Tutorial 7 - Lift + Cruise eVTOL Aircraft Simulation#
Welcome to this tutorial on simulating a lift + cruise eVTOL aircraft using RCAIDE. This guide will walk you through the code, explain its components, and highlight where modifications can be made to customize the simulation for different vehicle designs.
Header and Imports#
The Imports section is divided into two parts: simulation-specific libraries and general-purpose Python libraries.
The RCAIDE Imports section includes the core modules needed for the simulation. These libraries provide specialized classes and tools for building, analyzing, and running aircraft models.
[1]:
#----------------------------------------------------------------------
# Imports
# ---------------------------------------------------------------------
import RCAIDE
from RCAIDE.Framework.Core import Units
from RCAIDE.Library.Methods.Powertrain.Propulsors.Electric_Rotor import design_electric_rotor
from RCAIDE.Library.Plots import *
from RCAIDE.Input_Output import load as load_propulsor
from RCAIDE.Input_Output import save as save_propulsor
import os
import numpy as np
from copy import deepcopy
import matplotlib.pyplot as plt
import pickle
import sys
sys.path.insert(0,(os.path.dirname(os.getcwd())))
Vehicle Setup#
The ``vehicle_setup`` function defines the baseline configuration of the aircraft. This section builds the vehicle step-by-step by specifying its components, geometric properties, and high-level parameters.
1. Creating the Vehicle Instance#
The setup begins by creating a vehicle instance and assigning it a tag. The tag is a unique string identifier used to reference the vehicle during analysis or in post-processing steps.
2. Defining High-Level Vehicle Parameters#
The high-level parameters describe the aircraft’s key operational characteristics, such as:
Maximum Takeoff Weight: The heaviest allowable weight of the aircraft for safe flight.
Operating Empty Weight: The aircraft weight without fuel, passengers, or payload.
Payload: The weight of cargo and passengers.
Max Zero Fuel Weight: The maximum weight of the aircraft excluding fuel.
Units for these parameters can be converted automatically using the Units module to ensure consistency and reduce errors.
3. Defining the Landing Gear#
Landing gear parameters, such as the number of main and nose wheels, are set for the aircraft. While not used in this tutorial, these values can be applied in advanced analyses, such as ground loads or noise prediction.
4. Main Wing Setup#
The main wing is added using the ``Main_Wing`` class. This designation ensures that the primary lifting surface is recognized correctly by the analysis tools. Key properties of the wing include:
Area: The total wing surface area.
Span: The length of the wing from tip to tip.
Aspect Ratio: A ratio of span to average chord, determining wing efficiency.
Segments: Divisions of the wing geometry (e.g., root and tip sections).
Control Surfaces: High-lift devices like flaps and ailerons, defined by span fractions and deflections.
5. Horizontal and Vertical Stabilizers#
The stabilizers provide stability and control for the aircraft:
Horizontal Stabilizer: Defined using the
Horizontal_Tailclass. It follows a similar setup to the main wing but acts as a stabilizing surface.Vertical Stabilizer: Defined using the
Vertical_Tailclass, with an additional option to designate the tail as a T-tail for weight calculations.
6. Fuselage Definition#
The fuselage is modeled by specifying its geometric parameters, such as:
Length: The overall length of the aircraft body.
Width: The widest part of the fuselage cross-section.
Height: The height of the fuselage.
These values influence drag calculations and overall structural weight.
7. Energy Network#
The energy network models the propulsion system. The energy network determines the engine’s thrust, bypass ratio, and fuel type. These parameters are essential for performance and fuel efficiency analyses.
[2]:
# ----------------------------------------------------------------------
# Build the Vehicle
# ----------------------------------------------------------------------
def vehicle_setup(new_regression=True) :
ospath = os.path.abspath(os.path.join('Notebook'))
separator = os.path.sep
rel_path = os.path.dirname(ospath) + separator + '..' + separator + '..' + separator + 'VnV' + separator + 'Vehicles' + separator
airfoil_path = rel_path
local_path = os.path.dirname(ospath) + separator
#------------------------------------------------------------------------------------------------------------------------------------
# ################################################# Vehicle-level Properties ########################################################
#------------------------------------------------------------------------------------------------------------------------------------
vehicle = RCAIDE.Vehicle()
vehicle.tag = 'Lift_Cruise'
vehicle.configuration = 'eVTOL'
# mass properties
vehicle.mass_properties.max_takeoff = 2700
vehicle.mass_properties.takeoff = vehicle.mass_properties.max_takeoff
vehicle.mass_properties.operating_empty = vehicle.mass_properties.max_takeoff
vehicle.flight_envelope.ultimate_load = 5.7
vehicle.flight_envelope.positive_limit_load = 3.
vehicle.passengers = 5
#------------------------------------------------------------------------------------------------------------------------------------
# ######################################################## Wings ####################################################################
#------------------------------------------------------------------------------------------------------------------------------------
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = RCAIDE.Library.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 8.95198 # will be overwritten
wing.sweeps.quarter_chord = 0.0
wing.thickness_to_chord = 0.14
wing.taper = 0.292
wing.spans.projected = 11.82855
wing.chords.root = 1.75
wing.total_length = 1.75
wing.chords.tip = 1.0
wing.chords.mean_aerodynamic = 1.5
wing.dihedral = 0.0
wing.areas.reference = 15.629
wing.twists.root = 4. * Units.degrees
wing.twists.tip = 0.
wing.origin = [[1.5, 0., 0.991]]
wing.aerodynamic_center = [ 1.567, 0., 0.991]
wing.winglet_fraction = 0.0
wing.symmetric = True
wing.vertical = False
# Segment
segment = RCAIDE.Library.Components.Wings.Segments.Segment()
segment.tag = 'Section_1'
segment.percent_span_location = 0.0
segment.twist = 4. * Units.degrees
segment.root_chord_percent = 1.
