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Merge branch 'develop' into feature-detailed_tutorials
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# Conflicts:
#	tut_mission_B737.py
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@timdmacdo
timdmacdo committed Aug 10, 2020
2 parents 19b8600 + 4ee2257 commit bd0f478
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205 changes: 29 additions & 176 deletions B737_AVL_Tutorial/tut_mission_B737_AVL.py
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# Imports
# ----------------------------------------------------------------------

# Python Imports
import numpy as np
import pylab as plt

# SUAVE Imports
import SUAVE
from SUAVE.Core import Data, Units
from SUAVE.Plots.Mission_Plots import *
from SUAVE.Methods.Propulsion.turbofan_sizing import turbofan_sizing
from SUAVE.Methods.Geometry.Two_Dimensional.Cross_Section.Propulsion import compute_turbofan_geometry
from SUAVE.Input_Output.Results import print_parasite_drag, \
Expand All @@ -21,6 +18,10 @@
print_mission_breakdown, \
print_weight_breakdown

# Python Imports
import numpy as np
import pylab as plt

# ----------------------------------------------------------------------
# Main
# ----------------------------------------------------------------------
Expand Down Expand Up @@ -100,20 +101,22 @@ def base_analysis(vehicle):

# ------------------------------------------------------------------
# Weights
weights = SUAVE.Analyses.Weights.Weights_Tube_Wing()
weights = SUAVE.Analyses.Weights.Weights_Transport()
weights.vehicle = vehicle
analyses.append(weights)

# ------------------------------------------------------------------
# Aerodynamics Analysis
aerodynamics = SUAVE.Analyses.Aerodynamics.AVL()
aerodynamics.process.compute.lift.inviscid.settings.filenames.avl_bin_name = 'CHANGE ME TO YOUR DIRECTORY'
#aerodynamics.process.compute.lift.inviscid.settings.spanwise_vortex_density = 3
aerodynamics.geometry = vehicle
analyses.append(aerodynamics)

# ------------------------------------------------------------------
# Stability Analysis
stability = SUAVE.Analyses.Stability.AVL()
stability.settings.filenames.avl_bin_name = 'CHANGE ME TO YOUR DIRECTORY'
#stability.settings.spanwise_vortex_density = 3
stability.geometry = vehicle
analyses.append(stability)
Expand Down Expand Up @@ -200,42 +203,41 @@ def vehicle_setup():
wing.sweeps.quarter_chord = 25 * Units.deg
wing.thickness_to_chord = 0.1
wing.taper = 0.1
wing.span_efficiency = 0.9
wing.spans.projected = 34.32 * Units.meter
wing.chords.root = 7.760 * Units.meter
wing.chords.tip = 0.782 * Units.meter
wing.chords.mean_aerodynamic = 4.235 * Units.meter
wing.areas.reference = 124.862 * Units['meters**2']
wing.twists.root = 4.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [13.61,0,-1.27] # meters
wing.origin = [[13.61 * Units.meter, 0, -1.27 * Units.meter]]
wing.vertical = False
wing.symmetric = True
wing.high_lift = True
wing.dynamic_pressure_ratio = 1.0

# add to vehicle
vehicle.append_component(wing)

# ------------------------------------------------------------------
# Horizontal Stabilizer
# ------------------------------------------------------------------

wing = SUAVE.Components.Wings.Wing()
wing = SUAVE.Components.Wings.Horizontal_Tail()
wing.tag = 'horizontal_stabilizer'

wing.aspect_ratio = 6.16
wing.sweeps.quarter_chord = 40 * Units.deg
wing.thickness_to_chord = 0.08
wing.taper = 0.2
wing.span_efficiency = 0.9
wing.spans.projected = 14.2 * Units.meter
wing.chords.root = 4.7 * Units.meter
wing.chords.tip = .955 * Units.meter
wing.chords.mean_aerodynamic = 8.0 * Units.meter
wing.chords.mean_aerodynamic = 3.0 * Units.meter
wing.areas.reference = 32.488 * Units['meters**2']
wing.twists.root = 3.0 * Units.degrees
wing.twists.tip = 3.0 * Units.degrees
wing.origin = [32.83,0,1.14] # meters
wing.origin = [[32.83 * Units.meter, 0 , 1.14 * Units.meter]]
wing.vertical = False
wing.symmetric = True
wing.dynamic_pressure_ratio = 0.9
Expand All @@ -247,22 +249,21 @@ def vehicle_setup():
# Vertical Stabilizer
# ------------------------------------------------------------------

