def initial_sizing(nexus):
for config in nexus.vehicle_configurations:
config.mass_properties.max_zero_fuel = nexus.MZFW_ratio * config.mass_properties.max_takeoff
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
#compute atmosphere conditions for turbofan sizing
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) #freestream conditions
mach_number = air_speed / freestream.speed_of_sound
freestream.mach_number = mach_number
freestream.velocity = air_speed
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
conditions.freestream = freestream
turbofan_sizing(config.propulsors['turbofan'], mach_number, altitude)
compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
# diff the new data
config.store_diff()
return nexus
def initial_sizing(nexus):
for config in nexus.vehicle_configurations:
config.mass_properties.max_zero_fuel = nexus.MZFW_ratio*config.mass_properties.max_takeoff
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
#compute atmosphere conditions for turbofan sizing
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) #freestream conditions
mach_number = air_speed/freestream.speed_of_sound
freestream.mach_number = mach_number
freestream.velocity = air_speed
freestream.gravity = 9.81
conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
conditions.freestream = freestream
turbofan_sizing(config.propulsors['turbofan'], mach_number, altitude)
compute_turbofan_geometry(config.propulsors['turbofan'], conditions)
# diff the new data
config.store_diff()
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.networks['turbofan'],
mach_number=mach_number,
altitude=altitude)
compute_turbofan_geometry(config.networks['turbofan'],
config.nacelles.nacelle_1)
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.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 vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'EMBRAER E190AR'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 51800.0 # kg
vehicle.mass_properties.operating_empty = 29100.0 # kg
vehicle.mass_properties.takeoff = 51800.0 # kg
## vehicle.mass_properties.max_zero_fuel = 0.9 * vehicle.mass_properties.max_takeoff
vehicle.mass_properties.cargo = 0. * Units.kilogram
vehicle.mass_properties.center_of_gravity = [60 * Units.feet, 0,
0] # Not correct
vehicle.mass_properties.moments_of_inertia.tensor = [[10**5, 0, 0],
[
0,
10**6,
0,
], [0, 0, 10**7]
] # Not Correct
# envelope properties
vehicle.envelope.ultimate_load = 3.5
vehicle.envelope.limit_load = 1.5
# basic parameters
vehicle.reference_area = 92.0
vehicle.passengers = 114
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 8.4
wing.sweeps.quarter_chord = 23.0 * Units.deg
wing.thickness_to_chord = 0.11
wing.taper = 0.28
wing.span_efficiency = 1.0
## wing.spans.projected = 27.8
##
## wing.chords.root = 5.203
## wing.chords.tip = 1.460
## wing.chords.mean_aerodynamic = 3.680
wing.areas.reference = 92.0
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
wing.twists.root = 2.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [13.2, 0, 0] # Need to fix
## wing.aerodynamic_center = [3,0,0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.flaps.type = "double_slotted"
wing.flaps.chord = 0.280
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
size_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 5.5
wing.sweeps.quarter_chord = 34.5 * Units.deg
wing.thickness_to_chord = 0.11
wing.taper = 0.11
wing.span_efficiency = 0.9
## wing.spans.projected = 11.958
##
## wing.chords.root = 3.030
## wing.chords.tip = 0.883
## wing.chords.mean_aerodynamic = 2.3840
wing.areas.reference = 26.0
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
wing.twists.root = 2.0 * Units.degrees
wing.twists.tip = 2.0 * Units.degrees
wing.origin = [31., 0, 0] # need to fix
## wing.aerodynamic_center = [2,0,0]
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
size_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 1.7 #
wing.sweeps.quarter_chord = 35 * Units.deg
wing.thickness_to_chord = 0.11
wing.taper = 0.31
wing.span_efficiency = 0.9
## wing.spans.projected = 5.270 #
##
## wing.chords.root = 4.70
## wing.chords.tip = 1.45
## wing.chords.mean_aerodynamic = 3.36
wing.areas.reference = 16.0 #
