = longitudinal.phugoid_natural_frequency
stability_model.phugoid.damping_ratio = longitudinal.phugoid_damping_ratio
# pack results
results = Data()
results.static = static_stability
results.dynamic = dynamic_stability
return results
def __call__(self, conditions):
""" Process vehicle to setup geometry, condititon and configuration
Assumptions:
None
Source:
N/A
Inputs:
conditions - DataDict() of aerodynamic conditions
results - DataDict() of moment coeffients and stability and body axis derivatives
Outputs:
None
Properties Used:
self.geometry
"""
# unpack
configuration = self.configuration
geometry = self.geometry
q = conditions.freestream.dynamic_pressure
Sref = geometry.reference_area
mach = conditions.freestream.mach_number
velocity = conditions.freestream.velocity
density = conditions.freestream.density
Span = geometry.wings['main_wing'].spans.projected
mac = geometry.wings['main_wing'].chords.mean_aerodynamic
aero = conditions.aerodynamics
# set up data structures
stability = Data()
stability.static = Data()
stability.dynamic = Data()
# Calculate CL_alpha
conditions.lift_curve_slope = datcom(geometry.wings['main_wing'], mach)
# Calculate change in downwash with respect to change in angle of attack
for surf in geometry.wings:
sref = surf.areas.reference
span = (surf.aspect_ratio * sref)**0.5
surf.CL_alpha = datcom(surf, mach)
surf.ep_alpha = Supporting_Functions.ep_alpha(
surf.CL_alpha, sref, span)
# Static Stability Methods
stability.static.Cm_alpha, stability.static.Cm0, stability.static.CM = taw_cmalpha(
geometry, mach, conditions, configuration)
if 'vertical_stabilizer' in geometry.wings:
stability.static.Cn_beta = taw_cnbeta(geometry, conditions,
configuration)
else:
stability.static.Cn_beta = np.zeros_like(mach)
# calculate the static margin
stability.static.static_margin = -stability.static.Cm_alpha / conditions.lift_curve_slope
# Dynamic Stability
if np.count_nonzero(
configuration.mass_properties.moments_of_inertia.tensor) > 0:
# Dynamic Stability Approximation Methods - valid for non-zero I tensor
# Derivative of yawing moment with respect to the rate of yaw
cDw = aero.drag_breakdown.parasite[
'main_wing'].parasite_drag_coefficient # Might not be the correct value
l_v = geometry.wings['vertical_stabilizer'].origin[
0] + geometry.wings['vertical_stabilizer'].aerodynamic_center[
0] - geometry.wings['main_wing'].origin[
0] - geometry.wings['main_wing'].aerodynamic_center[0]
stability.static.Cn_r = Supporting_Functions.cn_r(
cDw, geometry.wings['vertical_stabilizer'].areas.reference,
Sref, l_v, span,
geometry.wings['vertical_stabilizer'].dynamic_pressure_ratio,
geometry.wings['vertical_stabilizer'].CL_alpha)
# Derivative of rolling moment with respect to roll rate
stability.static.Cl_p = Supporting_Functions.cl_p(
conditions.lift_curve_slope, geometry)
# Derivative of roll rate with respect to sideslip (dihedral effect)
if 'dihedral' in geometry.wings['main_wing']:
stability.static.Cl_beta = Supporting_Functions.cl_beta(
geometry, stability.static.Cl_p)
else:
stability.static.Cl_beta = np.zeros(1)
stability.static.Cy_beta = 0
# Derivative of pitching moment with respect to pitch rate
l_t = geometry.wings['horizontal_stabilizer'].origin[
0] + geometry.wings['horizontal_stabilizer'].aerodynamic_center[
0] - geometry.wings['main_wing'].origin[
0] - geometry.wings['main_wing'].aerodynamic_center[
0] #Need to check this is the length of the horizontal tail moment arm
stability.static.Cm_q = Supporting_Functions.cm_q(
conditions.lift_curve_slope, l_t,
mac) # Need to check Cm_i versus Cm_alpha
# Derivative of pitching rate with respect to d(alpha)/d(t)
stability.static.Cm_alpha_dot = Supporting_Functions.cm_alphadot(
stability.static.Cm_alpha,
geometry.wings['horizontal_stabilizer'].ep_alpha, l_t,
mac) # Need to check Cm_i versus Cm_alpha
# Derivative of Z-axis force with respect to angle of attack
stability.static.Cz_alpha = Supporting_Functions.cz_alpha(
