def simple_sizing(nexus):
configs = nexus.vehicle_configurations
analyses = nexus.analyses
base = configs.base
m_guess = nexus.m_guess #take in sizing inputs
base.mass_properties.max_takeoff = m_guess
#find conditions
air_speed = nexus.missions.base.segments['cruise'].air_speed
altitude = nexus.missions.base.segments['climb_3'].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
freestream0 = atmosphere.compute_values(6000. * Units.ft) #cabin altitude
diff_pressure = np.max(freestream0.pressure - freestream.pressure, 0)
fuselage = base.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#now size engine
mach_number = air_speed / freestream.speed_of_sound
#now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = SUAVE.Attributes.Planets.Earth().sea_level_gravity
#assign conditions in form for propulsor sizing
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
conditions.freestream = freestream
nose_load_fraction = .06
#now evaluate all of the vehicle configurations
for config in configs:
config.wings.horizontal_stabilizer.areas.reference = (
26.0 / 92.0) * config.wings.main_wing.areas.reference
for wing in config.wings:
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(
wing)
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
fuselage = config.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
# now evaluate weights
# diff the new data
config.mass_properties.max_takeoff = m_guess #take in parameters
config.mass_properties.takeoff = m_guess
config.mass_properties.max_zero_fuel = base.mass_properties.max_zero_fuel
config.store_diff()
# now evaluate the weights
weights = analyses.base.weights.evaluate()
# update zfw
empty_weight = base.mass_properties.operating_empty
payload = base.mass_properties.max_payload
zfw = empty_weight + payload
base.max_zero_fuel = zfw
base.store_diff()
for config in configs:
config.pull_base()
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.landing = base.mass_properties.max_zero_fuel
# Landing CL_max
altitude = nexus.missions.base.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
mu = freestream.dynamic_viscosity
rho = freestream.density
landing_conditions.freestream.velocity = nexus.missions.base.segments[
'descent_3'].air_speed
landing_conditions.freestream.density = rho
landing_conditions.freestream.dynamic_viscosity = mu / rho
CL_max_landing, CDi = compute_max_lift_coeff(landing, landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
# diff the new data
landing.store_diff()
#Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
mu = freestream.dynamic_viscosity
rho = freestream.density
takeoff_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
takeoff_conditions.freestream.density = rho
takeoff_conditions.freestream.dynamic_viscosity = mu / rho
max_CL_takeoff, CDi = compute_max_lift_coeff(takeoff, takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
takeoff.store_diff()
#Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
mu = freestream.dynamic_viscosity
rho = freestream.density
base_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
base_conditions.freestream.density = rho
base_conditions.freestream.dynamic_viscosity = mu / rho
max_CL_base, CDi = compute_max_lift_coeff(base, base_conditions)
base.maximum_lift_coefficient = max_CL_base
base.store_diff()
# done!
return nexus
def simple_sizing(nexus):
configs = nexus.vehicle_configurations
base = configs.base
#find conditions
air_speed = nexus.missions.base.segments['cruise'].air_speed
altitude = nexus.missions.base.segments['climb_5'].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
freestream0 = atmosphere.compute_values(6000. * Units.ft) #cabin altitude
diff_pressure = np.max(freestream0.pressure - freestream.pressure, 0)
fuselage = base.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#now size engine
mach_number = air_speed / freestream.speed_of_sound
#now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
) #assign conditions in form for propulsor sizing
conditions.freestream = freestream
for config in configs:
config.wings.horizontal_stabilizer.areas.reference = (
26.0 / 92.0) * config.wings.main_wing.areas.reference
for wing in config.wings:
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(
wing)
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
fuselage = config.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
turbofan_sizing(config.propulsors['turbofan'], mach_number, altitude)
compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
# diff the new data
#config.store_diff()
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.landing = 0.85 * config.mass_properties.takeoff
# Landing CL_max
altitude = nexus.missions.base.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(altitude)
landing_conditions.freestream.velocity = nexus.missions.base.segments[
'descent_3'].air_speed
landing_conditions.freestream.density = rho
landing_conditions.freestream.dynamic_viscosity = mu / rho
CL_max_landing, CDi = compute_max_lift_coeff(landing, landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
# diff the new data
#landing.store_diff()
#Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(altitude)
takeoff_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
takeoff_conditions.freestream.density = rho
takeoff_conditions.freestream.dynamic_viscosity = mu / rho
max_CL_takeoff, CDi = compute_max_lift_coeff(takeoff, takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
#takeoff.store_diff()
#Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(altitude)
base_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
base_conditions.freestream.density = rho
base_conditions.freestream.dynamic_viscosity = mu / rho
max_CL_base, CDi = compute_max_lift_coeff(base, base_conditions)
base.maximum_lift_coefficient = max_CL_base
#base.store_diff()
