''' Concorde.py Created: Jan 2026, M Clarke ''' # ---------------------------------------------------------------------------------------------------------------------- # IMPORT # ---------------------------------------------------------------------------------------------------------------------- # RCAIDE imports import RCAIDE from RCAIDE.Framework.Core import Units , Data from RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet import design_turbojet from RCAIDE.Library.Plots import * # python imports import numpy as np from copy import deepcopy import matplotlib.pyplot as plt import os import sys def main(): # Step 1: design a vehicle vehicle = vehicle_setup() # Step 2: create aircraft configuration based on vehicle configs = configs_setup(vehicle) # Step 3: set up analysis analyses = analyses_setup(configs) # Step 4: set up a flight mission mission = mission_setup(analyses) missions = missions_setup(mission) # Step 5: execute flight profile results = missions.base_mission.evaluate() # Step 6: plot results plot_mission(results) return def vehicle_setup(): # ------------------------------------------------------------------ # File Paths # ------------------------------------------------------------------ ospath = os.path.abspath(__file__) separator = os.path.sep local_path = os.path.dirname(ospath) + separator + '..' airfoil_file_path = local_path + separator + 'Airfoils_and_Polars' + separator polar_file_path = local_path + separator + 'Airfoils_and_Polars' + separator + 'Polars' + separator #------------------------------------------------------------------------------------------------------------------------------------ # ################################################# Vehicle-level Properties ######################################################## #------------------------------------------------------------------------------------------------------------------------------------ vehicle = RCAIDE.Vehicle() vehicle.tag = 'Concorde' # mass properties vehicle.mass_properties.max_takeoff = 185000. # kg vehicle.mass_properties.operating_empty = 78700. # kg vehicle.mass_properties.takeoff = 185000 # kg, adjusted due to significant fuel burn on runway vehicle.mass_properties.cargo = 1000. * Units.kilogram vehicle.mass_properties.max_zero_fuel = 92000. vehicle.mass_properties.max_fuel = 95680 # envelope properties vehicle.flight_envelope.design_cruise_altitude = 40000 * Units.feet vehicle.flight_envelope.design_dynamic_pressure = 55169 vehicle.flight_envelope.design_mach_number = 2.05 vehicle.flight_envelope.ultimate_load = 5.7 vehicle.flight_envelope.limit_load = 3.8 vehicle.flight_envelope.positive_limit_load = 2.5 vehicle.flight_envelope.design_range = 4488 * Units.nmi # basic parameters vehicle.reference_area = 358.25 vehicle.number_of_passengers = 100 vehicle.systems.control = "fully powered" vehicle.systems.accessories = "sst" vehicle.maximum_cross_sectional_area = 13.9 vehicle.total_length = 61.66 vehicle.design_mach_number = 2.02 vehicle.design_range = 4505 * Units.miles vehicle.design_cruise_alt = 60000.0 * Units.ft #------------------------------------------------------------------------------------------------------------------------------------ # ##################################################### Landing Gear ################################################################ #------------------------------------------------------------------------------------------------------------------------------------ main_gear = RCAIDE.Library.Components.Landing_Gear.Main_Landing_Gear() main_gear.origin = [[17, 0, -11]] main_gear.tire_diameter = 47 * Units.inches main_gear.rim_diameter = 22.1 * Units.inches main_gear.tire_width = 15.75 * Units.inches main_gear.strut_length = 12 * Units.ft main_gear.wheels = 8 main_gear.number_of_gear_types_in_tandem = 2 main_gear.number_of_wheels_in_gear_type = 2 main_gear.xz_plane_symmetric = True vehicle.append_component(main_gear) nose_gear = RCAIDE.Library.Components.Landing_Gear.Nose_Landing_Gear() nose_gear.origin = [[35, 0, -11]] nose_gear.tire_diameter = 31 * Units.inches nose_gear.rim_diameter = 14 * Units.inches nose_gear.tire_width = 10.75 * Units.inches nose_gear.strut_length = 12 * Units.ft nose_gear.wheels = 2 nose_gear.number_of_gear_types_in_tandem = 1 