aero.drag_coefficient, conditions.lift_curve_slope)
dutch_roll = Approximations.dutch_roll(
velocity, stability.static.Cn_beta, Sref, density, Span,
configuration.mass_properties.moments_of_inertia.tensor[2][2],
stability.static.Cn_r)
stability.dynamic.dutchRollFreqHz = dutch_roll.natural_frequency
stability.dynamic.dutchRollDamping = dutch_roll.damping_ratio
if stability.static.Cl_p.all() != 0:
stability.dynamic.rollSubsistenceTimeConstant = Approximations.roll(
configuration.mass_properties.moments_of_inertia.tensor[2]
[2], Sref, density, velocity, Span, stability.static.Cl_p)
stability.static.Cy_phi = Supporting_Functions.cy_phi(
aero.lift_coefficient)
stability.static.Cl_r = Supporting_Functions.cl_r(
aero.lift_coefficient) # Will need to be changed
stability.dynamic.spiralSubsistenceTimeConstant = Approximations.spiral(conditions.weights.total_mass, velocity, density, Sref, stability.static.Cl_p, stability.static.Cn_beta, stability.static.Cy_phi,\
stability.static.Cl_beta, stability.static.Cn_r, stability.static.Cl_r)
short_period_res = Approximations.short_period(velocity, density, Sref, mac, stability.static.Cm_q, stability.static.Cz_alpha, conditions.weights.total_mass, stability.static.Cm_alpha,\
configuration.mass_properties.moments_of_inertia.tensor[1][1], stability.static.Cm_alpha_dot)
stability.dynamic.shortPeriodFreqHz = short_period_res.natural_frequency
stability.dynamic.shortPeriodDamp = short_period_res.damping_ratio
phugoid_res = Approximations.phugoid(
conditions.freestream.gravity, conditions.freestream.velocity,
aero.drag_coefficient, aero.lift_coefficient)
stability.dynamic.phugoidFreqHz = phugoid_res.natural_frequency
stability.dynamic.phugoidDamp = phugoid_res.damping_ratio
# Dynamic Stability Full Linearized Methods
if stability.static.Cy_beta != 0 and stability.static.Cl_p.all(
) != 0 and stability.static.Cl_beta.all() != 0:
theta = conditions.frames.wind.body_rotations[:, 1]
stability.static.Cy_psi = Supporting_Functions.cy_psi(
aero.lift_coefficient, theta)
stability.static.CL_u = 0
stability.static.Cz_u = Supporting_Functions.cz_u(
aero.lift_coefficient, velocity, stability.static.cL_u)
stability.static.Cz_alpha_dot = Supporting_Functions.cz_alphadot(
stability.static.Cm_alpha,
geometry.wings['horizontal_stabilizer'].ep_alpha)
stability.static.Cz_q = Supporting_Functions.cz_q(
stability.static.Cm_alpha)
stability.static.Cx_u = Supporting_Functions.cx_u(
aero.drag_coefficient)
stability.static.Cx_alpha = Supporting_Functions.cx_alpha(
aero.lift_coefficient, conditions.lift_curve_slope)
lateral_directional = Full_Linearized_Equations.lateral_directional(velocity, stability.static.cn_beta , Sref, density, Span, configuration.mass_properties.moments_of_inertia.tensor[2][2], stability.static.Cn_r,\
configuration.mass_properties.moments_of_inertia.tensor[0][0], stability.static.Cl_p, configuration.mass_properties.moments_of_inertia.tensor[0][2], stability.static.Cl_r, stability.static.Cl_beta,\
stability.static.Cn_p, stability.static.Cy_phi, stability.static.Cy_psi, stability.static.Cy_beta, conditions.weights.total_mass)
longitudinal = Full_Linearized_Equations.longitudinal(velocity, density, Sref, mac, stability.static.Cm_q, stability.static.Cz_alpha, conditions.weights.total_mass, stability.static.Cm_alpha, \
configuration.mass_properties.moments_of_inertia.tensor[1][1], stability.static.Cm_alpha_dot, stability.static.Cz_u, stability.static.Cz_alpha_dot, stability.static.Cz_q, -aero.lift_coefficient,\
theta, stability.static.Cx_u, stability.static.Cx_alpha)
stability.dynamic.dutchRollFreqHz = lateral_directional.dutch_natural_frequency
stability.dynamic.dutchRollDamping = lateral_directional.dutch_damping_ratio
stability.dynamic.spiralSubsistenceTimeConstant = lateral_directional.spiral_tau
stability.dynamic.rollSubsistenceTimeConstant = lateral_directional.roll_tau