segment.dihedral_outboard = 8 * Units.degrees
segment.sweeps.quarter_chord = 0.9 * Units.degrees
segment.thickness_to_chord = 0.16
wing.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Wings.Segments.Segment()
segment.tag = 'Section_2'
segment.percent_span_location = 3.5/wing.spans.projected
segment.twist = 3. * Units.degrees
segment.root_chord_percent = 1.4000/1.7500
segment.dihedral_outboard = 0.0 * Units.degrees
segment.sweeps.quarter_chord = 1.27273 * Units.degrees
segment.thickness_to_chord = 0.16
wing.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Wings.Segments.Segment()
segment.tag = 'Section_3'
segment.percent_span_location = 11.3/wing.spans.projected
segment.twist = 2.0 * Units.degrees
segment.root_chord_percent = 1.000/1.7500
segment.dihedral_outboard = 35.000* Units.degrees
segment.sweeps.quarter_chord = 45.000* Units.degrees
segment.thickness_to_chord = 0.16
wing.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Wings.Segments.Segment()
segment.tag = 'Section_4'
segment.percent_span_location = 11.6/wing.spans.projected
segment.twist = 0.0 * Units.degrees
segment.root_chord_percent = 0.9/1.7500
segment.dihedral_outboard = 60. * Units.degrees
segment.sweeps.quarter_chord = 70.0 * Units.degrees
segment.thickness_to_chord = 0.16
wing.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Wings.Segments.Segment()
segment.tag = 'Section_5'
segment.percent_span_location = 1.0
segment.twist = 0.0 * Units.degrees
segment.root_chord_percent = 0.35/1.7500
segment.dihedral_outboard = 0 * Units.degrees
segment.sweeps.quarter_chord = 0 * Units.degrees
segment.thickness_to_chord = 0.16
wing.append_segment(segment)
vehicle.reference_area = wing.areas.reference
wing.areas.wetted = wing.areas.reference * 2
wing.areas.exposed = wing.areas.reference * 2
# control surfaces -------------------------------------------
flap = RCAIDE.Library.Components.Wings.Control_Surfaces.Flap()
flap.tag = 'flap'
flap.span_fraction_start = 0.2
flap.span_fraction_end = 0.5
flap.deflection = 0.0 * Units.degrees
flap.chord_fraction = 0.20
wing.append_control_surface(flap)
aileron = RCAIDE.Library.Components.Wings.Control_Surfaces.Aileron()
aileron.tag = 'aileron'
aileron.span_fraction_start = 0.7
aileron.span_fraction_end = 0.9
aileron.deflection = 0.0 * Units.degrees
aileron.chord_fraction = 0.2
wing.append_control_surface(aileron)
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------------------------------------------------------------------------
# Horizontal Tail
#------------------------------------------------------------------------------------------------------------------------------------
wing = RCAIDE.Library.Components.Wings.Horizontal_Tail()
wing.aspect_ratio = 3.04444
wing.sweeps.quarter_chord = 17. * Units.degrees
wing.thickness_to_chord = 0.12
wing.spans.projected = 2.71805
wing.chords.root = 0.94940
wing.total_length = 0.94940
wing.chords.tip = 0.62731
wing.chords.mean_aerodynamic = 0.809
wing.dihedral = 20 *Units.degrees
wing.taper = wing.chords.tip / wing.chords.root
wing.areas.reference = 2.14279
wing.areas.wetted = 2.14279 * 2
wing.areas.exposed = 2.14279 * 2
wing.twists.root = 0.0
wing.twists.tip = 0.0
wing.origin = [[ 5.374 ,0.0 , 0.596]]
wing.aerodynamic_center = [ 5.374, 0.0, 0.596]
wing.winglet_fraction = 0.0
wing.symmetric = True
elevator = RCAIDE.Library.Components.Wings.Control_Surfaces.Elevator()
elevator.tag = 'elevator'
elevator.span_fraction_start = 0.6
elevator.span_fraction_end = 0.9
elevator.deflection = 0.0 * Units.deg
elevator.chord_fraction = 0.4
wing.append_control_surface(elevator)
rudder = RCAIDE.Library.Components.Wings.Control_Surfaces.Rudder()
rudder.tag = 'rudder'
rudder.span_fraction_start = 0.1
rudder.span_fraction_end = 0.5
rudder.deflection = 0.0 * Units.deg
rudder.chord_fraction = 0.4
wing.append_control_surface(rudder)
# add to vehicle
vehicle.append_component(wing)
#------------------------------------------------------------------------------------------------------------------------------------
# ########################################################## Fuselage ############################################################
#------------------------------------------------------------------------------------------------------------------------------------
fuselage = RCAIDE.Library.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
fuselage.seats_abreast = 2.
fuselage.seat_pitch = 3.
fuselage.fineness.nose = 0.88
fuselage.fineness.tail = 1.13
fuselage.lengths.nose = 0.5
fuselage.lengths.tail = 1.5
fuselage.lengths.cabin = 4.46
fuselage.lengths.total = 6.46
fuselage.width = 5.85 * Units.feet # change
fuselage.heights.maximum = 4.65 * Units.feet # change
fuselage.heights.at_quarter_length = 3.75 * Units.feet # change
fuselage.heights.at_wing_root_quarter_chord = 4.65 * Units.feet # change
fuselage.heights.at_three_quarters_length = 4.26 * Units.feet # change
fuselage.areas.wetted = 236. * Units.feet**2 # change
fuselage.areas.front_projected = 0.14 * Units.feet**2 # change
fuselage.effective_diameter = 1.276 # change
fuselage.differential_pressure = 0.
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_0'
segment.percent_x_location = 0.0
segment.percent_z_location = 0. # change
segment.height = 0.049
segment.width = 0.032
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.10912/fuselage.lengths.total
segment.percent_z_location = 0.00849
segment.height = 0.481
segment.width = 0.553
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.47804/fuselage.lengths.total
segment.percent_z_location = 0.02874
segment.height = 1.00
segment.width = 0.912
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.161
segment.percent_z_location = 0.04348
segment.height = 1.41
segment.width = 1.174
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 0.284
segment.percent_z_location = 0.05435
segment.height = 1.62
segment.width = 1.276
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_5'
segment.percent_x_location = 3.43026/fuselage.lengths.total
segment.percent_z_location = 0.31483/fuselage.lengths.total
segment.height = 1.409
segment.width = 1.121
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_6'
segment.percent_x_location = 4.20546/fuselage.lengths.total
segment.percent_z_location = 0.32216/fuselage.lengths.total
segment.height = 1.11
segment.width = 0.833
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_7'
segment.percent_x_location = 4.99358/fuselage.lengths.total
segment.percent_z_location = 0.37815/fuselage.lengths.total
segment.height = 0.78
segment.width = 0.512
fuselage.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Fuselages.Segments.Segment()
segment.tag = 'segment_8'
segment.percent_x_location = 1.
