Tutorial 2 - Turbojet Aircraft Simulation#
Welcome to this tutorial on simulating a turbojet 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]:
# RCAIDE imports
import RCAIDE
from RCAIDE.Framework.Core import Units , Data
from RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet import design_turbojet
from RCAIDE.Library.Methods.Geometry.Planform import wing_segmented_planform
from RCAIDE.Library.Plots import *
# python imports
import numpy as np
from copy import deepcopy
import matplotlib.pyplot as plt
import os
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: turbojet Engine#
The energy network models the propulsion system, in this case, a turbojet engine. The turbojet network determines the engine’s thrust, bypass ratio, and fuel type. These parameters are essential for performance and fuel efficiency analyses.
More detailed information about turbojet behavior and configurations is provided in the turbojet Modeling Tutorial.
[2]:
def vehicle_setup():
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 = 183000. # kg, adjusted due to significant fuel burn on runway
vehicle.mass_properties.cargo = 1000. * Units.kilogram
vehicle.mass_properties.max_zero_fuel = 92000.
# envelope properties
vehicle.flight_envelope.ultimate_load = 3.75
vehicle.flight_envelope.positive_limit_load = 2.5
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.limit_load = 3.8
vehicle.flight_envelope.design_range = 4488 * Units.nmi
# basic parameters
vehicle.reference_area = 358.25
vehicle.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
#------------------------------------------------------------------------------------------------------------------------------------
# ######################################################## 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.symmetric = True
wing.high_lift = True
wing.vortex_lift = True
wing.high_mach = True
wing.dynamic_pressure_ratio = 1.0
wing_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil()
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
wing_airfoil.coordinate_file = rel_path + 'Airfoils' + separator + 'NACA65_203.txt'
wing.append_airfoil(wing_airfoil)
# 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
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
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
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
segment.append_airfoil(wing_airfoil)
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# 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.symmetric = False
wing.t_tail = False
wing.high_mach = True
wing.dynamic_pressure_ratio = 1.0
tail_airfoil = RCAIDE.Library.Components.Airfoils.Airfoil()
tail_airfoil.coordinate_file = rel_path + 'Airfoils' + separator + 'supersonic_tail.txt'
wing.append_airfoil(tail_airfoil)
# 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
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.spans.projected
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
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
segment.append_airfoil(tail_airfoil)
wing.append_segment(segment)
# Fill out more segment properties automatically
wing = wing_segmented_planform(wing)
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = RCAIDE.Library.Components.Fuselages.Tube_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.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
fuselage.OpenVSP_values = Data() # VSP uses degrees directly
fuselage.OpenVSP_values.nose = Data()
fuselage.OpenVSP_values.nose.top = Data()
fuselage.OpenVSP_values.nose.side = Data()
fuselage.OpenVSP_values.nose.top.angle = 20.0
fuselage.OpenVSP_values.nose.top.strength = 0.75
fuselage.OpenVSP_values.nose.side.angle = 20.0
fuselage.OpenVSP_values.nose.side.strength = 0.75
fuselage.OpenVSP_values.nose.TB_Sym = True
fuselage.OpenVSP_values.nose.z_pos = -.01
fuselage.OpenVSP_values.tail = Data()
fuselage.OpenVSP_values.tail.top = Data()
fuselage.OpenVSP_values.tail.side = Data()
fuselage.OpenVSP_values.tail.bottom = Data()
fuselage.OpenVSP_values.tail.top.angle = 0.0
fuselage.OpenVSP_values.tail.top.strength = 0.0
# 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
segment.height = 0.00100
segment.width = 0.00100
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 = False # for regression
#------------------------------------------------------------------------------------------------------------------------------------
# Fuel Distrubition 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 = 0.94
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 = 0.9
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.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 = 5
nacelle.append_segment(nac_segment)
nac_segment = RCAIDE.Library.Components.Nacelles.Segments.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 = 5
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_selector_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_selector_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_selector_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_selector_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_selector_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_selector_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_selector_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_selector_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
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. Cruise Configuration#
The cruise configuration demonstrates that new configurations can inherit properties directly from existing configurations (e.g., the base config). This avoids redundancy and ensures consistency across configurations.
