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TecnamElectric.py
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# ----------------------------------------------------------------------
# Imports
# ----------------------------------------------------------------------
# Python Imports
import numpy as np
import pylab as plt
from subprocess import call
# SUAVE Imports
import SUAVE
from SUAVE.Core import Data, Units
from SUAVE.Components.Energy.Networks.Battery_Propeller import Battery_Propeller
from SUAVE.Methods.Propulsion import propeller_design
from SUAVE.Input_Output.Results import print_parasite_drag, \
print_compress_drag, \
print_engine_data, \
print_mission_breakdown, \
print_weight_breakdown
from SUAVE.Methods.Power.Battery.Sizing import initialize_from_energy_and_power, initialize_from_mass
from SUAVE.Input_Output.OpenVSP import vsp_write
from SUAVE.Input_Output.OpenVSP import vspaero
# ----------------------------------------------------------------------
# Main
# ----------------------------------------------------------------------
def main():
# build the vehicle, configs, and analyses
configs, analyses = full_setup()
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# plot results
plot_mission(results)
return
# ----------------------------------------------------------------------
# Analysis Setup
# ----------------------------------------------------------------------
def full_setup():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
# mission analyses
mission = mission_setup(configs_analyses,vehicle)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
return configs, analyses
# ----------------------------------------------------------------------
# Define the Vehicle Analyses
# ----------------------------------------------------------------------
def analyses_setup(configs):
analyses = SUAVE.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
def base_analysis(vehicle):
# ------------------------------------------------------------------
# Initialize the Analyses
# ------------------------------------------------------------------
analyses = SUAVE.Analyses.Vehicle()
# ------------------------------------------------------------------
# Basic Geometry Relations
sizing = SUAVE.Analyses.Sizing.Sizing()
sizing.features.vehicle = vehicle
analyses.append(sizing)
# ------------------------------------------------------------------
# Weights
weights = SUAVE.Analyses.Weights.Weights_UAV()
weights.settings.empty_weight_method = \
SUAVE.Methods.Weights.Correlations.Human_Powered.empty
weights.vehicle = vehicle
analyses.append(weights)
# ------------------------------------------------------------------
# Aerodynamics Analysis
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.geometry = vehicle
aerodynamics.settings.drag_coefficient_increment = 0.0000
analyses.append(aerodynamics)
# ------------------------------------------------------------------
# Energy
energy = SUAVE.Analyses.Energy.Energy()
energy.network = vehicle.propulsors #what is called throughout the mission (at every time step))
analyses.append(energy)
# ------------------------------------------------------------------
# Planet Analysis
planet = SUAVE.Analyses.Planets.Planet()
analyses.append(planet)
# ------------------------------------------------------------------
# Atmosphere Analysis
atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
atmosphere.features.planet = planet.features
analyses.append(atmosphere)
# done!
return analyses
# ----------------------------------------------------------------------
# Define the Vehicle
# ----------------------------------------------------------------------
def vehicle_setup():
# ------------------------------------------------------------------
# Initialize the Vehicle
# ------------------------------------------------------------------
vehicle = SUAVE.Vehicle()
vehicle.tag = 'Tecnam_P2006TElectric'