wing = SUAVE.Components.Wings.Wing()
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.tag = 'vertical_stabilizer'

wing.aspect_ratio = 1.91
wing.sweeps.quarter_chord = 25. * Units.deg
wing.thickness_to_chord = 0.08
wing.taper = 0.25
wing.span_efficiency = 0.9
wing.spans.projected = 7.777 * Units.meter
wing.chords.root = 8.19 * Units.meter
wing.chords.tip = 0.95 * Units.meter
wing.chords.mean_aerodynamic = 4.0 * Units.meter
wing.areas.reference = 27.316 * Units['meters**2']
wing.twists.root = 0.0 * Units.degrees
wing.twists.tip = 0.0 * Units.degrees
wing.origin = [28.79,0,1.54] # meters
wing.origin = [[28.79 * Units.meter, 0, 1.54 * Units.meter]] # meters
wing.vertical = True
wing.symmetric = False
wing.t_tail = False
Expand Down Expand Up @@ -317,7 +318,7 @@ def vehicle_setup():
turbofan.bypass_ratio = 5.4
turbofan.engine_length = 2.71 * Units.meter
turbofan.nacelle_diameter = 2.05 * Units.meter
turbofan.origin = [[13.72, 4.86,-1.9],[13.72, -4.86,-1.9]] # meters
turbofan.origin = [[13.72, 4.86,-1.9],[13.72, -4.86,-1.9]]

#compute engine areas
turbofan.areas.wetted = 1.1*np.pi*turbofan.nacelle_diameter*turbofan.engine_length
Expand Down Expand Up @@ -415,6 +416,7 @@ def vehicle_setup():

# setup
combustor.efficiency = 0.99
combustor.alphac = 1.0
combustor.turbine_inlet_temperature = 1450 # K
combustor.pressure_ratio = 0.95
combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
Expand Down Expand Up @@ -493,6 +495,7 @@ def vehicle_setup():

return vehicle


# ----------------------------------------------------------------------
# Define the Configurations
# ---------------------------------------------------------------------
Expand Down Expand Up @@ -717,167 +720,17 @@ def missions_setup(base_mission):

def plot_mission(results,line_style='bo-'):

axis_font = {'fontname':'Arial', 'size':'14'}

# ------------------------------------------------------------------
# Aerodynamics
# ------------------------------------------------------------------


fig = plt.figure("Aerodynamic Forces",figsize=(8,6))
for segment in results.segments.values():

time = segment.conditions.frames.inertial.time[:,0] / Units.min
Thrust = segment.conditions.frames.body.thrust_force_vector[:,0] / Units.lbf
eta = segment.conditions.propulsion.throttle[:,0]

axes = fig.add_subplot(2,1,1)
axes.plot( time , Thrust , line_style )
axes.set_ylabel('Thrust (lbf)',axis_font)
axes.grid(True)

axes = fig.add_subplot(2,1,2)
axes.plot( time , eta , line_style )
axes.set_xlabel('Time (min)',axis_font)
axes.set_ylabel('Throttle',axis_font)
axes.grid(True)

plt.savefig("B737_engine.pdf")
plt.savefig("B737_engine.png")

# ------------------------------------------------------------------
# Aerodynamics 2
# ------------------------------------------------------------------
fig = plt.figure("Aerodynamic Coefficients",figsize=(8,10))
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]
aoa = segment.conditions.aerodynamics.angle_of_attack[:,0] / Units.deg
l_d = CLift/CDrag

axes = fig.add_subplot(3,1,1)
axes.plot( time , CLift , line_style )
axes.set_ylabel('Lift Coefficient',axis_font)
axes.grid(True)

axes = fig.add_subplot(3,1,2)
axes.plot( time , l_d , line_style )
axes.set_ylabel('L/D',axis_font)
axes.grid(True)

axes = fig.add_subplot(3,1,3)
axes.plot( time , aoa , 'ro-' )
axes.set_xlabel('Time (min)',axis_font)
axes.set_ylabel('AOA (deg)',axis_font)
axes.grid(True)

plt.savefig("B737_aero.pdf")
plt.savefig("B737_aero.png")