# done!
return nexus
def simple_sizing(nexus):
configs = nexus.vehicle_configurations
#find conditions
air_speed = nexus.missions.base.segments['cruise'].air_speed
altitude = nexus.missions.base.segments['climb_5'].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
#now size engine
mach_number = air_speed/freestream.speed_of_sound
#now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics() #assign conditions in form for propulsor sizing
conditions.freestream = freestream
HT_volume = 1.42542 * Units.less # E170
VT_volume = 0.11458 * Units.less # E170
lht = 14.24 * Units.m
lvt = 13.54 * Units.m
for config in configs:
wing_planform(config.wings.main_wing)
config.wings.horizontal_stabilizer.areas.reference = (HT_volume/lht)*(config.wings.main_wing.areas.reference *
3.194)
config.wings.vertical_stabilizer.areas.reference = (VT_volume/lvt)*(config.wings.main_wing.areas.reference *
26.0)
SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_fuel_volume(config.wings.main_wing)
nexus.summary.available_fuel = config.wings.main_wing.fuel_volume * 803.0 * 1.0197
for wing in config.wings:
wing_planform(wing)
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
turbofan_sizing(config.propulsors['turbofan'], mach_number=mach_number, altitude=altitude)
compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
config.propulsors['turbofan'].nacelle_diameter = config.propulsors['turbofan'].nacelle_diameter * 1.1462135
config.propulsors['turbofan'].engine_length = config.propulsors['turbofan'].engine_length * 1.24868
config.propulsors['turbofan'].areas.wetted = 1.1*np.pi * (config.propulsors['turbofan'].engine_length *
config.propulsors['turbofan'].nacelle_diameter)
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.landing = 0.863 * config.mass_properties.takeoff
# Landing CL_max
altitude = nexus.missions.base.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream_landing = atmosphere.compute_values(altitude)
landing_conditions.freestream.velocity = nexus.missions.base.segments['descent_3'].air_speed
landing_conditions.freestream.density = freestream_landing.density
landing_conditions.freestream.dynamic_viscosity = freestream_landing.dynamic_viscosity
CL_max_landing,CDi = compute_max_lift_coeff(landing, landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
#Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
freestream_takeoff = atmosphere.compute_values(altitude)
takeoff_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
takeoff_conditions.freestream.density = freestream_takeoff.density
takeoff_conditions.freestream.dynamic_viscosity = freestream_takeoff.dynamic_viscosity
max_CL_takeoff, CDi = compute_max_lift_coeff(takeoff, takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
#Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = takeoff_conditions.freestream
max_CL_base, CDi = compute_max_lift_coeff(base, base_conditions)
base.maximum_lift_coefficient = max_CL_base
return nexus
def simple_sizing(nexus):
configs = nexus.vehicle_configurations
base = configs.base
#find conditions
air_speed = nexus.missions.mission.segments['cruise'].air_speed
altitude = nexus.missions.mission.segments['climb_2'].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
freestream0 = atmosphere.compute_values(6000. * Units.ft) #cabin altitude
diff_pressure = np.max(freestream0.pressure - freestream.pressure, 0)
fuselage = base.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#now size engine
mach_number = air_speed / freestream.speed_of_sound
#now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
) #assign conditions in form for propulsor sizing
conditions.freestream = freestream
for config in configs:
# fuselage
fuselage = config.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
# Main Wing
main_wing_span = config.wings.main_wing.spans.projected
main_wing_root_chord = config.wings.main_wing.chords.root