nose_gear.number_of_wheels_in_gear_type = 2 vehicle.append_component(nose_gear) #------------------------------------------------------------------------------------------------------------------------------------ # ######################################################## Wings #################################################################### #------------------------------------------------------------------------------------------------------------------------------------ # ------------------------------------------------------------------ # Main Wing # ------------------------------------------------------------------ wing = RCAIDE.Library.Components.Wings.Main_Wing() wing.tag = 'main_wing' wing.aspect_ratio = 1.83 wing.sweeps.quarter_chord = 59.5 * Units.deg wing.sweeps.leading_edge = 66.5 * Units.deg wing.thickness_to_chord = 0.03 wing.taper = 0. wing.spans.projected = 25.6 wing.chords.root = 33.8 wing.total_length = 33.8 wing.chords.tip = 1.1 wing.chords.mean_aerodynamic = 18.4 wing.areas.reference = 358.25 wing.areas.wetted = 601. wing.areas.exposed = 326.5 wing.areas.affected = .6*wing.areas.reference wing.twists.root = 0.0 * Units.degrees wing.twists.tip = 0.0 * Units.degrees wing.origin = [[14,0,-.8]] wing.aerodynamic_center = [35,0,0] wing.vertical = False wing.xz_plane_symmetric = True wing.vortex_lift = True wing.high_mach = True wing.dynamic_pressure_ratio = 1.0 # set root sweep with inner section segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'section_1' segment.percent_span_location = 0. segment.twist = 0. * Units.deg segment.root_chord_percent = 1 segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 67. * Units.deg segment.thickness_to_chord = 0.03 wing_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() wing_airfoil.coordinate_file = airfoil_file_path + 'NACA65_203.txt' segment.append_airfoil(wing_airfoil) wing.append_segment(segment) # set section 2 start point segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'section_2' segment.percent_span_location = (6.15 * 2) /wing.spans.projected segment.twist = 0. * Units.deg segment.root_chord_percent = 13.8/wing.chords.root segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 48. * Units.deg segment.thickness_to_chord = 0.03 wing_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() wing_airfoil.coordinate_file = airfoil_file_path + 'NACA65_203.txt' segment.append_airfoil(wing_airfoil) wing.append_segment(segment) # set section 3 start point segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'section_3' segment.percent_span_location = (12.1 *2) /wing.spans.projected segment.twist = 0. * Units.deg segment.root_chord_percent = 4.4/wing.chords.root segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 71. * Units.deg segment.thickness_to_chord = 0.03 wing_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() wing_airfoil.coordinate_file = airfoil_file_path + 'NACA65_203.txt' segment.append_airfoil(wing_airfoil) wing.append_segment(segment) # set tip segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'tip' segment.percent_span_location = 1. segment.twist = 0. * Units.deg segment.root_chord_percent = 1.1/wing.chords.root segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 0. segment.thickness_to_chord = 0.03 wing_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() wing_airfoil.coordinate_file = airfoil_file_path + 'NACA65_203.txt' segment.append_airfoil(wing_airfoil) wing.append_segment(segment) # control surfaces ------------------------------------------- slat = RCAIDE.Library.Components.Wings.Control_Surfaces.Slat() slat.tag = 'slat' slat.span_fraction_start = 0.2 slat.span_fraction_end = 0.963 slat.deflection = 0.0 * Units.degrees slat.chord_fraction = 0.075 wing.append_control_surface(slat) flap = RCAIDE.Library.Components.Wings.Control_Surfaces.Flap() flap.tag = 'flap' flap.span_fraction_start = 0.42 flap.span_fraction_end = 0.62 flap.deflection = 0.0 * Units.degrees flap.configuration_type = 'single_slotted' flap.chord_fraction = 0.14 wing.append_control_surface(flap) aileron = RCAIDE.Library.Components.Wings.Control_Surfaces.Aileron() aileron.tag = 'aileron' aileron.span_fraction_start = 0.62 aileron.span_fraction_end = 0.83 aileron.deflection = 0.0 * Units.degrees aileron.chord_fraction = 0.22 