def taw_cnbeta(geometry, conditions, configuration):
""" CnBeta = SUAVE.Methods.Flight_Dynamics.Static_Stability.Approximations.Tube_Wing.taw_cnbeta(configuration,conditions)
This method computes the static directional stability derivative for a
standard Tube-and-Wing aircraft configuration.
CAUTION: The correlations used in this method do not account for the
destabilizing moments due to propellers. This can lead to higher-than-
expected values of CnBeta, particularly for smaller prop-driven aircraft
Inputs:
geometry - aircraft geometrical features: a data dictionary with the fields:
wings['Main Wing'] - the aircraft's main wing
areas.reference - wing reference area [meters**2]
spans.projected - span of the wing [meters]
sweep - sweep of the wing leading edge [radians]
aspect_ratio - wing aspect ratio [dimensionless]
origin - the position of the wing root in the aircraft body frame [meters]
wings['Vertical Stabilizer']
spans.projected - projected span (height for a vertical tail) of
the exposed surface [meters]
areas.reference - area of the reference vertical tail [meters**2]
sweep - leading edge sweep of the aerodynamic surface [radians]
chords.root - chord length at the junction between the tail and
the fuselage [meters]
chords.tip - chord length at the tip of the aerodynamic surface
[meters]
symmetric - Is the wing symmetric across the fuselage centerline?
origin - the position of the vertical tail root in the aircraft body frame [meters]
exposed_root_chord_offset - the displacement from the fuselage
centerline to the exposed area's physical root chordline [meters]
fuselages.Fuselage - a data dictionary with the fields:
areas.side_projected - fuselage body side area [meters**2]
lengths.total - length of the fuselage [meters]
heights.maximum - maximum height of the fuselage [meters]
width - maximum width of the fuselage [meters]
heights.at_quarter_length - fuselage height at 1/4 of the fuselage length [meters]
heights.at_three_quarters_length - fuselage height at 3/4 of fuselage
length [meters]
heights.at_vertical_root_quarter_chord - fuselage height at the quarter
chord of the vertical tail root [meters]
vertical - a data dictionary with the fields below:
NOTE: This vertical tail geometry will be used to define a reference
vertical tail that extends to the fuselage centerline.
x_ac_LE - the x-coordinate of the vertical tail aerodynamic
center measured relative to the tail root leading edge (root
of reference tail area - at fuselage centerline)
leading edge, relative to the nose [meters]
sweep_le - leading edge sweep of the vertical tail [radians]
span - height of the vertical tail [meters]
taper - vertical tail taper ratio [dimensionless]
aspect_ratio - vertical tail AR: bv/(Sv)^2 [dimensionless]
effective_aspect_ratio - effective aspect ratio considering
the effects of fuselage and horizontal tail [dimensionless]
symmetric - indicates whether the vertical panel is symmetric
about the fuselage centerline [Boolean]
other_bodies - an list of data dictionaries containing bodies
such as nacelles if these are large enough to strongly influence
stability. Each body data dictionary contains the same fields as
the fuselage data dictionary (described above), except no value
is needed for 'height_at_vroot_quarter_chord'. CAN BE EMPTY LIST
x_front - This is the only new field needed: the x-coordinate
of the nose of the body relative to the fuselage nose
conditions - a data dictionary with the fields:
v_inf - true airspeed [meters/second]
M - flight Mach number
rho - air density [kg/meters**3]
mew - air dynamic dynamic_viscosity [kg/meter/second]
configuration - a data dictionary with the fields:
mass_properties - a data dictionary with the field:
center_of_gravity - A vector in 3-space indicating CG position [meters]
other - a dictionary of aerodynamic bodies, other than the fuselage,
whose effect on directional stability is to be included in the analysis
Outputs:
CnBeta - a single float value: The static directional stability
derivative
Assumptions:
-Assumes a tube-and-wing configuration with a single centered
vertical tail
-Uses vertical tail effective aspect ratio, currently calculated by
hand, using methods from USAF Stability and Control DATCOM
-The validity of correlations for KN is questionable for sqrt(h1/h2)
greater than about 4 or h_max/w_max outside [0.3,2].