stability.dynamic.shortPeriodFreqHz = longitudinal.short_natural_frequency
stability.dynamic.shortPeriodDamp = longitudinal.short_damping_ratio
stability.dynamic.phugoidFreqHz = longitudinal.phugoid_natural_frequency
stability.dynamic.phugoidDamp = longitudinal.phugoid_damping_ratio
# pack results
results = Data()
results = stability
return results
def __call__(self,conditions):
""" process vehicle to setup geometry, condititon and configuration
Inputs:
conditions - DataDict() of aerodynamic conditions
results - DataDict() of
Outputs:
Assumptions:
"""
# unpack
configuration = self.configuration
geometry = self.geometry
q = conditions.freestream.dynamic_pressure
Sref = geometry.reference_area
mach = conditions.freestream.mach_number
velocity = conditions.freestream.velocity
density = conditions.freestream.density
Span = geometry.Wings['Main Wing'].span
mac = geometry.Wings['Main Wing'].chord_mac
aero = conditions.aerodynamics
# Calculate CL_alpha
if not conditions.has_key('lift_curve_slope'):
conditions.lift_curve_slope = (datcom(geometry.Wings['Main Wing'],mach))
# Calculate change in downwash with respect to change in angle of attack
for surf in geometry.Wings:
e = surf.e
sref = surf.sref
span = (surf.ar * sref ) ** 0.5
surf.CL_alpha = datcom(surf,mach)
surf.ep_alpha = Supporting_Functions.ep_alpha(surf.CL_alpha, sref, span, e)
# Static Stability Methods
aero.cm_alpha = taw_cmalpha(geometry,mach,conditions,configuration)
aero.cn_beta = taw_cnbeta(geometry,conditions,configuration)
if np.count_nonzero(configuration.mass_props.I_cg) > 0:
# Dynamic Stability Approximation Methods
if not aero.has_key('cn_r'):
cDw = aero.drag_breakdown.parasite['Main Wing'].parasite_drag_coefficient # Might not be the correct value
l_v = geometry.Wings['Vertical Stabilizer'].origin[0] + geometry.Wings['Vertical Stabilizer'].aero_center[0] - geometry.Wings['Main Wing'].origin[0] - geometry.Wings['Main Wing'].aero_center[0]
aero.cn_r = Supporting_Functions.cn_r(cDw, geometry.Wings['Vertical Stabilizer'].sref, Sref, l_v, span, geometry.Wings['Vertical Stabilizer'].eta, geometry.Wings['Vertical Stabilizer'].CL_alpha)
if not aero.has_key('cl_p'):
aero.cl_p = 0 # Need to see if there is a low fidelity way to calculate cl_p
if not aero.has_key('cl_beta'):
aero.cl_beta = 0 # Need to see if there is a low fidelity way to calculate cl_beta
l_t = geometry.Wings['Horizontal Stabilizer'].origin[0] + geometry.Wings['Horizontal Stabilizer'].aero_center[0] - geometry.Wings['Main Wing'].origin[0] - geometry.Wings['Main Wing'].aero_center[0] #Need to check this is the length of the horizontal tail moment arm
if not aero.has_key('cm_q'):
aero.cm_q = Supporting_Functions.cm_q(conditions.lift_curve_slope, l_t,mac) # Need to check Cm_i versus Cm_alpha
if not aero.has_key('cm_alpha_dot'):
aero.cm_alpha_dot = Supporting_Functions.cm_alphadot(aero.cm_alpha, geometry.Wings['Horizontal Stabilizer'].ep_alpha, l_t, mac) # Need to check Cm_i versus Cm_alpha
if not aero.has_key('cz_alpha'):
aero.cz_alpha = Supporting_Functions.cz_alpha(aero.drag_coefficient,conditions.lift_curve_slope)
conditions.dutch_roll = Approximations.dutch_roll(velocity, aero.cn_beta, Sref, density, Span, configuration.mass_props.I_cg[2][2], aero.cn_r)
if aero.cl_p != 0:
roll_tau = Approximations.roll(configuration.mass_props.I_cg[2][2], Sref, density, velocity, Span, aero.cl_p)
if aero.cl_beta != 0:
aero.cy_phi = Supporting_Functions.cy_phi(aero.lift_coefficient)
aero.cl_r = Supporting_Functions.cl_r( aero.lift_coefficient) # Will need to be changed
spiral_tau = Approximations.spiral(conditions.weights.total_mass, velocity, density, Sref, aero.cl_p, aero.cn_beta, aero.cy_phi, aero.cl_beta, aero.cn_r, aero.cl_r)
conditions.short_period = Approximations.short_period(velocity, density, Sref, mac, aero.cm_q, aero.cz_alpha, conditions.weights.total_mass, aero.cm_alpha, configuration.mass_props.I_cg[1][1], aero.cm_alpha_dot)
