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Merged
merged 7 commits into from
Aug 11, 2020
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Develop #20

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3b4f497
getting ready for 2.2, need to finish bwb
planes Aug 4, 2020
481c99b
new solar uav initials
planes Aug 4, 2020
19b8600
updated 737 tutorial
timdmacdo Aug 4, 2020
fa10e92
updated BWB
planes Aug 5, 2020
4ee2257
solar updates
planes Aug 6, 2020
bd0f478
Merge branch 'develop' into feature-detailed_tutorials
timdmacdo Aug 10, 2020
171cb61
Merge pull request #19 from suavecode/feature-detailed_tutorials
planes Aug 10, 2020
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205 changes: 29 additions & 176 deletions B737_AVL_Tutorial/tut_mission_B737_AVL.py
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Original file line number Diff line number Diff line change
Expand Up @@ -6,13 +6,10 @@
# 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

Expand Down
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