segment.percent_z_location = 0.55/fuselage.lengths.total
segment.height = 0.195
segment.width = 0.130
fuselage.append_segment(segment)
vehicle.append_component(fuselage)
#------------------------------------------------------------------------------------------------------------------------------------
# ########################################################## Booms ################################################################
#------------------------------------------------------------------------------------------------------------------------------------
boom = RCAIDE.Library.Components.Booms.Boom()
boom.tag = 'boom_1r'
boom.configuration = 'boom'
boom.origin = [[ 0.036, 1.950, 1]]
boom.seats_abreast = 0.
boom.seat_pitch = 0.0
boom.fineness.nose = 0.950
boom.fineness.tail = 1.029
boom.lengths.nose = 0.2
boom.lengths.tail = 0.2
boom.lengths.cabin = 4.15
boom.lengths.total = 4.2
boom.width = 0.15
boom.heights.maximum = 0.15
boom.heights.at_quarter_length = 0.15
boom.heights.at_three_quarters_length = 0.15
boom.heights.at_wing_root_quarter_chord = 0.15
boom.areas.wetted = 0.018
boom.areas.front_projected = 0.018
boom.effective_diameter = 0.15
boom.differential_pressure = 0.
boom.symmetric = True
boom.index = 1
# Segment
segment = RCAIDE.Library.Components.Booms.Segments.Segment()
segment.tag = 'segment_1'
segment.percent_x_location = 0.
segment.percent_z_location = 0.0
segment.height = 0.05
segment.width = 0.05
boom.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Booms.Segments.Segment()
segment.tag = 'segment_2'
segment.percent_x_location = 0.03
segment.percent_z_location = 0.
segment.height = 0.15
segment.width = 0.15
boom.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Booms.Segments.Segment()
segment.tag = 'segment_3'
segment.percent_x_location = 0.97
segment.percent_z_location = 0.
segment.height = 0.15
segment.width = 0.15
boom.append_segment(segment)
# Segment
segment = RCAIDE.Library.Components.Booms.Segments.Segment()
segment.tag = 'segment_4'
segment.percent_x_location = 1.
segment.percent_z_location = 0.
segment.height = 0.05
segment.width = 0.05
boom.append_segment(segment)
# add to vehicle
vehicle.append_component(boom)
# add left long boom
boom = deepcopy(vehicle.booms.boom_1r)
boom.origin[0][1] = -boom.origin[0][1]
boom.tag = 'boom_1l'
vehicle.append_component(boom)
# add left long boom
boom = deepcopy(vehicle.booms.boom_1r)
boom.origin = [[ 0.110, 4.891, 1.050]]
boom.tag = 'boom_2r'
boom.lengths.total = 4.16
vehicle.append_component(boom)
# add inner left boom
boom = deepcopy(vehicle.booms.boom_1r)
boom.origin = [[ 0.110, - 4.891, 1.050 ]]
boom.lengths.total = 4.16
boom.tag = 'boom_2l'
vehicle.append_component(boom)
#------------------------------------------------------------------------------------------------------------------------------------
# ######################################################## Energy Network #########################################################
#------------------------------------------------------------------------------------------------------------------------------------
network = RCAIDE.Framework.Networks.Electric()
#====================================================================================================================================
# Forward Bus
#====================================================================================================================================
cruise_bus = RCAIDE.Library.Components.Powertrain.Distributors.Electrical_Bus()
cruise_bus.tag = 'cruise_bus'
cruise_bus.number_of_battery_modules = 1
#------------------------------------------------------------------------------------------------------------------------------------
# Bus Battery
#------------------------------------------------------------------------------------------------------------------------------------
battery_module = RCAIDE.Library.Components.Powertrain.Sources.Battery_Modules.Lithium_Ion_NMC()
battery_module.tag = 'cruise_bus_battery'
battery_module.electrical_configuration.series = 140
battery_module.electrical_configuration.parallel = 60
battery_module.geometric_configuration.normal_count = 140
battery_module.geometric_configuration.parallel_count = 60
for _ in range( cruise_bus.number_of_battery_modules):
cruise_bus.battery_modules.append(deepcopy(battery_module))
cruise_bus.initialize_bus_properties()
#------------------------------------------------------------------------------------------------------------------------------------
# Forward Bus Propulsors
#------------------------------------------------------------------------------------------------------------------------------------
# Define Forward Propulsor Container
cruise_propulsor_1 = RCAIDE.Library.Components.Powertrain.Propulsors.Electric_Rotor()
cruise_propulsor_1.tag = 'cruise_propulsor_1'
cruise_propulsor_1.wing_mounted = False
# Electronic Speed Controller
propeller_esc = RCAIDE.Library.Components.Powertrain.Modulators.Electronic_Speed_Controller()
propeller_esc.efficiency = 0.95
propeller_esc.bus_voltage = cruise_bus.voltage
propeller_esc.origin = [[6.583, 1.300, 1.092 ]]
propeller_esc.tag = 'propeller_esc_1'
cruise_propulsor_1.electronic_speed_controller = propeller_esc
# Propeller
g = 9.81 # gravitational acceleration
speed_of_sound = 340 # speed of sound
Hover_Load = vehicle.mass_properties.takeoff*g *1.1 # hover load
propeller = RCAIDE.Library.Components.Powertrain.Converters.Propeller()
propeller.number_of_blades = 3
propeller.tag = 'propeller_1'
propeller.origin = [[6.583, 1.300, 1.092 ]]
propeller.tip_radius = 1.15
propeller.hub_radius = 0.1 * propeller.tip_radius
propeller.cruise.design_freestream_velocity = 130.* Units['mph']
propeller.cruise.design_tip_mach = 0.65
propeller.cruise.design_angular_velocity = propeller.cruise.design_tip_mach *speed_of_sound/propeller.tip_radius
propeller.cruise.design_lift_coefficient = 0.7
propeller.cruise.design_altitude = 1500 * Units.feet
propeller.cruise.design_thrust = 3150
propeller.rotation = 1
propeller.variable_pitch = True
airfoil = RCAIDE.Library.Components.Airfoils.Airfoil()
airfoil.coordinate_file = airfoil_path + 'Airfoils' + separator + 'NACA_4412.txt'
airfoil.polar_files = [airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_50000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_100000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_200000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_500000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_1000000.txt',
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_3500000.txt',
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_5000000.txt',
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_7500000.txt' ]
propeller.append_airfoil(airfoil)
propeller.airfoil_polar_stations = [0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0]
cruise_propulsor_1.rotor = propeller
# Propeller Motor
propeller_motor = RCAIDE.Library.Components.Powertrain.Converters.DC_Motor()
propeller_motor.efficiency = 0.95
propeller_motor.tag = 'propeller_motor_1'
propeller_motor.origin = [[6.583, 1.300, 1.092 ]]
propeller_motor.nominal_voltage = cruise_bus.voltage
propeller_motor.no_load_current = 0.001
propeller_motor.wing_tag = 'horizontal_tail'
cruise_propulsor_1.motor = propeller_motor
# rear propeller nacelle
propeller_nacelle = RCAIDE.Library.Components.Nacelles.Stack_Nacelle()
propeller_nacelle.tag = 'propeller_nacelle'
propeller_nacelle.length = 1.24
propeller_nacelle.diameter = 0.4