conditions.phugoid = Approximations.phugoid(conditions.freestream.gravity, conditions.freestream.velocity, aero.drag_coefficient, aero.lift_coefficient)
# Dynamic Stability Full Linearized Methods
if aero.has_key('cy_beta') and aero.cl_p != 0 and aero.cl_beta != 0:
if not aero.has_key('cy_psi'):
theta = conditions.frames.wind.body_rotations[:,1]
aero.cl_psi = Supporting_Functions.cy_psi(aero.lift_coefficient, theta)
if not aero.has_key('cz_u'):
if not aero.has_key('cL_u'):
aero.cL_u = 0
aero.cz_u = Supporting_Functions.cz_u(aero.lift_coefficient,velocity,aero.cL_u)
if not aero.has_key('cz_alpha_dot'):
aero.cz_alpha_dot = Supporting_Functions.cz_alphadot(aero.cm_alpha, geometry.Wings['Horizontal Stabilizer'].ep_alpha)
if not aero.has_key('cz_q'):
aero.cz_q = Supporting_Functions.cz_q(aero.cm_alpha)
if not aero.has_key('cx_u'):
aero.cx_u = Supporting_Functions.cx_u(aero.drag_coefficient)
if not aero.has_key('cx_alpha'):
aero.cx_alpha = Supporting_Functions.cx_alpha(aero.lift_coefficient, conditions.lift_curve_slope)
lateral_directional = Full_Linearized_Equations.lateral_directional(velocity, aero.cn_beta , Sref, density, Span, configuration.mass_props.I_cg[2][2], aero.cn_r, configuration.mass_props.I_cg[0][0], aero.cl_p, configuration.mass_props.I_cg[0][2], aero.cl_r, aero.cl_beta, aero.cn_p, aero.cy_phi, aero.cy_psi, aero.cy_beta, conditions.weights.total_mass)
longitudinal = Full_Linearized_Equations.longitudinal(velocity, density, Sref, mac, aero.cm_q, aero.cz_alpha, conditions.weights.total_mass, aero.cm_alpha, configuration.mass_props.I_cg[1][1], aero.cm_alpha_dot, aero.cz_u, aero.cz_alpha_dot, aero.cz_q, -aero.lift_coefficient, theta, aero.cx_u, aero.cx_alpha)
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.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 __call__(self,conditions):
""" process vehicle to setup geometry, condititon and configuration
Inputs:
conditions - DataDict() of aerodynamic conditions
results - DataDict() of
Outputs:
Assumptions:
"""
# unpack
configuration = self.configuration
geometry = self.geometry
stability_model = self.stability_model
q = conditions.freestream.dynamic_pressure
Sref = geometry.reference_area
mach = conditions.freestream.mach_number
velocity = conditions.freestream.velocity
density = conditions.freestream.density
Span = geometry.wings['main_wing'].spans.projected
mac = geometry.wings['main_wing'].chords.mean_aerodynamic
aero = conditions.aerodynamics
# set up data structures
static_stability = Data()
dynamic_stability = Data()
# Calculate CL_alpha
if not conditions.has_key('lift_curve_slope'):
conditions.lift_curve_slope = datcom(geometry.wings['main_wing'],mach)
# Calculate change in downwash with respect to change in angle of attack
for surf in geometry.wings:
sref = surf.areas.reference
span = (surf.aspect_ratio * sref ) ** 0.5
surf.CL_alpha = datcom(surf,mach)
surf.ep_alpha = Supporting_Functions.ep_alpha(surf.CL_alpha, sref, span)
# Static Stability Methods
static_stability.cm_alpha = taw_cmalpha(geometry,mach,conditions,configuration)
static_stability.cn_beta = taw_cnbeta(geometry,conditions,configuration)
# Dynamic Stability
if np.count_nonzero(configuration.mass_properties.moments_of_inertia.tensor) > 0:
# Dynamic Stability Approximation Methods - valid for non-zero I tensor
# Derivative of yawing moment with respect to the rate of yaw
cDw = aero.drag_breakdown.parasite['main_wing'].parasite_drag_coefficient # Might not be the correct value
l_v = geometry.wings['vertical_stabilizer'].origin[0] + geometry.wings['vertical_stabilizer'].aerodynamic_center[0] - geometry.wings['main_wing'].origin[0] - geometry.wings['main_wing'].aerodynamic_center[0]
dynamic_stability.cn_r = Supporting_Functions.cn_r(cDw, geometry.wings['vertical_stabilizer'].areas.reference, Sref, l_v, span, geometry.wings['vertical_stabilizer'].dynamic_pressure_ratio, geometry.wings['vertical_stabilizer'].CL_alpha)