-This method assumes a small angle of attack, so the vertical tail AC
z-position does not affect the sideslip derivative.
Correlations:
-Correlations are taken from Roskam's Airplane Design, Part VI.
"""
try:
configuration.other
except AttributeError:
configuration.other = 0
CnBeta_other = []
# Unpack inputs
S = geometry.wings['main_wing'].areas.reference
b = geometry.wings['main_wing'].spans.projected
sweep = geometry.wings['main_wing'].sweeps.quarter_chord
AR = geometry.wings['main_wing'].aspect_ratio
z_w = geometry.wings['main_wing'].origin[2]
S_bs = geometry.fuselages['fuselage'].areas.side_projected
l_f = geometry.fuselages['fuselage'].lengths.total
h_max = geometry.fuselages['fuselage'].heights.maximum
w_max = geometry.fuselages['fuselage'].width
h1 = geometry.fuselages['fuselage'].heights.at_quarter_length
h2 = geometry.fuselages['fuselage'].heights.at_three_quarters_length
d_i = geometry.fuselages['fuselage'].heights.at_wing_root_quarter_chord
other = configuration.other
vert = extend_to_ref_area(geometry.wings['vertical_stabilizer'])
S_v = vert.extended.areas.reference
x_v = vert.extended.origin[0]
b_v = vert.extended.spans.projected
ac_vLE = vert.aerodynamic_center[0]
x_cg = configuration.mass_properties.center_of_gravity[0]
v_inf = conditions.freestream.velocity
mu = conditions.freestream.dynamic_viscosity
rho = conditions.freestream.density
M = conditions.freestream.mach_number
#Compute wing contribution to Cn_beta
CnBeta_w = 0.0 #The wing contribution is assumed to be zero except at very
#high angles of attack.
#Compute fuselage contribution to Cn_beta
Re_fuse = rho * v_inf * l_f / mu
x1 = x_cg / l_f
x2 = l_f * l_f / S_bs
x3 = np.sqrt(h1 / h2)
x4 = h_max / w_max
kN_1 = 3.2413 * x1 - 0.663345 + 6.1086 * np.exp(-0.22 * x2)
kN_2 = (-0.2023 + 1.3422 * x3 - 0.1454 * x3 * x3) * kN_1
kN_3 = (0.7870 + 0.1038 * x4 + 0.1834 * x4 * x4 -
2.811 * np.exp(-4.0 * x4))
K_N = (-0.47899 + kN_3 * kN_2) * 0.001
K_Rel = 1.0 + 0.8 * np.log(Re_fuse / 1.0E6) / np.log(50.)
#K_Rel: Correction for fuselage Reynolds number. Roskam VI, page 400.
CnBeta_f = -57.3 * K_N * K_Rel * S_bs * l_f / S / b
#Compute contributions of other bodies on CnBeta
if other > 0:
for body in other:
#Unpack inputs
S_bs = body.areas.side_projected
x_le = body.origin[0]
l_b = body.lengths.total
h_max = body.heights.maximum
w_max = body.width
h1 = body.heights.at_quarter_length
h2 = body.heights.at_three_quarters_length
#Compute body contribution to Cn_beta
x_cg_on_body = (x_cg - x_le) / l_b
Re_body = rho * v_inf * l_b / mew
x1 = x_cg_on_body / l_b
x2 = l_b * l_b / S_bs
x3 = np.sqrt(h1 / h2)
x4 = h_max / w_max
kN_1 = 3.2413 * x1 - 0.663345 + 6.1086 * np.exp(-0.22 * x2)
kN_2 = (-0.2023 + 1.3422 * x3 - 0.1454 * x3 * x3) * kN_1
kN_3 = (0.7870 + 0.1038 * x4 + 0.1834 * x4 * x4 -
2.811 * np.exp(-4.0 * x4))
K_N = (-0.47899 + kN_3 * kN_2) * 0.001
#K_Rel: Correction for fuselage Reynolds number. Roskam VI, page 400.