propeller_nacelle.origin = [[5.583, 1.300 , 1.092]]
propeller_nacelle.flow_through = False
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_1'
nac_segment.percent_x_location = 0.0
nac_segment.height = 0.0
nac_segment.width = 0.0
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_2'
nac_segment.percent_x_location = 0.10/propeller_nacelle.length
nac_segment.height = 0.2
nac_segment.width = 0.2
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_2'
nac_segment.percent_x_location = 0.15 /propeller_nacelle.length
nac_segment.height = 0.25
nac_segment.width = 0.25
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_4'
nac_segment.percent_x_location = 0.2/propeller_nacelle.length
nac_segment.height = 0.3
nac_segment.width = 0.3
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_5'
nac_segment.percent_x_location = 0.25/propeller_nacelle.length
nac_segment.height = 0.35
nac_segment.width = 0.35
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_6'
nac_segment.percent_x_location = 0.5/propeller_nacelle.length
nac_segment.height = 0.4
nac_segment.width = 0.4
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_7'
nac_segment.percent_x_location = 0.75/propeller_nacelle.length
nac_segment.height = 0.35
nac_segment.width = 0.35
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_8'
nac_segment.percent_x_location = 0.98/propeller_nacelle.length
nac_segment.height = 0.3
nac_segment.width = 0.3
propeller_nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.Segment()
nac_segment.tag = 'segment_9'
nac_segment.percent_x_location = 1.0
nac_segment.height = 0.0
nac_segment.width = 0.0
propeller_nacelle.append_segment(nac_segment)
cruise_propulsor_1.nacelle = propeller_nacelle
design_electric_rotor(cruise_propulsor_1)
network.propulsors.append(cruise_propulsor_1)
# make and append copy of forward propulsor (efficient coding)
cruise_propulsor_2 = deepcopy(cruise_propulsor_1)
cruise_propulsor_2.tag = 'cruise_propulsor_2'
cruise_propulsor_2.rotor.origin = [[6.583, -1.300, 1.092 ]]
propeller_nacelle_2 = deepcopy(propeller_nacelle)
propeller_nacelle_2.tag = 'propeller_nacelle_2'
propeller_nacelle_2.origin = [[5.583, - 1.300, 1.092]]
cruise_propulsor_2.nacelle = propeller_nacelle_2
network.propulsors.append(cruise_propulsor_2)
cruise_bus.assigned_propulsors = [['cruise_propulsor_1','cruise_propulsor_2' ]]
#------------------------------------------------------------------------------------------------------------------------------------
# Additional Bus Loads
#------------------------------------------------------------------------------------------------------------------------------------
# Avionics
avionics = RCAIDE.Library.Components.Powertrain.Systems.Avionics()
avionics.power_draw = 10. # Watts
avionics.mass_properties.mass = 1.0 * Units.kg
cruise_bus.avionics = avionics
# append forward bus
network.busses.append(cruise_bus)
#====================================================================================================================================
# Lift Bus
#====================================================================================================================================
lift_bus = RCAIDE.Library.Components.Powertrain.Distributors.Electrical_Bus()
lift_bus.tag = 'lift_bus'
lift_bus.number_of_battery_modules = 1
#------------------------------------------------------------------------------------------------------------------------------------
# Bus Battery
#------------------------------------------------------------------------------------------------------------------------------------
battery_module = RCAIDE.Library.Components.Powertrain.Sources.Battery_Modules.Lithium_Ion_NMC()
battery_module.tag = 'lift_bus_battery'
battery_module.electrical_configuration.series = 140
battery_module.origin = [[4.2, 0.0, 0.0]]
battery_module.electrical_configuration.parallel = 20
battery_module.geometric_configuration.normal_count = 140
battery_module.geometric_configuration.parallel_count = 20
for _ in range( lift_bus.number_of_battery_modules):
lift_bus.battery_modules.append(deepcopy(battery_module))
lift_bus.initialize_bus_properties()
#------------------------------------------------------------------------------------------------------------------------------------
# Lift Propulsors
#------------------------------------------------------------------------------------------------------------------------------------
# Define Lift Propulsor Container
lift_propulsor_1 = RCAIDE.Library.Components.Powertrain.Propulsors.Electric_Rotor()
lift_propulsor_1.wing_mounted = True
# Electronic Speed Controller
lift_rotor_esc = RCAIDE.Library.Components.Powertrain.Modulators.Electronic_Speed_Controller()
lift_rotor_esc.efficiency = 0.95
lift_rotor_esc.origin = [[-0.073 , 1.950 , 1.2]]
lift_rotor_esc.bus_voltage = lift_bus.voltage
lift_propulsor_1.electronic_speed_controller = lift_rotor_esc
# Lift Rotor Design
lift_rotor = RCAIDE.Library.Components.Powertrain.Converters.Lift_Rotor()
lift_rotor.origin = [[-0.073 , 1.950 , 1.2]]
lift_rotor.active = True
lift_rotor.tip_radius = 2.8/2
lift_rotor.hub_radius = 0.1
lift_rotor.number_of_blades = 3
lift_rotor.hover.design_altitude = 40 * Units.feet
lift_rotor.hover.design_thrust = Hover_Load/8
lift_rotor.hover.design_freestream_velocity = np.sqrt(lift_rotor.hover.design_thrust/(2*1.2*np.pi*(lift_rotor.tip_radius**2)))
lift_rotor.oei.design_altitude = 40 * Units.feet
lift_rotor.oei.design_thrust = Hover_Load/7
lift_rotor.oei.design_freestream_velocity = np.sqrt(lift_rotor.oei.design_thrust/(2*1.2*np.pi*(lift_rotor.tip_radius**2)))
airfoil = RCAIDE.Library.Components.Airfoils.Airfoil()
airfoil.coordinate_file = airfoil_path + 'Airfoils' + separator + 'NACA_4412.txt'
airfoil.polar_files = [airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_50000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_100000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_200000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_500000.txt' ,
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_1000000.txt',
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_3500000.txt',
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_5000000.txt',
airfoil_path + 'Airfoils' + separator + 'Polars' + separator + 'NACA_4412_polar_Re_7500000.txt' ]
lift_rotor.append_airfoil(airfoil)
lift_rotor.airfoil_polar_stations = [0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0]
lift_propulsor_1.rotor = lift_rotor
#------------------------------------------------------------------------------------------------------------------------------------
# Lift Rotor Motor
#------------------------------------------------------------------------------------------------------------------------------------
lift_rotor_motor = RCAIDE.Library.Components.Powertrain.Converters.DC_Motor()
lift_rotor_motor.efficiency = 0.9
lift_rotor_motor.nominal_voltage = lift_bus.voltage*3/4
lift_rotor_motor.propeller_radius = lift_rotor.tip_radius
lift_rotor_motor.no_load_current = 0.01
lift_propulsor_1.motor = lift_rotor_motor
#------------------------------------------------------------------------------------------------------------------------------------
# Lift Rotor Nacelle
#------------------------------------------------------------------------------------------------------------------------------------
nacelle = RCAIDE.Library.Components.Nacelles.Nacelle()
nacelle.length = 0.45
nacelle.diameter = 0.3
nacelle.orientation_euler_angles = [0,-90*Units.degrees,0.]