# ------------------------------------------------------------------
# Vehicle-level Properties
# ------------------------------------------------------------------
# mass properties
vehicle.mass_properties.max_takeoff = 1230 * Units.kilogram
vehicle.mass_properties.takeoff = 1230 * Units.kilogram
vehicle.mass_properties.operating_empty = 819 * Units.kilogram
vehicle.mass_properties.takeoff = 1230 * Units.kilogram
vehicle.mass_properties.max_zero_fuel = 1145 * Units.kilogram
vehicle.mass_properties.cargo = 80 * Units.kilogram
# envelope properties
vehicle.envelope.ultimate_load = 5.7
vehicle.envelope.limit_load = 3.8
# basic parameters
vehicle.reference_area = 73 * Units['meters**2']
vehicle.passengers = 4
vehicle.systems.control = "fully powered"
vehicle.systems.accessories = "medium range"
# ------------------------------------------------------------------
# Landing Gear
# ------------------------------------------------------------------
# used for noise calculations
landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
landing_gear.tag = "main_landing_gear"
landing_gear.main_tire_diameter = 0.423 * Units.m
landing_gear.nose_tire_diameter = 0.3625 * Units.m
landing_gear.main_strut_length = 0.4833 * Units.m
landing_gear.nose_strut_length = 0.3625 * Units.m
landing_gear.main_units = 2 #number of main landing gear units
landing_gear.nose_units = 1 #number of nose landing gear
landing_gear.main_wheels = 1 #number of wheels on the main landing gear
landing_gear.nose_wheels = 1 #number of wheels on the nose landing gear
vehicle.landing_gear = landing_gear
# ------------------------------------------------------------------
# Main Wing
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Main_Wing()
wing.tag = 'main_wing'
wing.aspect_ratio = 8.80
#wing.sweeps.quarter_chord = 0 * Units.deg
wing.thickness_to_chord = 0.15
wing.taper = 0.621
wing.span_efficiency = 0.965
wing.spans.projected = 11.4 * Units.meter
wing.chords.root = 1.45 * Units.meter
wing.chords.tip = 0.90 * Units.meter
wing.chords.mean_aerodynamic = 1.34 * Units.meter
wing.areas.reference = 14.80 * Units['meters**2']
wing.twists.root = 0 * Units.degrees
wing.twists.tip = 0 * Units.degrees
wing.dihedral= 1 * Units.degrees
wing.origin = [2.986,0,1.077] # meters
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'root'
segment.percent_span_location = 0.0
segment.twist = 0. * Units.deg
segment.root_chord_percent = 1.
segment.dihedral_outboard = 1. * Units.degrees
segment.sweeps.quarter_chord = 0. * Units.degrees
segment.thickness_to_chord = 0.15
#segment airfoil
airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
segment.append_airfoil(airfoil)
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'mid'
segment.percent_span_location = 0.53
segment.twist = 0. * Units.deg
segment.root_chord_percent = 1.
segment.dihedral_outboard = 1. * Units.degrees
segment.sweeps.quarter_chord = 0. * Units.degrees
segment.thickness_to_chord = 0.15
#segment airfoil
airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
segment.append_airfoil(airfoil)
wing.Segments.append(segment)
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'tip'
segment.percent_span_location = 1.0
segment.twist = 0. * Units.deg
segment.root_chord_percent = 0.621
segment.dihedral_outboard = 1. * Units.degrees
segment.sweeps.quarter_chord = 0. * Units.degrees
segment.thickness_to_chord = 0.15
#segment airfoil
airfoil = SUAVE.Components.Wings.Airfoils.Airfoil()
airfoil.coordinate_file = '/Users/Bruno/Documents/Delft/Courses/2016-2017/Thesis/Code/Airfoils/naca642415.dat'
segment.append_airfoil(airfoil)
wing.Segments.append(segment)
# ------------------------------------------------------------------
# Flaps
# ------------------------------------------------------------------
wing.flaps.chord = 0.20
wing.flaps.span_start = 0.1053
wing.flaps.span_end = 0.6842
wing.flaps.type = 'single_slotted'
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'horizontal_stabilizer'
wing.aspect_ratio = 4.193
wing.sweeps.quarter_chord = 0.0 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 1.0
wing.span_efficiency = 0.733