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [29.5, 0, 0]
## wing.aerodynamic_center = [2,0,0]
wing.vertical = True
wing.symmetric = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
size_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Turbofan Network
# ------------------------------------------------------------------
#initialize the gas turbine network
gt_engine = SUAVE.Components.Energy.Networks.Turbofan()
gt_engine.tag = 'turbofan'
gt_engine.number_of_engines = 2.0
gt_engine.bypass_ratio = 5.4
gt_engine.engine_length = 2.71
gt_engine.nacelle_diameter = 2.05
gt_engine.position[1] = 4.50
#compute engine areas
Awet = 1.1 * np.pi * gt_engine.nacelle_diameter * gt_engine.engine_length
#assign engine areas
gt_engine.areas.wetted = Awet
#set the working fluid for the network
working_fluid = SUAVE.Attributes.Gases.Air
#add working fluid to the network
gt_engine.working_fluid = working_fluid
#Component 1 : ram, to convert freestream static to stagnation quantities
ram = SUAVE.Components.Energy.Converters.Ram()
ram.tag = 'ram'
#add ram to the network
gt_engine.ram = ram
#Component 2 : inlet nozzle
inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
inlet_nozzle.tag = 'inlet nozzle'
inlet_nozzle.polytropic_efficiency = 0.98
inlet_nozzle.pressure_ratio = 0.98 # turbofan.fan_nozzle_pressure_ratio = 0.98 #0.98
#add inlet nozzle to the network
gt_engine.inlet_nozzle = inlet_nozzle
#Component 3 :low pressure compressor
low_pressure_compressor = SUAVE.Components.Energy.Converters.Compressor()
low_pressure_compressor.tag = 'lpc'
low_pressure_compressor.polytropic_efficiency = 0.91
low_pressure_compressor.pressure_ratio = 1.9
#add low pressure compressor to the network
gt_engine.low_pressure_compressor = low_pressure_compressor
#Component 4 :high pressure compressor
high_pressure_compressor = SUAVE.Components.Energy.Converters.Compressor()
high_pressure_compressor.tag = 'hpc'
high_pressure_compressor.polytropic_efficiency = 0.91
high_pressure_compressor.pressure_ratio = 10.0
#add the high pressure compressor to the network
gt_engine.high_pressure_compressor = high_pressure_compressor
#Component 5 :low pressure turbine
low_pressure_turbine = SUAVE.Components.Energy.Converters.Turbine()
low_pressure_turbine.tag = 'lpt'
low_pressure_turbine.mechanical_efficiency = 0.99
low_pressure_turbine.polytropic_efficiency = 0.93
#add low pressure turbine to the network
gt_engine.low_pressure_turbine = low_pressure_turbine
#Component 5 :high pressure turbine
high_pressure_turbine = SUAVE.Components.Energy.Converters.Turbine()
high_pressure_turbine.tag = 'hpt'
high_pressure_turbine.mechanical_efficiency = 0.99
high_pressure_turbine.polytropic_efficiency = 0.93
#add the high pressure turbine to the network
gt_engine.high_pressure_turbine = high_pressure_turbine
#Component 6 :combustor
combustor = SUAVE.Components.Energy.Converters.Combustor()
combustor.tag = 'Comb'
combustor.efficiency = 0.99
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1500
combustor.pressure_ratio = 0.95
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
#add the combustor to the network
gt_engine.combustor = combustor
#Component 7 :core nozzle
core_nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
core_nozzle.tag = 'core nozzle'
core_nozzle.polytropic_efficiency = 0.95
core_nozzle.pressure_ratio = 0.99
#add the core nozzle to the network
gt_engine.core_nozzle = core_nozzle
#Component 8 :fan nozzle
fan_nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
fan_nozzle.tag = 'fan nozzle'
fan_nozzle.polytropic_efficiency = 0.95
fan_nozzle.pressure_ratio = 0.99
#add the fan nozzle to the network
gt_engine.fan_nozzle = fan_nozzle
#Component 9 : fan
fan = SUAVE.Components.Energy.Converters.Fan()
fan.tag = 'fan'
fan.polytropic_efficiency = 0.93
fan.pressure_ratio = 1.7
#add the fan to the network
gt_engine.fan = 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)
thrust.total_design = 37278.0 * Units.N #Newtons
#design sizing conditions
altitude = 35000.0 * Units.ft
mach_number = 0.78
isa_deviation = 0.