wing.append_control_surface(aileron) spoiler = RCAIDE.Library.Components.Wings.Control_Surfaces.Elevator() spoiler.tag = 'elevator' spoiler.span_fraction_start = 0.08 spoiler.span_fraction_end = 0.24 spoiler.deflection = 0.0 * Units.degrees spoiler.chord_fraction = 0.1 wing.append_control_surface(spoiler) # add to vehicle vehicle.append_component(wing) # ------------------------------------------------------------------ # Vertical Stabilizer # ------------------------------------------------------------------ wing = RCAIDE.Library.Components.Wings.Vertical_Tail() wing.tag = 'vertical_stabilizer' wing.aspect_ratio = 0.74 wing.sweeps.quarter_chord = 60 * Units.deg wing.thickness_to_chord = 0.04 wing.taper = 0.14 wing.spans.projected = 6.0 wing.chords.root = 14.5 wing.total_length = 14.5 wing.chords.tip = 2.7 wing.chords.mean_aerodynamic = 8.66 wing.areas.reference = 33.91 wing.areas.wetted = 76. wing.areas.exposed = 38. wing.areas.affected = 33.91 wing.twists.root = 0.0 * Units.degrees wing.twists.tip = 0.0 * Units.degrees wing.origin = [[42.,0,1.]] wing.aerodynamic_center = [50,0,0] wing.vertical = True wing.xz_plane_symmetric = False wing.t_tail = False wing.high_mach = True wing.dynamic_pressure_ratio = 1.0 # set root sweep with inner section segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'section_1' segment.percent_span_location = 0.0 segment.twist = 0. * Units.deg segment.root_chord_percent = 14.5/14.5 segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 63. * Units.deg segment.thickness_to_chord = 0.04 tail_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() tail_airfoil.coordinate_file = airfoil_file_path + 'supersonic_tail.txt' segment.append_airfoil(tail_airfoil) wing.append_segment(segment) # set mid section start point segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'section_2' segment.percent_span_location = 2.4/(6.0) + wing.segments['section_1'].percent_span_location segment.twist = 0. * Units.deg segment.root_chord_percent = 7.5/14.5 segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 40. * Units.deg segment.thickness_to_chord = 0.04 tail_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() tail_airfoil.coordinate_file = airfoil_file_path + 'supersonic_tail.txt' segment.append_airfoil(tail_airfoil) wing.append_segment(segment) # set tip segment = RCAIDE.Library.Components.Wings.Segments.Segment() segment.tag = 'tip' segment.percent_span_location = 1. segment.twist = 0. * Units.deg segment.root_chord_percent = 2.7/14.5 segment.dihedral_outboard = 0. segment.sweeps.quarter_chord = 0. segment.thickness_to_chord = 0.04 tail_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil() tail_airfoil.coordinate_file = airfoil_file_path + 'supersonic_tail.txt' segment.append_airfoil(tail_airfoil) wing.append_segment(segment) # add to vehicle vehicle.append_component(wing) # ------------------------------------------------------------------ # Fuselage # ------------------------------------------------------------------ fuselage = RCAIDE.Library.Components.Fuselages.Fuselage() fuselage.seats_abreast = 4 fuselage.seat_pitch = 38. * Units.inches fuselage.fineness.nose = 4.3 fuselage.fineness.tail = 6.4 fuselage.lengths.total = 61.66 fuselage.width = 2.88 fuselage.heights.maximum = 3.32 fuselage.heights.maximum = 3.32 fuselage.supersonic = True fuselage.heights.at_quarter_length = 3.32 fuselage.heights.at_wing_root_quarter_chord = 3.32 fuselage.heights.at_three_quarters_length = 3.32 fuselage.areas.wetted = 442. fuselage.areas.front_projected = 11.9 fuselage.effective_diameter = 3.1 fuselage.differential_pressure = 7.4e4 * Units.pascal # Maximum differential pressure cabin = RCAIDE.Library.Components.Fuselages.Cabins.Cabin() cabin.origin = [[8.5, 0, -0.5]] business_class = RCAIDE.Library.Components.Fuselages.Cabins.Classes.Business() business_class.number_of_seats_abrest = 4 business_class.number_of_rows = 32 business_class.seat_arm_rest_width = 3 * Units.inches business_class.seat_width = 17 * Units.inches business_class.aisle_width = 13 * Units.inches business_class.galley_lavatory_percent_x_locations = [0, 0.5, 0.51, 1] business_class.type_A_exit_percent_x_locations = [0.02, 0.4, 1] cabin.append_cabin_class(business_class) fuselage.append_cabin(cabin) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_0' segment.percent_x_location = 0.0000 segment.percent_z_location = -0.61 /fuselage.lengths.total fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_1' segment.percent_x_location = 3.02870/fuselage.lengths.total segment.percent_z_location = -0.3583/fuselage.lengths.total segment.height = 1.4502 segment.width = 1.567 fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_2' segment.percent_x_location = 5.7742/fuselage.lengths.total segment.percent_z_location = -0.1500/fuselage.lengths.total segment.height = 2.356 segment.width = 2.3429 fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_3' segment.percent_x_location = 9.0791/fuselage.lengths.total segment.percent_z_location = 0 segment.height = 3.0581 segment.width = 2.741 fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_4' segment.percent_x_location = 12.384/fuselage.lengths.total segment.percent_z_location = 0 segment.height = 3.3200 segment.width = 2.880 fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_6' segment.percent_x_location = 43.228 /fuselage.lengths.total segment.percent_z_location = 0 segment.height = 3.3200 segment.width = 2.8800 fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_6' segment.percent_x_location = 47.5354/fuselage.lengths.total segment.percent_z_location = 0.100/fuselage.lengths.total segment.height = 2.952 segment.width = 2.8800 fuselage.append_segment(segment) # Segment segment = RCAIDE.Library.Components.Fuselages.Segments.Segment() segment.tag = 'segment_7' segment.percent_x_location = 1 segment.percent_z_location = 1.2332/fuselage.lengths.total segment.height = 0.00100 segment.width = 0.00100 fuselage.append_segment(segment) vehicle.append_component(fuselage) #------------------------------------------------------------------------------------------------------------------------------------ # ########################################################## Energy Network ######################################################### #------------------------------------------------------------------------------------------------------------------------------------ #initialize the fuel network net = RCAIDE.Framework.Networks.Fuel() net.identical_propulsors = True #------------------------------------------------------------------------------------------------------------------------------------ # Fuel Distribution Line #------------------------------------------------------------------------------------------------------------------------------------ fuel_line = RCAIDE.Library.Components.Powertrain.Distributors.Fuel_Line() #------------------------------------------------------------------------------------------------------------------------------------ # Inner Right Propulsor #------------------------------------------------------------------------------------------------------------------------------------ outer_right_turbojet = RCAIDE.Library.Components.Powertrain.Propulsors.Turbojet() outer_right_turbojet.tag = 'outer_right_turbojet' outer_right_turbojet.length = 4.039 outer_right_turbojet.nacelle_diameter = 1.3 outer_right_turbojet.inlet_diameter = 1.212 outer_right_turbojet.areas.wetted = 30 outer_right_turbojet.design_altitude = 60000.0*Units.ft outer_right_turbojet.design_mach_number = 2.02 outer_right_turbojet.design_thrust = 10000. * Units.lbf outer_right_turbojet.origin = [[37.,5.5,-1.6]] outer_right_turbojet.working_fluid = RCAIDE.Library.Attributes.Gases.Air() # Ram ram = RCAIDE.Library.Components.Powertrain.Converters.Ram() ram.tag = 'ram' outer_right_turbojet.ram = ram # Inlet Nozzle inlet_nozzle = RCAIDE.Library.Components.Powertrain.Converters.Compression_Nozzle() inlet_nozzle.tag = 'inlet_nozzle' inlet_nozzle.polytropic_efficiency = 1.0 inlet_nozzle.pressure_ratio = 1.0 inlet_nozzle.pressure_recovery = 0.94 outer_right_turbojet.inlet_nozzle = inlet_nozzle # Low Pressure Compressor lp_compressor = RCAIDE.Library.Components.Powertrain.Converters.Compressor() lp_compressor.tag = 'low_pressure_compressor' lp_compressor.polytropic_efficiency = 0.88 lp_compressor.pressure_ratio = 3.1 outer_right_turbojet.low_pressure_compressor = lp_compressor # High Pressure Compressor hp_compressor = RCAIDE.Library.Components.Powertrain.Converters.Compressor() hp_compressor.tag = 'high_pressure_compressor' hp_compressor.polytropic_efficiency = 0.88 hp_compressor.pressure_ratio = 5.0 outer_right_turbojet.high_pressure_compressor = hp_compressor # Low Pressure Turbine lp_turbine = RCAIDE.Library.Components.Powertrain.Converters.Turbine() lp_turbine.tag ='low_pressure_turbine' lp_turbine.mechanical_efficiency = 0.99 lp_turbine.polytropic_efficiency = 0.89 outer_right_turbojet.low_pressure_turbine = lp_turbine # High Pressure Turbine hp_turbine = RCAIDE.Library.Components.Powertrain.Converters.Turbine() hp_turbine.tag ='high_pressure_turbine' hp_turbine.mechanical_efficiency = 0.99 hp_turbine.polytropic_efficiency = 0.87 outer_right_turbojet.high_pressure_turbine = hp_turbine # Combustor combustor = RCAIDE.Library.Components.Powertrain.Converters.Combustor() combustor.tag = 'combustor' combustor.efficiency = 0.96 combustor.alphac = 1.0 combustor.turbine_inlet_temperature = 1440. combustor.pressure_ratio = 0.92 combustor.fuel_data = RCAIDE.Library.Attributes.Propellants.Jet_A() outer_right_turbojet.combustor = combustor # Afterburner afterburner = RCAIDE.Library.Components.Powertrain.Converters.Combustor() afterburner.tag = 'afterburner' afterburner.efficiency = 0.95 afterburner.alphac = 1.0 afterburner.turbine_inlet_temperature = 1500 afterburner.pressure_ratio = 1.0 afterburner.fuel_data = RCAIDE.Library.Attributes.Propellants.Jet_A() outer_right_turbojet.afterburner = afterburner # Core Nozzle nozzle = RCAIDE.Library.Components.Powertrain.Converters.Supersonic_Nozzle() nozzle.tag = 'core_nozzle' nozzle.pressure_recovery = 0.95 nozzle.pressure_ratio = 1. outer_right_turbojet.core_nozzle = nozzle # design turbojet design_turbojet(outer_right_turbojet) nacelle = RCAIDE.Library.Components.Nacelles.Stack_Nacelle() nacelle.diameter = 1.3 nacelle.tag = 'nacelle_1' nacelle.origin = [[37.,5.5,-1.6]] nacelle.length = 10 nacelle.inlet_diameter = 1.1 nacelle.areas.wetted = 30. nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Rounded_Rectangle_Segment() nac_segment.tag = 'segment_1' nac_segment.orientation_euler_angles = [0., -45*Units.degrees,0.] nac_segment.percent_x_location = 0.0 nac_segment.height = 2.12 nac_segment.width = 1.5 nac_segment.curvature = 10 nacelle.append_segment(nac_segment) nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Rounded_Rectangle_Segment() nac_segment.tag = 'segment_2' nac_segment.percent_x_location = 1.0 nac_segment.height = 1.5 nac_segment.width = 1.5 nac_segment.curvature = 10 nacelle.append_segment(nac_segment) outer_right_turbojet.nacelle = nacelle net.propulsors.append(outer_right_turbojet) #------------------------------------------------------------------------------------------------------------------------------------ # Inner Right Propulsor #------------------------------------------------------------------------------------------------------------------------------------ inner_right_turbojet = deepcopy(outer_right_turbojet) inner_right_turbojet.tag = 'inner_right_turbojet' inner_right_turbojet.origin = [[37.,4,-1.6]] nacelle_2 = deepcopy(nacelle) nacelle_2.tag = 'nacelle_2' nacelle_2.origin = [[37.,4,-1.6]] inner_right_turbojet.nacelle = nacelle_2 net.propulsors.append(inner_right_turbojet) #------------------------------------------------------------------------------------------------------------------------------------ # Inner Right Propulsor #------------------------------------------------------------------------------------------------------------------------------------ inner_left_turbojet = deepcopy(outer_right_turbojet) inner_left_turbojet.tag = 'inner_left_turbojet' inner_left_turbojet.origin = [[37.,-4,-1.6]] nacelle_3 = deepcopy(nacelle) nacelle_3.tag = 'nacelle_3' nacelle_3.origin = [[37.,-4,-1.6]] inner_left_turbojet.nacelle = nacelle_3 net.propulsors.append(inner_left_turbojet) #------------------------------------------------------------------------------------------------------------------------------------ # Inner Left Propulsor #------------------------------------------------------------------------------------------------------------------------------------ outer_left_turbojet = deepcopy(outer_right_turbojet) outer_left_turbojet.tag = 'outer_left_turbojet' outer_left_turbojet.origin = [[37.,-5.5,-1.6]] nacelle_4 = deepcopy(nacelle) nacelle_4.tag = 'nacelle_4' nacelle_4.origin = [[37.,-5.5,-1.6]] outer_left_turbojet.nacelle = nacelle_4 net.propulsors.append(outer_left_turbojet) #------------------------------------------------------------------------------------------------------------------------------------ # Fuel Tank & Fuel #------------------------------------------------------------------------------------------------------------------------------------ fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_9' fuel_tank.mass_properties.center_of_gravity = np.array([[26.5,0,0]]) fuel_tank.mass_properties.mass = 11096 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_10' fuel_tank.mass_properties.center_of_gravity = np.array([[28.7,0,0]]) fuel_tank.mass_properties.mass = 11943 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_1_and_4' fuel_tank.mass_properties.center_of_gravity = np.array([[31.0,0,0]]) fuel_tank.mass_properties.mass = 4198+4198 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_5_and_8' fuel_tank.mass_properties.center_of_gravity = np.array([[32.9,0,0]]) fuel_tank.mass_properties.mass = 7200+12838 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_6_and_7' fuel_tank.mass_properties.center_of_gravity = np.array([[37.4,0,0]]) fuel_tank.mass_properties.mass = 11587+7405 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_5A_and_7A' fuel_tank.mass_properties.center_of_gravity = np.array([[40.2,0,0]]) fuel_tank.mass_properties.mass = 2225+2225 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_2_and_3' fuel_tank.mass_properties.center_of_gravity = np.array([[40.2,0,0]]) fuel_tank.mass_properties.mass = 4570+4570 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) fuel_tank = RCAIDE.Library.Components.Powertrain.Sources.Fuel_Tanks.Fuel_Tank() fuel_tank.tag = 'tank_11' fuel_tank.mass_properties.center_of_gravity = np.array([[49.8,0,0]]) fuel_tank.mass_properties.mass = 10415 fuel_tank.fuel_flow_split_ratio = 1/8 fuel_tank.fuel = RCAIDE.Library.Attributes.Propellants.Jet_A() fuel_line.fuel_tanks.append(fuel_tank) #------------------------------------------------------------------------------------------------------------------------------------ # Assign propulsors to fuel line to network fuel_line.assigned_propulsors = [[outer_left_turbojet.tag,inner_left_turbojet.tag, outer_right_turbojet.tag, inner_right_turbojet.tag]] # Append fuel line to network net.fuel_lines.append(fuel_line) #------------------------------------------------------------------------------------------------------------------------------------ # Append energy network to aircraft vehicle.append_energy_network(net) return vehicle # ---------------------------------------------------------------------- # Define the Configurations # --------------------------------------------------------------------- def configs_setup(vehicle): # ------------------------------------------------------------------ # Initialize Configurations # ------------------------------------------------------------------ configs = RCAIDE.Library.Components.Configs.Config.Container() base_config = RCAIDE.Library.Components.Configs.Config(vehicle) base_config.tag = 'base' configs.append(base_config) # ------------------------------------------------------------------ # Cruise Configuration # ------------------------------------------------------------------ config = RCAIDE.Library.Components.Configs.Config(base_config) config.tag = 'cruise' configs.append(config) # ------------------------------------------------------------------ # Afterburner Climb Configuration # ------------------------------------------------------------------ config = RCAIDE.Library.Components.Configs.Config(base_config) config.tag = 'climb' config.wings['main_wing'].control_surfaces.flap.deflection = 5. * Units.deg for propulsor in config.networks.fuel.propulsors: propulsor.afterburner_active = True configs.append(config) # ------------------------------------------------------------------ # Takeoff Configuration # ------------------------------------------------------------------ config = RCAIDE.Library.Components.Configs.Config(base_config) config.tag = 'takeoff' config.wings['main_wing'].control_surfaces.flap.deflection = 15. * Units.deg for propulsor in config.networks.fuel.propulsors: propulsor.afterburner_active = True