wing.spans.projected = 3.3 * Units.meter
wing.chords.root = 0.787 * Units.meter
wing.chords.tip = 0.787 * Units.meter
wing.chords.mean_aerodynamic = (wing.chords.root*(2.0/3.0)*((1.0+wing.taper+wing.taper**2.0)/(1.0+wing.taper))) * Units.meter
wing.areas.reference = 2.5971 * Units['meters**2']
wing.areas.exposed = 4.0 * Units['meters**2']
wing.areas.wetted = 4.0 * Units['meters**2']
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [7.789,0.0,0.3314] # meters
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Vertical Stabilizer
# ------------------------------------------------------------------
wing = SUAVE.Components.Wings.Wing()
wing.tag = 'vertical_stabilizer'
wing.aspect_ratio = 1.407
wing.sweeps.quarter_chord = 38.75 * Units.deg
wing.thickness_to_chord = 0.12
wing.taper = 0.414
wing.span_efficiency = -0.107
wing.spans.projected = 1.574 * Units.meter
wing.chords.root = 1.2 * Units.meter
wing.chords.tip = 0.497 * Units.meter
wing.chords.mean_aerodynamic = (wing.chords.root*(2.0/3.0)*((1.0+wing.taper+wing.taper**2.0)/(1.0+wing.taper))) * Units.meter
wing.areas.reference = 1.761 * Units['meters**2']
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [7.25,0,0.497] # meters
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
wing.dynamic_pressure_ratio = 1.0
# add to vehicle
vehicle.append_component(wing)
# ------------------------------------------------------------------
# Fuselage
# ------------------------------------------------------------------
fuselage = SUAVE.Components.Fuselages.Fuselage()
fuselage.tag = 'fuselage'
#fuselage.aerodynamic_center= [2.986,0,1.077]
fuselage.number_coach_seats = vehicle.passengers
fuselage.seats_abreast = 2
fuselage.seat_pitch = 0.995 * Units.meter
fuselage.fineness.nose = 1.27
fuselage.fineness.tail = 1 #3.31
fuselage.lengths.nose = 1.16 * Units.meter
fuselage.lengths.tail = 4.637 * Units.meter
fuselage.lengths.cabin = 2.653 * Units.meter
fuselage.lengths.total = 8.45 * Units.meter
fuselage.lengths.fore_space = 0.0 * Units.meter
fuselage.lengths.aft_space = 0.0 * Units.meter
fuselage.width = 1.22 * Units.meter
fuselage.heights.maximum = 1.41 * Units.meter
fuselage.effective_diameter = 2 * Units.meter
fuselage.areas.side_projected = 7.46 * Units['meters**2']
fuselage.areas.wetted = 25.0 * Units['meters**2']
fuselage.areas.front_projected = 1.54 * Units['meters**2']
fuselage.differential_pressure = 0.0 * Units.pascal # Maximum differential pressure
fuselage.heights.at_quarter_length = 1.077 * Units.meter
fuselage.heights.at_three_quarters_length = 0.5 * Units.meter #0.621 * Units.meter
fuselage.heights.at_wing_root_quarter_chord = 1.41 * Units.meter
## OpenVSP Design
fuselage.OpenVSP_values = Data() # VSP uses degrees directly
#Nose Section
#fuselage.OpenVSP_values.nose = Data()
#fuselage.OpenVSP_values.nose.top = Data()
#fuselage.OpenVSP_values.nose.side = Data()
#fuselage.OpenVSP_values.nose.top.angle = 75.0
#fuselage.OpenVSP_values.nose.top.strength = 0.40
#fuselage.OpenVSP_values.nose.side.angle = 45.0
#fuselage.OpenVSP_values.nose.side.strength = 0.75
#fuselage.OpenVSP_values.nose.TB_Sym = True
#fuselage.OpenVSP_values.nose.z_pos = -.015
#MidFuselage1 Section
fuselage.OpenVSP_values.midfus1 = Data()
fuselage.OpenVSP_values.midfus1.z_pos=0.03
#MidFuselage2 Section
fuselage.OpenVSP_values.midfus2 = Data()
fuselage.OpenVSP_values.midfus2.z_pos=0.06
#MidFuselage3 Section
fuselage.OpenVSP_values.midfus3 = Data()
fuselage.OpenVSP_values.midfus3.z_pos=0.04
#Tail Section
fuselage.OpenVSP_values.tail = Data()
#fuselage.OpenVSP_values.tail.bottom = Data()
fuselage.OpenVSP_values.tail.z_pos = 0.039
#fuselage.OpenVSP_values.tail.bottom.angle = -20.0
#fuselage.OpenVSP_values.tail.bottom.strength = 1
## Sections for fuselage
# Section 1
#body=SUAVE.Components.Lofted_Body
#body.points=[[0,1,1],[0,-1,-1],[0,1,-1],[0,-1,1]]
#body.points.append()
#fuselage.Section.append(body)
# Section 2
#curve=SUAVE.Components.Lofted_Body.Curve
#curve.points=[[7,1,1],[7,-1,-1],[7,1,-1],[7,-1,1]]
#section=SUAVE.Components.Lofted_Body