# add thrust to the network
gt_engine.thrust = thrust
#size the turbofan
#create conditions object for sizing the turbofan
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
conditions = atmosphere.compute_values(altitude)
turbofan_sizing(gt_engine, mach_number, altitude)
compute_turbofan_geometry(gt_engine, conditions)
# add gas turbine network gt_engine to the vehicle
vehicle.append_component(gt_engine)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'EMBRAER E190AR'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 51800.0 # kg
vehicle.mass_properties.operating_empty = 29100.0 # kg
vehicle.mass_properties.takeoff = 51800.0 # kg
## vehicle.mass_properties.max_zero_fuel = 0.9 * vehicle.mass_properties.max_takeoff
vehicle.mass_properties.cargo = 0. * Units.kilogram
vehicle.mass_properties.center_of_gravity = [60 * Units.feet, 0, 0] # Not correct
vehicle.mass_properties.moments_of_inertia.tensor = [[10 ** 5, 0, 0],[0, 10 ** 6, 0,],[0,0, 10 ** 7]] # Not Correct
# envelope properties
vehicle.envelope.ultimate_load = 3.5
vehicle.envelope.limit_load = 1.5
# basic parameters
vehicle.reference_area = 92.0
vehicle.passengers = 114
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 8.4
wing.sweep = 23.0 * Units.deg
wing.thickness_to_chord = 0.11
wing.taper = 0.28
wing.span_efficiency = 1.0
## wing.spans.projected = 27.8
##
## wing.chords.root = 5.203
## wing.chords.tip = 1.460
## wing.chords.mean_aerodynamic = 3.680
wing.areas.reference = 92.0
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
wing.twists.root = 2.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [13.2,0,0] # Need to fix
## wing.aerodynamic_center = [3,0,0]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.flaps.type = "double_slotted"
wing.flaps.chord = 0.280
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
size_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 5.5
wing.sweep = 34.5 * Units.deg
wing.thickness_to_chord = 0.11
wing.taper = 0.11
wing.span_efficiency = 0.9
## wing.spans.projected = 11.958
##
## wing.chords.root = 3.030
## wing.chords.tip = 0.883
## wing.chords.mean_aerodynamic = 2.3840
wing.areas.reference = 26.0
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
wing.twists.root = 2.0 * Units.degrees
wing.twists.tip = 2.0 * Units.degrees
wing.origin = [31.,0,0] # need to fix
## wing.aerodynamic_center = [2,0,0]
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
size_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 1.7 #
wing.sweep = 35 * Units.deg
wing.thickness_to_chord = 0.11
wing.taper = 0.31
wing.span_efficiency = 0.9
## wing.spans.projected = 5.270 #
##
## wing.chords.root = 4.70
## wing.chords.tip = 1.45
## wing.chords.mean_aerodynamic = 3.36
wing.areas.reference = 16.0 #
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.reference
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [29.5,0,0]
## wing.aerodynamic_center = [2,0,0]
wing.vertical = True
wing.symmetric = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
size_planform(wing)
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Turbofan Network
# ------------------------------------------------------------------
#initialize the gas turbine network
gt_engine = SUAVE.Components.Energy.Networks.Turbofan()
gt_engine.tag = 'turbofan'
gt_engine.number_of_engines = 2.0
gt_engine.bypass_ratio = 5.4
gt_engine.engine_length = 2.71
gt_engine.nacelle_diameter = 2.05
gt_engine.position[1] = 4.50
#compute engine areas
Awet = 1.1*np.pi*gt_engine.nacelle_diameter*gt_engine.engine_length
#assign engine areas
gt_engine.areas.wetted = Awet
#set the working fluid for the network
working_fluid = SUAVE.Attributes.Gases.Air