config.wings.main_wing.chords.tip = config.wings.main_wing.chords.root * config.wings.main_wing.taper
config.wings.main_wing.chords.mean_aerodynamic = (
main_wing_root_chord * (2.0 / 3.0) *
((1.0 + config.wings.main_wing.taper +
config.wings.main_wing.taper**2.0) /
(1.0 + config.wings.main_wing.taper)))
config.wings.main_wing.areas.reference = (
main_wing_root_chord +
config.wings.main_wing.chords.tip) * main_wing_span / 2.0
config.wings.main_wing.areas.wetted = 2.0 * config.wings.main_wing.areas.reference
config.wings.main_wing.areas.exposed = config.wings.main_wing.areas.wetted
config.wings.main_wing.areas.affected = config.wings.main_wing.areas.wetted
config.wings.main_wing.aspect_ratio = (
main_wing_span *
main_wing_span) / config.wings.main_wing.areas.reference
config.wings.main_wing.flaps.chord = main_wing_root_chord * 0.15
config.wings.main_wing.flaps.span_start = 0.3 * main_wing_span
config.wings.main_wing.flaps.span_end = 0.8 * main_wing_span
config.wings.main_wing.flaps.area = config.wings.main_wing.flaps.chord * (
config.wings.main_wing.flaps.span_end -
config.wings.main_wing.flaps.span_start)
# Horizontal Stabilizer
config.wings.horizontal_stabilizer.areas.reference = 0.145 * config.wings.main_wing.areas.reference
config.wings.horizontal_stabilizer.spans.projected = np.sqrt(
config.wings.horizontal_stabilizer.aspect_ratio *
config.wings.horizontal_stabilizer.areas.reference)
config.wings.horizontal_stabilizer.chords.root = config.wings.horizontal_stabilizer.areas.reference / config.wings.horizontal_stabilizer.spans.projected
config.wings.horizontal_stabilizer.chords.tip = config.wings.horizontal_stabilizer.chords.root
config.wings.horizontal_stabilizer.chords.mean_aerodynamic = (
config.wings.horizontal_stabilizer.chords.root * (2.0 / 3.0) *
((1.0 + config.wings.horizontal_stabilizer.taper +
config.wings.horizontal_stabilizer.taper**2.0) /
(1.0 + config.wings.horizontal_stabilizer.taper)))
config.wings.horizontal_stabilizer.areas.wetted = 2.0 * config.wings.horizontal_stabilizer.areas.reference
config.wings.horizontal_stabilizer.areas.exposed = config.wings.horizontal_stabilizer.areas.wetted
config.wings.horizontal_stabilizer.areas.affected = config.wings.horizontal_stabilizer.areas.wetted
# Vertical Stabilizer
config.wings.vertical_stabilizer.areas.reference = 0.099 * config.wings.main_wing.areas.reference
config.wings.vertical_stabilizer.areas.wetted = 2.0 * config.wings.vertical_stabilizer.areas.reference
config.wings.vertical_stabilizer.areas.exposed = config.wings.vertical_stabilizer.areas.wetted
config.wings.vertical_stabilizer.areas.affected = config.wings.vertical_stabilizer.areas.wetted
config.wings.vertical_stabilizer.spans.projected = np.sqrt(
config.wings.vertical_stabilizer.aspect_ratio *
config.wings.vertical_stabilizer.areas.reference)
config.wings.vertical_stabilizer.chords.root = (
2.0 * config.wings.vertical_stabilizer.areas.reference) / (
config.wings.vertical_stabilizer.spans.projected *
(1 + config.wings.vertical_stabilizer.taper))
config.wings.vertical_stabilizer.chords.tip = config.wings.vertical_stabilizer.chords.root * config.wings.vertical_stabilizer.taper
config.wings.vertical_stabilizer.chords.mean_aerodynamic = (
config.wings.vertical_stabilizer.chords.root * (2.0 / 3.0) *
((1.0 + config.wings.vertical_stabilizer.taper +
config.wings.vertical_stabilizer.taper**2.0) /
(1.0 + config.wings.vertical_stabilizer.taper)))
# Resize the motor
kv = config.propulsors.propulsor.motor_forward.speed_constant
config.propulsors.propulsor.motor_forward = size_from_kv(
config.propulsors.propulsor.motor_forward, kv)
kv = config.propulsors.propulsor.motor_lift.speed_constant
config.propulsors.propulsor.motor_lift = size_from_kv(
config.propulsors.propulsor.motor_lift, kv)
# diff the new data
config.store_diff()
#turbofan_sizing(config.propulsors['turbofan'], mach_number = mach_number, altitude = altitude)
#compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.landing = 1.0 * config.mass_properties.takeoff