#section.Curves=curve
#fuselage.Section.append(section)
# add to vehicle
vehicle.append_component(fuselage)
# ------------------------------------------------------------------
# Propellers Powered By Batteries
# ------------------------------------------------------------------
# build network
net = Battery_Propeller()
net.number_of_engines = 6
net.nacelle_diameter = 0.58 * Units.meters
net.engine_length = 1.74 * Units.meters
net.thrust_angle = 0.0 * Units.degrees
net.voltage = 40.0
net.areas = Data()
net.areas.wetted = 1.1*np.pi*net.nacelle_diameter*net.engine_length
# Component 3 the ESC
esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
esc.efficiency = 0.95 # Gundlach for brushless motors
net.esc = esc
# Component 4 - Propeller
# Design the Propeller
prop_attributes = Data()
#prop_attributes.tag = 'propeller_variable_pitch'
prop_attributes.number_blades = 2.0
prop_attributes.freestream_velocity = 1. * Units['m/s'] # freestream m/s
prop_attributes.angular_velocity = 2750 * Units['rpm'] # For 20x10 prop
prop_attributes.tip_radius = 0.89 * Units.meter
prop_attributes.hub_radius = 0.124 * Units.meter
prop_attributes.design_Cl = 0.8
prop_attributes.design_altitude = 0.0 * Units.meter
prop_attributes.design_thrust = 0.0
prop_attributes.design_power = 74000. * Units.watts
w=3
h=net.number_of_engines/2
Matrix = [[0 for x in range(w)] for y in range(h)]
Matrix[0][0]=2.486
Matrix[0][1]=1.5
Matrix[0][2]=0.911
Matrix[1][0]=2.486
Matrix[1][1]=3.5
Matrix[1][2]=0.911
Matrix[2][0]=2.486
Matrix[2][1]=5.5
Matrix[2][2]=0.911
prop_attributes.position = Matrix
prop_attributes = propeller_design(prop_attributes)
prop = SUAVE.Components.Energy.Converters.Propeller_variable_pitch()
prop.prop_attributes = prop_attributes
net.propeller = prop
# Component 5 the Motor
motor = SUAVE.Components.Energy.Converters.Motor()
motor.resistance = 0.008
motor.no_load_current = 4.5 * Units.ampere
motor.speed_constant = 5800. * Units['rpm'] # RPM/volt converted to (rad/s)/volt
motor.propeller_radius = prop.prop_attributes.tip_radius
motor.propeller_Cp = prop.prop_attributes.Cp
motor.gear_ratio = 12. # Gear ratio
motor.gearbox_efficiency = .98 # Gear box efficiency
motor.expected_current = 160. # Expected current
motor.mass_properties.mass = 2.0 * Units.kg
net.motor = motor
# Component 6 the Payload
payload = SUAVE.Components.Energy.Peripherals.Payload()
payload.power_draw = 50. * Units.watts
payload.mass_properties.mass = 5.0 * Units.kg
net.payload = payload
# Component 7 the Avionics
avionics = SUAVE.Components.Energy.Peripherals.Avionics()
avionics.power_draw = 50. * Units.watts
net.avionics = avionics
# Component 8 the Battery
bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
bat.mass_properties.mass = 55.0 * Units.kg
bat.specific_energy = 450. * Units.Wh/Units.kg
bat.resistance = 0.05
initialize_from_mass(bat,bat.mass_properties.mass)
net.battery = bat
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
# add the energy network to the vehicle
vehicle.append_component(net)
#print vehicle
return vehicle
# ----------------------------------------------------------------------
# Define the Configurations
# ---------------------------------------------------------------------
def configs_setup(vehicle):
# ------------------------------------------------------------------
# Initialize Configurations
# ------------------------------------------------------------------
configs = SUAVE.Components.Configs.Config.Container()
base_config = SUAVE.Components.Configs.Config(vehicle)
base_config.tag = 'base'
configs.append(base_config)
# ------------------------------------------------------------------
# Cruise Configuration
# ------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'cruise'
configs.append(config)
# ------------------------------------------------------------------
# Takeoff Configuration
# ------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'takeoff'
config.wings['main_wing'].flaps.angle = 30. * Units.deg
config.max_lift_coefficient_factor = 1.