configs.append(config) # ------------------------------------------------------------------ # Landing Configuration # ------------------------------------------------------------------ config = RCAIDE.Library.Components.Configs.Config(base_config) config.tag = 'landing' config.wings['main_wing'].control_surfaces.flap.deflection = 30. * Units.deg config.wings['main_wing'].control_surfaces.slat.deflection = 25. * Units.deg configs.append(config) # done! return configs # ---------------------------------------------------------------------- # Define the Vehicle Analyses # ---------------------------------------------------------------------- def analyses_setup(configs): analyses = RCAIDE.Framework.Analyses.Analysis.Container() # build a base analysis for each config for tag,config in list(configs.items()): analysis = base_analysis(config) analyses[tag] = analysis return analyses def base_analysis(vehicle): # ------------------------------------------------------------------ # Initialize the Analyses # ------------------------------------------------------------------ analyses = RCAIDE.Framework.Analyses.Vehicle() analyses.vehicle = vehicle # Geometry geometry = RCAIDE.Framework.Analyses.Geometry.Geometry() analyses.append(geometry) # ------------------------------------------------------------------ # Weights weights = RCAIDE.Framework.Analyses.Weights.Conventional_Transport() analyses.append(weights) # ------------------------------------------------------------------ # Aerodynamics Analysis aerodynamics = RCAIDE.Framework.Analyses.Aerodynamics.Vortex_Lattice_Method() aerodynamics.vortex_lift = True analyses.append(aerodynamics) # ------------------------------------------------------------------ # Emissions emissions = RCAIDE.Framework.Analyses.Emissions.Emission_Index_Correlation_Method() analyses.append(emissions) # ------------------------------------------------------------------ # Energy energy= RCAIDE.Framework.Analyses.Energy.Energy() analyses.append(energy) # ------------------------------------------------------------------ # Planet Analysis planet = RCAIDE.Framework.Analyses.Planets.Earth() analyses.append(planet) # ------------------------------------------------------------------ # Atmosphere Analysis atmosphere = RCAIDE.Framework.Analyses.Atmospheric.US_Standard_1976() analyses.append(atmosphere) # done! return analyses # ---------------------------------------------------------------------- # Plot Mission # ---------------------------------------------------------------------- def plot_mission(results): plot_altitude_sfc_weight(results) plot_flight_conditions(results) plot_aerodynamic_coefficients(results) plot_aircraft_velocities(results) plot_drag_components(results) plot_emission_indices(results) plot_emission_species_masses(results) plot_CO2e_emissions(results) plot_aerodynamic_forces(results) return # ---------------------------------------------------------------------- # Define the Mission # ---------------------------------------------------------------------- def mission_setup(analyses): # ------------------------------------------------------------------ # Initialize the Mission # ------------------------------------------------------------------ mission = RCAIDE.Framework.Mission.Sequential_Segments() mission.tag = 'the_mission' # unpack Segments module Segments = RCAIDE.Framework.Mission.Segments base_segment = Segments.Segment() # ------------------------------------------------------------------ # First Climb Segment: constant Mach, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Linear_Speed_Constant_Rate(base_segment) segment.tag = "climb_1" segment.analyses.extend( analyses.takeoff ) segment.altitude_start = 0.0 * Units.km segment.altitude_end = 4000. * Units.ft segment.air_speed_start = 250. * Units.kts segment.air_speed_end = 350. * Units.kts segment.climb_rate = 4000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Second Climb Segment: constant Speed, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Linear_Mach_Constant_Rate(base_segment) segment.tag = "climb_2" segment.analyses.extend( analyses.climb ) segment.altitude_end = 8000. * Units.ft segment.mach_number_end = 0.45 segment.climb_rate = 2000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Second Climb Segment: constant