K_Rel = 1.0 + 0.8 * np.log(Re_body / 1.0E6) / np.log(50.)
CnBeta_b = -57.3 * K_N * K_Rel * S_bs * l_b / S / b
CnBeta_other.append(CnBeta_b)
#Compute vertical tail contribution
l_v = x_v + ac_vLE - x_cg
try:
iter(M)
except TypeError:
M = [M]
CLa_v = datcom(vert, M)
#k_v correlated from Roskam Fig. 10.12. NOT SMOOTH.
bf = b_v / d_i
if bf < 2.0:
k_v = 0.76
elif bf < 3.5:
k_v = 0.76 + 0.24 * (bf - 2.0) / 1.5
else:
k_v = 1.0
quarter_chord_sweep = convert_sweep(geometry.wings['main_wing'])
k_sweep = (1.0 + np.cos(quarter_chord_sweep))
dsdb_e = 0.724 + 3.06 * (
(S_v / S) / k_sweep) + 0.4 * z_w / h_max + 0.009 * AR
Cy_bv = -k_v * CLa_v * dsdb_e * (S_v / S) #ASSUMING SINGLE VERTICAL TAIL
CnBeta_v = -Cy_bv * l_v / b
CnBeta = CnBeta_w + CnBeta_f + CnBeta_v + sum(CnBeta_other)
return CnBeta
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 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
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
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
fuselage.side_area = 4696.16 * Units.feet**2 fuselage.length = 229.7 * Units.feet fuselage.h_max = 26.9 * Units.feet fuselage.w_max = 20.9 * Units.feet fuselage.height_at_vroot_quarter_chord = 15.8 * Units.feet fuselage.height_at_quarter_length = 26 * Units.feet fuselage.height_at_three_quarters_length = 19.7 * Units.feet 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 = 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
def taw_cnbeta(geometry,conditions,configuration):
""" CnBeta = SUAVE.Methods.Flight_Dynamics.Static_Stability.Approximations.Tube_Wing.taw_cnbeta(configuration,conditions)
This method computes the static directional stability derivative for a
standard Tube-and-Wing aircraft configuration.
CAUTION: The correlations used in this method do not account for the
destabilizing moments due to propellers. This can lead to higher-than-
expected values of CnBeta, particularly for smaller prop-driven aircraft
Inputs:
geometry - aircraft geometrical features: a data dictionary with the fields:
wings['Main Wing'] - the aircraft's main wing
areas.reference - wing reference area [meters**2]
spans.projected - span of the wing [meters]
sweep - sweep of the wing leading edge [radians]
aspect_ratio - wing aspect ratio [dimensionless]
origin - the position of the wing root in the aircraft body frame [meters]
wings['Vertical Stabilizer']
spans.projected - projected span (height for a vertical tail) of
the exposed surface [meters]
areas.reference - area of the reference vertical tail [meters**2]
sweep - leading edge sweep of the aerodynamic surface [radians]
chords.root - chord length at the junction between the tail and
the fuselage [meters]
chords.tip - chord length at the tip of the aerodynamic surface
[meters]
symmetric - Is the wing symmetric across the fuselage centerline?
origin - the position of the vertical tail root in the aircraft body frame [meters]
exposed_root_chord_offset - the displacement from the fuselage
centerline to the exposed area's physical root chordline [meters]
x_v = vert.origin[0]
b_v = vert.spans.projected
ac_vLE = vert.aerodynamic_center[0]
fuselages.Fuselage - a data dictionary with the fields:
areas.side_projected - fuselage body side area [meters**2]
lengths.total - length of the fuselage [meters]
heights.maximum - maximum height of the fuselage [meters]
width - maximum width of the fuselage [meters]
heights.at_quarter_length - fuselage height at 1/4 of the fuselage length [meters]
heights.at_three_quarters_length - fuselage height at 3/4 of fuselage
length [meters]
heights.at_vertical_root_quarter_chord - fuselage height at the quarter
chord of the vertical tail root [meters]
vertical - a data dictionary with the fields below:
NOTE: This vertical tail geometry will be used to define a reference
vertical tail that extends to the fuselage centerline.