#add working fluid to the network
gt_engine.working_fluid = working_fluid
#Component 1 : ram, to convert freestream static to stagnation quantities
ram = SUAVE.Components.Energy.Converters.Ram()
ram.tag = 'ram'
#add ram to the network
gt_engine.ram = ram
#Component 2 : inlet nozzle
inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
inlet_nozzle.tag = 'inlet nozzle'
inlet_nozzle.polytropic_efficiency = 0.98
inlet_nozzle.pressure_ratio = 0.98 # turbofan.fan_nozzle_pressure_ratio = 0.98 #0.98
#add inlet nozzle to the network
gt_engine.inlet_nozzle = inlet_nozzle
#Component 3 :low pressure compressor
low_pressure_compressor = SUAVE.Components.Energy.Converters.Compressor()
low_pressure_compressor.tag = 'lpc'
low_pressure_compressor.polytropic_efficiency = 0.91
low_pressure_compressor.pressure_ratio = 1.9
#add low pressure compressor to the network
gt_engine.low_pressure_compressor = low_pressure_compressor
#Component 4 :high pressure compressor
high_pressure_compressor = SUAVE.Components.Energy.Converters.Compressor()
high_pressure_compressor.tag = 'hpc'
high_pressure_compressor.polytropic_efficiency = 0.91
high_pressure_compressor.pressure_ratio = 10.0
#add the high pressure compressor to the network
gt_engine.high_pressure_compressor = high_pressure_compressor
#Component 5 :low pressure turbine
low_pressure_turbine = SUAVE.Components.Energy.Converters.Turbine()
low_pressure_turbine.tag='lpt'
low_pressure_turbine.mechanical_efficiency = 0.99
low_pressure_turbine.polytropic_efficiency = 0.93
#add low pressure turbine to the network
gt_engine.low_pressure_turbine = low_pressure_turbine
#Component 5 :high pressure turbine
high_pressure_turbine = SUAVE.Components.Energy.Converters.Turbine()
high_pressure_turbine.tag='hpt'
high_pressure_turbine.mechanical_efficiency = 0.99
high_pressure_turbine.polytropic_efficiency = 0.93
#add the high pressure turbine to the network
gt_engine.high_pressure_turbine = high_pressure_turbine
#Component 6 :combustor
combustor = SUAVE.Components.Energy.Converters.Combustor()
combustor.tag = 'Comb'
combustor.efficiency = 0.99
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1500
combustor.pressure_ratio = 0.95
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
#add the combustor to the network
gt_engine.combustor = combustor
#Component 7 :core nozzle
core_nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
core_nozzle.tag = 'core nozzle'
core_nozzle.polytropic_efficiency = 0.95
core_nozzle.pressure_ratio = 0.99
#add the core nozzle to the network
gt_engine.core_nozzle = core_nozzle
#Component 8 :fan nozzle
fan_nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
fan_nozzle.tag = 'fan nozzle'
fan_nozzle.polytropic_efficiency = 0.95
fan_nozzle.pressure_ratio = 0.99
#add the fan nozzle to the network
gt_engine.fan_nozzle = fan_nozzle
#Component 9 : fan
fan = SUAVE.Components.Energy.Converters.Fan()
fan.tag = 'fan'
fan.polytropic_efficiency = 0.93
fan.pressure_ratio = 1.7
#add the fan to the network
gt_engine.fan = 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)
thrust.total_design = 37278.0* Units.N #Newtons
#design sizing conditions
altitude = 35000.0*Units.ft
mach_number = 0.78
isa_deviation = 0.
# add thrust to the network
gt_engine.thrust = thrust
#size the turbofan
#create conditions object for sizing the turbofan
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
conditions = atmosphere.compute_values(altitude)
turbofan_sizing(gt_engine,mach_number,altitude)
compute_turbofan_geometry(gt_engine, conditions)
# add gas turbine network gt_engine to the vehicle
vehicle.append_component(gt_engine)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
return vehicle