nacelle.flow_through = False
nacelle.origin = [[ -0.073, 1.950, 1.2]]
lift_propulsor_1.nacelle = nacelle
design_electric_rotor(lift_propulsor_1)
network.propulsors.append(lift_propulsor_1)
# Front Rotors Locations
origins = [[ -0.073, 1.950, 1.2], [-0.073 , -1.950 , 1.2],[ 4.440 , 1.950 , 1.2], [ 4.440 , -1.950 , 1.2],
[ 0.219 , 4.891 , 1.2], [ 0.219 , - 4.891 , 1.2], [ 4.196 , 4.891 , 1.2], [ 4.196 , - 4.891 , 1.2]]
orientation_euler_angles = [[10.0*Units.degrees,np.pi/2.,0.],[-10.0* Units.degrees,np.pi/2.,0.], [10.0* Units.degrees,np.pi/2.,0.], [-10.0* Units.degrees,np.pi/2.,0.],
[10.0* Units.degrees,np.pi/2.,0.], [-10.0* Units.degrees,np.pi/2.,0.], [10.0* Units.degrees,np.pi/2.,0.], [-10.0* Units.degrees,np.pi/2.,0.]] # vector of angles defining default orientation of rotor
assigned_propulsor_list = []
for i in range(len(origins)):
propulsor_i = deepcopy(lift_propulsor_1)
propulsor_i.tag = 'lift_propulsor_' + str(i + 1)
propulsor_i.rotor.tag = 'lift_rotor_' + str(i + 1)
propulsor_i.rotor.origin = [origins[i]]
propulsor_i.rotor.orientation_euler_angle = orientation_euler_angles[i]
propulsor_i.motor.tag = 'lift_rotor_motor_' + str(i + 1)
propulsor_i.motor.origin = [origins[i]]
propulsor_i.electronic_speed_controller.tag = 'lift_rotor_esc_' + str(i + 1)
propulsor_i.electronic_speed_controller.origin = [origins[i]]
propulsor_i.nacelle.tag = 'lift_rotor_nacelle_' + str(i + 1)
propulsor_i.nacelle.origin = [origins[i]]
network.propulsors.append(propulsor_i)
assigned_propulsor_list.append(propulsor_i.tag)
lift_bus.assigned_propulsors = [assigned_propulsor_list]
#------------------------------------------------------------------------------------------------------------------------------------
# Additional Bus Loads
#------------------------------------------------------------------------------------------------------------------------------------ = payload
# Avionics
avionics = RCAIDE.Library.Components.Powertrain.Systems.Avionics()
avionics.power_draw = 10. # Watts
avionics.mass_properties.mass = 1.0 * Units.kg
lift_bus.avionics = avionics
network.busses.append(lift_bus)
# append energy network
vehicle.append_energy_network(network)
return vehicle
Configurations Setup#
The ``configs_setup`` function defines the different vehicle configurations (referred to as configs) used during the simulation. Configurations allow for modifications to the baseline vehicle, such as altering control surface settings, without redefining the entire vehicle.
1. Base Configuration#
The base configuration serves as the foundation for all other configurations. It is defined to match the baseline vehicle created in the vehicle_setup function. Configurations in RCAIDE are created as containers using RCAIDE Data classes. These classes provide additional functionality, such as the ability to append new configurations or modifications.
2. Remaining Configurations#
The remaining configurations, such as forward, transition, and vertical, follow a similar pattern:
Forward
Transition
Vertical
Each configuration is built upon the previous one or the base configuration, ensuring modularity and easy customization.
[3]:
def configs_setup(vehicle):
configs = RCAIDE.Library.Components.Configs.Config.Container()
base_config = RCAIDE.Library.Components.Configs.Config(vehicle)
base_config.tag = 'base'
configs.append(base_config)
forward_config = RCAIDE.Library.Components.Configs.Config(vehicle)
forward_config.tag = 'forward_flight'
forward_config.networks.electric.busses['lift_bus'].active = False
configs.append(forward_config)
transition_config = RCAIDE.Library.Components.Configs.Config(vehicle)
transition_config.tag = 'transition_flight'
configs.append(transition_config)
vertical_config = RCAIDE.Library.Components.Configs.Config(vehicle)
vertical_config.tag = 'vertical_flight'
vertical_config.networks.electric.busses['cruise_bus'].active = False
configs.append(vertical_config)
return configs
Base Analysis#
The ``base_analysis`` function defines the analyses required for evaluating the aircraft. Each analysis addresses a specific aspect of the vehicle’s performance or characteristics. Below are the key analyses, their purpose, and considerations for their use.
1. Weights Analysis#
The weights analysis calculates the distribution of the aircraft’s weight across various components. This method is based on empirical correlations designed for tube-and-wing transport aircraft configurations.
Provides a breakdown of component weights (e.g., wings, fuselage, engines).
While informative, the results of this analysis are not directly used in the performance evaluation.