# ------------------------------------------------------------------
# Aerodynamics 2
# ------------------------------------------------------------------
fig = plt.figure("Drag Components",figsize=(8,10))
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]

if line_style == 'bo-':
axes.plot( time , cdp , 'ko-', label='CD parasite' )
axes.plot( time , cdi , 'bo-', label='CD induced' )
axes.plot( time , cdc , 'go-', label='CD compressibility' )
axes.plot( time , cdm , 'yo-', label='CD miscellaneous' )
axes.plot( time , cd , 'ro-', label='CD total' )
if i == 0:
axes.legend(loc='upper center')
else:
axes.plot( time , cdp , line_style )
axes.plot( time , cdi , line_style )
axes.plot( time , cdc , line_style )
axes.plot( time , cdm , line_style )
axes.plot( time , cd , line_style )

axes.set_xlabel('Time (min)')
axes.set_ylabel('CD')
axes.grid(True)
plt.savefig("B737_drag.pdf")
plt.savefig("B737_drag.png")

# ------------------------------------------------------------------
# Altitude, sfc, vehicle weight
# ------------------------------------------------------------------

fig = plt.figure("Altitude_sfc_weight",figsize=(8,10))
for segment in results.segments.values():

time = segment.conditions.frames.inertial.time[:,0] / Units.min
aoa = segment.conditions.aerodynamics.angle_of_attack[:,0] / Units.deg
mass = segment.conditions.weights.total_mass[:,0] / Units.lb
altitude = segment.conditions.freestream.altitude[:,0] / Units.ft
mdot = segment.conditions.weights.vehicle_mass_rate[:,0]
thrust = segment.conditions.frames.body.thrust_force_vector[:,0]
sfc = (mdot / Units.lb) / (thrust /Units.lbf) * Units.hr

axes = fig.add_subplot(3,1,1)
axes.plot( time , altitude , line_style )
axes.set_ylabel('Altitude (ft)',axis_font)
axes.grid(True)

axes = fig.add_subplot(3,1,3)
axes.plot( time , sfc , line_style )
axes.set_xlabel('Time (min)',axis_font)
axes.set_ylabel('sfc (lb/lbf-hr)',axis_font)
axes.grid(True)

axes = fig.add_subplot(3,1,2)
axes.plot( time , mass , 'ro-' )
axes.set_ylabel('Weight (lb)',axis_font)
axes.grid(True)

plt.savefig("B737_mission.pdf")
plt.savefig("B737_mission.png")

# ------------------------------------------------------------------
# Velocities
# ------------------------------------------------------------------
fig = plt.figure("Velocities",figsize=(8,10))
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] / Units.lbf
Thrust = segment.conditions.frames.body.thrust_force_vector[:,0] / Units.lb
velocity = segment.conditions.freestream.velocity[:,0]
pressure = segment.conditions.freestream.pressure[:,0]
density = segment.conditions.freestream.density[:,0]
EAS = velocity * np.sqrt(density/1.225)
mach = segment.conditions.freestream.mach_number[:,0]

axes = fig.add_subplot(3,1,1)
axes.plot( time , velocity / Units.kts, line_style )
axes.set_ylabel('velocity (kts)',axis_font)
axes.grid(True)

axes = fig.add_subplot(3,1,2)
axes.plot( time , EAS / Units.kts, line_style )
axes.set_xlabel('Time (min)',axis_font)
axes.set_ylabel('Equivalent Airspeed',axis_font)
axes.grid(True)

axes = fig.add_subplot(3,1,3)
axes.plot( time , mach , line_style )
axes.set_xlabel('Time (min)',axis_font)
axes.set_ylabel('Mach',axis_font)
axes.grid(True)
# Plot Aerodynamic Forces
plot_aerodynamic_forces(results, line_style)

# Plot Aerodynamic Coefficients
plot_aerodynamic_coefficients(results, line_style)

# Drag Components
plot_drag_components(results, line_style)

# Plot Velocities
plot_aircraft_velocities(results, line_style)

return

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