configs.append(config)
# ------------------------------------------------------------------
# Cutback Configuration
# ------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'cutback'
config.wings['main_wing'].flaps.angle = 30. * Units.deg
config.max_lift_coefficient_factor = 1. #0.95
configs.append(config)
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'landing'
config.wings['main_wing'].flaps.angle = 40. * Units.deg
config.max_lift_coefficient_factor = 1. #0.95
configs.append(config)
# ------------------------------------------------------------------
# Short Field Takeoff Configuration
# ------------------------------------------------------------------
config = SUAVE.Components.Configs.Config(base_config)
config.tag = 'short_field_takeoff'
config.wings['main_wing'].flaps.angle = 30. * Units.deg
config.max_lift_coefficient_factor = 1. #0.95
configs.append(config)
return configs
def simple_sizing(configs):
base = configs.base
base.pull_base()
# zero fuel weight
base.mass_properties.max_zero_fuel = 0.9 * base.mass_properties.max_takeoff
# wing areas
for wing in base.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# diff the new data
base.store_diff()
# ------------------------------------------------------------------
# Landing Configuration
# ------------------------------------------------------------------
landing = configs.landing
# make sure base data is current
landing.pull_base()
# landing weight
landing.mass_properties.landing = 0.85 * base.mass_properties.takeoff
# diff the new data
landing.store_diff()
return
# ----------------------------------------------------------------------
# Define the Mission
# ----------------------------------------------------------------------
def mission_setup(analyses,vehicle):
# ------------------------------------------------------------------
# Initialize the Mission
# ------------------------------------------------------------------
mission = SUAVE.Analyses.Mission.Sequential_Segments()
mission.tag = 'The Test Mission'
mission.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
mission.planet = SUAVE.Attributes.Planets.Earth()
# unpack Segments module
Segments = SUAVE.Analyses.Mission.Segments
# base segment
base_segment = Segments.Segment()
ones_row = base_segment.state.ones_row
base_segment.process.iterate.unknowns.network = vehicle.propulsors.network.unpack_unknowns
base_segment.process.iterate.residuals.network = vehicle.propulsors.network.residuals
base_segment.state.unknowns.battery_voltage_under_load = vehicle.propulsors.network.battery.max_voltage * ones_row(1)
base_segment.process.iterate.initials.initialize_battery = SUAVE.Methods.Missions.Segments.Common.Energy.initialize_battery
base_segment.state.unknowns.propeller_power_coefficient = vehicle.propulsors.network.propeller.prop_attributes.Cp * ones_row(1)/15.
base_segment.state.residuals.network = 0. * ones_row(2)
# ------------------------------------------------------------------
# First Climb Segment: Constant Speed, Constant Rate
# ------------------------------------------------------------------
segment = Segments.Climb.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "climb_1"
segment.analyses.extend( analyses.takeoff )
segment.state.numerics.number_control_points = 64
segment.altitude_start = 0.0 * Units.meter
segment.altitude_end = 91.4402 * Units.meter
segment.air_speed = 64 * Units.knots
segment.battery_energy = vehicle.propulsors.network.battery.max_energy*0.2 #Charge the battery to start
segment.climb_rate = 1.168 * Units['m/s']
# add to misison
mission.append_segment(segment)
# ------------------------------------------------------------------
# Second Climb Segment: Constant Speed, Constant Rate
# ------------------------------------------------------------------
segment = Segments.Climb.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "climb_2"
segment.analyses.extend( analyses.cruise )
segment.state.numerics.number_control_points = 64
segment.altitude_end = 3048.0 * Units.meter
segment.air_speed = 80 * Units.knots
segment.climb_rate = 1.016 * Units['m/s']
# add to misison
mission.append_segment(segment)
# ------------------------------------------------------------------
# Cruise Segment: constant speed, constant altitude
# ------------------------------------------------------------------
segment = SUAVE.Analyses.Mission.Segments.Cruise.Constant_Speed_Constant_Altitude(base_segment)