2. Aerodynamics Analysis#
The aerodynamics analysis evaluates the aerodynamic performance of the aircraft. It uses RCAIDE’s fidelity zero method:
Fidelity Zero: This is RCAIDE’s baseline aerodynamic analysis method, suitable for subsonic transport aircraft.
Similar to aerodynamic methods found in conceptual design texts.
Provides estimates for lift, drag, and other aerodynamic coefficients.
Note: Higher-fidelity aerodynamic methods are available for more detailed analyses if needed.
3. Stability Analysis#
The stability analysis calculates stability derivatives for the aircraft. While it is not used in the current mission setup, it can be run post-mission for checks or additional analysis.
Like the aerodynamic method, it uses fidelity zero for baseline stability analysis.
Applicable for basic stability checks of subsonic transport aircraft.
4. Energy Analysis#
The energy analysis runs the energy network attached to the vehicle. For this turboprop-powered aircraft:
The analysis evaluates the turboprop energy network.
Ensures the propulsion system behavior, such as thrust and fuel consumption, is accounted for.
5. Planet Analysis#
The planet analysis defines the planetary environment the vehicle operates in. This setup allows for the attachment of an atmospheric model.
6. Atmosphere Analysis#
The atmosphere analysis sets the atmospheric conditions for the simulation. A common choice is the US 1976 Standard Atmosphere, which provides:
Standard temperature, pressure, and density profiles with altitude.
Consistent atmospheric conditions for performance evaluations.
[4]:
def base_analysis(vehicle):
# ------------------------------------------------------------------
# Initialize the Analyses
# ------------------------------------------------------------------
analyses = RCAIDE.Framework.Analyses.Vehicle()
analyses.vehicle = vehicle
# Geometry
geometry = RCAIDE.Framework.Analyses.Geometry.Geometry()
geometry.settings.overwrite_reference = True
geometry.settings.update_wing_properties = True
analyses.append(geometry)
# ------------------------------------------------------------------
# Weights
weights = RCAIDE.Framework.Analyses.Weights.Electric_VTOL()
weights.aircraft_type = "VTOL"
analyses.append(weights)
# ------------------------------------------------------------------
# Aerodynamics Analysis
aerodynamics = RCAIDE.Framework.Analyses.Aerodynamics.Vortex_Lattice_Method()
aerodynamics.training.Mach = np.array([0.1 ,0.3, 0.5, 0.65 , 0.85, 1.3])
analyses.append(aerodynamics)
# ------------------------------------------------------------------
# Stability Analysis
stability = RCAIDE.Framework.Analyses.Stability.Vortex_Lattice_Method()
analyses.append(stability)
# ------------------------------------------------------------------
# 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
Analyses Setup#
The ``analyses_setup`` function assigns a set of analyses to each vehicle configuration. Analyses are used to evaluate the aircraft’s performance, aerodynamics, energy systems, and other characteristics for a given configuration.
1. Overview of Analyses Assignment#
In this tutorial, all configurations share the same set of analyses. However, this function provides the flexibility to assign a unique set of analyses to any specific configuration.
2. Purpose of Analyses Assignment#
The analyses ensure that the defined vehicle configurations (e.g., cruise, takeoff, landing) are evaluated correctly during the simulation. Each configuration can have:
Common Analyses: Shared across multiple configurations for simplicity.
Custom Analyses: Tailored to a specific phase of flight or performance evaluation.
3. Typical Analyses Included#
The following analyses are typically assigned to each configuration:
Weights Analysis: Computes weight distribution across components.
Aerodynamics Analysis: Estimates lift, drag, and aerodynamic coefficients.
Stability Analysis: Evaluates stability derivatives for flight control assessments.
Energy Analysis: Runs the energy network (e.g., turboprop engine) for thrust and fuel performance.
Atmosphere Analysis: Sets atmospheric conditions using standard atmospheric models.
By assigning these analyses, the vehicle’s behavior under different configurations (e.g., cruise, takeoff, landing) can be comprehensively evaluated.
4. Customizing Analyses#
To assign a custom analysis set for a specific configuration:
Define a new analysis function tailored to the desired evaluation.
Replace the default analyses for the target configuration by calling the custom function.
For example, the takeoff configuration might use a modified aerodynamic analysis to account for flap and slat deployment.
[5]:
def analyses_setup(configs):
analyses = RCAIDE.Framework.Analyses.Analysis.Container()
# build a base analysis for each config
for tag,config in configs.items():
analysis = base_analysis(config)
analyses[tag] = analysis
return analyses
Mission Setup#
The ``mission_setup`` function defines the mission profile used to compute the aircraft’s performance. A mission profile consists of sequential segments that represent different phases of flight, such as climb, cruise, and descent.
1. Mission Profile Overview#
A mission profile is made up of individual flight segments. Each segment specifies the aircraft’s flight conditions, such as:
Altitude
Speed
Range
Time
These segments are simulated sequentially, allowing for a detailed performance analysis of the vehicle across all phases of flight.
2. Segments in the Mission Profile#
Common segments in a mission profile include:
Taxi: Ground movement of the aircraft before takeoff and after landing.
Takeoff: Acceleration and lift-off phase with high-lift devices deployed.
Climb: Gradual ascent to cruise altitude, often with reduced flap/slat deployment.
Cruise: Level flight at a constant altitude and speed for fuel-efficient operation.
Descent: Controlled reduction in altitude as the aircraft prepares for landing.
Landing: Final phase of flight with maximum flap and slat deployment for touchdown.
Each segment defines specific performance conditions and parameters, such as speed, altitude, and duration.
For more information on the mission solver and its implementation, refer to the relevant RCAIDE documentation.