# Derivative of rolling moment with respect to roll rate
dynamic_stability.cl_p = 0 # Need to see if there is a low fidelity way to calculate cl_p
# Derivative of roll rate with respect to sideslip (dihedral effect)
dynamic_stability.cl_beta = 0 # Need to see if there is a low fidelity way to calculate cl_beta
# Derivative of pitching moment with respect to pitch rate
l_t = geometry.wings['horizontal_stabilizer'].origin[0] + geometry.wings['horizontal_stabilizer'].aerodynamic_center[0] - geometry.wings['main_wing'].origin[0] - geometry.wings['main_wing'].aerodynamic_center[0] #Need to check this is the length of the horizontal tail moment arm
dynamic_stability.cm_q = Supporting_Functions.cm_q(conditions.lift_curve_slope, l_t,mac) # Need to check Cm_i versus Cm_alpha
# Derivative of pitching rate with respect to d(alpha)/d(t)
dynamic_stability.cm_alpha_dot = Supporting_Functions.cm_alphadot(static_stability.cm_alpha, geometry.wings['horizontal_stabilizer'].ep_alpha, l_t, mac) # Need to check Cm_i versus Cm_alpha
# Derivative of Z-axis force with respect to angle of attack
dynamic_stability.cz_alpha = Supporting_Functions.cz_alpha(aero.drag_coefficient,conditions.lift_curve_slope)
stability_model.dutch_roll = Approximations.dutch_roll(velocity, static_stability.cn_beta, Sref, density, Span, configuration.mass_properties.moments_of_inertia.tensor[2][2], dynamic_stability.cn_r)
if dynamic_stability.cl_p != 0:
stability_model.roll_tau = Approximations.roll(configuration.mass_properties.momen[2][2], Sref, density, velocity, Span, dynamic_stability.cl_p)
dynamic_stability.cy_phi = Supporting_Functions.cy_phi(aero.lift_coefficient)
dynamic_stability.cl_r = Supporting_Functions.cl_r( aero.lift_coefficient) # Will need to be changed
stability_model.spiral_tau = Approximations.spiral(conditions.weights.total_mass, velocity, density, Sref, dynamic_stability.cl_p, static_stability.cn_beta, dynamic_stability.cy_phi, dynamic_stability.cl_beta, dynamic_stability.cn_r, dynamic_stability.cl_r)
stability_model.short_period = Approximations.short_period(velocity, density, Sref, mac, dynamic_stability.cm_q, dynamic_stability.cz_alpha, conditions.weights.total_mass, static_stability.cm_alpha, configuration.mass_properties.moments_of_inertia.tensor[1][1], dynamic_stability.cm_alpha_dot)
stability_model.phugoid = Approximations.phugoid(conditions.freestream.gravity, conditions.freestream.velocity, aero.drag_coefficient, aero.lift_coefficient)
# Dynamic Stability Full Linearized Methods
if aero.has_key('cy_beta') and dynamic_stability.cl_p != 0 and dynamic_stability.cl_beta != 0:
theta = conditions.frames.wind.body_rotations[:,1]
dynamic_stability.cy_beta = aero.cy_beta
dynamic_stability.cl_psi = Supporting_Functions.cy_psi(aero.lift_coefficient, theta)
dynamic_stability.cL_u = 0
dynamic_stability.cz_u = Supporting_Functions.cz_u(aero.lift_coefficient,velocity,dynamic_stability.cL_u)
dynamic_stability.cz_alpha_dot = Supporting_Functions.cz_alphadot(static_stability.cm_alpha, geometry.wings['horizontal_stabilizer'].ep_alpha)
dynamic_stability.cz_q = Supporting_Functions.cz_q(static_stability.cm_alpha)
dynamic_stability.cx_u = Supporting_Functions.cx_u(aero.drag_coefficient)
dynamic_stability.cx_alpha = Supporting_Functions.cx_alpha(aero.lift_coefficient, conditions.lift_curve_slope)
lateral_directional = Full_Linearized_Equations.lateral_directional(velocity, static_stability.cn_beta , Sref, density, Span, configuration.mass_properties.moments_of_inertia.tensor[2][2], dynamic_stability.cn_r, configuration.mass_properties.Moments_Of_Inertia.tensor[0][0], dynamic_stability.cl_p, configuration.mass_properties.moments_of_inertia.tensor[0][2], dynamic_stability.cl_r, dynamic_stability.cl_beta, dynamic_stability.cn_p, dynamic_stability.cy_phi, dynamic_stability.cy_psi, dynamic_stability.cy_beta, conditions.weights.total_mass)