Speed, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Linear_Mach_Constant_Rate(base_segment) segment.tag = "climb_2" segment.analyses.extend( analyses.cruise ) segment.altitude_start = 8000. * Units.ft segment.altitude_end = 33000. * Units.ft segment.mach_number_end = 0.95 segment.climb_rate = 3000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Third Climb Segment: linear Mach, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Linear_Mach_Constant_Rate(base_segment) segment.tag = "climb_3" segment.analyses.extend( analyses.cruise ) segment.altitude_start = 33000. * Units.ft segment.altitude_end = 34000. * Units.ft segment.mach_number_start = 0.95 segment.mach_number_end = 1.1 segment.climb_rate = 3000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Third Climb Segment: linear Mach, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Linear_Mach_Constant_Rate(base_segment) segment.tag = "climb_4" segment.analyses.extend( analyses.cruise ) segment.altitude_end = 40000. * Units.ft segment.mach_number_start = 1.1 segment.mach_number_end = 1.7 segment.climb_rate = 2000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Fourth Climb Segment: linear Mach, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Linear_Mach_Constant_Rate(base_segment) segment.tag = "climb_5" segment.analyses.extend( analyses.cruise ) segment.altitude_end = 50000. * Units.ft segment.mach_number_start = 1.7 segment.mach_number_end = 2.02 segment.climb_rate = 750. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Fourth Climb Segment: linear Mach, constant segment angle # ------------------------------------------------------------------ segment = Segments.Climb.Constant_Mach_Constant_Rate(base_segment) segment.tag = "climbing_cruise" segment.analyses.extend( analyses.cruise ) segment.altitude_start = 50000. * Units.ft segment.altitude_end = 60000. * Units.ft segment.mach_number = 2.02 segment.climb_rate = 100. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Cruise Segment: constant speed # ------------------------------------------------------------------ segment = Segments.Cruise.Constant_Mach_Constant_Altitude(base_segment) segment.tag = "level_cruise" segment.analyses.extend( analyses.cruise ) segment.altitude = 60000. * Units.ft segment.mach_number = 2.02 segment.distance = 1635. * Units.nmi # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # First Descent Segment # ------------------------------------------------------------------ segment = Segments.Descent.Linear_Speed_Constant_Rate(base_segment) segment.tag = "descent_2" segment.analyses.extend( analyses.cruise ) segment.altitude_end = 41000. * Units.ft segment.air_speed_end = 800 * Units.mph segment.descent_rate = 1000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # First Descent Segment # ------------------------------------------------------------------ segment = Segments.Descent.Linear_Mach_Constant_Rate(base_segment) segment.tag = "descent_3" segment.analyses.extend( analyses.cruise ) segment.altitude_end = 10000. * Units.ft segment.mach_number_end = 0.4 segment.descent_rate = 1000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # First Descent Segment # ------------------------------------------------------------------ segment = Segments.Descent.Constant_Speed_Constant_Rate(base_segment) segment.tag = "descent_4" segment.analyses.extend( analyses.landing ) segment.altitude_end = 0. * Units.ft segment.descent_rate = 1000. * Units['ft/min'] # define flight dynamics to model segment.flight_dynamics.force_x = True segment.flight_dynamics.force_z = True # define flight controls segment.assigned_control_variables.throttle.active = True segment.assigned_control_variables.throttle.assigned_propulsors = [['inner_right_turbojet','outer_right_turbojet','outer_left_turbojet','inner_left_turbojet']] segment.assigned_control_variables.pitch_angle.active = True mission.append_segment(segment) # ------------------------------------------------------------------ # Mission definition complete # ------------------------------------------------------------------ return mission def missions_setup(mission): missions = RCAIDE.Framework.Mission.Missions() mission.tag = 'base_mission' missions.append(mission) # done! return missions if __name__ == '__main__': main() plt.show()