def estimate_take_off_field_length(vehicle,analyses,airport,compute_2nd_seg_climb = 0):
""" Computes the takeoff field length for a given vehicle configuration in a given airport.
Also optionally computes the second segment climb gradient.
Assumptions:
For second segment climb gradient:
One engine inoperative
Only validated for two engine aircraft
Source:
http://adg.stanford.edu/aa241/AircraftDesign.html
Inputs:
analyses.base.atmosphere [SUAVE data type]
airport.
altitude [m]
delta_isa [K]
vehicle.
mass_properties.takeoff [kg]
reference_area [m^2]
V2_VS_ratio (optional) [Unitless]
maximum_lift_coefficient (optional) [Unitless]
propulsors.*.number_of_engines [Unitless]
Outputs:
takeoff_field_length [m]
Properties Used:
N/A
"""
# ==============================================
# Unpack
# ==============================================
atmo = analyses.base.atmosphere
altitude = airport.altitude * Units.ft
delta_isa = airport.delta_isa
weight = vehicle.mass_properties.takeoff
reference_area = vehicle.reference_area
try:
V2_VS_ratio = vehicle.V2_VS_ratio
except:
V2_VS_ratio = 1.20
# ==============================================
# Computing atmospheric conditions
# ==============================================
atmo_values = atmo.compute_values(altitude,delta_isa)
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
p = atmo_values.pressure
T = atmo_values.temperature
rho = atmo_values.density
a = atmo_values.speed_of_sound
mu = atmo_values.dynamic_viscosity
sea_level_gravity = atmo.planet.sea_level_gravity
# ==============================================
# Determining vehicle maximum lift coefficient
# ==============================================
try: # aircraft maximum lift informed by user
maximum_lift_coefficient = vehicle.maximum_lift_coefficient
except:
# Using semi-empirical method for maximum lift coefficient calculation
from SUAVE.Methods.Aerodynamics.Fidelity_Zero.Lift import compute_max_lift_coeff
# Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
conditions = atmo.compute_values(10000. * Units.ft)
conditions.freestream=Data()
conditions.freestream.density = conditions.density
conditions.freestream.dynamic_viscosity = conditions.dynamic_viscosity
conditions.freestream.velocity = 90. * Units.knots
try:
maximum_lift_coefficient, induced_drag_high_lift = compute_max_lift_coeff(vehicle,conditions)
vehicle.maximum_lift_coefficient = maximum_lift_coefficient
except:
raise ValueError, "Maximum lift coefficient calculation error. Please, check inputs"
# ==============================================
# Computing speeds (Vs, V2, 0.7*V2)
# ==============================================
stall_speed = (2 * weight * sea_level_gravity / (rho * reference_area * maximum_lift_coefficient)) ** 0.5
V2_speed = V2_VS_ratio * stall_speed
speed_for_thrust = 0.70 * V2_speed
# ==============================================
# Determining vehicle number of engines
# ==============================================
engine_number = 0.
for propulsor in vehicle.propulsors : # may have than one propulsor
engine_number += propulsor.number_of_engines
if engine_number == 0:
raise ValueError, "No engine found in the vehicle"
# ==============================================
# Getting engine thrust
# ==============================================
state = Data()
state.conditions = Aerodynamics()
state.numerics = Numerics()
conditions = state.conditions
conditions.freestream.dynamic_pressure = np.array(np.atleast_1d(0.5 * rho * speed_for_thrust**2))
conditions.freestream.gravity = np.array([np.atleast_1d(sea_level_gravity)])
conditions.freestream.velocity = np.array(np.atleast_1d(speed_for_thrust))
conditions.freestream.mach_number = np.array(np.atleast_1d(speed_for_thrust/ a))
conditions.freestream.speed_of_sound = np.array(a)
conditions.freestream.temperature = np.array(np.atleast_1d(T))
conditions.freestream.pressure = np.array(np.atleast_1d(p))
conditions.propulsion.throttle = np.array(np.atleast_1d(1.))
results = vehicle.propulsors.evaluate_thrust(state) # total thrust
thrust = results.thrust_force_vector
# ==============================================
# Calculate takeoff distance
# ==============================================
# Defining takeoff distance equations coefficients
try:
takeoff_constants = vehicle.takeoff_constants # user defined
except: # default values
takeoff_constants = np.zeros(3)
if engine_number == 2:
takeoff_constants[0] = 857.4
takeoff_constants[1] = 2.476
takeoff_constants[2] = 0.00014
elif engine_number == 3:
takeoff_constants[0] = 667.9
takeoff_constants[1] = 2.343
takeoff_constants[2] = 0.000093
elif engine_number == 4:
takeoff_constants[0] = 486.7
takeoff_constants[1] = 2.282
takeoff_constants[2] = 0.0000705
elif engine_number > 4:
takeoff_constants[0] = 486.7
takeoff_constants[1] = 2.282
takeoff_constants[2] = 0.0000705
print 'The vehicle has more than 4 engines. Using 4 engine correlation. Result may not be correct.'