longitudinal = Full_Linearized_Equations.longitudinal(velocity, density, Sref, mac, dynamic_stability.cm_q, dynamic_stability.cz_alpha, conditions.weights.total_mass, static_stability.cm_alpha, configuration.mass_properties.moments_of_inertia.tensor[1][1], dynamic_stability.cm_alpha_dot, dynamic_stability.cz_u, dynamic_stability.cz_alpha_dot, dynamic_stability.cz_q, -aero.lift_coefficient, theta, dynamic_stability.cx_u, dynamic_stability.cx_alpha)
stability_model.dutch_roll.natural_frequency = lateral_directional.dutch_natural_frequency
stability_model.dutch_roll.damping_ratio = lateral_directional.dutch_damping_ratio
stability_model.spiral_tau = lateral_directional.spiral_tau
stability_model.roll_tau = lateral_directional.roll_tau
stability_model.short_period.natural_frequency = longitudinal.short_natural_frequency
stability_model.short_period.damping_ratio = longitudinal.short_damping_ratio
stability_model.phugoid.natural_frequency = longitudinal.phugoid_natural_frequency
stability_model.phugoid.damping_ratio = longitudinal.phugoid_damping_ratio
# pack results
results = Data()
results.static = static_stability
results.dynamic = dynamic_stability
return results
def __call__(self, conditions):
""" process vehicle to setup geometry, condititon and configuration
Inputs:
conditions - DataDict() of aerodynamic conditions
results - DataDict() of
Outputs:
Assumptions:
"""
# unpack
configuration = self.configuration
geometry = self.geometry
stability_model = self.stability_model
q = conditions.freestream.dynamic_pressure
Sref = geometry.reference_area
mach = conditions.freestream.mach_number
velocity = conditions.freestream.velocity
density = conditions.freestream.density
Span = geometry.wings['Main Wing'].spans.projected
mac = geometry.wings['Main Wing'].chords.mean_aerodynamic
aero = conditions.aerodynamics
# Calculate CL_alpha
if not conditions.has_key('lift_curve_slope'):
conditions.lift_curve_slope = datcom(geometry.wings['Main Wing'],
mach)
# Calculate change in downwash with respect to change in angle of attack
for surf in geometry.wings:
e = surf.span_efficiency
sref = surf.areas.reference
span = (surf.aspect_ratio * sref)**0.5
surf.CL_alpha = datcom(surf, mach)
surf.ep_alpha = Supporting_Functions.ep_alpha(
surf.CL_alpha, sref, span, e)
# Static Stability Methods
aero.cm_alpha = taw_cmalpha(geometry, mach, conditions, configuration)
aero.cn_beta = taw_cnbeta(geometry, conditions, configuration)
if np.count_nonzero(
configuration.mass_properties.moments_of_inertia.tensor) > 0:
# Dynamic Stability Approximation Methods
if not aero.has_key('cn_r'):
cDw = aero.drag_breakdown.parasite[
'Main Wing'].parasite_drag_coefficient # Might not be the correct value
l_v = geometry.wings['Vertical Stabilizer'].origin[
0] + geometry.wings[
'Vertical Stabilizer'].aerodynamic_center[
0] - geometry.wings['Main Wing'].origin[
0] - geometry.wings[
'Main Wing'].aerodynamic_center[0]
aero.cn_r = Supporting_Functions.cn_r(
cDw, geometry.wings['Vertical Stabilizer'].areas.reference,
Sref, l_v, span, geometry.wings['Vertical Stabilizer'].eta,
geometry.wings['Vertical Stabilizer'].CL_alpha)
if not aero.has_key('cl_p'):
aero.cl_p = 0 # Need to see if there is a low fidelity way to calculate cl_p
if not aero.has_key('cl_beta'):
aero.cl_beta = 0 # Need to see if there is a low fidelity way to calculate cl_beta
l_t = geometry.wings['Horizontal Stabilizer'].origin[
0] + geometry.wings['Horizontal Stabilizer'].aerodynamic_center[
0] - geometry.wings['Main Wing'].origin[
0] - geometry.wings['Main Wing'].aerodynamic_center[
0] #Need to check this is the length of the horizontal tail moment arm
if not aero.has_key('cm_q'):
aero.cm_q = Supporting_Functions.cm_q(
conditions.lift_curve_slope, l_t,
mac) # Need to check Cm_i versus Cm_alpha
if not aero.has_key('cm_alpha_dot'):
aero.cm_alpha_dot = Supporting_Functions.cm_alphadot(
aero.cm_alpha,
geometry.wings['Horizontal Stabilizer'].ep_alpha, l_t,
mac) # Need to check Cm_i versus Cm_alpha
if not aero.has_key('cz_alpha'):
aero.cz_alpha = Supporting_Functions.cz_alpha(
aero.drag_coefficient, conditions.lift_curve_slope)
stability_model.dutch_roll = Approximations.dutch_roll(
velocity, aero.cn_beta, Sref, density, Span,
configuration.mass_properties.moments_of_inertia.tensor[2][2],
aero.cn_r)
if aero.cl_p != 0:
stability_model.roll_tau = Approximations.roll(
configuration.mass_properties.momen[2][2], Sref, density,
velocity, Span, aero.cl_p)
if aero.cl_beta != 0:
aero.cy_phi = Supporting_Functions.cy_phi(
aero.lift_coefficient)
aero.cl_r = Supporting_Functions.cl_r(
aero.lift_coefficient) # Will need to be changed
stability_model.spiral_tau = Approximations.spiral(
conditions.weights.total_mass, velocity, density, Sref,
aero.cl_p, aero.cn_beta, aero.cy_phi, aero.cl_beta,
aero.cn_r, aero.cl_r)
stability_model.short_period = Approximations.short_period(
velocity, density, Sref, mac, aero.cm_q, aero.cz_alpha,
conditions.weights.total_mass, aero.cm_alpha,
configuration.mass_properties.moments_of_inertia.tensor[1][1],
aero.cm_alpha_dot)
stability_model.phugoid = Approximations.phugoid(
conditions.freestream.gravity, conditions.freestream.velocity,
aero.drag_coefficient, aero.lift_coefficient)
# Dynamic Stability Full Linearized Methods
if aero.has_key(
'cy_beta') and aero.cl_p != 0 and aero.cl_beta != 0:
if not aero.has_key('cy_psi'):
theta = conditions.frames.wind.body_rotations[:, 1]
aero.cl_psi = Supporting_Functions.cy_psi(
aero.lift_coefficient, theta)
if not aero.has_key('cz_u'):
if not aero.has_key('cL_u'):
aero.cL_u = 0
aero.cz_u = Supporting_Functions.cz_u(
aero.lift_coefficient, velocity, aero.cL_u)
if not aero.has_key('cz_alpha_dot'):
aero.cz_alpha_dot = Supporting_Functions.cz_alphadot(
aero.cm_alpha,
geometry.wings['Horizontal Stabilizer'].ep_alpha)
if not aero.has_key('cz_q'):
aero.cz_q = Supporting_Functions.cz_q(aero.cm_alpha)
if not aero.has_key('cx_u'):
aero.cx_u = Supporting_Functions.cx_u(
aero.drag_coefficient)
if not aero.has_key('cx_alpha'):
aero.cx_alpha = Supporting_Functions.cx_alpha(
aero.lift_coefficient, conditions.lift_curve_slope)
lateral_directional = Full_Linearized_Equations.lateral_directional(
velocity, aero.cn_beta, Sref, density, Span,
configuration.mass_properties.moments_of_inertia.tensor[2]
[2], aero.cn_r,
configuration.mass_properties.Moments_Of_Inertia.tensor[0]
[0], aero.cl_p,
configuration.mass_properties.moments_of_inertia.tensor[0]
[2], aero.cl_r, aero.cl_beta, aero.cn_p, aero.cy_phi,
aero.cy_psi, aero.cy_beta, conditions.weights.total_mass)
longitudinal = Full_Linearized_Equations.longitudinal(
velocity, density, Sref, mac, aero.cm_q, aero.cz_alpha,
conditions.weights.total_mass, aero.cm_alpha,
configuration.mass_properties.moments_of_inertia.tensor[1]
[1], aero.cm_alpha_dot, aero.cz_u, aero.cz_alpha_dot,
aero.cz_q, -aero.lift_coefficient, theta, aero.cx_u,
aero.cx_alpha)
stability_model.dutch_roll.natural_frequency = lateral_directional.dutch_natural_frequency
stability_model.dutch_roll.damping_ratio = lateral_directional.dutch_damping_ratio
stability_model.spiral_tau = lateral_directional.spiral_tau
stability_model.roll_tau = lateral_directional.roll_tau
stability_model.short_period.natural_frequency = longitudinal.short_natural_frequency
stability_model.short_period.damping_ratio = longitudinal.short_damping_ratio
stability_model.phugoid.natural_frequency = longitudinal.phugoid_natural_frequency
stability_model.phugoid.damping_ratio = longitudinal.phugoid_damping_ratio
return
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
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
aircraft.Mass_Props.pos_cg[0] = 112. * Units.feet
#Method Test
print '<<Test run of the taw_cmalpha() method>>'
print '--Boeing 747 at Mach {0}--'.format(Mach)
cm_a = taw_cmalpha(aircraft,Mach)
expected = -1.45
print 'Cm_alpha = {0:.4f}'.format(cm_a)
print 'Expected value = {}'.format(expected)