# Landing CL_max
altitude = nexus.missions.mission.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream_landing = atmosphere.compute_values(0.)
landing_conditions.freestream.velocity = nexus.missions.mission.segments[
'descent_2'].air_speed
landing_conditions.freestream.density = freestream_landing.density
landing_conditions.freestream.dynamic_viscosity = freestream_landing.dynamic_viscosity
CL_max_landing, CDi = compute_max_lift_coeff(landing, landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
#Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.mission.airport.altitude
freestream_takeoff = atmosphere.compute_values(altitude)
takeoff_conditions.freestream.velocity = nexus.missions.mission.segments.climb_1.air_speed
takeoff_conditions.freestream.density = freestream_takeoff.density
takeoff_conditions.freestream.dynamic_viscosity = freestream_takeoff.dynamic_viscosity
max_CL_takeoff, CDi = compute_max_lift_coeff(takeoff, takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
#Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = takeoff_conditions.freestream
max_CL_base, CDi = compute_max_lift_coeff(base, base_conditions)
base.maximum_lift_coefficient = max_CL_base
return nexus
def simple_sizing(nexus):
configs = nexus.vehicle_configurations
base = configs.base
# find conditions
air_speed = nexus.missions.base.segments['cruise_2'].air_speed
altitude = 18.5 * Units.km #nexus.missions.base.segments['climb_8'].altitude_end #FIXME
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
freestream0 = atmosphere.compute_values(6000. * Units.ft) # cabin altitude
diff_pressure = np.max(freestream0.pressure - freestream.pressure, 0)
fuselage = base.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
# now size engine
mach_number = air_speed / freestream.speed_of_sound
# now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics() # assign conditions in form for propulsor sizing
conditions.freestream = freestream
# conditions.weights.vehicle_mass_rate = -200 * Units['kg/s']
for config in configs:
config.wings.horizontal_stabilizer.areas.reference = (26.0 / 92.0) * config.wings.main_wing.areas.reference
for wing in config.wings:
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
fuselage = config.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
turbofan_sizing(config.propulsors['turbofan'], mach_number, altitude)
compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
# engine_length, nacelle_diameter, areas.wette
# diff the new data
config.store_diff()
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
# landing.mass_properties.landing = 0.85 * config.mass_properties.takeoff
# Landing CL_max
altitude = nexus.missions.base.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(altitude)
landing_conditions.freestream.velocity = nexus.missions.base.segments['descent_final'].air_speed
landing_conditions.freestream.density = rho
landing_conditions.freestream.dynamic_viscosity = mu / rho
CL_max_landing, CDi = compute_max_lift_coeff(landing, landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
# diff the new data
landing.store_diff()
# Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(altitude)
takeoff_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
takeoff_conditions.freestream.density = rho
takeoff_conditions.freestream.dynamic_viscosity = mu / rho
max_CL_takeoff, CDi = compute_max_lift_coeff(takeoff, takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
takeoff.store_diff()
# Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
p, T, rho, a, mu = atmosphere.compute_values(altitude)
base_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
base_conditions.freestream.density = rho
base_conditions.freestream.dynamic_viscosity = mu / rho
max_CL_base, CDi = compute_max_lift_coeff(base, base_conditions)
base.maximum_lift_coefficient = max_CL_base
base.store_diff()