segment.tag = "cruise1"
# connect vehicle configuration
segment.analyses.extend( analyses.cruise)
# segment attributes
segment.state.numerics.number_control_points = 64
segment.altitude = 3048.0 * Units.meter
segment.air_speed = 50. * Units.knots
segment.distance = 100.0 * Units.nautical_miles
mission.append_segment(segment)
# ------------------------------------------------------------------
# First Descent Segment: Constant Speed, Constant Rate
# ------------------------------------------------------------------
segment = Segments.Descent.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "descent_1"
segment.analyses.extend( analyses.cruise )
# segment attributes
segment.state.numerics.number_control_points = 64
segment.altitude_end = 2000.0 * Units.meter
segment.air_speed = 100.0 * Units.knots
segment.descent_rate = 1.016 * Units['m/s']
# add to mission
mission.append_segment(segment)
# ------------------------------------------------------------------
# Second Descent Segment: Constant Speed, Constant Rate
# ------------------------------------------------------------------
segment = Segments.Descent.Constant_Speed_Constant_Rate(base_segment)
segment.tag = "descent_2"
segment.analyses.extend( analyses.landing )
analyses.landing.aerodynamics.settings.spoiler_drag_increment = 0.00
# segment attributes
segment.state.numerics.number_control_points = 64
segment.altitude_end = 0.0 * Units.meter
segment.air_speed = 64 * Units.knots
segment.descent_rate = 1.168 * Units['m/s']
# add to mission
mission.append_segment(segment)
# ------------------------------------------------------------------
# Mission definition complete
# ------------------------------------------------------------------
return mission
def missions_setup(base_mission):
# the mission container
missions = SUAVE.Analyses.Mission.Mission.Container()
# ------------------------------------------------------------------
# Base Mission
# ------------------------------------------------------------------
missions.base = base_mission
return missions
# ----------------------------------------------------------------------
# Plot Mission
# ----------------------------------------------------------------------
def plot_mission(results):
# ------------------------------------------------------------------
# Throttle
# ------------------------------------------------------------------
plt.figure("Throttle History")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min
eta = results.segments[i].conditions.propulsion.throttle[:,0]
axes.plot(time, eta, 'bo-')
axes.set_xlabel('Time (mins)')
axes.set_ylabel('Throttle')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
plt.ylim((0,1))
axes.grid(True)
# ------------------------------------------------------------------
# Altitude
# ------------------------------------------------------------------
plt.figure("Altitude")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min
altitude = results.segments[i].conditions.freestream.altitude[:,0] / Units.km
axes.plot(time, altitude, 'bo-')
axes.set_xlabel('Time (mins)')
axes.set_ylabel('Altitude (km)')
axes.grid(True)
# ------------------------------------------------------------------
# Aerodynamics
# ------------------------------------------------------------------
fig = plt.figure("Aerodynamic Forces")
for segment in results.segments.values():
time = segment.conditions.frames.inertial.time[:,0] / Units.min
Lift = -segment.conditions.frames.wind.lift_force_vector[:,2]
Drag = -segment.conditions.frames.wind.drag_force_vector[:,0]
Thrust = segment.conditions.frames.body.thrust_force_vector[:,0]
axes = fig.add_subplot(3,1,1)
axes.plot( time , Lift , 'bo-' )
axes.set_xlabel('Time (min)')
axes.set_ylabel('Lift (N)')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
axes.grid(True)
axes = fig.add_subplot(3,1,2)
axes.plot( time , Drag , 'bo-' )
axes.set_xlabel('Time (min)')
axes.set_ylabel('Drag (N)')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
axes.grid(True)
axes = fig.add_subplot(3,1,3)
axes.plot( time , Thrust , 'bo-' )
axes.set_xlabel('Time (min)')
axes.set_ylabel('Thrust (N)')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
plt.ylim((0,50))
axes.grid(True)