x_ac_LE - the x-coordinate of the vertical tail aerodynamic
center measured relative to the tail root leading edge (root
of reference tail area - at fuselage centerline)
leading edge, relative to the nose [meters]
sweep_le - leading edge sweep of the vertical tail [radians]
span - height of the vertical tail [meters]
taper - vertical tail taper ratio [dimensionless]
aspect_ratio - vertical tail AR: bv/(Sv)^2 [dimensionless]
effective_aspect_ratio - effective aspect ratio considering
the effects of fuselage and horizontal tail [dimensionless]
symmetric - indicates whether the vertical panel is symmetric
about the fuselage centerline [Boolean]
other_bodies - an list of data dictionaries containing bodies
such as nacelles if these are large enough to strongly influence
stability. Each body data dictionary contains the same fields as
the fuselage data dictionary (described above), except no value
is needed for 'height_at_vroot_quarter_chord'. CAN BE EMPTY LIST
x_front - This is the only new field needed: the x-coordinate
of the nose of the body relative to the fuselage nose
conditions - a data dictionary with the fields:
v_inf - true airspeed [meters/second]
M - flight Mach number
rho - air density [kg/meters**3]
mew - air dynamic viscosity [kg/meter/second]
configuration - a data dictionary with the fields:
mass_properties - a data dictionary with the field:
center_of_gravity - A vector in 3-space indicating CG position [meters]
other - a dictionary of aerodynamic bodies, other than the fuselage,
whose effect on directional stability is to be included in the analysis
Outputs:
CnBeta - a single float value: The static directional stability
derivative
Assumptions:
-Assumes a tube-and-wing configuration with a single centered
vertical tail
-Uses vertical tail effective aspect ratio, currently calculated by
hand, using methods from USAF Stability and Control DATCOM
-The validity of correlations for KN is questionable for sqrt(h1/h2)
greater than about 4 or h_max/w_max outside [0.3,2].
-This method assumes a small angle of attack, so the vertical tail AC
z-position does not affect the sideslip derivative.
Correlations:
-Correlations are taken from Roskam's Airplane Design, Part VI.
"""
try:
configuration.other
except AttributeError:
configuration.other = 0
CnBeta_other = []
# Unpack inputs
S = geometry.wings['Main Wing'].areas.reference
b = geometry.wings['Main Wing'].spans.projected
sweep = geometry.wings['Main Wing'].sweep
AR = geometry.wings['Main Wing'].aspect_ratio
z_w = geometry.wings['Main Wing'].origin[2]
S_bs = geometry.fuselages.Fuselage.areas.side_projected
l_f = geometry.fuselages.Fuselage.lengths.total
h_max = geometry.fuselages.Fuselage.heights.maximum
w_max = geometry.fuselages.Fuselage.width
h1 = geometry.fuselages.Fuselage.heights.at_quarter_length
h2 = geometry.fuselages.Fuselage.heights.at_three_quarters_length
d_i = geometry.fuselages.Fuselage.heights.at_vertical_root_quarter_chord
other = configuration.other
vert = extend_to_ref_area(geometry.wings['Vertical Stabilizer'])
S_v = vert.areas.reference
x_v = vert.origin[0]
b_v = vert.spans.projected
ac_vLE = vert.aerodynamic_center[0]
x_cg = configuration.mass_properties.center_of_gravity[0]
v_inf = conditions.freestream.velocity
mu = conditions.freestream.viscosity
rho = conditions.freestream.density
M = conditions.freestream.mach_number
#Compute wing contribution to Cn_beta
CnBeta_w = 0.0 #The wing contribution is assumed to be zero except at very
#high angles of attack.