# done!
return nexus
def simple_sizing(nexus):
configs = nexus.vehicle_configurations
base = configs.base
#find conditions
air_speed = nexus.missions.base.segments['cruise'].air_speed
altitude = nexus.missions.base.segments['climb_2'].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
freestream0 = atmosphere.compute_values(6000. * Units.ft) #cabin altitude
diff_pressure = np.max(freestream0.pressure - freestream.pressure, 0)
fuselage = base.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#now size engine
mach_number = air_speed / freestream.speed_of_sound
#now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
) #assign conditions in form for propulsor sizing
conditions.freestream = freestream
for config in configs:
config.wings.horizontal_stabilizer.areas.reference = (
26.0 / 92.0) * config.wings.main_wing.areas.reference
for wing in config.wings:
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(
wing)
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
fuselage = config.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#turbofan_sizing(config.propulsors['turbofan'], mach_number = mach_number, altitude = altitude)
#compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
# Change the dynamic pressure based on the, add a factor of safety
base.envelope.maximum_dynamic_pressure = nexus.missions.mission.segments.cruise.dynamic_pressure * 1.2
# Scale the horizontal and vertical tails based on the main wing area
#For more or less "normal" looking designs the stabilizer and elevators together should be between 15 to 20% of the wing area.
#The fin and rudder together should be around 12 to 15%
base.wings.horizontal_stabilizer.areas.reference = 0.17 * base.reference_area
base.wings.vertical_stabilizer.areas.reference = 0.13 * base.reference_area
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.landing = 1.0 * config.mass_properties.takeoff
# Landing CL_max
altitude = nexus.missions.base.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream_landing = atmosphere.compute_values(0.)
landing_conditions.freestream.velocity = nexus.missions.base.segments[
'descent_2'].air_speed
landing_conditions.freestream.density = freestream_landing.density
landing_conditions.freestream.dynamic_viscosity = freestream_landing.dynamic_viscosity
CL_max_landing, CDi = compute_max_lift_coeff(landing, landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
#Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
freestream_takeoff = atmosphere.compute_values(altitude)
takeoff_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
takeoff_conditions.freestream.density = freestream_takeoff.density
takeoff_conditions.freestream.dynamic_viscosity = freestream_takeoff.dynamic_viscosity
max_CL_takeoff, CDi = compute_max_lift_coeff(takeoff, takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
#Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = takeoff_conditions.freestream
max_CL_base, CDi = compute_max_lift_coeff(base, base_conditions)
base.maximum_lift_coefficient = max_CL_base
return nexus
def estimate_clmax(nexus):
# Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
state = Data()
state.conditions = Data()
state.conditions.freestream = Data()
state.conditions.freestream.density = 0.90477283
state.conditions.freestream.dynamic_viscosity = 1.69220918e-05
state.conditions.freestream.velocity = 90. * Units.knots
#Takeoff CL_max
config = nexus.vehicle_configurations.takeoff
settings = nexus.analyses.takeoff.aerodynamics.settings
maximum_lift_coefficient, CDi = compute_max_lift_coeff(
state, settings, config)
config.maximum_lift_coefficient = maximum_lift_coefficient
# diff the new data
config.store_diff()
#Takeoff CL_max - for short field config
config = nexus.vehicle_configurations.short_field_takeoff
settings = nexus.analyses.short_field_takeoff.aerodynamics.settings