[6]:
# ------------------------------------------------------------------
# Baseline Mission Setup
# ------------------------------------------------------------------
def mission_setup(analyses):
# ------------------------------------------------------------------
# Initialize the Mission
# ------------------------------------------------------------------
mission = RCAIDE.Framework.Mission.Sequential_Segments()
mission.tag = 'baseline_mission'
# unpack Segments module
Segments = RCAIDE.Framework.Mission.Segments
# base segment
base_segment = Segments.Segment()
Vstall = 80 * Units.kts
#------------------------------------------------------------------------------------------------------------------------------------
# Vertical Climb
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Vertical_Flight.Climb(base_segment)
segment.tag = "Vertical_Climb"
segment.analyses.extend( analyses.vertical_flight )
segment.altitude_start = 0.0 * Units.ft
segment.altitude_end = 50. * Units.ft
segment.initial_battery_state_of_charge = 1.0
segment.climb_rate = 500. * Units['ft/min']
# define flight dynamics to model
segment.flight_dynamics.force_z = True
# define flight controls
segment.assigned_control_variables.throttle.active = True
segment.assigned_control_variables.throttle.assigned_propulsors = [['lift_propulsor_1','lift_propulsor_2','lift_propulsor_3','lift_propulsor_4',
'lift_propulsor_5','lift_propulsor_6','lift_propulsor_7','lift_propulsor_8']]
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# High-Speed Climbing Transition
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Climb.Constant_Acceleration_Constant_Pitchrate_Constant_Angle(base_segment)
segment.tag = "High_Speed_Climbing_Transition"
segment.analyses.extend( analyses.transition_flight)
segment.true_course = 0 * Units.degree
segment.altitude_start = 50.0 * Units.ft
segment.altitude_end = 500.0 * Units.ft
segment.climb_angle = 3 * Units.degrees
segment.acceleration = 0.25 * Units['m/s/s']
segment.pitch_initial = 2. * Units.degrees
segment.pitch_final = 7. * Units.degrees
# 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 = [['cruise_propulsor_1','cruise_propulsor_2'],
['lift_propulsor_1','lift_propulsor_2','lift_propulsor_3','lift_propulsor_4',
'lift_propulsor_5','lift_propulsor_6','lift_propulsor_7','lift_propulsor_8']]
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# Second Climb
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Climb.Linear_Speed_Constant_Rate(base_segment)
segment.tag = "Climb"
segment.analyses.extend( analyses.forward_flight)
segment.altitude_start = 500.0 * Units.ft
segment.altitude_end = 1000. * Units.ft
segment.climb_rate = 300. * Units['ft/min']
segment.air_speed_end = 110 * Units['mph']
# 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 = [['cruise_propulsor_1','cruise_propulsor_2']]
segment.assigned_control_variables.pitch_angle.active = True
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# Cruise
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Cruise.Constant_Speed_Constant_Altitude(base_segment)
segment.tag = "Cruise"
segment.analyses.extend( analyses.forward_flight )
segment.altitude = 1000.0 * Units.ft
segment.air_speed = 110. * Units['mph']
segment.distance = 40 *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 = [['cruise_propulsor_1','cruise_propulsor_2']]
segment.assigned_control_variables.pitch_angle.active = True
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# Descent
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Climb.Linear_Speed_Constant_Rate(base_segment)
segment.tag = "Descent"
segment.analyses.extend(analyses.forward_flight)
segment.altitude_start = 1000.0 * Units.ft
segment.altitude_end = 500. * Units.ft
segment.climb_rate = -200. * Units['ft/min']
segment.air_speed_start = 90 * Units.kts
segment.air_speed_end = Vstall
# 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 = [['cruise_propulsor_1','cruise_propulsor_2']]
segment.assigned_control_variables.pitch_angle.active = True
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# High-Speed Descending Transition
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Climb.Constant_Acceleration_Constant_Pitchrate_Constant_Angle(base_segment)
segment.tag = "Descending_Transition"
segment.analyses.extend( analyses.transition_flight )
segment.altitude_start = 500.0 * Units.ft
segment.altitude_end = 50.0 * Units.ft
segment.climb_angle = 3. * Units.degrees
segment.acceleration = -0.5
segment.pitch_initial = 2 * Units.degrees
segment.pitch_final = 5. * Units.degrees
# 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 = [['cruise_propulsor_1','cruise_propulsor_2'],
['lift_propulsor_1','lift_propulsor_2','lift_propulsor_3','lift_propulsor_4',
'lift_propulsor_5','lift_propulsor_6','lift_propulsor_7','lift_propulsor_8'] ]
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# Low Speed Approach Transition
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Cruise.Constant_Acceleration_Constant_Pitchrate_Constant_Altitude(base_segment)
segment.tag = "Low_Speed_Approach_Transition"
segment.analyses.extend( analyses.transition_flight )
segment.altitude = 50. * Units.ft
segment.air_speed_start = 36.5 * Units['mph']
segment.air_speed_end = 50. * Units['ft/min']
segment.acceleration = -0.25 * Units['m/s/s']
segment.pitch_initial = 5. * Units.degrees
segment.pitch_final = 2. * Units.degrees
# 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 = [['cruise_propulsor_1','cruise_propulsor_2'],
['lift_propulsor_1','lift_propulsor_2','lift_propulsor_3','lift_propulsor_4',
'lift_propulsor_5','lift_propulsor_6','lift_propulsor_7','lift_propulsor_8']]
mission.append_segment(segment)
#------------------------------------------------------------------------------------------------------------------------------------
# Vertical Descent
#------------------------------------------------------------------------------------------------------------------------------------
segment = Segments.Vertical_Flight.Descent(base_segment)
segment.tag = "Vertical_Descent"
segment.analyses.extend( analyses.vertical_flight)
segment.altitude_start = 50.0 * Units.ft
segment.altitude_end = 0. * Units.ft
segment.descent_rate = 300. * Units['ft/min']
# define flight dynamics to model
segment.flight_dynamics.force_z = True
# define flight controls
segment.assigned_control_variables.throttle.active = True
segment.assigned_control_variables.throttle.assigned_propulsors = [['lift_propulsor_1','lift_propulsor_2','lift_propulsor_3','lift_propulsor_4',
'lift_propulsor_5','lift_propulsor_6','lift_propulsor_7','lift_propulsor_8']]
mission.append_segment(segment)
return mission
Missions Setup#
The missions_setup function is responsible for setting up a list of missions. This allows multiple missions to be incorporated if desired, but only one is used here.
Initialize Missions Object: It creates an empty
Missionsobject from theRCAIDE.Framework.Missionmodule.Tag the Mission: It assigns the tag
'base_mission'to the providedmissionobject. This tag is used to identify the mission.Add Mission to List: It adds the tagged
missionto theMissionsobject.Return Missions Object: Finally, it returns the
Missionsobject, which now contains the tagged mission.
[7]:
def missions_setup(mission):
missions = RCAIDE.Framework.Mission.Missions()
# base mission
mission.tag = 'base_mission'
missions.append(mission)
return missions
Plot Mission#
The last function in this file is used to plot the performance results from the mission evaluation. The results shown are not an exhaustive list of RCAIDE outputs, and custom plotting routines can be created.