else:
takeoff_constants[0] = 857.4
takeoff_constants[1] = 2.476
takeoff_constants[2] = 0.00014
print 'Incorrect number of engines: {0:.1f}. Using twin engine correlation.'.format(engine_number)
# Define takeoff index (V2^2 / (T/W)
takeoff_index = V2_speed**2. / (thrust[0][0] / weight)
# Calculating takeoff field length
takeoff_field_length = 0.
for idx,constant in enumerate(takeoff_constants):
takeoff_field_length += constant * takeoff_index**idx
takeoff_field_length = takeoff_field_length * Units.ft
# calculating second segment climb gradient, if required by user input
if compute_2nd_seg_climb:
# Getting engine thrust at V2 (update only speed related conditions)
state.conditions.freestream.dynamic_pressure = np.array(np.atleast_1d(0.5 * rho * V2_speed**2))
state.conditions.freestream.velocity = np.array(np.atleast_1d(V2_speed))
state.conditions.freestream.mach_number = np.array(np.atleast_1d(V2_speed/ a))
results = vehicle.propulsors['turbofan'].engine_out(state)
thrust = results.thrust_force_vector[0][0]
# Compute windmilling drag
windmilling_drag_coefficient = windmilling_drag(vehicle,state)
# Compute asymmetry drag
asymmetry_drag_coefficient = asymmetry_drag(state, vehicle, windmilling_drag_coefficient)
# Compute l over d ratio for takeoff condition, NO engine failure
l_over_d = estimate_2ndseg_lift_drag_ratio(vehicle)
# Compute L over D ratio for takeoff condition, WITH engine failure
clv2 = maximum_lift_coefficient / (V2_VS_ratio) **2
cdv2_all_engine = clv2 / l_over_d
cdv2 = cdv2_all_engine + asymmetry_drag_coefficient + windmilling_drag_coefficient
l_over_d_v2 = clv2 / cdv2
# Compute 2nd segment climb gradient
second_seg_climb_gradient = thrust / (weight*sea_level_gravity) - 1. / l_over_d_v2
return takeoff_field_length, second_seg_climb_gradient
else:
# return only takeoff_field_length
return takeoff_field_length
def estimate_take_off_field_length(vehicle,
analyses,
airport,
compute_2nd_seg_climb=0):
""" Computes the takeoff field length for a given vehicle configuration in a given airport.
Also optionally computes the second segment climb gradient.
Assumptions:
For second segment climb gradient:
One engine inoperative
Only validated for two engine aircraft
Source:
http://adg.stanford.edu/aa241/AircraftDesign.html
Inputs:
analyses.base.atmosphere [SUAVE data type]
airport.
altitude [m]
delta_isa [K]
vehicle.
mass_properties.takeoff [kg]
reference_area [m^2]
V2_VS_ratio (optional) [Unitless]
maximum_lift_coefficient (optional) [Unitless]
propulsors.*.number_of_engines [Unitless]
Outputs:
takeoff_field_length [m]
Properties Used:
N/A
"""
# ==============================================
# Unpack
# ==============================================
atmo = analyses.base.atmosphere
altitude = airport.altitude * Units.ft
delta_isa = airport.delta_isa
weight = vehicle.mass_properties.takeoff
reference_area = vehicle.reference_area
try:
V2_VS_ratio = vehicle.V2_VS_ratio
except:
V2_VS_ratio = 1.20
# ==============================================
# Computing atmospheric conditions
# ==============================================
atmo_values = atmo.compute_values(altitude, delta_isa)
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
p = atmo_values.pressure
T = atmo_values.temperature
rho = atmo_values.density
a = atmo_values.speed_of_sound
mu = atmo_values.dynamic_viscosity
sea_level_gravity = atmo.planet.sea_level_gravity
# ==============================================
# Determining vehicle maximum lift coefficient
# ==============================================
try: # aircraft maximum lift informed by user
maximum_lift_coefficient = vehicle.maximum_lift_coefficient
except:
# Using semi-empirical method for maximum lift coefficient calculation
from SUAVE.Methods.Aerodynamics.Fidelity_Zero.Lift import compute_max_lift_coeff
# Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
conditions = atmo.compute_values(10000. * Units.ft)
conditions.freestream = Data()
conditions.freestream.density = conditions.density
conditions.freestream.dynamic_viscosity = conditions.dynamic_viscosity
conditions.freestream.velocity = 90. * Units.knots
try:
maximum_lift_coefficient, induced_drag_high_lift = compute_max_lift_coeff(
vehicle, conditions)
vehicle.maximum_lift_coefficient = maximum_lift_coefficient
except:
raise ValueError, "Maximum lift coefficient calculation error. Please, check inputs"
# ==============================================
# Computing speeds (Vs, V2, 0.7*V2)
# ==============================================
stall_speed = (2 * weight * sea_level_gravity /
(rho * reference_area * maximum_lift_coefficient))**0.5
V2_speed = V2_VS_ratio * stall_speed
speed_for_thrust = 0.70 * V2_speed
# ==============================================
# Determining vehicle number of engines
# ==============================================
engine_number = 0.