print 'Percent Error = {0:.2f}%'.format(100.0*(cm_a-expected)/expected)
print 'Static Margin = {0:.4f}'.format(-cm_a/reference.CL_alpha_wing)
print ' '
#Parameters Required
#Using values for a Beech 99
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
def main():
from Boeing_747 import vehicle_setup, configs_setup
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
Mach = np.array([0.198])
#conditions object used to create mission-like structure
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = configs.base.wings['main_wing'].CL_alpha
conditions.weights.total_mass = np.array([[vehicle.mass_properties.max_takeoff]])
conditions.aerodynamics = Data()
conditions.aerodynamics.angle_of_attack = 0.
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configs.base)[0]
expected = -1.56222373 #Should be -1.45
error = Data()
error.cm_a_747 = (cm_a - expected)/expected
from Beech_99 import vehicle_setup, configs_setup
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
Mach = np.array([0.152])
#conditions object used to create mission-like structure
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = configs.base.wings['main_wing'].CL_alpha
conditions.weights.total_mass = np.array([[vehicle.mass_properties.max_takeoff]])
conditions.aerodynamics = Data()
conditions.aerodynamics.angle_of_attack = 0.
#Method Test
#print configuration
cm_a = taw_cmalpha(vehicle,Mach,conditions,configs.base)[0]
expected = -2.48843437 #Should be -2.08
error.cm_a_beech_99 = (cm_a - expected)/expected
from SIAI_Marchetti_S211 import vehicle_setup, configs_setup
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
Mach = np.array([0.111])
#conditions object used to create mission-like structure
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = configs.base.wings['main_wing'].CL_alpha
conditions.weights.total_mass = np.array([[vehicle.mass_properties.max_takeoff]])
conditions.aerodynamics = Data()
conditions.aerodynamics.angle_of_attack = 0.
cm_a = taw_cmalpha(vehicle,Mach,conditions,configs.base)[0]
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 main():
from Boeing_747 import vehicle_setup, configs_setup
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
Mach = np.array([0.198])
#conditions object used to create mission-like structure
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = configs.base.wings['main_wing'].CL_alpha
conditions.weights.total_mass = np.array(
[[vehicle.mass_properties.max_takeoff]])
conditions.aerodynamics = Data()
conditions.aerodynamics.angle_of_attack = 0.
#print configuration
cm_a = taw_cmalpha(vehicle, Mach, conditions, configs.base)[0]
expected = -1.56222373 #Should be -1.45
error = Data()
error.cm_a_747 = (cm_a - expected) / expected
from Beech_99 import vehicle_setup, configs_setup
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
Mach = np.array([0.152])
#conditions object used to create mission-like structure
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = configs.base.wings['main_wing'].CL_alpha
conditions.weights.total_mass = np.array(
[[vehicle.mass_properties.max_takeoff]])
conditions.aerodynamics = Data()
conditions.aerodynamics.angle_of_attack = 0.
#Method Test
#print configuration
cm_a = taw_cmalpha(vehicle, Mach, conditions, configs.base)[0]
expected = -2.48843437 #Should be -2.08
error.cm_a_beech_99 = (cm_a - expected) / expected
from SIAI_Marchetti_S211 import vehicle_setup, configs_setup
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
Mach = np.array([0.111])
#conditions object used to create mission-like structure
conditions = Data()
conditions.weights = Data()
conditions.lift_curve_slope = configs.base.wings['main_wing'].CL_alpha
conditions.weights.total_mass = np.array(
[[vehicle.mass_properties.max_takeoff]])
conditions.aerodynamics = Data()
conditions.aerodynamics.angle_of_attack = 0.
cm_a = taw_cmalpha(vehicle, Mach, conditions, configs.base)[0]
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