maximum_lift_coefficient, CDi = compute_max_lift_coeff(
state, settings, config)
config.maximum_lift_coefficient = maximum_lift_coefficient
# diff the new data
config.store_diff()
# compute V2 speed for noise, based in MTOW
config = nexus.vehicle_configurations.takeoff
weight = config.mass_properties.max_takeoff
reference_area = config.wings.main_wing.areas.reference
max_CL_takeoff = config.maximum_lift_coefficient
stall_speed = (2 * 9.81 * weight /
(1.225 * reference_area * max_CL_takeoff))**0.5
V2_speed = 1.20 * stall_speed
speed_for_noise = V2_speed + nexus.noise_V2_increase
#nexus.missions.takeoff_initialization.segments.climb.air_speed = speed_for_noise
nexus.missions.takeoff.segments.climb.air_speed = speed_for_noise
nexus.missions.takeoff.segments.cutback.air_speed = speed_for_noise
nexus.missions.sideline_takeoff.segments.climb.air_speed = speed_for_noise
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.takeoff = 0.85 * config.mass_properties.takeoff
landing.mass_properties.landing = 0.85 * config.mass_properties.takeoff
# Landing CL_max
settings = nexus.analyses.landing.aerodynamics.settings
maximum_lift_coefficient, CDi = compute_max_lift_coeff(
state, settings, landing)
landing.maximum_lift_coefficient = maximum_lift_coefficient
# compute approach speed
weight = landing.mass_properties.landing
stall_speed = (2 * 9.81 * weight /
(1.225 * reference_area * maximum_lift_coefficient))**0.5
Vref_speed = 1.23 * stall_speed
nexus.missions.landing.segments.descent.air_speed = Vref_speed + 10. * Units.knots
# diff the new data
landing.store_diff()
return nexus
def estimate_clmax(nexus):
# Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
conditions = Data()
conditions.freestream = Data()
conditions.freestream.density = 0.90477283
conditions.freestream.dynamic_viscosity = 1.69220918e-05
conditions.freestream.velocity = 90. * Units.knots
#Takeoff CL_max
config = nexus.vehicle_configurations.takeoff
maximum_lift_coefficient,CDi = compute_max_lift_coeff(config,conditions)
config.maximum_lift_coefficient = maximum_lift_coefficient
# diff the new data
config.store_diff()
#Takeoff CL_max - for short field config
config = nexus.vehicle_configurations.short_field_takeoff
maximum_lift_coefficient,CDi = compute_max_lift_coeff(config,conditions)
config.maximum_lift_coefficient = maximum_lift_coefficient
# diff the new data
config.store_diff()
# compute V2 speed for noise, based in MTOW
config = nexus.vehicle_configurations.takeoff
weight = config.mass_properties.max_takeoff
reference_area = config.wings.main_wing.areas.reference
max_CL_takeoff = config.maximum_lift_coefficient
stall_speed = (2 * 9.81 * weight / (1.225 * reference_area * max_CL_takeoff)) ** 0.5
V2_speed = 1.20 * stall_speed
speed_for_noise = V2_speed + nexus.noise_V2_increase
#nexus.missions.takeoff_initialization.segments.climb.air_speed = speed_for_noise
nexus.missions.takeoff.segments.climb.air_speed = speed_for_noise
nexus.missions.takeoff.segments.cutback.air_speed = speed_for_noise
nexus.missions.sideline_takeoff.segments.climb.air_speed = speed_for_noise
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.takeoff = 0.85 * config.mass_properties.takeoff
landing.mass_properties.landing = 0.85 * config.mass_properties.takeoff
# Landing CL_max
maximum_lift_coefficient,CDi = compute_max_lift_coeff(landing,conditions)
landing.maximum_lift_coefficient = maximum_lift_coefficient
# compute approach speed
weight = landing.mass_properties.landing
stall_speed = (2 * 9.81 * weight / (1.225 * reference_area * maximum_lift_coefficient)) ** 0.5
Vref_speed = 1.23 * stall_speed
nexus.missions.landing.segments.descent.air_speed = Vref_speed + 10. * Units.knots
# diff the new data
landing.store_diff()
return nexus
def simple_sizing(nexus):
configs = nexus.vehicle_configurations