for propulsor in vehicle.propulsors: # may have than one propulsor
engine_number += propulsor.number_of_engines
if engine_number == 0:
raise ValueError, "No engine found in the vehicle"
# ==============================================
# Getting engine thrust
# ==============================================
state = Data()
state.conditions = Aerodynamics()
state.numerics = Numerics()
conditions = state.conditions
conditions.freestream.dynamic_pressure = np.array(
np.atleast_1d(0.5 * rho * speed_for_thrust**2))
conditions.freestream.gravity = np.array(
[np.atleast_1d(sea_level_gravity)])
conditions.freestream.velocity = np.array(np.atleast_1d(speed_for_thrust))
conditions.freestream.mach_number = np.array(
np.atleast_1d(speed_for_thrust / a))
conditions.freestream.speed_of_sound = np.array(a)
conditions.freestream.temperature = np.array(np.atleast_1d(T))
conditions.freestream.pressure = np.array(np.atleast_1d(p))
conditions.propulsion.throttle = np.array(np.atleast_1d(1.))
results = vehicle.propulsors.evaluate_thrust(state) # total thrust
thrust = results.thrust_force_vector
# ==============================================
# Calculate takeoff distance
# ==============================================
# Defining takeoff distance equations coefficients
try:
takeoff_constants = vehicle.takeoff_constants # user defined
except: # default values
takeoff_constants = np.zeros(3)
if engine_number == 2:
takeoff_constants[0] = 857.4
takeoff_constants[1] = 2.476
takeoff_constants[2] = 0.00014
elif engine_number == 3:
takeoff_constants[0] = 667.9
takeoff_constants[1] = 2.343
takeoff_constants[2] = 0.000093
elif engine_number == 4:
takeoff_constants[0] = 486.7
takeoff_constants[1] = 2.282
takeoff_constants[2] = 0.0000705
elif engine_number > 4:
takeoff_constants[0] = 486.7
takeoff_constants[1] = 2.282
takeoff_constants[2] = 0.0000705
print 'The vehicle has more than 4 engines. Using 4 engine correlation. Result may not be correct.'
else:
takeoff_constants[0] = 857.4
takeoff_constants[1] = 2.476
takeoff_constants[2] = 0.00014
print 'Incorrect number of engines: {0:.1f}. Using twin engine correlation.'.format(
engine_number)
# Define takeoff index (V2^2 / (T/W)
takeoff_index = V2_speed**2. / (thrust[0][0] / weight)
# Calculating takeoff field length
takeoff_field_length = 0.
for idx, constant in enumerate(takeoff_constants):
takeoff_field_length += constant * takeoff_index**idx
takeoff_field_length = takeoff_field_length * Units.ft
# calculating second segment climb gradient, if required by user input
if compute_2nd_seg_climb:
# Getting engine thrust at V2 (update only speed related conditions)
state.conditions.freestream.dynamic_pressure = np.array(
np.atleast_1d(0.5 * rho * V2_speed**2))
state.conditions.freestream.velocity = np.array(
np.atleast_1d(V2_speed))
state.conditions.freestream.mach_number = np.array(
np.atleast_1d(V2_speed / a))
results = vehicle.propulsors['turbofan'].engine_out(state)
thrust = results.thrust_force_vector[0][0]
# Compute windmilling drag
windmilling_drag_coefficient = windmilling_drag(vehicle, state)
# Compute asymmetry drag
asymmetry_drag_coefficient = asymmetry_drag(
state, vehicle, windmilling_drag_coefficient)
# Compute l over d ratio for takeoff condition, NO engine failure
l_over_d = estimate_2ndseg_lift_drag_ratio(vehicle)
# Compute L over D ratio for takeoff condition, WITH engine failure
clv2 = maximum_lift_coefficient / (V2_VS_ratio)**2
cdv2_all_engine = clv2 / l_over_d
cdv2 = cdv2_all_engine + asymmetry_drag_coefficient + windmilling_drag_coefficient
l_over_d_v2 = clv2 / cdv2
# Compute 2nd segment climb gradient
second_seg_climb_gradient = thrust / (
weight * sea_level_gravity) - 1. / l_over_d_v2
return takeoff_field_length, second_seg_climb_gradient
else:
# return only takeoff_field_length
return takeoff_field_length