[8]:
def plot_results(results):
# Plots fligh conditions
plot_flight_conditions(results)
# Plot arcraft trajectory
plot_flight_trajectory(results)
# Plot Aerodynamic Coefficients
plot_aerodynamic_coefficients(results)
# Plot Aircraft Stability
plot_longitudinal_stability(results)
# Plot Aircraft Electronics
plot_battery_temperature(results)
plot_battery_cell_conditions(results)
plot_battery_degradation(results)
plot_electric_propulsor_efficiencies(results)
# Plot Propeller Conditions
plot_rotor_conditions(results)
# Plot Battery Degradation
plot_battery_degradation(results)
return
Save Aircraft Geometry#
``save_aircraft_geometry`` is a function that saves an aircraft geometry to a pickle (.pkl) file. This is useful for saving the geometry to a file for later use.
[9]:
def save_aircraft_geometry(geometry,filename):
pickle_file = filename + '.pkl'
with open(pickle_file, 'wb') as file:
pickle.dump(geometry, file)
return
Load Aircraft Geometry#
``load_aircraft_geometry`` is a function that loads an aircraft geometry from a pickle (.pkl)file. This geometry may have been created earlier or in a separate file.
[10]:
def load_aircraft_geometry(filename):
load_file = filename + '.pkl'
with open(load_file, 'rb') as file:
results = pickle.load(file)
return results
Load Rotor#
``load_rotor`` is a function that loads a rotor from a file. This rotor may have been created earlier or in a separate file. Documentation for the load function can be found in the RCAIDE documentation.
[11]:
def load_rotor(filename):
rotor = load(filename)
return rotor
Save Rotor#
``save_rotor`` is a function that saves a rotor to a file. This is useful for saving the rotor to a file for later use.
[12]:
def save_rotor(rotor, filename):
save(rotor, filename)
return
Main Script#
The main script is used to call each of the functions defined above to execute the mission. A main script is used to run the functions for increased readability and maintainability.
[13]:
# vehicle data
vehicle = vehicle_setup()
# Set up configs
configs = configs_setup(vehicle)
# vehicle analyses
analyses = analyses_setup(configs)
# mission analyses
mission = mission_setup(analyses)
missions = missions_setup(mission)
results = missions.base_mission.evaluate()
# plot the results
plot_results(results)
# plot vehicle
plot_3d_vehicle(vehicle,
show_figure = False
)
Solving Vertical_Climb: | | 0/8 segs [00:00<?, ?seg/s]
Lift-rotor Optimization Successful
Simulation Time: 0.18 mins
Optimum motor design failed. Using slack constraints
Warning: Max Fuel or Max Zero Fuel not defined. Iterating to find these values.
Performing Weights Analysis
--------------------------------------------------------
Propulsion Architecture: Electric
Aircraft Type : VTOL
Method : Physics_Based
=== WEIGHT BREAKDOWN REPORT ===
Structural Components:
Component Weight (kg)
----------------------------------------
Wings 142.26
Fuselage 142.31
Empennage 0.00
Landing Gear 45.86
Nacelle 0.00
Booms 155.19
Paint 0.00
----------------------------------------
Total 485.62
Propulsion Components:
Component Weight (kg)
----------------------------------------
Engines 69.40
Thrust Reversers 0.00
Miscellaneous 35.20
Fuel System 0.00
Fuel Tanks 0.00
Electrical Cabling 45.44
Thermal Management 0.00
Battery 839.73
Motors 592.05
----------------------------------------
Total 1581.82
Systems:
Component Weight (kg)
----------------------------------------
Control Systems 0.00
Apu 0.00
Electrical 0.00
Avionics 16.50
Hydraulics 0.00
Furnishings 0.00
Air Conditioner 0.00
Instruments 0.00
----------------------------------------
Total 16.50
Payload Breakdown:
Component Weight (kg)
----------------------------------------
Passengers 0.00
Baggage 0.00
Cargo 0.00
----------------------------------------
Total 0.00
Operational Items Breakdown:
Component Weight (kg)
----------------------------------------
Misc 0.00
Flight Crew 0.00
Flight Attendants 0.00
Passenger Service 0.00
----------------------------------------
Total 0.00
Overall Summary:
Metric Weight (kg)
----------------------------------------
Operating Empty Weight 2083.94
Payload Weight 0.00
Fuel Weight 0.00
Takeoff Weight 2083.94
Zero Fuel Weight 2083.94
Max Takeoff Weight 2700.00
===============================
Creating aerodynamic surrogate ...
Computing neutral point ...
+----------------------------------------------------+
| MISSION SOLVER INITIATED |
+----------------------------------------------------+
Solving Vertical_Descent: |██████████| 8/8 segs [06:32<00:00, 49.05s/seg]
2026-07-17 23:57:04.335 ( 452.370s) [ 7F2EE89CEB80]vtkXOpenGLRenderWindow.:1460 WARN| bad X server connection. DISPLAY=
2026-07-17 23:57:04.335 ( 452.370s) [ 7F2EE89CEB80]vtkOpenGLRenderWindow.c:645 WARN| Failed to load EGL! Please install the EGL library from your distribution's package manager.
2026-07-17 23:57:04.336 ( 452.371s) [ 7F2EE89CEB80]vtkOSOpenGLRenderWindow:149 WARN| libOSMesa not found. It appears that OSMesa is not installed in your system. Please install the OSMesa library from your distribution's package manager.
2026-07-17 23:57:04.337 ( 452.372s) [ 7F2EE89CEB80]vtkOpenGLRenderWindow.c:645 WARN| Failed to load EGL! Please install the EGL library from your distribution's package manager.
2026-07-17 23:57:04.337 ( 452.372s) [ 7F2EE89CEB80]vtkOSOpenGLRenderWindow:149 WARN| libOSMesa not found. It appears that OSMesa is not installed in your system. Please install the OSMesa library from your distribution's package manager.
[13]:
<pyvista.plotting.plotter.Plotter at 0x7f2e7d212850>
findfont: Failed to find font weight heavy, now using 700.
Segment did not converge. Segment Tag: Descending_Transition
Error Message:
Singular matrix C in LSQ subproblem
Segment did not converge. Segment Tag: Low_Speed_Approach_Transition
Error Message:
Singular matrix C in LSQ subproblem
Segment did not converge. Segment Tag: Vertical_Descent
Error Message:
Singular matrix C in LSQ subproblem