The cruise configuration typically reflects the clean flight condition, with no high-lift devices like flaps or slats deployed.
3. Takeoff Configuration#
The takeoff configuration is the first configuration that introduces changes to the baseline vehicle. It shows how specific vehicle parameters, such as flap and slat settings, can be modified. For example:
Flap Deflection: Flaps are deployed to increase lift during takeoff.
Slat Deployment: Slats may also be deployed to improve low-speed aerodynamic performance.
This highlights the flexibility of vehicle configurations for different phases of flight.
4. Remaining Configurations#
The remaining configurations, such as climb, approach, and landing, follow a similar pattern:
Afterburner Climb: Partial deployment of flaps/slats to optimize lift and drag during ascent.
Landing: Maximum flap and slat deflection for increased lift and drag, enabling a controlled descent and touchdown.
Each configuration is built upon the previous one or the base configuration, ensuring modularity and easy customization.
[3]:
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'
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.maximum_lift_coefficient = 2.
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'].flaps_angle = 0. * Units.deg
config.wings['main_wing'].slats_angle = 0. * Units.deg
config.maximum_lift_coefficient = 2.
configs.append(config)
# done!
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()
# ------------------------------------------------------------------
# Weights
weights = RCAIDE.Framework.Analyses.Weights.Conventional()
weights.aircraft_type = "Transport"
weights.vehicle = vehicle
analyses.append(weights)
# ------------------------------------------------------------------
# Aerodynamics Analysis
aerodynamics = RCAIDE.Framework.Analyses.Aerodynamics.Vortex_Lattice_Method()
aerodynamics.vehicle = vehicle
aerodynamics.settings.number_of_spanwise_vortices = 5
aerodynamics.settings.number_of_chordwise_vortices = 2
aerodynamics.settings.model_fuselage = True
aerodynamics.settings.drag_coefficient_increment = 0.0000
analyses.append(aerodynamics)
# ------------------------------------------------------------------
# Emissions
emissions = RCAIDE.Framework.Analyses.Emissions.Emission_Index_Correlation_Method()
emissions.vehicle = vehicle
analyses.append(emissions)
# ------------------------------------------------------------------
# Energy
energy= RCAIDE.Framework.Analyses.Energy.Energy()
energy.vehicle = vehicle
analyses.append(energy)
# ------------------------------------------------------------------
# Planet Analysis
planet = RCAIDE.Framework.Analyses.Planets.Earth()
analyses.append(planet)
# ------------------------------------------------------------------
# Atmosphere Analysis
atmosphere = RCAIDE.Framework.Analyses.Atmospheric.US_Standard_1976()
atmosphere.features.planet = planet.features
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]:
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.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "climb_1"
segment.analyses.extend( analyses.climb )
segment.altitude_start = 0.0 * Units.km
segment.altitude_end = 4000. * Units.ft
segment.air_speed = 250. * Units.kts
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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# Second Climb Segment: constant Speed, constant segment angle
# ------------------------------------------------------------------
segment = Segments.Climb.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "climb_2"
segment.analyses.extend( analyses.cruise )
segment.altitude_end = 8000. * Units.ft
segment.air_speed = 250. * Units.kts
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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
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_end = 33000. * Units.ft
segment.mach_number_start = .45
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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
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.climb )
segment.altitude_end = 34000. * Units.ft
segment.mach_number_start = 0.95
segment.mach_number_end = 1.1
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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
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.climb )
segment.altitude_end = 40000. * Units.ft
segment.mach_number_start = 1.1
segment.mach_number_end = 1.7
segment.climb_rate = 1750. * 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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
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_end = 56500. * Units.ft
segment.mach_number = 2.02
segment.climb_rate = 50. * 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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