# ------------------------------------------------------------------
# Aerodynamics 2
# ------------------------------------------------------------------
fig = plt.figure("Aerodynamic Coefficients")
for segment in results.segments.values():
time = segment.conditions.frames.inertial.time[:,0] / Units.min
CLift = segment.conditions.aerodynamics.lift_coefficient[:,0]
CDrag = segment.conditions.aerodynamics.drag_coefficient[:,0]
Drag = -segment.conditions.frames.wind.drag_force_vector[:,0]
Thrust = segment.conditions.frames.body.thrust_force_vector[:,0]
axes = fig.add_subplot(3,1,1)
axes.plot( time , CLift , 'bo-' )
axes.set_xlabel('Time (min)')
axes.set_ylabel('CL')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
axes.grid(True)
axes = fig.add_subplot(3,1,2)
axes.plot( time , CDrag , 'bo-' )
axes.set_xlabel('Time (min)')
axes.set_ylabel('CD')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
axes.grid(True)
axes = fig.add_subplot(3,1,3)
axes.plot( time , Drag , 'bo-' )
axes.plot( time , Thrust , 'ro-' )
axes.set_xlabel('Time (min)')
axes.set_ylabel('Drag and Thrust (N)')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
axes.grid(True)
# ------------------------------------------------------------------
# Aerodynamics 2
# ------------------------------------------------------------------
fig = plt.figure("Drag Components")
axes = plt.gca()
for i, segment in enumerate(results.segments.values()):
time = segment.conditions.frames.inertial.time[:,0] / Units.min
drag_breakdown = segment.conditions.aerodynamics.drag_breakdown
cdp = drag_breakdown.parasite.total[:,0]
cdi = drag_breakdown.induced.total[:,0]
cdc = drag_breakdown.compressible.total[:,0]
cdm = drag_breakdown.miscellaneous.total[:,0]
cd = drag_breakdown.total[:,0]
axes.plot( time , cdp , 'ko-', label='CD_P' )
axes.plot( time , cdi , 'bo-', label='CD_I' )
axes.plot( time , cdc , 'go-', label='CD_C' )
axes.plot( time , cdm , 'yo-', label='CD_M' )
axes.plot( time , cd , 'ro-', label='CD' )
if i == 0:
axes.legend(loc='upper center')
axes.set_xlabel('Time (min)')
axes.set_ylabel('CD')
axes.grid(True)
# ------------------------------------------------------------------
# Battery Energy
# ------------------------------------------------------------------
plt.figure("Battery Energy")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min
energy = results.segments[i].conditions.propulsion.battery_energy[:,0]
axes.plot(time, energy, 'bo-')
axes.set_xlabel('Time (mins)')
axes.set_ylabel('Battery Energy (J)')
axes.grid(True)
# ------------------------------------------------------------------
# Current Draw
# ------------------------------------------------------------------
plt.figure("Current Draw")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min
energy = results.segments[i].conditions.propulsion.current[:,0]
axes.plot(time, energy, 'bo-')
axes.set_xlabel('Time (mins)')
axes.set_ylabel('Current Draw (Amps)')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
plt.ylim((0,200))
axes.grid(True)
# ------------------------------------------------------------------
# Motor RPM
# ------------------------------------------------------------------
plt.figure("Motor RPM")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min
energy = results.segments[i].conditions.propulsion.rpm[:,0]
axes.plot(time, energy, 'bo-')
axes.set_xlabel('Time (mins)')
axes.set_ylabel('Motor RPM')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
plt.ylim((0,200))
axes.grid(True)
# ------------------------------------------------------------------
# Battery Draw
# ------------------------------------------------------------------
plt.figure("Battery Charging")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min
energy = results.segments[i].conditions.propulsion.battery_draw[:,0]
axes.plot(time, energy, 'bo-')
axes.set_xlabel('Time (mins)')
axes.set_ylabel('Battery Charging (Watts)')
axes.get_yaxis().get_major_formatter().set_scientific(False)
axes.get_yaxis().get_major_formatter().set_useOffset(False)
axes.grid(True)
# ------------------------------------------------------------------
# Propulsive efficiency
# ------------------------------------------------------------------
plt.figure("Propeller Efficiency")
axes = plt.gca()
for i in range(len(results.segments)):
time = results.segments[i].conditions.frames.inertial.time[:,0] / Units.min