analyses = nexus.analyses
base = configs.base
m_guess = nexus.m_guess #take in sizing inputs
base.mass_properties.max_takeoff = m_guess
#find conditions
air_speed = nexus.missions.base.segments['cruise'].air_speed
altitude = nexus.missions.base.segments['climb_3'].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
freestream0 = atmosphere.compute_values(6000.*Units.ft) #cabin altitude
diff_pressure = np.max(freestream0.pressure-freestream.pressure,0)
fuselage = base.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#now size engine
mach_number = air_speed/freestream.speed_of_sound
#now add to freestream data object
freestream.velocity = air_speed
freestream.mach_number = mach_number
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics() #assign conditions in form for propulsor sizing
conditions.freestream = freestream
nose_load_fraction = .06
#now evaluate all of the vehicle configurations
for config in configs:
config.wings.horizontal_stabilizer.areas.reference = (26.0/92.0)*config.wings.main_wing.areas.reference
for wing in config.wings:
wing = SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_planform(wing)
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
fuselage = config.fuselages['fuselage']
fuselage.differential_pressure = diff_pressure
#now evaluate weights
# diff the new data
config.mass_properties.max_takeoff = m_guess #take in parameters
config.mass_properties.takeoff = m_guess
config.mass_properties.max_zero_fuel = base.mass_properties.max_zero_fuel
config.store_diff()
#now evaluate the weights
weights = analyses.base.weights.evaluate() #base.weights.evaluate()
#update zfw
empty_weight = base.mass_properties.operating_empty
payload = base.mass_properties.max_payload
zfw = empty_weight + payload
base.max_zero_fuel = zfw
base.store_diff()
for config in configs:
config.pull_base()
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = nexus.vehicle_configurations.landing
landing_conditions = Data()
landing_conditions.freestream = Data()
# landing weight
landing.mass_properties.landing = base.mass_properties.max_zero_fuel
# Landing CL_max
altitude = nexus.missions.base.segments[-1].altitude_end
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
mu = freestream.dynamic_viscosity
rho = freestream.density
landing_conditions.freestream.velocity = nexus.missions.base.segments['descent_3'].air_speed
landing_conditions.freestream.density = rho
landing_conditions.freestream.dynamic_viscosity = mu/rho
CL_max_landing,CDi = compute_max_lift_coeff(landing,landing_conditions)
landing.maximum_lift_coefficient = CL_max_landing
# diff the new data
landing.store_diff()
#Takeoff CL_max
takeoff = nexus.vehicle_configurations.takeoff
takeoff_conditions = Data()
takeoff_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
mu = freestream.dynamic_viscosity
rho = freestream.density
takeoff_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
takeoff_conditions.freestream.density = rho
takeoff_conditions.freestream.dynamic_viscosity = mu/rho
max_CL_takeoff,CDi = compute_max_lift_coeff(takeoff,takeoff_conditions)
takeoff.maximum_lift_coefficient = max_CL_takeoff
takeoff.store_diff()
#Base config CL_max
base = nexus.vehicle_configurations.base
base_conditions = Data()
base_conditions.freestream = Data()
altitude = nexus.missions.base.airport.altitude
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
freestream = atmosphere.compute_values(altitude)
mu = freestream.dynamic_viscosity
rho = freestream.density
base_conditions.freestream.velocity = nexus.missions.base.segments.climb_1.air_speed
base_conditions.freestream.density = rho
base_conditions.freestream.dynamic_viscosity = mu/rho
max_CL_base,CDi = compute_max_lift_coeff(base,base_conditions)
base.maximum_lift_coefficient = max_CL_base
base.store_diff()
# done!
return nexus