#Compute fuselage contribution to Cn_beta
Re_fuse = rho*v_inf*l_f/mu
x1 = x_cg/l_f
x2 = l_f**2.0/S_bs
x3 = np.sqrt(h1/h2)
x4 = h_max/w_max
kN_1 = 3.2413*x1 - 0.663345 + 6.1086*np.exp(-0.22*x2)
kN_2 = (-0.2023 + 1.3422*x3 - 0.1454*x3**2)*kN_1
kN_3 = (0.7870 + 0.1038*x4 + 0.1834*x4**2 - 2.811*np.exp(-4.0*x4))
K_N = (-0.47899 + kN_3*kN_2)*0.001
K_Rel = 1.0+0.8*np.log(Re_fuse/1.0E6)/np.log(50.)
#K_Rel: Correction for fuselage Reynolds number. Roskam VI, page 400.
CnBeta_f = -57.3*K_N*K_Rel*S_bs*l_f/S/b
#Compute contributions of other bodies on CnBeta
if other > 0:
for body in other:
#Unpack inputs
S_bs = body.areas.side_projected
x_le = body.origin[0]
l_b = body.lengths.total
h_max = body.heights.maximum
w_max = body.width
h1 = body.heights.at_quarter_length
h2 = body.heights.at_three_quarters_length
#Compute body contribution to Cn_beta
x_cg_on_body = (x_cg-x_le)/l_b
Re_body = rho*v_inf*l_b/mew
x1 = x_cg_on_body/l_b
x2 = l_b**2.0/S_bs
x3 = np.sqrt(h1/h2)
x4 = h_max/w_max
kN_1 = 3.2413*x1 - 0.663345 + 6.1086*np.exp(-0.22*x2)
kN_2 = (-0.2023 + 1.3422*x3 - 0.1454*x3**2)*kN_1
kN_3 = (0.7870 + 0.1038*x4 + 0.1834*x4**2 - 2.811*np.exp(-4.0*x4))
K_N = (-0.47899 + kN_3*kN_2)*0.001
#K_Rel: Correction for fuselage Reynolds number. Roskam VI, page 400.
K_Rel = 1.0+0.8*np.log(Re_body/1.0E6)/np.log(50.)
CnBeta_b = -57.3*K_N*K_Rel*S_bs*l_b/S/b
CnBeta_other.append(CnBeta_b)
#Compute vertical tail contribution
l_v = x_v + ac_vLE - x_cg
#try:
# CLa_v = geometry.wings['Vertical Stabilizer'].CL_alpha
#except AttributeError:
# CLa_v = datcom(geometry.wings['Vertical Stabilizer'], [M])
try:
iter(M)
except TypeError:
M = [M]
CLa_v = datcom(vert,M)
#k_v correlated from Roskam Fig. 10.12. NOT SMOOTH.
bf = b_v/d_i
if bf < 2.0:
k_v = 0.76
elif bf < 3.5:
k_v = 0.76 + 0.24*(bf-2.0)/1.5
else:
k_v = 1.0
quarter_chord_sweep = convert_sweep(geometry.wings['Main Wing'])
k_sweep = (1.0+np.cos(quarter_chord_sweep))
dsdb_e = 0.724 + 3.06*((S_v/S)/k_sweep) + 0.4*z_w/h_max + 0.009*AR
Cy_bv = -k_v*CLa_v*dsdb_e*(S_v/S) #ASSUMING SINGLE VERTICAL TAIL
CnBeta_v = -Cy_bv*l_v/b
CnBeta = CnBeta_w + CnBeta_f + CnBeta_v + sum(CnBeta_other)
##print "Wing: {} Fuse: {} Vert: {} Othr: {}".format(CnBeta_w,CnBeta_f,CnBeta_v,sum(CnBeta_other))
return CnBeta
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