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.mach_number = 2.02
segment.distance = 1. * Units.nmi
segment.state.numerics.number_control_points = 4
# 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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# First Descent Segment: decceleration
# ------------------------------------------------------------------
segment = Segments.Cruise.Constant_Acceleration_Constant_Altitude(base_segment)
segment.tag = "decel_1"
segment.analyses.extend( analyses.cruise )
segment.acceleration = -.5 * Units['m/s/s']
segment.air_speed_end = 1.5*573. * Units.kts
# 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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# First Descent Segment
# ------------------------------------------------------------------
segment = Segments.Descent.Linear_Mach_Constant_Rate(base_segment)
segment.tag = "descent_1"
segment.analyses.extend( analyses.cruise )
segment.altitude_end = 41000. * Units.ft
segment.mach_number_end = 1.3
segment.descent_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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# First Descent Segment: decceleration
# ------------------------------------------------------------------
segment = Segments.Cruise.Constant_Acceleration_Constant_Altitude(base_segment)
segment.tag = "decel_2"
segment.analyses.extend( analyses.cruise )
segment.acceleration = -.5 * Units['m/s/s']
segment.air_speed_end = 0.95*573. * Units.kts
# 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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# First Descent Segment
# ------------------------------------------------------------------
segment = Segments.Descent.Linear_Mach_Constant_Rate(base_segment)
segment.tag = "descent_2"
segment.analyses.extend( analyses.cruise )
segment.altitude_end = 10000. * Units.ft
segment.mach_number_end = 250./638.
segment.descent_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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# First Descent Segment
# ------------------------------------------------------------------
segment = Segments.Descent.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "descent_3"
segment.analyses.extend( analyses.cruise )
segment.altitude_end = 0. * Units.ft
segment.air_speed = 250. * Units.kts
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.throttle.initial_values = [[0.5]]
segment.assigned_control_variables.body_angle.active = True
segment.assigned_control_variables.body_angle.initial_values = [[3*Units.degrees]]
mission.append_segment(segment)
# ------------------------------------------------------------------
# Mission definition complete
# ------------------------------------------------------------------
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()
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_mission(results):
# Plot Flight Conditions
plot_flight_conditions(results)
# Plot Aerodynamic Forces
plot_aerodynamic_forces(results)
# Plot Aerodynamic Coefficients
plot_aerodynamic_coefficients(results)
# Drag Components
plot_drag_components(results)
# Plot Altitude, sfc, vehicle weight
plot_altitude_sfc_weight(results)
# Plot Velocities
plot_aircraft_velocities(results)
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.
[9]:
# ----------------------------------------------------------------------
# Main
# ----------------------------------------------------------------------
# vehicle data
vehicle = vehicle_setup()
# plot vehicle
plot_3d_vehicle(vehicle,
min_x_axis_limit = 0,
max_x_axis_limit = 60,
min_y_axis_limit = -30,
max_y_axis_limit = 30,
min_z_axis_limit = -30,
max_z_axis_limit = 30,
)
# Set up vehicle configs
configs = configs_setup(vehicle)
# create analyses
analyses = analyses_setup(configs)
# mission analyses
mission = mission_setup(analyses)
# create mission instances (for multiple types of missions)
missions = missions_setup(mission)
# mission analysis
results = missions.base_mission.evaluate()
# plot the results
plot_mission(results)
Plotting vehicle
Performing Weights Analysis
--------------------------------------------------------
Propulsion Architecture: Conventional
Aircraft Type : Transport
Method : FLOPS
Aircraft operating empty weight will be overwritten
Aircraft center of gravity location will be overwritten
Aircraft moment of intertia tensor will be overwritten
Mission Solver Initiated
0%| | 0/100 [00:00<?, ?it/s]
Solving climb_1 segment.
108it [03:01, 1.68s/it]
Solving climb_2 segment.
Solving climb_3 segment.
Solving climb_4 segment.
Solving climb_5 segment.
Solving climbing_cruise segment.
Solving level_cruise segment.
Solving decel_1 segment.
Solving descent_1 segment.
Solving decel_2 segment.
Solving descent_2 segment.
Solving descent_3 segment.
[ ]: