diff --git a/Airfoils/B737a.txt b/Airfoils/B737a.txt
new file mode 100644
index 0000000..714ca3d
--- /dev/null
+++ b/Airfoils/B737a.txt
@@ -0,0 +1,50 @@
+BOEING 737 ROOT AIRFOIL
+ 23. 23.
+
+ 0.000000 0.017700
+ 0.002300 0.030900
+ 0.005000 0.037200
+ 0.007600 0.041500
+ 0.014300 0.049900
+ 0.024900 0.058200
+ 0.049500 0.073000
+ 0.074000 0.081400
+ 0.099000 0.086600
+ 0.153000 0.090700
+ 0.196100 0.090500
+ 0.250400 0.088700
+ 0.309400 0.085800
+ 0.352000 0.083300
+ 0.391900 0.080400
+ 0.447700 0.075600
+ 0.503400 0.069600
+ 0.559300 0.062600
+ 0.596500 0.057500
+ 0.648800 0.049800
+ 0.835100 0.022400
+ 0.910900 0.013200
+ 1.000000 0.000000
+
+ 0.000000 0.017700
+ 0.002200 0.003800
+ 0.004900 -0.001800
+ 0.007200 -0.005300
+ 0.011900 -0.010600
+ 0.024300 -0.020400
+ 0.048600 -0.034200
+ 0.071600 -0.045700
+ 0.097900 -0.051600
+ 0.148800 -0.060700
+ 0.195300 -0.063200
+ 0.250100 -0.063200
+ 0.294500 -0.062600
+ 0.357900 -0.061000
+ 0.396500 -0.059500
+ 0.454300 -0.056300
+ 0.505000 -0.052700
+ 0.555600 -0.048200
+ 0.606300 -0.042700
+ 0.648500 -0.037500
+ 0.831700 -0.014900
+ 0.941000 -0.005300
+ 1.000000 0.000000
\ No newline at end of file
diff --git a/Airfoils/B737b.txt b/Airfoils/B737b.txt
new file mode 100644
index 0000000..e7aebdc
--- /dev/null
+++ b/Airfoils/B737b.txt
@@ -0,0 +1,50 @@
+BOEING 737 MIDSPAN AIRFOIL
+ 23. 23.
+
+ 0.000000 0.008800
+ 0.002600 0.019800
+ 0.004700 0.023600
+ 0.007500 0.027500
+ 0.012900 0.033200
+ 0.022900 0.040800
+ 0.053000 0.055200
+ 0.073600 0.061300
+ 0.099600 0.066600
+ 0.151300 0.072800
+ 0.208000 0.075600
+ 0.250000 0.076100
+ 0.297200 0.075600
+ 0.360200 0.073400
+ 0.407500 0.070500
+ 0.454700 0.066700
+ 0.510100 0.061400
+ 0.552500 0.056700
+ 0.600100 0.051200
+ 0.700300 0.038800
+ 0.826600 0.023300
+ 0.902100 0.014200
+ 1.000000 0.000000
+
+ 0.000000 0.008800
+ 0.002100 0.000400
+ 0.005100 -0.003700
+ 0.007800 -0.006200
+ 0.013900 -0.010300
+ 0.023000 -0.014700
+ 0.050900 -0.024400
+ 0.072500 -0.030100
+ 0.096100 -0.035200
+ 0.151300 -0.043200
+ 0.208000 -0.047700
+ 0.250000 -0.049300
+ 0.309500 -0.050000
+ 0.344900 -0.049800
+ 0.398100 -0.048600
+ 0.451200 -0.046300
+ 0.506600 -0.042800
+ 0.549000 -0.039700
+ 0.596600 -0.035700
+ 0.688900 -0.027500
+ 0.850500 -0.013100
+ 0.931300 -0.006000
+ 1.000000 0.000000
\ No newline at end of file
diff --git a/Airfoils/B737c.txt b/Airfoils/B737c.txt
new file mode 100644
index 0000000..cf4a965
--- /dev/null
+++ b/Airfoils/B737c.txt
@@ -0,0 +1,50 @@
+BOEING 737 MIDSPAN AIRFOIL
+ 23. 23.
+
+ 0.000000 0.000000
+ 0.001900 0.005700
+ 0.006000 0.010500
+ 0.007800 0.012000
+ 0.012900 0.015700
+ 0.023700 0.021800
+ 0.050500 0.032400
+ 0.079400 0.040200
+ 0.097500 0.044100
+ 0.151100 0.052400
+ 0.203600 0.057600
+ 0.250000 0.060600
+ 0.296400 0.062300
+ 0.346800 0.062900
+ 0.399000 0.062500
+ 0.444700 0.060800
+ 0.490400 0.059700
+ 0.548600 0.052900
+ 0.609100 0.047000
+ 0.657100 0.042200
+ 0.874100 0.018300
+ 0.947000 0.007600
+ 1.000000 0.000000
+
+ 0.000000 0.000000
+ 0.002200 -0.004000
+ 0.004900 -0.005700
+ 0.007000 -0.006600
+ 0.012500 -0.007600
+ 0.020900 -0.010100
+ 0.055500 -0.014700
+ 0.081600 -0.017400
+ 0.107800 -0.020100
+ 0.157100 -0.024700
+ 0.203600 -0.028600
+ 0.250000 -0.031900
+ 0.296400 -0.034700
+ 0.354900 -0.037000
+ 0.399600 -0.037500
+ 0.453200 -0.036600
+ 0.491100 -0.035200
+ 0.543400 -0.032800
+ 0.626900 -0.028200
+ 0.739600 -0.022000
+ 0.783200 -0.019000
+ 0.935400 -0.005600
+ 1.000000 0.000000
\ No newline at end of file
diff --git a/Airfoils/B737d.txt b/Airfoils/B737d.txt
new file mode 100644
index 0000000..f056bd9
--- /dev/null
+++ b/Airfoils/B737d.txt
@@ -0,0 +1,50 @@
+BOEING 737 OUTBOARD AIRFOIL
+ 23. 23.
+
+ 0.000000 0.000000
+ 0.002500 0.007000
+ 0.005000 0.010000
+ 0.007500 0.012300
+ 0.012500 0.016000
+ 0.025000 0.023200
+ 0.050000 0.033500
+ 0.075000 0.041000
+ 0.100000 0.046800
+ 0.150000 0.054900
+ 0.200000 0.060600
+ 0.250000 0.064300
+ 0.300000 0.066200
+ 0.350000 0.067800
+ 0.400000 0.067800
+ 0.450000 0.066800
+ 0.500000 0.064600
+ 0.550000 0.061100
+ 0.600000 0.056300
+ 0.700000 0.043700
+ 0.800000 0.029100
+ 0.900000 0.014500
+ 1.000000 0.000000
+
+ 0.000000 0.000000
+ 0.002500 -0.005100
+ 0.005000 -0.006600
+ 0.007500 -0.007700
+ 0.012500 -0.009100
+ 0.025000 -0.011600
+ 0.050000 -0.014800
+ 0.075000 -0.017400
+ 0.100000 -0.020000
+ 0.150000 -0.024600
+ 0.200000 -0.029100
+ 0.250000 -0.033100
+ 0.300000 -0.035900
+ 0.350000 -0.038800
+ 0.400000 -0.040200
+ 0.450000 -0.040400
+ 0.500000 -0.039300
+ 0.550000 -0.037100
+ 0.600000 -0.033900
+ 0.700000 -0.025700
+ 0.800000 -0.017200
+ 0.900000 -0.008600
+ 1.000000 0.000000
\ No newline at end of file
diff --git a/Airfoils/Clark_y.txt b/Airfoils/Clark_y.txt
new file mode 100644
index 0000000..196ed91
--- /dev/null
+++ b/Airfoils/Clark_y.txt
@@ -0,0 +1,126 @@
+CLARK Y AIRFOIL
+ 61. 61.
+
+ 0.000000 0.000000
+ 0.000500 0.002339
+ 0.001000 0.003727
+ 0.002000 0.005803
+ 0.004000 0.008924
+ 0.008000 0.013735
+ 0.012000 0.017858
+ 0.020000 0.025374
+ 0.030000 0.033022
+ 0.040000 0.039128
+ 0.050000 0.044275
+ 0.060000 0.048757
+ 0.080000 0.056431
+ 0.100000 0.062998
+ 0.120000 0.068620
+ 0.140000 0.073436
+ 0.160000 0.077571
+ 0.180000 0.081069
+ 0.200000 0.083920
+ 0.220000 0.086143
+ 0.240000 0.087831
+ 0.260000 0.089084
+ 0.280000 0.090002
+ 0.300000 0.090680
+ 0.320000 0.091186
+ 0.340000 0.091508
+ 0.360000 0.091627
+ 0.380000 0.091521
+ 0.400000 0.091171
+ 0.420000 0.090566
+ 0.440000 0.089718
+ 0.460000 0.088643
+ 0.480000 0.087357
+ 0.500000 0.085877
+ 0.520000 0.084214
+ 0.540000 0.082371
+ 0.560000 0.080348
+ 0.580000 0.078145
+ 0.600000 0.075763
+ 0.620000 0.073206
+ 0.640000 0.070482
+ 0.660000 0.067605
+ 0.680000 0.064584
+ 0.700000 0.061433
+ 0.720000 0.058160
+ 0.740000 0.054767
+ 0.760000 0.051257
+ 0.780000 0.047628
+ 0.800000 0.043884
+ 0.820000 0.040024
+ 0.840000 0.036054
+ 0.860000 0.031974
+ 0.880000 0.027789
+ 0.900000 0.023502
+ 0.920000 0.019116
+ 0.940000 0.014624
+ 0.960000 0.010023
+ 0.970000 0.007687
+ 0.980000 0.005333
+ 0.990000 0.002969
+ 1.000000 0.000599
+
+ 0.000000 0.000000
+ 0.000500 -0.004670
+ 0.001000 -0.005942
+ 0.002000 -0.007811
+ 0.004000 -0.010513
+ 0.008000 -0.014286
+ 0.012000 -0.016973
+ 0.020000 -0.020272
+ 0.030000 -0.022606
+ 0.040000 -0.024521
+ 0.050000 -0.026045
+ 0.060000 -0.027128
+ 0.080000 -0.028459
+ 0.100000 -0.029379
+ 0.120000 -0.029963
+ 0.140000 -0.030240
+ 0.160000 -0.030255
+ 0.180000 -0.030049
+ 0.200000 -0.029666
+ 0.220000 -0.029145
+ 0.240000 -0.028518
+ 0.260000 -0.027816
+ 0.280000 -0.027070
+ 0.300000 -0.026308
+ 0.320000 -0.025556
+ 0.340000 -0.024818
+ 0.360000 -0.024087
+ 0.380000 -0.023361
+ 0.400000 -0.022634
+ 0.420000 -0.021904
+ 0.440000 -0.021171
+ 0.460000 -0.020435
+ 0.480000 -0.019699
+ 0.500000 -0.018962
+ 0.520000 -0.018226
+ 0.540000 -0.017491
+ 0.560000 -0.016757
+ 0.580000 -0.016023
+ 0.600000 -0.015289
+ 0.620000 -0.014555
+ 0.640000 -0.013821
+ 0.660000 -0.013086
+ 0.680000 -0.012351
+ 0.700000 -0.011617
+ 0.720000 -0.010882
+ 0.740000 -0.010148
+ 0.760000 -0.009413
+ 0.780000 -0.008679
+ 0.800000 -0.007944
+ 0.820000 -0.007210
+ 0.840000 -0.006475
+ 0.860000 -0.005741
+ 0.880000 -0.005006
+ 0.900000 -0.004272
+ 0.920000 -0.003537
+ 0.940000 -0.002803
+ 0.960000 -0.002068
+ 0.970000 -0.001701
+ 0.980000 -0.001334
+ 0.990000 -0.000967
+ 1.000000 -0.000599
diff --git a/Airfoils/E190.py b/Airfoils/E190.py
new file mode 100644
index 0000000..71280a8
--- /dev/null
+++ b/Airfoils/E190.py
@@ -0,0 +1,954 @@
+# Embraer_190
+#
+# Created: Jan 2020 M. Clarke
+# Modified:
+
+""" setup file for the E190 vehicle
+"""
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+import SUAVE
+from SUAVE.Core import Units
+
+import numpy as np
+import pylab as plt
+import pickle
+import copy, time
+
+from SUAVE.Core import (
+Data, Container
+)
+import vsp
+from SUAVE.Input_Output.OpenVSP.vsp_write import write
+from SUAVE.Input_Output.OpenVSP.get_vsp_areas import get_vsp_areas
+from SUAVE.Methods.Geometry.Two_Dimensional.Planform import wing_planform
+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.Methods.Center_of_Gravity.compute_component_centers_of_gravity import compute_component_centers_of_gravity
+from SUAVE.Methods.Center_of_Gravity.compute_aircraft_center_of_gravity import compute_aircraft_center_of_gravity
+
+# ----------------------------------------------------------------------
+# Main
+# ----------------------------------------------------------------------
+
+def main():
+
+ 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()
+
+ #save_results(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)
+ 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_Transport()
+ weights.vehicle = vehicle
+ analyses.append(weights)
+
+ # ------------------------------------------------------------------
+ # Aerodynamics Analysis
+ aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
+ #aerodynamics.process.compute.lift.inviscid.keep_files = True
+ aerodynamics.geometry = vehicle
+
+ aerodynamics.settings.drag_coefficient_increment = 0.0000
+ analyses.append(aerodynamics)
+
+ # ------------------------------------------------------------------
+ # Stability Analysis
+ stability = SUAVE.Analyses.Stability.Fidelity_Zero()
+ stability.geometry = vehicle
+ analyses.append(stability)
+
+ # ------------------------------------------------------------------
+ # Energy
+ energy= SUAVE.Analyses.Energy.Energy()
+ energy.network = vehicle.propulsors
+ 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 = 'Embraer_E190AR'
+
+ # ------------------------------------------------------------------
+ # Vehicle-level Properties
+ # ------------------------------------------------------------------
+
+ # mass properties (http://www.embraercommercialaviation.com/AircraftPDF/E190_Weights.pdf)
+ vehicle.mass_properties.max_takeoff = 51800. # kg
+ vehicle.mass_properties.operating_empty = 27837. # kg
+ vehicle.mass_properties.takeoff = 51800. # kg
+ vehicle.mass_properties.max_zero_fuel = 40900. # kg
+ vehicle.mass_properties.max_payload = 13063. # kg
+ vehicle.mass_properties.max_fuel = 12971. # kg
+ vehicle.mass_properties.cargo = 0.0 # kg
+
+ vehicle.mass_properties.center_of_gravity = [16.8, 0, 1.6]#[[60 * Units.feet, 0, 0]] # Not correct
+ vehicle.mass_properties.moments_of_inertia.tensor = [[10 ** 5, 0, 0],[0, 10 ** 6, 0,],[0,0, 10 ** 7]] # Not Correct
+
+ # envelope properties
+ vehicle.envelope.ultimate_load = 3.5
+ vehicle.envelope.limit_load = 1.5
+
+ # basic parameters
+ vehicle.reference_area = 92.
+ vehicle.passengers = 114
+ vehicle.systems.control = "fully powered"
+ vehicle.systems.accessories = "medium range"
+
+ # ------------------------------------------------------------------
+ # Main Wing
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Main_Wing()
+ wing.tag = 'main_wing'
+
+ wing.areas.reference = 92.0
+ wing.aspect_ratio = 8.4
+ wing.sweeps.quarter_chord = 23.0 * Units.deg
+ wing.thickness_to_chord = 0.11
+ wing.taper = 0.28
+ wing.dihedral = 5.00
+
+ wing.origin = [13,0,0]
+ wing.vertical = False
+ wing.symmetric = True
+ wing.high_lift = True
+ wing.flaps.type = 'double_slotted'
+ wing.flaps.chord = 0.28
+ wing.flaps.span_start = 0.11
+ wing.flaps.span_end = 0.85
+
+ wing = wing_planform(wing)
+ wing.areas.exposed = 0.80 * wing.areas.wetted
+
+ wing.twists.root = 2.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.span_efficiency = 1.0
+ wing.dynamic_pressure_ratio = 1.0
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+ # ------------------------------------------------------------------
+ # Horizontal Stabilizer
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Wing()
+ wing.tag = 'horizontal_stabilizer'
+
+ wing.areas.reference = 26.0
+ wing.aspect_ratio = 5.5
+ wing.sweeps.quarter_chord = 34.5 * Units.deg
+ wing.thickness_to_chord = 0.11
+ wing.taper = 0.11
+ wing.dihedral = 8.00
+
+ wing.origin = [32,0,0]
+ wing.vertical = False
+ wing.symmetric = True
+ wing.high_lift = False
+
+ wing = wing_planform(wing)
+ wing.areas.exposed = 0.9 * wing.areas.wetted
+
+ wing.twists.root = 2.0 * Units.degrees
+ wing.twists.tip = 2.0 * Units.degrees
+ wing.span_efficiency = 0.90
+ wing.dynamic_pressure_ratio = 0.90
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+ # ------------------------------------------------------------------
+ # Vertical Stabilizer
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Wing()
+ wing.tag = 'vertical_stabilizer'
+
+ wing.areas.reference = 16.0
+ wing.aspect_ratio = 1.7
+ wing.sweeps.quarter_chord = 35. * Units.deg
+ wing.thickness_to_chord = 0.11
+ wing.taper = 0.31
+ wing.dihedral = 0.00
+
+ wing.origin = [32,0,0]
+ wing.vertical = True
+ wing.symmetric = False
+ wing.high_lift = False
+
+ wing = wing_planform(wing)
+ wing.areas.exposed = 0.9 * wing.areas.wetted
+
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.span_efficiency = 0.90
+ wing.dynamic_pressure_ratio = 1.00
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+ # ------------------------------------------------------------------
+ # Fuselage
+ # ------------------------------------------------------------------
+
+ fuselage = SUAVE.Components.Fuselages.Fuselage()
+ fuselage.tag = 'fuselage'
+ fuselage.origin = [[0,0,0]]
+ fuselage.number_coach_seats = vehicle.passengers
+ fuselage.seats_abreast = 4
+ fuselage.seat_pitch = 0.7455
+
+ fuselage.fineness.nose = 2.0
+ fuselage.fineness.tail = 3.0
+
+ fuselage.lengths.nose = 6.0
+ fuselage.lengths.tail = 9.0
+ fuselage.lengths.cabin = 21.24
+ fuselage.lengths.total = 36.24
+ fuselage.lengths.fore_space = 0.
+ fuselage.lengths.aft_space = 0.
+
+ fuselage.width = 3.18
+
+ fuselage.heights.maximum = 4.18
+ fuselage.heights.at_quarter_length = 3.18
+ fuselage.heights.at_three_quarters_length = 3.18
+ fuselage.heights.at_wing_root_quarter_chord = 4.00
+
+ fuselage.areas.side_projected = 239.20
+ fuselage.areas.wetted = 327.01
+ fuselage.areas.front_projected = 8.0110
+
+ fuselage.effective_diameter = 3.18
+
+ fuselage.differential_pressure = 10**5 * Units.pascal # Maximum differential pressure
+
+ # add to vehicle
+ vehicle.append_component(fuselage)
+
+ # ------------------------------------------------------------------
+ # Turbofan Network
+ # ------------------------------------------------------------------
+
+
+ #initialize the gas turbine network
+ gt_engine = SUAVE.Components.Energy.Networks.Turbofan()
+ gt_engine.tag = 'turbofan'
+
+ gt_engine.number_of_engines = 2.0
+ gt_engine.bypass_ratio = 5.4
+ gt_engine.engine_length = 2.71
+ gt_engine.nacelle_diameter = 2.05
+
+ #compute engine areas)
+ Amax = (np.pi/4.)*gt_engine.nacelle_diameter**2.
+ Awet = 1.1*np.pi*gt_engine.nacelle_diameter*gt_engine.engine_length # 1.1 is simple coefficient
+
+ #Assign engine areas
+
+ gt_engine.areas.wetted = Awet
+
+ #set the working fluid for the network
+ working_fluid = SUAVE.Attributes.Gases.Air()
+
+ #add working fluid to the network
+ gt_engine.working_fluid = working_fluid
+
+
+ #Component 1 : ram, to convert freestream static to stagnation quantities
+ ram = SUAVE.Components.Energy.Converters.Ram()
+ ram.tag = 'ram'
+
+ #add ram to the network
+ gt_engine.ram = ram
+
+
+ #Component 2 : inlet nozzle
+ inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
+ inlet_nozzle.tag = 'inlet nozzle'
+
+ inlet_nozzle.polytropic_efficiency = 0.98
+ inlet_nozzle.pressure_ratio = 0.98
+
+ #add inlet nozzle to the network
+ gt_engine.inlet_nozzle = inlet_nozzle
+
+
+ #Component 3 :low pressure compressor
+ low_pressure_compressor = SUAVE.Components.Energy.Converters.Compressor()
+ low_pressure_compressor.tag = 'lpc'
+
+ low_pressure_compressor.polytropic_efficiency = 0.91
+ low_pressure_compressor.pressure_ratio = 1.9
+
+ #add low pressure compressor to the network
+ gt_engine.low_pressure_compressor = low_pressure_compressor
+
+ #Component 4 :high pressure compressor
+ high_pressure_compressor = SUAVE.Components.Energy.Converters.Compressor()
+ high_pressure_compressor.tag = 'hpc'
+
+ high_pressure_compressor.polytropic_efficiency = 0.91
+ high_pressure_compressor.pressure_ratio = 10.0
+
+ #add the high pressure compressor to the network
+ gt_engine.high_pressure_compressor = high_pressure_compressor
+
+ #Component 5 :low pressure turbine
+ low_pressure_turbine = SUAVE.Components.Energy.Converters.Turbine()
+ low_pressure_turbine.tag='lpt'
+
+ low_pressure_turbine.mechanical_efficiency = 0.99
+ low_pressure_turbine.polytropic_efficiency = 0.93
+
+ #add low pressure turbine to the network
+ gt_engine.low_pressure_turbine = low_pressure_turbine
+
+ #Component 5 :high pressure turbine
+ high_pressure_turbine = SUAVE.Components.Energy.Converters.Turbine()
+ high_pressure_turbine.tag='hpt'
+
+ high_pressure_turbine.mechanical_efficiency = 0.99
+ high_pressure_turbine.polytropic_efficiency = 0.93
+
+ #add the high pressure turbine to the network
+ gt_engine.high_pressure_turbine = high_pressure_turbine
+
+ #Component 6 :combustor
+ combustor = SUAVE.Components.Energy.Converters.Combustor()
+ combustor.tag = 'Comb'
+
+ combustor.efficiency = 0.99
+ combustor.alphac = 1.0
+ combustor.turbine_inlet_temperature = 1500
+ combustor.pressure_ratio = 0.95
+ combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
+
+ #add the combustor to the network
+ gt_engine.combustor = combustor
+
+ #Component 7 :core nozzle
+ core_nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
+ core_nozzle.tag = 'core nozzle'
+
+ core_nozzle.polytropic_efficiency = 0.95
+ core_nozzle.pressure_ratio = 0.99
+
+ #add the core nozzle to the network
+ gt_engine.core_nozzle = core_nozzle
+
+ #Component 8 :fan nozzle
+ fan_nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
+ fan_nozzle.tag = 'fan nozzle'
+
+ fan_nozzle.polytropic_efficiency = 0.95
+ fan_nozzle.pressure_ratio = 0.99
+
+ #add the fan nozzle to the network
+ gt_engine.fan_nozzle = fan_nozzle
+
+ #Component 9 : fan
+ fan = SUAVE.Components.Energy.Converters.Fan()
+ fan.tag = 'fan'
+
+ fan.polytropic_efficiency = 0.93
+ fan.pressure_ratio = 1.7
+
+ #add the fan to the network
+ gt_engine.fan = fan
+
+ #Component 10 : thrust (to compute the thrust)
+ thrust = SUAVE.Components.Energy.Processes.Thrust()
+ thrust.tag ='compute_thrust'
+
+ #total design thrust (includes all the engines)
+ thrust.total_design = 37278.0* Units.N #Newtons
+
+ #design sizing conditions
+ altitude = 35000.0*Units.ft
+ mach_number = 0.78
+ isa_deviation = 0.
+
+ # add thrust to the network
+ gt_engine.thrust = thrust
+
+ #size the turbofan
+ turbofan_sizing(gt_engine,mach_number,altitude)
+
+ # add gas turbine network gt_engine to the vehicle
+ vehicle.append_component(gt_engine)
+
+ fuel =SUAVE.Components.Physical_Component()
+ vehicle.fuel =fuel
+
+ fuel.mass_properties.mass =vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_fuel
+ fuel.origin =vehicle.wings.main_wing.mass_properties.center_of_gravity
+ fuel.mass_properties.center_of_gravity=vehicle.wings.main_wing.aerodynamic_center
+ # ------------------------------------------------------------------
+ # Vehicle Definition Complete
+ # ------------------------------------------------------------------
+
+ 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 = 20. * Units.deg
+ config.wings['main_wing'].slats.angle = 25. * Units.deg
+
+ config.V2_VS_ratio = 1.21
+ configs.append(config)
+
+ # ------------------------------------------------------------------
+ # Landing Configuration
+ # ------------------------------------------------------------------
+
+ config = SUAVE.Components.Configs.Config(base_config)
+ config.tag = 'landing'
+
+ config.wings['main_wing'].flaps.angle = 30. * Units.deg
+ config.wings['main_wing'].slats.angle = 25. * Units.deg
+
+ config.Vref_VS_ratio = 1.23
+ 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 = 20. * Units.deg
+ config.wings['main_wing'].slats.angle = 25. * Units.deg
+
+ config.V2_VS_ratio = 1.21
+
+ # payload?
+
+ configs.append(config)
+
+ # done!
+ return configs
+
+
+
+# ----------------------------------------------------------------------
+# Define the Mission
+# ----------------------------------------------------------------------
+def mission_setup(analyses):
+
+ # ------------------------------------------------------------------
+ # Initialize the Mission
+ # ------------------------------------------------------------------
+
+ mission = SUAVE.Analyses.Mission.Sequential_Segments()
+ mission.tag = 'the_mission'
+
+ #airport
+ airport = SUAVE.Attributes.Airports.Airport()
+ airport.altitude = 0.0 * Units.ft
+ airport.delta_isa = 0.0
+ airport.atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+
+ mission.airport = airport
+
+ # unpack Segments module
+ Segments = SUAVE.Analyses.Mission.Segments
+
+ # base segment
+ base_segment = Segments.Segment()
+ atmosphere=SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
+ planet = SUAVE.Attributes.Planets.Earth()
+ # ------------------------------------------------------------------
+ # First Climb Segment: Constant Speed, Constant Throttle
+ # ------------------------------------------------------------------
+
+ segment = Segments.Climb.Constant_Speed_Constant_Rate()
+ segment.tag = "climb_1_mr"
+ ones_row = segment.state.ones_row
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.takeoff )
+
+ # define segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_start = 0.0 * Units.km
+ segment.altitude_end = 3.048 * Units.km
+ segment.air_speed = 138.0 * Units['m/s']
+ segment.climb_rate = 2900. * Units['ft/min']
+ segment.state.unknowns.throttle = 0.75 * ones_row(1)
+ segment.state.unknowns.body_angle = ones_row(1) * 5.0 * Units.degrees
+
+ # add to misison
+ mission.append_segment(segment)
+
+ # ------------------------------------------------------------------
+ # Second Climb Segment: Constant Speed, Constant Rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Climb.Constant_Speed_Constant_Rate()
+ segment.tag = "climb_2_mr"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 3.657 * Units.km
+ segment.air_speed = 168.0 * Units['m/s']
+ segment.climb_rate = 2500. * Units['ft/min']
+ segment.state.unknowns.throttle = 0.75 * ones_row(1)
+
+ # add to mission
+ mission.append_segment(segment)
+
+ # ------------------------------------------------------------------
+ # Third Climb Segment: Constant Speed, Constant Climb Rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Climb.Constant_Speed_Constant_Rate()
+ segment.tag = "climb_3_mr"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 25000. * Units.ft
+ segment.air_speed = 200.0 * Units['m/s']
+ segment.climb_rate = 1700. * Units['ft/min']
+ segment.state.unknowns.throttle = 0.75 * ones_row(1)
+ # add to mission
+ mission.append_segment(segment)
+
+ # ------------------------------------------------------------------
+ # Fourth Climb Segment: Constant Speed, Constant Rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Climb.Constant_Speed_Constant_Rate()
+ segment.tag = "climb_4_mr"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 32000. * Units.ft
+ segment.air_speed = 225.0* Units['m/s']
+ segment.climb_rate = 800. * Units['ft/min']
+ segment.state.unknowns.throttle = 0.75 * ones_row(1)
+ # add to mission
+ mission.append_segment(segment)
+
+ # ------------------------------------------------------------------
+ # Fifth Climb Segment: Constant Speed, Constant Rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Climb.Constant_Speed_Constant_Rate()
+ segment.tag = "climb_5"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 36999. * Units.ft
+ segment.air_speed = 230.0 * Units['m/s']
+ segment.climb_rate = 300. * Units['ft/min']
+
+ # add to mission
+ mission.append_segment(segment)
+
+
+ # ------------------------------------------------------------------
+ # Cruise Segment: constant speed, constant altitude
+ # ------------------------------------------------------------------
+
+ segment = Segments.Cruise.Constant_Speed_Constant_Altitude()
+ segment.tag = "cruise"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.air_speed = 450. * Units.knots
+ segment.distance = 2050. * Units.nmi
+ segment.state.unknowns.throttle = 0.75 * ones_row(1)
+
+ # add to mission
+ mission.append_segment(segment)
+
+ # ------------------------------------------------------------------
+ # First Descent Segment: consant speed, constant segment rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Descent.Constant_Speed_Constant_Rate()
+ segment.tag = "descent_1_mr"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 9.31 * Units.km
+ segment.air_speed = 440.0 * Units.knots
+ segment.descent_rate = 2600. * Units['ft/min']
+
+ # add to mission
+ mission.append_segment(segment)
+
+
+ # ------------------------------------------------------------------
+ # Second Descent Segment: consant speed, constant segment rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Descent.Constant_Speed_Constant_Rate()
+ segment.tag = "descent_2_mr"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 3.657 * Units.km
+ segment.air_speed = 365.0 * Units.knots
+ segment.descent_rate = 2300. * Units['ft/min']
+ segment.state.unknowns.throttle = 0.1 * ones_row(1)
+ segment.state.unknowns.body_angle = ones_row(1) * 5.0 * Units.degrees
+
+ # append to mission
+ mission.append_segment(segment)
+
+
+ # ------------------------------------------------------------------
+ # Third Descent Segment: consant speed, constant segment rate
+ # ------------------------------------------------------------------
+
+ segment = Segments.Descent.Constant_Speed_Constant_Rate()
+ segment.tag = "descent_3_mr"
+
+ # connect vehicle configuration
+ segment.analyses.extend( analyses.cruise )
+
+ # segment attributes
+ segment.atmosphere = atmosphere
+ segment.planet = planet
+
+ segment.altitude_end = 0.0 * Units.km
+ segment.air_speed = 250.0 * Units.knots
+ segment.descent_rate = 1500. * Units['ft/min']
+ segment.state.unknowns.throttle = 0.1 * ones_row(1)
+ segment.state.unknowns.body_angle = ones_row(1) * 10.0 * Units.degrees
+
+ # append to mission
+ mission.append_segment(segment)
+
+ # ------------------------------------------------------------------
+ # Mission definition complete
+ # ------------------------------------------------------------------
+
+
+ ##------------------------------------------------------------------
+ #### Reserve mission
+ ##------------------------------------------------------------------
+
+ ## ------------------------------------------------------------------
+ ## First Climb Segment: constant Mach, constant segment angle
+ ## ------------------------------------------------------------------
+
+ ##segment = Segments.Climb.Constant_Speed_Constant_Rate(base_segment)
+ #segment = Segments.Climb.Linear_Mach_Constant_Rate(base_segment)
+ #segment.tag = "reserve_climb_1_mr"
+
+ #segment.analyses.extend( analyses.cruise )
+
+ #segment.altitude_start = 0.0 * Units.km
+ #segment.altitude_end = 15000.0 * Units.ft
+ #segment.climb_rate = 1800. * Units['ft/min']
+ #segment.mach_end = 0.3
+ #segment.mach_start = 0.2
+ #segment.state.unknowns.throttle = 0.7 * ones_row(1) # 0.65 0.6 0.
+ ## add to misison
+ #mission.append_segment(segment)
+
+
+
+ ## ------------------------------------------------------------------
+ ## Cruise Segment: constant speed, constant altitude
+ ## ------------------------------------------------------------------
+
+ #segment = Segments.Cruise.Constant_Speed_Constant_Altitude(base_segment)
+ #segment.tag = "reserve_cruise_mr"
+
+ #segment.analyses.extend( analyses.cruise )
+
+ #segment.air_speed = 96.66 * Units['m/s']
+ #segment.distance = 100.0 * Units.nautical_mile
+ #mission.append_segment(segment)
+
+ ## ------------------------------------------------------------------
+ ## Loiter Segment: constant mach, constant time
+ ## ------------------------------------------------------------------
+
+ #segment = Segments.Cruise.Constant_Mach_Constant_Altitude_Loiter(base_segment)
+ #segment.tag = "reserve_loiter_mr"
+
+ #segment.analyses.extend( analyses.cruise )
+
+ #segment.mach = 0.5
+ #segment.time = 30.0 * Units.minutes
+
+ #mission.append_segment(segment)
+
+
+ ## ------------------------------------------------------------------
+ ## Fifth Descent Segment: consant speed, constant segment rate
+ ## ------------------------------------------------------------------
+
+ #segment = Segments.Descent.Linear_Mach_Constant_Rate(base_segment)
+ #segment.tag = "reserve_descent_1_mr"
+
+ #segment.analyses.extend( analyses.landing )
+ #analyses.landing.aerodynamics.settings.spoiler_drag_increment = 0.00
+
+
+ #segment.altitude_end = 0.0 * Units.km
+ #segment.descent_rate = 3.0 * Units['m/s']
+
+
+ #segment.mach_end = 0.24
+ #segment.mach_start = 0.3
+
+ ## append to mission
+ #mission.append_segment(segment)
+
+ ##------------------------------------------------------------------
+ #### Reserve mission completed
+ ##------------------------------------------------------------------
+
+
+ return mission
+
+
+def missions_setup(base_mission):
+
+ # the mission container
+ missions = SUAVE.Analyses.Mission.Mission.Container()
+
+ # ------------------------------------------------------------------
+ # Base Mission
+ # ------------------------------------------------------------------
+
+ missions.base = base_mission
+
+
+ # ------------------------------------------------------------------
+ # Mission for Constrained Fuel
+ # ------------------------------------------------------------------
+ fuel_mission = SUAVE.Analyses.Mission.Mission() #Fuel_Constrained()
+ fuel_mission.tag = 'fuel'
+ fuel_mission.range = 1277. * Units.nautical_mile
+ fuel_mission.payload = 19000.
+ missions.append(fuel_mission)
+
+
+ # ------------------------------------------------------------------
+ # Mission for Constrained Short Field
+ # ------------------------------------------------------------------
+ short_field = SUAVE.Analyses.Mission.Mission(base_mission) #Short_Field_Constrained()
+ short_field.tag = 'short_field'
+
+ #airport
+ airport = SUAVE.Attributes.Airports.Airport()
+ airport.altitude = 0.0 * Units.ft
+ airport.delta_isa = 0.0
+ airport.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
+ airport.available_tofl = 1500.
+ short_field.airport = airport
+ missions.append(short_field)
+
+ # ------------------------------------------------------------------
+ # Mission for Fixed Payload
+ # ------------------------------------------------------------------
+ payload = SUAVE.Analyses.Mission.Mission() #Payload_Constrained()
+ payload.tag = 'payload'
+ payload.range = 2316. * Units.nautical_mile
+ payload.payload = 19000.
+ missions.append(payload)
+
+
+ # done!
+ return missions
+
+
+# ----------------------------------------------------------------------
+# Plot Mission
+# ----------------------------------------------------------------------
+
+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
+
+
+if __name__ == '__main__':
+ main()
+ plt.show()
+
diff --git a/Airfoils/E63.txt b/Airfoils/E63.txt
new file mode 100644
index 0000000..fe072fd
--- /dev/null
+++ b/Airfoils/E63.txt
@@ -0,0 +1,63 @@
+E63 (4.25%)
+ 1.00000 0.00000
+ 0.99719 0.00121
+ 0.98938 0.00473
+ 0.97751 0.00986
+ 0.96173 0.01553
+ 0.94164 0.02126
+ 0.91717 0.02709
+ 0.88861 0.03301
+ 0.85624 0.03885
+ 0.82039 0.04451
+ 0.78141 0.04985
+ 0.73968 0.05480
+ 0.69562 0.05921
+ 0.64967 0.06304
+ 0.60229 0.06617
+ 0.55394 0.06857
+ 0.50509 0.07016
+ 0.45624 0.07094
+ 0.40786 0.07084
+ 0.36043 0.06990
+ 0.31441 0.06809
+ 0.27026 0.06545
+ 0.22840 0.06198
+ 0.18920 0.05775
+ 0.15304 0.05280
+ 0.12023 0.04723
+ 0.09103 0.04111
+ 0.06568 0.03457
+ 0.04435 0.02775
+ 0.02714 0.02083
+ 0.01416 0.01404
+ 0.00536 0.00766
+ 0.00076 0.00218
+ 0.00055 -0.00141
+ 0.0 -0.00123
+ 0.00557 -0.00306
+ 0.01651 -0.00330
+ 0.03316 -0.00227
+ 0.05550 -0.00004
+ 0.08342 0.00315
+ 0.11671 0.00708
+ 0.15504 0.01151
+ 0.19800 0.01620
+ 0.24509 0.02093
+ 0.29574 0.02546
+ 0.34931 0.02962
+ 0.40513 0.03319
+ 0.46247 0.03605
+ 0.52056 0.03803
+ 0.57859 0.03907
+ 0.63576 0.03907
+ 0.69125 0.03806
+ 0.74430 0.03604
+ 0.79414 0.03310
+ 0.84004 0.02930
+ 0.88132 0.02482
+ 0.91735 0.01979
+ 0.94756 0.01439
+ 0.97115 0.00887
+ 0.98754 0.00410
+ 0.99695 0.00102
+ 1.00000 0.00000
\ No newline at end of file
diff --git a/Airfoils/E850.txt b/Airfoils/E850.txt
new file mode 100644
index 0000000..e031c17
--- /dev/null
+++ b/Airfoils/E850.txt
@@ -0,0 +1,72 @@
+EPPLER E850 AIRFOIL
+ 35. 33.
+
+ 0.000000 0.000000
+ 0.000130 0.000650
+ 0.000810 0.001790
+ 0.002440 0.003490
+ 0.006350 0.006210
+ 0.008450 0.007350
+ 0.020250 0.012240
+ 0.036920 0.017150
+ 0.058390 0.021960
+ 0.084450 0.026610
+ 0.114800 0.031050
+ 0.149130 0.035180
+ 0.187080 0.038940
+ 0.228220 0.042270
+ 0.272150 0.045110
+ 0.318390 0.047410
+ 0.366450 0.049140
+ 0.415830 0.050260
+ 0.466000 0.050750
+ 0.516440 0.050590
+ 0.566610 0.049780
+ 0.615970 0.048320
+ 0.664020 0.046180
+ 0.710230 0.043350
+ 0.754200 0.039760
+ 0.795630 0.035440
+ 0.834270 0.030510
+ 0.869890 0.025270
+ 0.902070 0.020070
+ 0.930270 0.015200
+ 0.953990 0.010700
+ 0.973020 0.006520
+ 0.987400 0.002970
+ 0.996720 0.000770
+ 1.000000 0.000080
+
+ 0.000000 0.000000
+ 0.000210 -0.000640
+ 0.001050 -0.001710
+ 0.002940 -0.003250
+ 0.007220 -0.005760
+ 0.015280 -0.008890
+ 0.030370 -0.012940
+ 0.050350 -0.016720
+ 0.075100 -0.020200
+ 0.104290 -0.023360
+ 0.137580 -0.026140
+ 0.174560 -0.028480
+ 0.214820 -0.030280
+ 0.257880 -0.031470
+ 0.303250 -0.031860
+ 0.350550 -0.031250
+ 0.399520 -0.029530
+ 0.449900 -0.026720
+ 0.501490 -0.023020
+ 0.553890 -0.018850
+ 0.606440 -0.014580
+ 0.658460 -0.010480
+ 0.709210 -0.006750
+ 0.757970 -0.003580
+ 0.804010 -0.001050
+ 0.846600 0.000740
+ 0.885080 0.001830
+ 0.918810 0.002250
+ 0.947220 0.002150
+ 0.969880 0.001690
+ 0.986450 0.001040
+ 0.996590 0.000380
+ 1.000000 0.000080
diff --git a/Airfoils/E851.txt b/Airfoils/E851.txt
new file mode 100644
index 0000000..f50cee2
--- /dev/null
+++ b/Airfoils/E851.txt
@@ -0,0 +1,72 @@
+EPPLER E851 AIRFOIL
+ 35. 33.
+
+ 0.000000 0.000000
+ 0.000090 0.000740
+ 0.000710 0.002050
+ 0.002260 0.004010
+ 0.006000 0.007250
+ 0.010580 0.010110
+ 0.023250 0.015960
+ 0.040660 0.021860
+ 0.062721 0.027660
+ 0.089231 0.033300
+ 0.119891 0.038670
+ 0.154362 0.043700
+ 0.192302 0.048271
+ 0.233302 0.052311
+ 0.276943 0.055741
+ 0.322753 0.058511
+ 0.370264 0.060561
+ 0.418984 0.061841
+ 0.468385 0.062311
+ 0.517955 0.061941
+ 0.567186 0.060691
+ 0.615526 0.058501
+ 0.662597 0.055251
+ 0.708097 0.050940
+ 0.751777 0.045691
+ 0.793428 0.039770
+ 0.832628 0.033620
+ 0.868809 0.027570
+ 0.901399 0.021830
+ 0.929879 0.016530
+ 0.953760 0.011670
+ 0.972870 0.007160
+ 0.987320 0.003330
+ 0.996690 0.000930
+ 1.000000 0.000170
+
+ 0.000000 0.000000
+ 0.000250 -0.000720
+ 0.001150 -0.001900
+ 0.003130 -0.003600
+ 0.007620 -0.006250
+ 0.012680 -0.008370
+ 0.026840 -0.012610
+ 0.045970 -0.016530
+ 0.069931 -0.020070
+ 0.098461 -0.023240
+ 0.131211 -0.025990
+ 0.167762 -0.028270
+ 0.207692 -0.030000
+ 0.250522 -0.031090
+ 0.295753 -0.031370
+ 0.343013 -0.030630
+ 0.392044 -0.028760
+ 0.442584 -0.025780
+ 0.494445 -0.021880
+ 0.547235 -0.017540
+ 0.600276 -0.013110
+ 0.652867 -0.008890
+ 0.704277 -0.005100
+ 0.753728 -0.001910
+ 0.800488 0.000570
+ 0.843798 0.002260
+ 0.882969 0.003180
+ 0.917329 0.003380
+ 0.946299 0.003020
+ 0.969410 0.002280
+ 0.986270 0.001360
+ 0.996540 0.000530
+ 1.000000 0.000170
diff --git a/Airfoils/E854.txt b/Airfoils/E854.txt
new file mode 100644
index 0000000..937b6e6
--- /dev/null
+++ b/Airfoils/E854.txt
@@ -0,0 +1,72 @@
+EPPLER E854 AIRFOIL
+ 35. 33.
+
+ 0.000000 0.000000
+ 0.000340 0.001990
+ 0.001520 0.004950
+ 0.004680 0.009850
+ 0.006190 0.011670
+ 0.016030 0.020680
+ 0.030281 0.030091
+ 0.048851 0.039571
+ 0.071591 0.048891
+ 0.098292 0.057901
+ 0.128633 0.066401
+ 0.162303 0.074211
+ 0.198944 0.081152
+ 0.238185 0.087032
+ 0.279606 0.091692
+ 0.322776 0.094862
+ 0.367427 0.096262
+ 0.413418 0.095792
+ 0.460549 0.093522
+ 0.508600 0.089612
+ 0.557221 0.084402
+ 0.605902 0.078212
+ 0.654083 0.071291
+ 0.701214 0.063891
+ 0.746685 0.056251
+ 0.789936 0.048561
+ 0.830367 0.041021
+ 0.867427 0.033771
+ 0.900588 0.026930
+ 0.929358 0.020570
+ 0.953279 0.014640
+ 0.972329 0.008940
+ 0.986860 0.003910
+ 0.996520 0.000630
+ 1.000000 -0.000420
+
+ 0.000000 0.000000
+ 0.000090 -0.001150
+ 0.000520 -0.002250
+ 0.001750 -0.003960
+ 0.004220 -0.006310
+ 0.009540 -0.009890
+ 0.019390 -0.014470
+ 0.036421 -0.019910
+ 0.058351 -0.024791
+ 0.085002 -0.029051
+ 0.116072 -0.032721
+ 0.151143 -0.035751
+ 0.189764 -0.038061
+ 0.231455 -0.039551
+ 0.275695 -0.040031
+ 0.322126 -0.039261
+ 0.370477 -0.037061
+ 0.420578 -0.033351
+ 0.472439 -0.028361
+ 0.525760 -0.022721
+ 0.579832 -0.016950
+ 0.633893 -0.011440
+ 0.687134 -0.006480
+ 0.738715 -0.002330
+ 0.787766 0.000860
+ 0.833447 0.003020
+ 0.874947 0.004130
+ 0.911508 0.004300
+ 0.942419 0.003700
+ 0.967149 0.002600
+ 0.985230 0.001280
+ 0.996280 0.000090
+ 1.000000 -0.000420
diff --git a/Airfoils/NACA65-203.txt b/Airfoils/NACA65-203.txt
new file mode 100644
index 0000000..00d2c16
--- /dev/null
+++ b/Airfoils/NACA65-203.txt
@@ -0,0 +1,56 @@
+NACA 65-203
+ 26. 26.
+
+ 0.00000 0.00010
+ 0.00460 0.00297
+ 0.00706 0.00366
+ 0.01200 0.00474
+ 0.02444 0.00673
+ 0.04939 0.00986
+ 0.07437 0.01237
+ 0.09936 0.01451
+ 0.14939 0.01797
+ 0.19945 0.02064
+ 0.24953 0.02270
+ 0.29962 0.02422
+ 0.34971 0.02525
+ 0.39981 0.02581
+ 0.44990 0.02589
+ 0.50000 0.02546
+ 0.55009 0.02451
+ 0.60016 0.02311
+ 0.65022 0.02133
+ 0.70026 0.01920
+ 0.75028 0.01675
+ 0.80027 0.01399
+ 0.85024 0.01097
+ 0.90018 0.00768
+ 0.95009 0.00413
+ 1.00000 0.00000
+
+ 0.00000 0.00010
+ 0.00540 -0.00183
+ 0.00794 -0.00212
+ 0.01300 -0.00249
+ 0.02556 -0.00292
+ 0.05061 -0.00343
+ 0.07563 -0.00377
+ 0.10064 -0.00403
+ 0.15061 -0.00435
+ 0.20055 -0.00454
+ 0.25047 -0.00461
+ 0.30038 -0.00457
+ 0.35029 -0.00444
+ 0.40019 -0.00418
+ 0.45010 -0.00378
+ 0.50000 -0.00319
+ 0.54991 -0.00241
+ 0.59984 -0.00149
+ 0.64978 -0.00054
+ 0.69974 0.00042
+ 0.74972 0.00132
+ 0.79973 0.00208
+ 0.84976 0.00262
+ 0.89982 0.00276
+ 0.94991 0.00225
+ 1.00000 0.00000
\ No newline at end of file
diff --git a/Airfoils/NACA65_203.txt b/Airfoils/NACA65_203.txt
new file mode 100644
index 0000000..00d2c16
--- /dev/null
+++ b/Airfoils/NACA65_203.txt
@@ -0,0 +1,56 @@
+NACA 65-203
+ 26. 26.
+
+ 0.00000 0.00010
+ 0.00460 0.00297
+ 0.00706 0.00366
+ 0.01200 0.00474
+ 0.02444 0.00673
+ 0.04939 0.00986
+ 0.07437 0.01237
+ 0.09936 0.01451
+ 0.14939 0.01797
+ 0.19945 0.02064
+ 0.24953 0.02270
+ 0.29962 0.02422
+ 0.34971 0.02525
+ 0.39981 0.02581
+ 0.44990 0.02589
+ 0.50000 0.02546
+ 0.55009 0.02451
+ 0.60016 0.02311
+ 0.65022 0.02133
+ 0.70026 0.01920
+ 0.75028 0.01675
+ 0.80027 0.01399
+ 0.85024 0.01097
+ 0.90018 0.00768
+ 0.95009 0.00413
+ 1.00000 0.00000
+
+ 0.00000 0.00010
+ 0.00540 -0.00183
+ 0.00794 -0.00212
+ 0.01300 -0.00249
+ 0.02556 -0.00292
+ 0.05061 -0.00343
+ 0.07563 -0.00377
+ 0.10064 -0.00403
+ 0.15061 -0.00435
+ 0.20055 -0.00454
+ 0.25047 -0.00461
+ 0.30038 -0.00457
+ 0.35029 -0.00444
+ 0.40019 -0.00418
+ 0.45010 -0.00378
+ 0.50000 -0.00319
+ 0.54991 -0.00241
+ 0.59984 -0.00149
+ 0.64978 -0.00054
+ 0.69974 0.00042
+ 0.74972 0.00132
+ 0.79973 0.00208
+ 0.84976 0.00262
+ 0.89982 0.00276
+ 0.94991 0.00225
+ 1.00000 0.00000
\ No newline at end of file
diff --git a/Airfoils/NACA_4412.txt b/Airfoils/NACA_4412.txt
new file mode 100644
index 0000000..efc9cf5
--- /dev/null
+++ b/Airfoils/NACA_4412.txt
@@ -0,0 +1,40 @@
+NACA 4412
+ 18. 18.
+
+ 0.000000 0.000000
+ 0.012500 0.024400
+ 0.025000 0.033900
+ 0.050000 0.047300
+ 0.075000 0.057600
+ 0.100000 0.065900
+ 0.150000 0.078900
+ 0.200000 0.088000
+ 0.250000 0.094100
+ 0.300000 0.097600
+ 0.400000 0.098000
+ 0.500000 0.091900
+ 0.600000 0.081400
+ 0.700000 0.066900
+ 0.800000 0.048900
+ 0.900000 0.027100
+ 0.950000 0.014700
+ 1.000000 0.001300
+
+ 0.000000 0.000000
+ 0.012500 -0.014300
+ 0.025000 -0.019500
+ 0.050000 -0.024900
+ 0.075000 -0.027400
+ 0.100000 -0.028600
+ 0.150000 -0.028800
+ 0.200000 -0.027400
+ 0.250000 -0.025000
+ 0.300000 -0.022600
+ 0.400000 -0.018000
+ 0.500000 -0.014000
+ 0.600000 -0.010000
+ 0.700000 -0.006500
+ 0.800000 -0.003900
+ 0.900000 -0.002200
+ 0.950000 -0.001600
+ 1.000000 -0.001300
diff --git a/Airfoils/NACA_63_412.txt b/Airfoils/NACA_63_412.txt
new file mode 100644
index 0000000..d04c4b5
--- /dev/null
+++ b/Airfoils/NACA_63_412.txt
@@ -0,0 +1,56 @@
+NACA 63-412 AIRFOIL
+ 26. 26.
+
+ 0.000000 0.000000
+ 0.003360 0.010710
+ 0.005670 0.013200
+ 0.010410 0.017190
+ 0.022570 0.024600
+ 0.047270 0.035440
+ 0.072180 0.043790
+ 0.097180 0.050630
+ 0.147350 0.061380
+ 0.197650 0.069290
+ 0.248000 0.074990
+ 0.298400 0.078720
+ 0.348820 0.080590
+ 0.399240 0.080620
+ 0.449640 0.078940
+ 0.500000 0.075670
+ 0.550310 0.071250
+ 0.600570 0.065620
+ 0.650760 0.058990
+ 0.700870 0.051530
+ 0.750890 0.043440
+ 0.800840 0.034920
+ 0.850700 0.026180
+ 0.900490 0.017390
+ 0.950230 0.008810
+ 1.000000 0.000000
+
+ 0.000000 0.000000
+ 0.006640 -0.008710
+ 0.009330 -0.010400
+ 0.014590 -0.012910
+ 0.027430 -0.017160
+ 0.052730 -0.022800
+ 0.077820 -0.026850
+ 0.102820 -0.029950
+ 0.152650 -0.034460
+ 0.202350 -0.037450
+ 0.252000 -0.039190
+ 0.301600 -0.039840
+ 0.351118 -0.039390
+ 0.400760 -0.037780
+ 0.450350 -0.035140
+ 0.500000 -0.031640
+ 0.549690 -0.027450
+ 0.599430 -0.022780
+ 0.649240 -0.017990
+ 0.699130 -0.012650
+ 0.749110 -0.007640
+ 0.799160 -0.003080
+ 0.849300 0.000740
+ 0.899510 0.003290
+ 0.949770 0.003300
+ 1.000000 0.000000
diff --git a/Airfoils/Polars/Clark_y_polar_Re_100000.txt b/Airfoils/Polars/Clark_y_polar_Re_100000.txt
new file mode 100644
index 0000000..81ec5f2
--- /dev/null
+++ b/Airfoils/Polars/Clark_y_polar_Re_100000.txt
@@ -0,0 +1,125 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: CLARK Y AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.000 -0.3474 0.10140 0.09638 -0.0367 1.0000 0.1273
+ -8.750 -0.3809 0.10037 0.09553 -0.0393 1.0000 0.1308
+ -8.500 -0.4214 0.09979 0.09515 -0.0389 1.0000 0.1313
+ -8.250 -0.3699 0.09349 0.08874 -0.0357 1.0000 0.1351
+ -8.000 -0.3685 0.09146 0.08675 -0.0335 1.0000 0.1389
+ -7.750 -0.3862 0.09004 0.08545 -0.0311 1.0000 0.1423
+ -7.500 -0.4177 0.08922 0.08477 -0.0278 1.0000 0.1440
+ -7.250 -0.4616 0.08774 0.08342 -0.0294 1.0000 0.1460
+ -7.000 -0.4818 0.08423 0.07996 -0.0300 1.0000 0.1476
+ -6.750 -0.4620 0.08250 0.07829 -0.0229 1.0000 0.1509
+ -6.500 -0.4660 0.08065 0.07647 -0.0207 1.0000 0.1552
+ -6.250 -0.4964 0.07643 0.07214 -0.0280 1.0000 0.1629
+ -6.000 -0.4869 0.07438 0.07022 -0.0227 1.0000 0.1653
+ -5.750 -0.4833 0.07245 0.06830 -0.0203 1.0000 0.1693
+ -5.500 -0.4804 0.06812 0.06384 -0.0253 0.9986 0.1800
+ -5.250 -0.4492 0.06395 0.05946 -0.0321 0.9922 0.1948
+ -5.000 -0.4217 0.06129 0.05687 -0.0322 0.9869 0.2008
+ -4.750 -0.3727 0.04267 0.03658 -0.0508 0.9819 0.1207
+ -4.500 -0.3344 0.03587 0.02837 -0.0541 0.9768 0.1009
+ -4.250 -0.2952 0.03304 0.02513 -0.0572 0.9721 0.0999
+ -4.000 -0.2584 0.03133 0.02290 -0.0593 0.9662 0.1012
+ -3.750 -0.2209 0.02948 0.02067 -0.0615 0.9603 0.1027
+ -3.500 -0.1775 0.02785 0.01885 -0.0648 0.9562 0.1046
+ -3.250 -0.1467 0.02687 0.01776 -0.0656 0.9484 0.1073
+ -3.000 -0.1049 0.02601 0.01674 -0.0683 0.9429 0.1129
+ -2.750 -0.0711 0.02520 0.01588 -0.0696 0.9359 0.1198
+ -2.500 -0.0332 0.02449 0.01516 -0.0716 0.9294 0.1296
+ -2.250 0.0075 0.02360 0.01442 -0.0740 0.9242 0.1486
+ -2.000 0.0372 0.02287 0.01400 -0.0746 0.9157 0.1943
+ -1.750 0.0798 0.02169 0.01361 -0.0777 0.9111 0.3339
+ -1.500 0.1048 0.02083 0.01363 -0.0774 0.9016 0.5186
+ -1.250 0.1696 0.01958 0.01355 -0.0816 0.8979 1.0000
+ -1.000 0.2100 0.01947 0.01315 -0.0838 0.8869 1.0000
+ -0.750 0.2685 0.01902 0.01241 -0.0888 0.8803 1.0000
+ -0.500 0.2971 0.01899 0.01222 -0.0887 0.8684 1.0000
+ -0.250 0.3365 0.01882 0.01188 -0.0905 0.8612 1.0000
+ 0.000 0.3674 0.01875 0.01169 -0.0908 0.8514 1.0000
+ 0.250 0.3960 0.01875 0.01157 -0.0906 0.8414 1.0000
+ 0.500 0.4341 0.01846 0.01117 -0.0918 0.8343 1.0000
+ 0.750 0.4587 0.01851 0.01114 -0.0909 0.8228 1.0000
+ 1.000 0.4950 0.01820 0.01074 -0.0917 0.8154 1.0000
+ 1.250 0.5218 0.01814 0.01061 -0.0910 0.8041 1.0000
+ 1.500 0.5474 0.01813 0.01054 -0.0901 0.7925 1.0000
+ 1.750 0.5797 0.01788 0.01022 -0.0901 0.7833 1.0000
+ 2.000 0.6085 0.01772 0.01001 -0.0896 0.7720 1.0000
+ 2.250 0.6339 0.01769 0.00993 -0.0886 0.7589 1.0000
+ 2.500 0.6607 0.01761 0.00981 -0.0877 0.7459 1.0000
+ 2.750 0.6885 0.01750 0.00966 -0.0870 0.7330 1.0000
+ 3.000 0.7166 0.01739 0.00949 -0.0864 0.7195 1.0000
+ 3.250 0.7438 0.01730 0.00936 -0.0855 0.7046 1.0000
+ 3.500 0.7699 0.01727 0.00928 -0.0846 0.6881 1.0000
+ 3.750 0.7954 0.01729 0.00925 -0.0836 0.6706 1.0000
+ 4.000 0.8211 0.01732 0.00924 -0.0826 0.6524 1.0000
+ 4.250 0.8471 0.01737 0.00921 -0.0817 0.6336 1.0000
+ 4.500 0.8738 0.01746 0.00919 -0.0809 0.6150 1.0000
+ 4.750 0.8972 0.01771 0.00940 -0.0798 0.5947 1.0000
+ 5.000 0.9213 0.01798 0.00962 -0.0788 0.5753 1.0000
+ 5.250 0.9460 0.01829 0.00985 -0.0780 0.5574 1.0000
+ 5.500 0.9706 0.01863 0.01012 -0.0772 0.5405 1.0000
+ 5.750 0.9949 0.01901 0.01045 -0.0763 0.5244 1.0000
+ 6.000 1.0188 0.01941 0.01081 -0.0755 0.5089 1.0000
+ 6.250 1.0422 0.01978 0.01114 -0.0745 0.4930 1.0000
+ 6.500 1.0650 0.02014 0.01146 -0.0735 0.4768 1.0000
+ 6.750 1.0873 0.02051 0.01180 -0.0724 0.4611 1.0000
+ 7.000 1.1094 0.02094 0.01224 -0.0713 0.4465 1.0000
+ 7.250 1.1315 0.02142 0.01276 -0.0703 0.4331 1.0000
+ 7.500 1.1536 0.02191 0.01329 -0.0693 0.4197 1.0000
+ 7.750 1.1755 0.02241 0.01382 -0.0683 0.4065 1.0000
+ 8.000 1.1971 0.02293 0.01435 -0.0672 0.3929 1.0000
+ 8.250 1.2182 0.02347 0.01492 -0.0661 0.3789 1.0000
+ 8.500 1.2388 0.02405 0.01551 -0.0649 0.3645 1.0000
+ 8.750 1.2581 0.02466 0.01615 -0.0635 0.3492 1.0000
+ 9.000 1.2755 0.02531 0.01685 -0.0619 0.3326 1.0000
+ 9.250 1.2908 0.02599 0.01760 -0.0599 0.3146 1.0000
+ 9.500 1.3041 0.02665 0.01830 -0.0576 0.2954 1.0000
+ 9.750 1.3159 0.02731 0.01893 -0.0552 0.2763 1.0000
+ 10.000 1.3249 0.02802 0.01965 -0.0524 0.2575 1.0000
+ 10.250 1.3309 0.02876 0.02046 -0.0492 0.2393 1.0000
+ 10.500 1.3363 0.02954 0.02127 -0.0461 0.2233 1.0000
+ 10.750 1.3413 0.03044 0.02219 -0.0430 0.2094 1.0000
+ 11.000 1.3473 0.03148 0.02322 -0.0403 0.1972 1.0000
+ 11.250 1.3539 0.03261 0.02428 -0.0378 0.1865 1.0000
+ 11.500 1.3594 0.03385 0.02564 -0.0355 0.1760 1.0000
+ 11.750 1.3649 0.03520 0.02707 -0.0333 0.1663 1.0000
+ 12.000 1.3683 0.03666 0.02850 -0.0311 0.1570 1.0000
+ 12.250 1.3687 0.03826 0.03022 -0.0290 0.1473 1.0000
+ 12.500 1.3694 0.04007 0.03217 -0.0271 0.1381 1.0000
+ 12.750 1.3687 0.04207 0.03420 -0.0254 0.1294 1.0000
+ 13.000 1.3644 0.04442 0.03666 -0.0238 0.1196 1.0000
+ 13.250 1.3568 0.04727 0.03965 -0.0225 0.1086 1.0000
+ 13.500 1.3456 0.05070 0.04316 -0.0215 0.0969 1.0000
+ 13.750 1.3323 0.05464 0.04712 -0.0208 0.0856 1.0000
+ 14.000 1.3196 0.05873 0.05120 -0.0202 0.0761 1.0000
+ 14.250 1.3103 0.06255 0.05499 -0.0199 0.0694 1.0000
+ 14.500 1.3050 0.06622 0.05879 -0.0194 0.0635 1.0000
+ 14.750 1.3011 0.06963 0.06219 -0.0193 0.0596 1.0000
+ 15.000 1.2996 0.07289 0.06556 -0.0188 0.0562 1.0000
+ 15.250 1.2970 0.07647 0.06929 -0.0189 0.0533 1.0000
+ 15.500 1.2952 0.07985 0.07272 -0.0191 0.0508 1.0000
+ 15.750 1.2975 0.08261 0.07541 -0.0187 0.0485 1.0000
+ 16.000 1.2921 0.08687 0.07995 -0.0194 0.0471 1.0000
+ 16.250 1.2873 0.09111 0.08441 -0.0200 0.0458 1.0000
+ 16.500 1.2822 0.09546 0.08896 -0.0209 0.0448 1.0000
+ 16.750 1.2755 0.10014 0.09382 -0.0221 0.0440 1.0000
+ 17.000 1.2673 0.10518 0.09905 -0.0238 0.0434 1.0000
+ 17.250 1.2568 0.11077 0.10483 -0.0260 0.0430 1.0000
+ 17.500 1.2419 0.11739 0.11167 -0.0291 0.0428 1.0000
+ 17.750 1.2181 0.12614 0.12069 -0.0340 0.0430 1.0000
+ 18.000 1.1716 0.14084 0.13576 -0.0436 0.0444 1.0000
+ 18.250 1.0657 0.17534 0.17052 -0.0663 0.0498 1.0000
+ 18.500 1.0634 0.18080 0.17597 -0.0691 0.0489 1.0000
+ 18.750 0.9950 0.21955 0.21443 -0.0905 0.0654 1.0000
+ 19.000 1.0031 0.22395 0.21884 -0.0919 0.0673 1.0000
diff --git a/Airfoils/Polars/Clark_y_polar_Re_1000000.txt b/Airfoils/Polars/Clark_y_polar_Re_1000000.txt
new file mode 100644
index 0000000..7bee3a7
--- /dev/null
+++ b/Airfoils/Polars/Clark_y_polar_Re_1000000.txt
@@ -0,0 +1,142 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: CLARK Y AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -14.750 -0.3156 0.14464 0.14298 -0.0315 1.0000 0.0162
+ -14.500 -1.0248 0.03605 0.03333 -0.0892 1.0000 0.0164
+ -14.250 -1.0371 0.03284 0.02999 -0.0896 1.0000 0.0166
+ -14.000 -1.0347 0.03166 0.02877 -0.0877 1.0000 0.0169
+ -13.750 -1.0214 0.03117 0.02828 -0.0864 1.0000 0.0171
+ -13.500 -1.0070 0.03090 0.02801 -0.0848 1.0000 0.0174
+ -13.250 -0.9948 0.03059 0.02769 -0.0828 1.0000 0.0176
+ -13.000 -0.9847 0.03027 0.02735 -0.0804 1.0000 0.0179
+ -12.750 -0.9771 0.02987 0.02691 -0.0777 1.0000 0.0182
+ -12.500 -0.9635 0.02930 0.02628 -0.0762 0.9996 0.0186
+ -12.250 -0.9347 0.02845 0.02533 -0.0778 0.9981 0.0192
+ -12.000 -0.9045 0.02781 0.02458 -0.0793 0.9966 0.0196
+ -11.750 -0.8800 0.02584 0.02237 -0.0812 0.9947 0.0201
+ -11.500 -0.8522 0.02475 0.02122 -0.0827 0.9932 0.0206
+ -11.250 -0.8218 0.02455 0.02103 -0.0837 0.9910 0.0210
+ -11.000 -0.7899 0.02444 0.02093 -0.0850 0.9891 0.0214
+ -10.750 -0.7579 0.02416 0.02062 -0.0864 0.9873 0.0219
+ -10.500 -0.7261 0.02364 0.02003 -0.0879 0.9859 0.0225
+ -10.250 -0.6946 0.02280 0.01907 -0.0896 0.9846 0.0232
+ -10.000 -0.6628 0.02209 0.01823 -0.0911 0.9833 0.0237
+ -9.750 -0.6348 0.02188 0.01793 -0.0915 0.9791 0.0241
+ -9.500 -0.6138 0.01942 0.01521 -0.0919 0.9744 0.0249
+ -9.250 -0.5853 0.01872 0.01449 -0.0926 0.9712 0.0254
+ -9.000 -0.5602 0.01823 0.01395 -0.0924 0.9655 0.0259
+ -8.750 -0.5348 0.01779 0.01345 -0.0921 0.9596 0.0264
+ -8.500 -0.5089 0.01733 0.01294 -0.0920 0.9542 0.0270
+ -8.250 -0.4853 0.01678 0.01229 -0.0913 0.9467 0.0276
+ -8.000 -0.4604 0.01615 0.01154 -0.0908 0.9405 0.0280
+ -7.750 -0.4358 0.01569 0.01100 -0.0903 0.9326 0.0286
+ -7.500 -0.4099 0.01537 0.01058 -0.0899 0.9259 0.0290
+ -7.250 -0.3842 0.01500 0.01011 -0.0895 0.9180 0.0293
+ -7.000 -0.3599 0.01406 0.00901 -0.0889 0.9107 0.0297
+ -6.750 -0.3365 0.01283 0.00769 -0.0884 0.9026 0.0304
+ -6.500 -0.3112 0.01219 0.00697 -0.0879 0.8951 0.0308
+ -6.250 -0.2852 0.01169 0.00643 -0.0876 0.8870 0.0313
+ -5.750 -0.2325 0.01090 0.00553 -0.0870 0.8711 0.0322
+ -5.500 -0.2059 0.01057 0.00512 -0.0867 0.8633 0.0328
+ -5.250 -0.1791 0.01026 0.00476 -0.0864 0.8543 0.0334
+ -5.000 -0.1523 0.00999 0.00443 -0.0862 0.8450 0.0341
+ -4.750 -0.1255 0.00972 0.00409 -0.0859 0.8348 0.0346
+ -4.500 -0.0985 0.00947 0.00378 -0.0856 0.8253 0.0350
+ -4.250 -0.0714 0.00926 0.00350 -0.0854 0.8165 0.0354
+ -4.000 -0.0440 0.00908 0.00327 -0.0852 0.8067 0.0357
+ -3.750 -0.0175 0.00870 0.00281 -0.0849 0.7972 0.0369
+ -3.500 0.0096 0.00847 0.00252 -0.0846 0.7879 0.0381
+ -3.250 0.0371 0.00830 0.00232 -0.0845 0.7787 0.0394
+ -3.000 0.0645 0.00818 0.00214 -0.0843 0.7689 0.0407
+ -2.750 0.0922 0.00806 0.00199 -0.0841 0.7587 0.0422
+ -2.500 0.1198 0.00796 0.00184 -0.0840 0.7486 0.0440
+ -2.250 0.1471 0.00784 0.00170 -0.0838 0.7388 0.0488
+ -2.000 0.1747 0.00767 0.00160 -0.0837 0.7297 0.0659
+ -1.750 0.2020 0.00752 0.00151 -0.0835 0.7209 0.0930
+ -1.500 0.2293 0.00737 0.00143 -0.0834 0.7112 0.1210
+ -1.250 0.2564 0.00718 0.00137 -0.0833 0.7019 0.1672
+ -1.000 0.2833 0.00701 0.00133 -0.0831 0.6927 0.2229
+ -0.750 0.3107 0.00686 0.00131 -0.0830 0.6834 0.2696
+ -0.500 0.3380 0.00677 0.00130 -0.0829 0.6741 0.3090
+ -0.250 0.3651 0.00664 0.00128 -0.0828 0.6638 0.3575
+ 0.000 0.3918 0.00645 0.00128 -0.0826 0.6533 0.4324
+ 0.250 0.4171 0.00614 0.00130 -0.0822 0.6422 0.5591
+ 0.500 0.4408 0.00578 0.00134 -0.0814 0.6307 0.7081
+ 0.750 0.4632 0.00547 0.00141 -0.0800 0.6194 0.8388
+ 1.000 0.4866 0.00538 0.00150 -0.0786 0.6079 0.9226
+ 1.250 0.5201 0.00546 0.00157 -0.0796 0.5950 0.9665
+ 1.500 0.5600 0.00558 0.00162 -0.0821 0.5805 0.9856
+ 1.750 0.6022 0.00572 0.00167 -0.0853 0.5639 0.9946
+ 2.000 0.6464 0.00585 0.00171 -0.0889 0.5456 1.0000
+ 2.250 0.6708 0.00598 0.00176 -0.0882 0.5297 1.0000
+ 2.500 0.6949 0.00612 0.00182 -0.0874 0.5128 1.0000
+ 2.750 0.7187 0.00628 0.00189 -0.0866 0.4935 1.0000
+ 3.000 0.7423 0.00647 0.00198 -0.0857 0.4720 1.0000
+ 3.250 0.7650 0.00672 0.00209 -0.0847 0.4436 1.0000
+ 3.500 0.7870 0.00702 0.00223 -0.0836 0.4086 1.0000
+ 3.750 0.8091 0.00733 0.00238 -0.0825 0.3791 1.0000
+ 4.000 0.8326 0.00757 0.00252 -0.0816 0.3604 1.0000
+ 4.250 0.8567 0.00779 0.00267 -0.0809 0.3459 1.0000
+ 4.500 0.8810 0.00801 0.00282 -0.0802 0.3335 1.0000
+ 4.750 0.9060 0.00819 0.00296 -0.0797 0.3242 1.0000
+ 5.000 0.9314 0.00836 0.00311 -0.0793 0.3174 1.0000
+ 5.250 0.9570 0.00852 0.00325 -0.0788 0.3109 1.0000
+ 5.500 0.9822 0.00872 0.00342 -0.0784 0.3038 1.0000
+ 5.750 1.0081 0.00886 0.00357 -0.0780 0.2974 1.0000
+ 6.000 1.0330 0.00909 0.00375 -0.0775 0.2894 1.0000
+ 6.250 1.0590 0.00923 0.00390 -0.0772 0.2817 1.0000
+ 6.500 1.0840 0.00945 0.00408 -0.0768 0.2718 1.0000
+ 6.750 1.1090 0.00966 0.00426 -0.0763 0.2614 1.0000
+ 7.000 1.1341 0.00988 0.00445 -0.0759 0.2500 1.0000
+ 7.250 1.1584 0.01014 0.00466 -0.0754 0.2364 1.0000
+ 7.500 1.1816 0.01048 0.00491 -0.0747 0.2176 1.0000
+ 7.750 1.2028 0.01097 0.00525 -0.0737 0.1896 1.0000
+ 8.250 1.2407 0.01224 0.00619 -0.0710 0.1327 1.0000
+ 8.500 1.2606 0.01279 0.00663 -0.0699 0.1136 1.0000
+ 8.750 1.2805 0.01333 0.00708 -0.0687 0.0988 1.0000
+ 9.000 1.3008 0.01381 0.00752 -0.0676 0.0880 1.0000
+ 9.250 1.3208 0.01430 0.00796 -0.0665 0.0790 1.0000
+ 9.500 1.3401 0.01481 0.00841 -0.0653 0.0698 1.0000
+ 9.750 1.3571 0.01539 0.00893 -0.0636 0.0573 1.0000
+ 10.000 1.3618 0.01664 0.00992 -0.0602 0.0254 1.0000
+ 10.250 1.3715 0.01761 0.01085 -0.0575 0.0168 1.0000
+ 10.500 1.3856 0.01831 0.01159 -0.0555 0.0149 1.0000
+ 10.750 1.4007 0.01897 0.01230 -0.0538 0.0139 1.0000
+ 11.000 1.4136 0.01976 0.01313 -0.0519 0.0130 1.0000
+ 11.250 1.4236 0.02077 0.01422 -0.0497 0.0121 1.0000
+ 11.500 1.4375 0.02154 0.01504 -0.0481 0.0118 1.0000
+ 11.750 1.4503 0.02239 0.01595 -0.0464 0.0113 1.0000
+ 12.000 1.4619 0.02335 0.01697 -0.0447 0.0109 1.0000
+ 12.250 1.4724 0.02441 0.01809 -0.0431 0.0105 1.0000
+ 12.500 1.4810 0.02565 0.01939 -0.0413 0.0101 1.0000
+ 12.750 1.4847 0.02733 0.02115 -0.0393 0.0097 1.0000
+ 13.000 1.4888 0.02903 0.02295 -0.0375 0.0094 1.0000
+ 13.250 1.4985 0.03033 0.02432 -0.0363 0.0092 1.0000
+ 13.500 1.5063 0.03185 0.02591 -0.0352 0.0089 1.0000
+ 13.750 1.5122 0.03356 0.02770 -0.0340 0.0087 1.0000
+ 14.000 1.5169 0.03546 0.02968 -0.0329 0.0085 1.0000
+ 14.250 1.5205 0.03753 0.03183 -0.0320 0.0083 1.0000
+ 14.500 1.5231 0.03975 0.03412 -0.0312 0.0081 1.0000
+ 14.750 1.5240 0.04222 0.03668 -0.0305 0.0079 1.0000
+ 15.000 1.5231 0.04500 0.03955 -0.0299 0.0078 1.0000
+ 15.250 1.5193 0.04819 0.04284 -0.0296 0.0076 1.0000
+ 15.500 1.5118 0.05198 0.04673 -0.0295 0.0075 1.0000
+ 15.750 1.5001 0.05647 0.05134 -0.0298 0.0074 1.0000
+ 16.000 1.4855 0.06151 0.05651 -0.0304 0.0073 1.0000
+ 16.250 1.4724 0.06656 0.06168 -0.0312 0.0072 1.0000
+ 16.500 1.4677 0.07059 0.06582 -0.0320 0.0072 1.0000
+ 16.750 1.4609 0.07500 0.07033 -0.0331 0.0071 1.0000
+ 17.000 1.4527 0.07967 0.07511 -0.0342 0.0071 1.0000
+ 17.250 1.4435 0.08463 0.08017 -0.0356 0.0070 1.0000
+ 17.500 1.4336 0.08975 0.08540 -0.0372 0.0069 1.0000
+ 17.750 1.4229 0.09504 0.09079 -0.0388 0.0069 1.0000
+ 18.000 1.4117 0.10050 0.09635 -0.0407 0.0068 1.0000
diff --git a/Airfoils/Polars/Clark_y_polar_Re_200000.txt b/Airfoils/Polars/Clark_y_polar_Re_200000.txt
new file mode 100644
index 0000000..7b2cd72
--- /dev/null
+++ b/Airfoils/Polars/Clark_y_polar_Re_200000.txt
@@ -0,0 +1,114 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: CLARK Y AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.000 -0.4082 0.09344 0.09005 -0.0468 1.0000 0.0739
+ -8.750 -0.4391 0.09151 0.08824 -0.0450 1.0000 0.0740
+ -8.500 -0.4670 0.08892 0.08571 -0.0443 1.0000 0.0741
+ -8.250 -0.4772 0.08456 0.08142 -0.0411 1.0000 0.0748
+ -8.000 -0.4670 0.08349 0.08039 -0.0352 1.0000 0.0758
+ -7.750 -0.4730 0.08209 0.07903 -0.0319 1.0000 0.0765
+ -7.500 -0.4528 0.07915 0.07607 -0.0338 0.9969 0.0785
+ -7.250 -0.4291 0.07367 0.07053 -0.0427 0.9910 0.0830
+ -7.000 -0.4158 0.06100 0.05751 -0.0619 0.9808 0.0895
+ -6.750 -0.3903 0.05905 0.05559 -0.0625 0.9760 0.0913
+ -6.500 -0.3615 0.05546 0.05190 -0.0668 0.9718 0.0951
+ -6.250 -0.3450 0.04808 0.04400 -0.0741 0.9623 0.1041
+ -6.000 -0.3240 0.03416 0.02886 -0.0789 0.9571 0.0704
+ -5.750 -0.3036 0.02800 0.02149 -0.0788 0.9496 0.0631
+ -5.500 -0.2702 0.02569 0.01892 -0.0806 0.9455 0.0638
+ -5.250 -0.2318 0.02391 0.01690 -0.0831 0.9428 0.0648
+ -5.000 -0.1910 0.02223 0.01497 -0.0858 0.9410 0.0653
+ -4.750 -0.1654 0.02118 0.01377 -0.0854 0.9327 0.0660
+ -4.500 -0.1273 0.02005 0.01249 -0.0874 0.9294 0.0673
+ -4.250 -0.0881 0.01908 0.01142 -0.0895 0.9268 0.0695
+ -4.000 -0.0604 0.01847 0.01072 -0.0893 0.9193 0.0719
+ -3.750 -0.0261 0.01778 0.00991 -0.0903 0.9145 0.0739
+ -3.500 0.0091 0.01659 0.00877 -0.0916 0.9112 0.0772
+ -3.250 0.0342 0.01609 0.00829 -0.0910 0.9027 0.0811
+ -3.000 0.0668 0.01553 0.00768 -0.0915 0.8975 0.0871
+ -2.750 0.0940 0.01485 0.00710 -0.0912 0.8900 0.0985
+ -2.500 0.1221 0.01403 0.00642 -0.0910 0.8825 0.1304
+ -2.250 0.1479 0.01328 0.00605 -0.0905 0.8741 0.2081
+ -2.000 0.1751 0.01271 0.00578 -0.0901 0.8662 0.2941
+ -1.750 0.1998 0.01226 0.00560 -0.0894 0.8568 0.3729
+ -1.500 0.2249 0.01156 0.00536 -0.0885 0.8491 0.4994
+ -1.250 0.2429 0.01087 0.00534 -0.0860 0.8383 0.6774
+ -1.000 0.2709 0.01036 0.00534 -0.0841 0.8309 0.8768
+ -0.750 0.3315 0.01030 0.00523 -0.0899 0.8231 0.9730
+ -0.500 0.3953 0.01014 0.00489 -0.0970 0.8170 1.0000
+ -0.250 0.4182 0.01015 0.00480 -0.0960 0.8055 1.0000
+ 0.000 0.4424 0.01016 0.00470 -0.0952 0.7954 1.0000
+ 0.250 0.4680 0.01015 0.00455 -0.0945 0.7864 1.0000
+ 0.500 0.4915 0.01020 0.00453 -0.0935 0.7750 1.0000
+ 0.750 0.5161 0.01024 0.00447 -0.0927 0.7644 1.0000
+ 1.000 0.5419 0.01026 0.00437 -0.0920 0.7544 1.0000
+ 1.250 0.5656 0.01032 0.00437 -0.0911 0.7423 1.0000
+ 1.500 0.5899 0.01039 0.00437 -0.0902 0.7308 1.0000
+ 1.750 0.6149 0.01046 0.00436 -0.0895 0.7196 1.0000
+ 2.000 0.6402 0.01053 0.00433 -0.0887 0.7080 1.0000
+ 2.250 0.6641 0.01061 0.00436 -0.0877 0.6945 1.0000
+ 2.500 0.6883 0.01071 0.00440 -0.0868 0.6807 1.0000
+ 2.750 0.7126 0.01081 0.00444 -0.0859 0.6666 1.0000
+ 3.000 0.7370 0.01092 0.00449 -0.0851 0.6519 1.0000
+ 3.250 0.7612 0.01104 0.00456 -0.0842 0.6360 1.0000
+ 3.500 0.7852 0.01117 0.00462 -0.0832 0.6187 1.0000
+ 3.750 0.8091 0.01133 0.00468 -0.0823 0.6003 1.0000
+ 4.000 0.8324 0.01152 0.00480 -0.0812 0.5801 1.0000
+ 4.250 0.8553 0.01174 0.00494 -0.0801 0.5580 1.0000
+ 4.500 0.8781 0.01201 0.00508 -0.0790 0.5358 1.0000
+ 4.750 0.9005 0.01231 0.00530 -0.0779 0.5124 1.0000
+ 5.000 0.9227 0.01266 0.00551 -0.0768 0.4911 1.0000
+ 5.250 0.9450 0.01300 0.00576 -0.0758 0.4700 1.0000
+ 5.500 0.9671 0.01335 0.00602 -0.0747 0.4502 1.0000
+ 5.750 0.9892 0.01371 0.00630 -0.0737 0.4321 1.0000
+ 6.000 1.0116 0.01408 0.00660 -0.0727 0.4164 1.0000
+ 6.250 1.0345 0.01444 0.00693 -0.0719 0.4033 1.0000
+ 6.500 1.0576 0.01483 0.00730 -0.0711 0.3917 1.0000
+ 6.750 1.0804 0.01525 0.00766 -0.0703 0.3808 1.0000
+ 7.000 1.1032 0.01563 0.00805 -0.0695 0.3699 1.0000
+ 7.250 1.1262 0.01602 0.00848 -0.0687 0.3598 1.0000
+ 7.500 1.1486 0.01648 0.00890 -0.0679 0.3499 1.0000
+ 7.750 1.1705 0.01686 0.00933 -0.0669 0.3391 1.0000
+ 8.000 1.1918 0.01726 0.00978 -0.0659 0.3278 1.0000
+ 8.250 1.2120 0.01769 0.01023 -0.0647 0.3155 1.0000
+ 8.500 1.2310 0.01811 0.01066 -0.0634 0.3022 1.0000
+ 8.750 1.2490 0.01852 0.01110 -0.0618 0.2876 1.0000
+ 9.000 1.2663 0.01891 0.01155 -0.0602 0.2720 1.0000
+ 9.250 1.2826 0.01932 0.01203 -0.0584 0.2533 1.0000
+ 9.500 1.2960 0.01984 0.01253 -0.0563 0.2317 1.0000
+ 9.750 1.3072 0.02048 0.01314 -0.0538 0.2074 1.0000
+ 10.000 1.3158 0.02138 0.01392 -0.0512 0.1848 1.0000
+ 10.250 1.3246 0.02239 0.01486 -0.0487 0.1663 1.0000
+ 10.500 1.3333 0.02347 0.01590 -0.0464 0.1524 1.0000
+ 10.750 1.3420 0.02458 0.01699 -0.0442 0.1408 1.0000
+ 11.000 1.3494 0.02581 0.01818 -0.0421 0.1314 1.0000
+ 11.250 1.3609 0.02682 0.01927 -0.0404 0.1222 1.0000
+ 11.500 1.3704 0.02799 0.02047 -0.0387 0.1126 1.0000
+ 11.750 1.3775 0.02935 0.02184 -0.0370 0.1040 1.0000
+ 12.000 1.3859 0.03069 0.02323 -0.0355 0.0937 1.0000
+ 12.250 1.3920 0.03227 0.02483 -0.0339 0.0811 1.0000
+ 12.500 1.3965 0.03407 0.02662 -0.0324 0.0629 1.0000
+ 12.750 1.3940 0.03655 0.02899 -0.0307 0.0468 1.0000
+ 13.000 1.3907 0.03921 0.03163 -0.0292 0.0389 1.0000
+ 13.250 1.3873 0.04198 0.03444 -0.0279 0.0350 1.0000
+ 13.500 1.3833 0.04495 0.03749 -0.0269 0.0328 1.0000
+ 13.750 1.3802 0.04793 0.04060 -0.0262 0.0311 1.0000
+ 14.000 1.3753 0.05122 0.04401 -0.0257 0.0299 1.0000
+ 14.250 1.3681 0.05491 0.04780 -0.0255 0.0289 1.0000
+ 14.500 1.3581 0.05908 0.05208 -0.0256 0.0281 1.0000
+ 14.750 1.3508 0.06309 0.05620 -0.0259 0.0275 1.0000
+ 15.000 1.3451 0.06702 0.06027 -0.0263 0.0268 1.0000
+ 15.250 1.3390 0.07109 0.06447 -0.0269 0.0262 1.0000
+ 15.500 1.3327 0.07526 0.06876 -0.0276 0.0255 1.0000
+ 15.750 1.3262 0.07955 0.07316 -0.0285 0.0250 1.0000
+ 16.000 1.3199 0.08387 0.07756 -0.0294 0.0244 1.0000
+ 16.250 1.3136 0.08815 0.08192 -0.0303 0.0239 1.0000
diff --git a/Airfoils/Polars/Clark_y_polar_Re_50000.txt b/Airfoils/Polars/Clark_y_polar_Re_50000.txt
new file mode 100644
index 0000000..dbb62f8
--- /dev/null
+++ b/Airfoils/Polars/Clark_y_polar_Re_50000.txt
@@ -0,0 +1,124 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: CLARK Y AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -10.750 -0.3683 0.12749 0.11999 -0.0321 1.0000 0.1294
+ -10.500 -0.3799 0.12552 0.11811 -0.0349 1.0000 0.1311
+ -10.250 -0.3871 0.12285 0.11552 -0.0371 1.0000 0.1315
+ -10.000 -0.3561 0.11742 0.11002 -0.0338 1.0000 0.1362
+ -9.750 -0.3865 0.10888 0.10159 -0.0427 1.0000 0.0878
+ -9.500 -0.3694 0.10555 0.09820 -0.0406 1.0000 0.0851
+ -9.250 -0.3674 0.10205 0.09476 -0.0408 1.0000 0.0829
+ -9.000 -0.3705 0.09846 0.09122 -0.0415 1.0000 0.0808
+ -8.750 -0.3776 0.09491 0.08777 -0.0421 1.0000 0.0795
+ -8.250 -0.4394 0.08464 0.07783 -0.0474 1.0000 0.0740
+ -8.000 -0.4506 0.08108 0.07432 -0.0472 1.0000 0.0739
+ -7.750 -0.4626 0.07740 0.07067 -0.0469 1.0000 0.0738
+ -7.500 -0.4743 0.07361 0.06686 -0.0465 1.0000 0.0738
+ -7.250 -0.4843 0.06974 0.06291 -0.0461 1.0000 0.0738
+ -7.000 -0.4913 0.06583 0.05888 -0.0457 1.0000 0.0738
+ -6.750 -0.4944 0.06200 0.05487 -0.0453 1.0000 0.0737
+ -6.500 -0.4936 0.05827 0.05092 -0.0448 1.0000 0.0736
+ -6.250 -0.4896 0.05457 0.04694 -0.0445 1.0000 0.0735
+ -6.000 -0.4821 0.05100 0.04301 -0.0443 1.0000 0.0736
+ -5.750 -0.4540 0.04675 0.03823 -0.0478 0.9942 0.0746
+ -5.500 -0.4248 0.04466 0.03605 -0.0499 0.9882 0.0770
+ -5.250 -0.3926 0.04209 0.03312 -0.0527 0.9827 0.0793
+ -5.000 -0.3627 0.03934 0.02988 -0.0547 0.9761 0.0807
+ -4.750 -0.3280 0.03687 0.02688 -0.0571 0.9709 0.0823
+ -4.500 -0.2966 0.03491 0.02444 -0.0586 0.9642 0.0845
+ -4.250 -0.2616 0.03329 0.02236 -0.0605 0.9584 0.0884
+ -4.000 -0.2293 0.03213 0.02111 -0.0619 0.9522 0.0922
+ -3.750 -0.1962 0.03099 0.01975 -0.0632 0.9456 0.0959
+ -3.500 -0.1577 0.02990 0.01834 -0.0652 0.9408 0.1008
+ -3.250 -0.1300 0.02912 0.01758 -0.0655 0.9325 0.1071
+ -3.000 -0.0927 0.02836 0.01668 -0.0674 0.9270 0.1181
+ -2.750 -0.0630 0.02772 0.01598 -0.0679 0.9193 0.1302
+ -2.500 -0.0276 0.02702 0.01537 -0.0696 0.9130 0.1521
+ -2.250 0.0039 0.02644 0.01494 -0.0707 0.9057 0.1918
+ -2.000 0.0363 0.02580 0.01472 -0.0721 0.8985 0.2686
+ -1.750 0.0685 0.02503 0.01447 -0.0733 0.8921 0.3781
+ -1.500 0.0939 0.02427 0.01439 -0.0728 0.8840 0.5183
+ -1.250 0.1299 0.02332 0.01447 -0.0719 0.8799 0.8331
+ -1.000 0.1803 0.02334 0.01424 -0.0763 0.8714 1.0000
+ -0.750 0.2176 0.02339 0.01398 -0.0781 0.8645 1.0000
+ -0.500 0.2416 0.02361 0.01395 -0.0777 0.8536 1.0000
+ -0.250 0.2742 0.02368 0.01379 -0.0785 0.8446 1.0000
+ 0.000 0.3084 0.02361 0.01351 -0.0793 0.8341 1.0000
+ 0.250 0.3367 0.02360 0.01332 -0.0790 0.8210 1.0000
+ 0.500 0.3660 0.02355 0.01311 -0.0789 0.8082 1.0000
+ 0.750 0.3972 0.02346 0.01288 -0.0789 0.7967 1.0000
+ 1.000 0.4301 0.02334 0.01263 -0.0793 0.7868 1.0000
+ 1.250 0.4543 0.02348 0.01268 -0.0785 0.7743 1.0000
+ 1.500 0.4808 0.02357 0.01268 -0.0779 0.7626 1.0000
+ 1.750 0.5119 0.02350 0.01252 -0.0780 0.7527 1.0000
+ 2.000 0.5391 0.02355 0.01250 -0.0775 0.7408 1.0000
+ 2.250 0.5637 0.02369 0.01260 -0.0766 0.7279 1.0000
+ 2.500 0.5897 0.02377 0.01264 -0.0759 0.7152 1.0000
+ 2.750 0.6171 0.02381 0.01263 -0.0754 0.7027 1.0000
+ 3.000 0.6462 0.02377 0.01256 -0.0750 0.6908 1.0000
+ 3.250 0.6728 0.02383 0.01260 -0.0743 0.6773 1.0000
+ 3.500 0.6976 0.02396 0.01272 -0.0734 0.6626 1.0000
+ 3.750 0.7225 0.02410 0.01286 -0.0725 0.6478 1.0000
+ 4.000 0.7476 0.02425 0.01300 -0.0717 0.6327 1.0000
+ 4.250 0.7728 0.02441 0.01315 -0.0708 0.6173 1.0000
+ 4.500 0.7979 0.02459 0.01334 -0.0700 0.6017 1.0000
+ 4.750 0.8229 0.02481 0.01355 -0.0691 0.5861 1.0000
+ 5.000 0.8478 0.02506 0.01380 -0.0683 0.5704 1.0000
+ 5.250 0.8724 0.02535 0.01411 -0.0675 0.5549 1.0000
+ 5.500 0.8967 0.02571 0.01447 -0.0667 0.5398 1.0000
+ 5.750 0.9206 0.02611 0.01489 -0.0658 0.5250 1.0000
+ 6.000 0.9442 0.02655 0.01537 -0.0650 0.5105 1.0000
+ 6.250 0.9676 0.02703 0.01589 -0.0642 0.4965 1.0000
+ 6.500 0.9908 0.02753 0.01642 -0.0633 0.4828 1.0000
+ 6.750 1.0137 0.02803 0.01695 -0.0624 0.4694 1.0000
+ 7.000 1.0365 0.02854 0.01751 -0.0615 0.4560 1.0000
+ 7.250 1.0589 0.02904 0.01803 -0.0605 0.4426 1.0000
+ 7.500 1.0802 0.02956 0.01859 -0.0593 0.4286 1.0000
+ 7.750 1.0990 0.03014 0.01923 -0.0579 0.4135 1.0000
+ 8.000 1.1171 0.03072 0.01985 -0.0563 0.3981 1.0000
+ 8.250 1.1344 0.03136 0.02052 -0.0547 0.3829 1.0000
+ 8.500 1.1507 0.03207 0.02129 -0.0530 0.3680 1.0000
+ 8.750 1.1663 0.03284 0.02218 -0.0513 0.3538 1.0000
+ 9.000 1.1812 0.03368 0.02312 -0.0496 0.3398 1.0000
+ 9.250 1.1953 0.03458 0.02412 -0.0478 0.3261 1.0000
+ 9.500 1.2082 0.03551 0.02516 -0.0459 0.3122 1.0000
+ 9.750 1.2198 0.03650 0.02626 -0.0438 0.2985 1.0000
+ 10.000 1.2292 0.03753 0.02741 -0.0415 0.2851 1.0000
+ 10.250 1.2370 0.03863 0.02860 -0.0392 0.2718 1.0000
+ 10.500 1.2441 0.03980 0.02984 -0.0368 0.2588 1.0000
+ 10.750 1.2481 0.04125 0.03147 -0.0345 0.2459 1.0000
+ 11.000 1.2524 0.04284 0.03320 -0.0324 0.2338 1.0000
+ 11.250 1.2570 0.04447 0.03496 -0.0304 0.2226 1.0000
+ 11.500 1.2624 0.04606 0.03662 -0.0286 0.2125 1.0000
+ 11.750 1.2666 0.04784 0.03850 -0.0269 0.2028 1.0000
+ 12.000 1.2690 0.04994 0.04076 -0.0254 0.1939 1.0000
+ 12.250 1.2747 0.05164 0.04246 -0.0240 0.1859 1.0000
+ 12.500 1.2723 0.05423 0.04524 -0.0227 0.1775 1.0000
+ 12.750 1.2707 0.05664 0.04770 -0.0216 0.1694 1.0000
+ 13.000 1.2669 0.05930 0.05044 -0.0208 0.1611 1.0000
+ 13.250 1.2598 0.06261 0.05389 -0.0203 0.1535 1.0000
+ 13.500 1.2563 0.06533 0.05660 -0.0198 0.1456 1.0000
+ 13.750 1.2448 0.06962 0.06115 -0.0201 0.1389 1.0000
+ 14.000 1.2399 0.07271 0.06416 -0.0202 0.1313 1.0000
+ 14.250 1.2245 0.07794 0.06970 -0.0214 0.1250 1.0000
+ 14.500 1.2176 0.08161 0.07330 -0.0221 0.1177 1.0000
+ 14.750 1.2014 0.08752 0.07952 -0.0240 0.1120 1.0000
+ 15.000 1.1960 0.09127 0.08321 -0.0250 0.1052 1.0000
+ 15.250 1.1795 0.09779 0.09003 -0.0275 0.1005 1.0000
+ 15.500 1.1711 0.10264 0.09495 -0.0292 0.0944 1.0000
+ 15.750 1.1586 0.10858 0.10103 -0.0316 0.0893 1.0000
+ 16.000 1.1435 0.11533 0.10796 -0.0346 0.0844 1.0000
+ 16.250 1.1380 0.11991 0.11252 -0.0365 0.0785 1.0000
+ 16.500 1.1148 0.12918 0.12205 -0.0412 0.0752 1.0000
+ 16.750 1.1211 0.13104 0.12374 -0.0416 0.0685 1.0000
+ 17.000 1.0920 0.14257 0.13553 -0.0479 0.0668 1.0000
+ 17.250 1.0525 0.15811 0.15118 -0.0568 0.0657 1.0000
diff --git a/Airfoils/Polars/Clark_y_polar_Re_500000.txt b/Airfoils/Polars/Clark_y_polar_Re_500000.txt
new file mode 100644
index 0000000..cfd2640
--- /dev/null
+++ b/Airfoils/Polars/Clark_y_polar_Re_500000.txt
@@ -0,0 +1,118 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: CLARK Y AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.250 -0.3940 0.08504 0.08280 -0.0444 1.0000 0.0381
+ -9.000 -0.4033 0.08275 0.08056 -0.0431 1.0000 0.0387
+ -8.750 -0.4929 0.04594 0.04332 -0.0802 0.9855 0.0348
+ -8.500 -0.4974 0.03456 0.03116 -0.0866 0.9772 0.0344
+ -8.250 -0.4759 0.02994 0.02595 -0.0893 0.9733 0.0352
+ -8.000 -0.4463 0.02768 0.02325 -0.0914 0.9710 0.0358
+ -7.750 -0.4301 0.02331 0.01843 -0.0916 0.9643 0.0365
+ -7.500 -0.4024 0.02155 0.01653 -0.0925 0.9600 0.0373
+ -7.250 -0.3722 0.02018 0.01500 -0.0935 0.9568 0.0379
+ -7.000 -0.3479 0.01919 0.01389 -0.0931 0.9498 0.0386
+ -6.750 -0.3205 0.01829 0.01288 -0.0932 0.9443 0.0394
+ -6.500 -0.2942 0.01724 0.01167 -0.0930 0.9387 0.0399
+ -6.250 -0.2696 0.01629 0.01057 -0.0923 0.9311 0.0403
+ -6.000 -0.2426 0.01543 0.00958 -0.0921 0.9257 0.0408
+ -5.750 -0.2181 0.01471 0.00876 -0.0913 0.9173 0.0413
+ -5.500 -0.1915 0.01408 0.00802 -0.0910 0.9110 0.0419
+ -5.250 -0.1658 0.01361 0.00747 -0.0904 0.9029 0.0426
+ -5.000 -0.1391 0.01312 0.00689 -0.0900 0.8960 0.0430
+ -4.750 -0.1139 0.01244 0.00613 -0.0894 0.8877 0.0436
+ -4.500 -0.0885 0.01171 0.00534 -0.0889 0.8804 0.0446
+ -4.250 -0.0629 0.01125 0.00485 -0.0884 0.8713 0.0457
+ -4.000 -0.0363 0.01091 0.00444 -0.0879 0.8630 0.0469
+ -3.750 -0.0102 0.01061 0.00411 -0.0875 0.8528 0.0482
+ -3.500 0.0167 0.01036 0.00381 -0.0871 0.8441 0.0499
+ -3.250 0.0438 0.01016 0.00354 -0.0868 0.8350 0.0516
+ -3.000 0.0703 0.00985 0.00319 -0.0864 0.8255 0.0547
+ -2.750 0.0975 0.00965 0.00296 -0.0861 0.8168 0.0589
+ -2.500 0.1243 0.00942 0.00275 -0.0857 0.8067 0.0669
+ -2.250 0.1510 0.00915 0.00254 -0.0854 0.7976 0.0936
+ -2.000 0.1774 0.00887 0.00238 -0.0851 0.7876 0.1350
+ -1.750 0.2037 0.00859 0.00228 -0.0848 0.7769 0.1935
+ -1.500 0.2303 0.00838 0.00221 -0.0846 0.7675 0.2491
+ -1.250 0.2571 0.00820 0.00215 -0.0844 0.7583 0.2986
+ -1.000 0.2836 0.00799 0.00210 -0.0841 0.7496 0.3576
+ -0.750 0.3095 0.00774 0.00206 -0.0838 0.7404 0.4367
+ -0.500 0.3338 0.00732 0.00205 -0.0831 0.7305 0.5667
+ -0.250 0.3551 0.00684 0.00209 -0.0816 0.7213 0.7347
+ 0.000 0.3760 0.00652 0.00219 -0.0794 0.7115 0.8796
+ 0.250 0.4107 0.00654 0.00224 -0.0804 0.7020 0.9451
+ 0.500 0.4534 0.00663 0.00225 -0.0833 0.6921 0.9765
+ 0.750 0.5065 0.00669 0.00224 -0.0886 0.6803 0.9953
+ 1.000 0.5439 0.00675 0.00222 -0.0907 0.6684 1.0000
+ 1.250 0.5687 0.00682 0.00222 -0.0901 0.6567 1.0000
+ 1.500 0.5934 0.00691 0.00222 -0.0894 0.6450 1.0000
+ 1.750 0.6180 0.00699 0.00225 -0.0887 0.6323 1.0000
+ 2.000 0.6426 0.00708 0.00228 -0.0880 0.6193 1.0000
+ 2.250 0.6670 0.00719 0.00232 -0.0872 0.6053 1.0000
+ 2.500 0.6912 0.00732 0.00237 -0.0864 0.5901 1.0000
+ 2.750 0.7151 0.00746 0.00244 -0.0855 0.5739 1.0000
+ 3.000 0.7389 0.00761 0.00251 -0.0847 0.5572 1.0000
+ 3.250 0.7624 0.00779 0.00261 -0.0838 0.5390 1.0000
+ 3.500 0.7857 0.00799 0.00271 -0.0828 0.5188 1.0000
+ 3.750 0.8090 0.00820 0.00284 -0.0819 0.4959 1.0000
+ 4.000 0.8315 0.00848 0.00299 -0.0808 0.4690 1.0000
+ 4.250 0.8536 0.00880 0.00315 -0.0797 0.4386 1.0000
+ 4.500 0.8757 0.00914 0.00334 -0.0787 0.4111 1.0000
+ 4.750 0.8987 0.00945 0.00355 -0.0778 0.3891 1.0000
+ 5.000 0.9220 0.00976 0.00376 -0.0770 0.3723 1.0000
+ 5.250 0.9457 0.01004 0.00398 -0.0763 0.3590 1.0000
+ 5.500 0.9697 0.01032 0.00421 -0.0756 0.3484 1.0000
+ 5.750 0.9940 0.01057 0.00444 -0.0750 0.3396 1.0000
+ 6.000 1.0185 0.01082 0.00468 -0.0745 0.3319 1.0000
+ 6.250 1.0426 0.01108 0.00492 -0.0739 0.3237 1.0000
+ 6.500 1.0671 0.01132 0.00517 -0.0733 0.3159 1.0000
+ 6.750 1.0910 0.01160 0.00543 -0.0727 0.3073 1.0000
+ 7.000 1.1155 0.01183 0.00567 -0.0721 0.2979 1.0000
+ 7.250 1.1388 0.01212 0.00594 -0.0715 0.2876 1.0000
+ 7.500 1.1627 0.01238 0.00619 -0.0708 0.2767 1.0000
+ 7.750 1.1866 0.01263 0.00644 -0.0702 0.2648 1.0000
+ 8.000 1.2096 0.01293 0.00672 -0.0695 0.2509 1.0000
+ 8.250 1.2317 0.01328 0.00703 -0.0687 0.2337 1.0000
+ 8.500 1.2520 0.01375 0.00740 -0.0676 0.2092 1.0000
+ 8.750 1.2691 0.01442 0.00789 -0.0660 0.1778 1.0000
+ 9.000 1.2847 0.01519 0.00850 -0.0643 0.1507 1.0000
+ 9.250 1.2999 0.01594 0.00914 -0.0624 0.1304 1.0000
+ 9.500 1.3150 0.01659 0.00973 -0.0605 0.1159 1.0000
+ 9.750 1.3299 0.01726 0.01035 -0.0587 0.1043 1.0000
+ 10.000 1.3447 0.01794 0.01099 -0.0568 0.0940 1.0000
+ 10.250 1.3602 0.01858 0.01163 -0.0551 0.0853 1.0000
+ 10.500 1.3747 0.01929 0.01232 -0.0534 0.0759 1.0000
+ 10.750 1.3868 0.02016 0.01312 -0.0514 0.0621 1.0000
+ 11.000 1.3917 0.02153 0.01429 -0.0487 0.0373 1.0000
+ 11.250 1.3923 0.02324 0.01587 -0.0456 0.0223 1.0000
+ 11.500 1.3998 0.02452 0.01720 -0.0434 0.0198 1.0000
+ 11.750 1.4079 0.02581 0.01857 -0.0415 0.0183 1.0000
+ 12.000 1.4163 0.02711 0.01997 -0.0397 0.0173 1.0000
+ 12.250 1.4223 0.02865 0.02159 -0.0380 0.0164 1.0000
+ 12.500 1.4251 0.03050 0.02353 -0.0361 0.0157 1.0000
+ 12.750 1.4237 0.03282 0.02595 -0.0343 0.0151 1.0000
+ 13.000 1.4287 0.03464 0.02789 -0.0330 0.0148 1.0000
+ 13.250 1.4317 0.03673 0.03008 -0.0319 0.0145 1.0000
+ 13.500 1.4329 0.03907 0.03254 -0.0308 0.0141 1.0000
+ 13.750 1.4323 0.04167 0.03524 -0.0299 0.0138 1.0000
+ 14.000 1.4303 0.04452 0.03820 -0.0293 0.0135 1.0000
+ 14.250 1.4269 0.04763 0.04141 -0.0288 0.0133 1.0000
+ 14.500 1.4220 0.05103 0.04492 -0.0285 0.0130 1.0000
+ 14.750 1.4153 0.05481 0.04879 -0.0286 0.0128 1.0000
+ 15.000 1.4066 0.05897 0.05306 -0.0288 0.0126 1.0000
+ 15.250 1.3959 0.06349 0.05768 -0.0293 0.0124 1.0000
+ 15.500 1.3832 0.06838 0.06268 -0.0300 0.0122 1.0000
+ 15.750 1.3730 0.07302 0.06742 -0.0307 0.0120 1.0000
+ 16.000 1.3700 0.07690 0.07141 -0.0315 0.0119 1.0000
+ 16.250 1.3662 0.08095 0.07557 -0.0325 0.0117 1.0000
+ 16.500 1.3615 0.08517 0.07990 -0.0335 0.0116 1.0000
+ 16.750 1.3564 0.08948 0.08431 -0.0346 0.0114 1.0000
+ 17.000 1.3510 0.09382 0.08876 -0.0358 0.0113 1.0000
diff --git a/Airfoils/Polars/E850_polar_Re_100000.txt b/Airfoils/Polars/E850_polar_Re_100000.txt
new file mode 100644
index 0000000..c126f40
--- /dev/null
+++ b/Airfoils/Polars/E850_polar_Re_100000.txt
@@ -0,0 +1,72 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E850 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -8.250 -0.5345 0.09099 0.08633 -0.0342 1.0000 0.0849
+ -8.000 -0.5570 0.08767 0.08312 -0.0366 1.0000 0.0853
+ -7.750 -0.5786 0.08375 0.07919 -0.0409 1.0000 0.0857
+ -7.500 -0.5555 0.08074 0.07631 -0.0327 1.0000 0.0900
+ -7.250 -0.5614 0.07696 0.07257 -0.0339 1.0000 0.0930
+ -7.000 -0.5747 0.07268 0.06823 -0.0372 1.0000 0.0967
+ -6.750 -0.5882 0.06808 0.06347 -0.0401 1.0000 0.1004
+ -6.500 -0.5777 0.06507 0.06058 -0.0371 1.0000 0.1051
+ -6.250 -0.5820 0.06076 0.05603 -0.0395 1.0000 0.1141
+ -6.000 -0.5725 0.05763 0.05292 -0.0381 1.0000 0.1205
+ -5.750 -0.5658 0.05394 0.04914 -0.0382 1.0000 0.1318
+ -5.500 -0.5561 0.05049 0.04541 -0.0384 1.0000 0.1467
+ -5.000 -0.4938 0.03583 0.02865 -0.0406 1.0000 0.0424
+ -4.750 -0.4669 0.03257 0.02463 -0.0394 1.0000 0.0334
+ -4.500 -0.4432 0.02908 0.02077 -0.0390 1.0000 0.0310
+ -4.250 -0.4175 0.02639 0.01766 -0.0383 1.0000 0.0289
+ -4.000 -0.3908 0.02407 0.01495 -0.0374 1.0000 0.0275
+ -3.750 -0.3648 0.02213 0.01278 -0.0364 1.0000 0.0273
+ -3.500 -0.3399 0.02053 0.01107 -0.0354 1.0000 0.0281
+ -3.250 -0.3152 0.01919 0.00963 -0.0346 1.0000 0.0306
+ -3.000 -0.2899 0.01806 0.00840 -0.0342 1.0000 0.0423
+ -2.750 -0.2679 0.01453 0.00762 -0.0334 1.0000 0.6059
+ -2.500 -0.2651 0.01506 0.00852 -0.0256 1.0000 0.7233
+ -2.250 -0.2499 0.01508 0.00844 -0.0221 1.0000 0.7588
+ -2.000 -0.2373 0.01504 0.00837 -0.0180 1.0000 0.7913
+ -1.750 -0.2225 0.01490 0.00803 -0.0149 1.0000 0.8186
+ -1.500 -0.2034 0.01472 0.00775 -0.0132 1.0000 0.8389
+ -1.250 -0.1837 0.01455 0.00748 -0.0117 1.0000 0.8598
+ -1.000 -0.1644 0.01436 0.00725 -0.0101 1.0000 0.8838
+ -0.750 -0.1426 0.01417 0.00704 -0.0091 1.0000 0.9154
+ -0.500 -0.1004 0.01399 0.00683 -0.0127 1.0000 0.9656
+ -0.250 -0.0658 0.01386 0.00656 -0.0160 1.0000 1.0000
+ 0.000 -0.0338 0.01396 0.00645 -0.0185 1.0000 1.0000
+ 0.250 -0.0044 0.01413 0.00648 -0.0201 1.0000 1.0000
+ 0.500 0.0230 0.01434 0.00659 -0.0211 1.0000 1.0000
+ 0.750 0.0489 0.01458 0.00675 -0.0216 1.0000 1.0000
+ 1.000 0.0737 0.01485 0.00696 -0.0219 1.0000 1.0000
+ 1.250 0.0977 0.01514 0.00723 -0.0220 1.0000 1.0000
+ 1.500 0.1212 0.01546 0.00754 -0.0220 1.0000 1.0000
+ 1.750 0.1441 0.01581 0.00789 -0.0220 1.0000 1.0000
+ 2.000 0.1665 0.01619 0.00829 -0.0218 1.0000 1.0000
+ 2.250 0.1885 0.01660 0.00875 -0.0217 1.0000 1.0000
+ 2.500 0.2100 0.01705 0.00927 -0.0215 1.0000 1.0000
+ 2.750 0.2310 0.01754 0.00985 -0.0213 1.0000 1.0000
+ 3.000 0.2660 0.01832 0.01078 -0.0239 0.9944 1.0000
+ 3.250 0.3245 0.01922 0.01213 -0.0308 0.9771 1.0000
+ 3.500 0.3976 0.01956 0.01290 -0.0395 0.9506 1.0000
+ 3.750 0.5639 0.01472 0.00934 -0.0571 0.8505 1.0000
+ 4.000 0.6097 0.01865 0.00916 -0.0555 0.0768 1.0000
+ 4.250 0.6314 0.02077 0.01125 -0.0535 0.0585 1.0000
+ 4.500 0.6578 0.02271 0.01316 -0.0529 0.0432 1.0000
+ 4.750 0.6897 0.02579 0.01625 -0.0530 0.0393 1.0000
+ 5.000 0.7211 0.02903 0.01979 -0.0527 0.0391 1.0000
+ 5.250 0.7494 0.03145 0.02276 -0.0510 0.0412 1.0000
+ 5.500 0.7738 0.03551 0.02754 -0.0486 0.0469 1.0000
+ 5.750 0.7971 0.04046 0.03314 -0.0458 0.0612 1.0000
+ 7.250 0.8692 0.06973 0.06482 -0.0330 0.1147 1.0000
+ 8.250 0.8799 0.08740 0.08262 -0.0290 0.0766 1.0000
+ 8.500 0.8623 0.08999 0.08548 -0.0275 0.0763 1.0000
+ 8.750 0.8437 0.09314 0.08877 -0.0266 0.0760 1.0000
diff --git a/Airfoils/Polars/E850_polar_Re_1000000.txt b/Airfoils/Polars/E850_polar_Re_1000000.txt
new file mode 100644
index 0000000..9ba7f51
--- /dev/null
+++ b/Airfoils/Polars/E850_polar_Re_1000000.txt
@@ -0,0 +1,82 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E850 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.500 -0.4385 0.09119 0.08970 -0.0267 1.0000 0.0056
+ -9.250 -0.4441 0.08738 0.08591 -0.0267 1.0000 0.0056
+ -9.000 -0.4507 0.08358 0.08212 -0.0266 1.0000 0.0057
+ -8.750 -0.4499 0.07815 0.07671 -0.0295 0.9994 0.0059
+ -8.500 -0.4495 0.07233 0.07089 -0.0331 0.9984 0.0059
+ -8.250 -0.4500 0.06602 0.06458 -0.0374 0.9972 0.0059
+ -8.000 -0.4531 0.05849 0.05706 -0.0436 0.9956 0.0058
+ -7.750 -0.4728 0.04663 0.04511 -0.0588 0.9913 0.0051
+ -7.500 -0.4714 0.04019 0.03850 -0.0649 0.9870 0.0052
+ -7.250 -0.4619 0.03469 0.03283 -0.0693 0.9845 0.0054
+ -7.000 -0.4465 0.02972 0.02766 -0.0729 0.9828 0.0057
+ -6.750 -0.4272 0.02518 0.02291 -0.0759 0.9817 0.0062
+ -6.500 -0.4139 0.02176 0.01928 -0.0762 0.9779 0.0068
+ -6.250 -0.3951 0.01844 0.01571 -0.0767 0.9748 0.0077
+ -6.000 -0.3701 0.01583 0.01287 -0.0775 0.9731 0.0091
+ -5.750 -0.3367 0.01625 0.01313 -0.0777 0.9722 0.0108
+ -5.500 -0.3105 0.01354 0.01011 -0.0788 0.9710 0.0109
+ -5.250 -0.3134 0.01934 0.01512 -0.0798 0.9704 0.0124
+ -5.000 -0.2838 0.01741 0.01305 -0.0810 0.9697 0.0144
+ -4.750 -0.2525 0.01602 0.01149 -0.0820 0.9692 0.0173
+ -4.500 -0.2191 0.01542 0.01077 -0.0829 0.9688 0.0212
+ -4.250 -1.2480 0.06364 0.06202 0.2358 0.9815 0.0050
+ -4.000 -0.1756 0.01028 0.00521 -0.0786 0.9622 0.0070
+ -3.750 -0.1452 0.00933 0.00409 -0.0790 0.9609 0.0048
+ -3.500 -0.1134 0.00891 0.00360 -0.0798 0.9600 0.0041
+ -3.250 -0.0812 0.00833 0.00283 -0.0808 0.9592 0.0038
+ -3.000 -0.0482 0.00809 0.00242 -0.0819 0.9586 0.0037
+ -2.750 -0.0149 0.00791 0.00220 -0.0831 0.9580 0.0040
+ -2.500 0.0178 0.00765 0.00202 -0.0844 0.9574 0.0335
+ -2.250 0.0492 0.00691 0.00179 -0.0858 0.9567 0.1845
+ -2.000 0.0787 0.00572 0.00165 -0.0873 0.9559 0.4918
+ -1.750 0.1011 0.00556 0.00171 -0.0862 0.9519 0.5611
+ -1.500 0.1307 0.00547 0.00165 -0.0866 0.9496 0.5755
+ -1.250 0.1626 0.00536 0.00156 -0.0876 0.9476 0.5883
+ -1.000 0.1962 0.00523 0.00145 -0.0888 0.9456 0.5989
+ -0.750 0.2300 0.00509 0.00134 -0.0902 0.9434 0.6075
+ -0.500 0.2563 0.00502 0.00130 -0.0898 0.9382 0.6159
+ -0.250 0.2870 0.00491 0.00121 -0.0904 0.9335 0.6251
+ 0.000 0.3178 0.00481 0.00114 -0.0910 0.9285 0.6343
+ 0.250 0.3449 0.00473 0.00110 -0.0908 0.9208 0.6440
+ 0.500 0.3732 0.00467 0.00106 -0.0908 0.9125 0.6544
+ 1.000 0.4269 0.00456 0.00100 -0.0901 0.8830 0.6774
+ 1.250 0.4529 0.00454 0.00098 -0.0895 0.8599 0.6898
+ 1.500 0.4803 0.00454 0.00111 -0.0894 0.8520 0.7034
+ 2.000 0.5178 0.00518 0.00123 -0.0851 0.6782 0.7331
+ 2.250 0.5299 0.00604 0.00152 -0.0819 0.5217 0.7505
+ 2.500 0.5434 0.00709 0.00187 -0.0794 0.3315 0.7701
+ 2.750 0.5603 0.00801 0.00223 -0.0775 0.1718 0.7920
+ 3.000 0.5806 0.00865 0.00256 -0.0763 0.0758 0.8173
+ 3.250 0.5995 0.00956 0.00334 -0.0742 0.0063 0.8476
+ 3.500 0.6197 0.01007 0.00415 -0.0723 0.0055 0.8886
+ 3.750 0.6438 0.01009 0.00433 -0.0714 0.0046 0.9752
+ 4.000 0.6694 0.01070 0.00503 -0.0709 0.0040 1.0000
+ 4.250 0.6936 0.01122 0.00560 -0.0702 0.0034 1.0000
+ 4.500 0.7181 0.01168 0.00607 -0.0696 0.0026 1.0000
+ 4.750 0.7362 0.01392 0.00854 -0.0673 0.0020 1.0000
+ 5.750 0.8012 0.03377 0.03031 -0.0544 0.0057 1.0000
+ 6.000 0.8311 0.03331 0.02990 -0.0540 0.0047 1.0000
+ 6.250 0.8486 0.03612 0.03292 -0.0517 0.0042 1.0000
+ 6.500 0.8615 0.03981 0.03686 -0.0490 0.0039 1.0000
+ 6.750 0.8717 0.04381 0.04111 -0.0462 0.0037 1.0000
+ 7.000 0.8797 0.04794 0.04546 -0.0436 0.0035 1.0000
+ 7.250 0.8850 0.05227 0.05000 -0.0410 0.0034 1.0000
+ 7.500 0.8876 0.05670 0.05461 -0.0385 0.0033 1.0000
+ 7.750 0.8874 0.06113 0.05920 -0.0363 0.0032 1.0000
+ 8.000 0.8833 0.06570 0.06392 -0.0343 0.0032 1.0000
+ 8.250 0.8768 0.06994 0.06828 -0.0326 0.0031 1.0000
+ 8.500 0.8629 0.07390 0.07232 -0.0303 0.0031 1.0000
+ 8.750 0.8448 0.07776 0.07626 -0.0284 0.0032 1.0000
+ 9.000 0.8268 0.08238 0.08093 -0.0289 0.0032 1.0000
diff --git a/Airfoils/Polars/E850_polar_Re_200000.txt b/Airfoils/Polars/E850_polar_Re_200000.txt
new file mode 100644
index 0000000..266104b
--- /dev/null
+++ b/Airfoils/Polars/E850_polar_Re_200000.txt
@@ -0,0 +1,74 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E850 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -8.500 -0.5186 0.08797 0.08457 -0.0367 1.0000 0.0341
+ -8.250 -0.5295 0.08381 0.08047 -0.0387 1.0000 0.0341
+ -8.000 -0.5454 0.08042 0.07712 -0.0392 1.0000 0.0341
+ -7.750 -0.5596 0.07695 0.07363 -0.0401 1.0000 0.0341
+ -7.500 -0.5701 0.07378 0.07038 -0.0404 1.0000 0.0342
+ -7.250 -0.5779 0.07073 0.06720 -0.0403 1.0000 0.0344
+ -7.000 -0.5819 0.06772 0.06402 -0.0399 1.0000 0.0345
+ -6.750 -0.5823 0.06459 0.06069 -0.0394 1.0000 0.0346
+ -6.500 -0.5879 0.05631 0.05245 -0.0400 1.0000 0.0359
+ -6.250 -0.5813 0.05269 0.04883 -0.0393 1.0000 0.0371
+ -6.000 -0.5725 0.04950 0.04556 -0.0388 1.0000 0.0387
+ -5.750 -0.5611 0.04617 0.04203 -0.0387 1.0000 0.0411
+ -5.500 -0.5400 0.04663 0.04166 -0.0378 1.0000 0.0471
+ -4.250 -0.4268 0.02445 0.01762 -0.0360 1.0000 0.0197
+ -4.000 -0.3996 0.02159 0.01435 -0.0355 1.0000 0.0171
+ -3.750 -0.3724 0.01934 0.01176 -0.0346 1.0000 0.0152
+ -3.500 -0.3464 0.01764 0.00987 -0.0337 1.0000 0.0140
+ -3.250 -0.3207 0.01633 0.00844 -0.0331 1.0000 0.0136
+ -3.000 -0.2948 0.01530 0.00728 -0.0328 1.0000 0.0139
+ -2.750 -0.2686 0.01456 0.00634 -0.0326 1.0000 0.0154
+ -2.500 -0.2426 0.01407 0.00563 -0.0322 1.0000 0.0201
+ -2.250 -0.2143 0.01173 0.00527 -0.0338 1.0000 0.4865
+ -2.000 -0.1964 0.01173 0.00586 -0.0314 1.0000 0.6476
+ -1.750 -0.1764 0.01186 0.00595 -0.0296 1.0000 0.6881
+ -1.500 -0.1557 0.01195 0.00607 -0.0281 1.0000 0.7159
+ -1.250 -0.1334 0.01199 0.00610 -0.0272 1.0000 0.7323
+ -1.000 -0.1109 0.01204 0.00614 -0.0264 1.0000 0.7460
+ -0.750 -0.0887 0.01209 0.00620 -0.0256 1.0000 0.7604
+ -0.500 -0.0668 0.01215 0.00629 -0.0248 1.0000 0.7760
+ -0.250 -0.0417 0.01226 0.00640 -0.0246 0.9989 0.7940
+ 0.000 -0.0069 0.01252 0.00673 -0.0264 0.9945 0.8144
+ 0.250 0.0253 0.01265 0.00694 -0.0276 0.9894 0.8398
+ 0.750 0.1081 0.01284 0.00744 -0.0338 0.9812 0.9795
+ 1.000 0.1437 0.01292 0.00748 -0.0362 0.9743 1.0000
+ 1.250 0.1870 0.01324 0.00777 -0.0400 0.9696 1.0000
+ 1.500 0.2227 0.01336 0.00790 -0.0421 0.9622 1.0000
+ 1.750 0.2683 0.01356 0.00815 -0.0460 0.9567 1.0000
+ 2.000 0.3061 0.01356 0.00822 -0.0483 0.9476 1.0000
+ 2.250 0.3526 0.01348 0.00825 -0.0520 0.9389 1.0000
+ 2.500 0.4104 0.01291 0.00799 -0.0572 0.9249 1.0000
+ 2.750 0.4809 0.01165 0.00700 -0.0641 0.9071 1.0000
+ 3.000 0.5500 0.00979 0.00544 -0.0697 0.8729 1.0000
+ 3.250 0.6070 0.00937 0.00414 -0.0723 0.5738 1.0000
+ 3.500 0.5943 0.01241 0.00497 -0.0650 0.1533 1.0000
+ 3.750 0.6061 0.01461 0.00647 -0.0620 0.0404 1.0000
+ 4.000 0.6246 0.01628 0.00814 -0.0599 0.0250 1.0000
+ 4.250 0.6482 0.01788 0.00983 -0.0585 0.0227 1.0000
+ 4.500 0.6762 0.01997 0.01204 -0.0578 0.0219 1.0000
+ 4.750 0.7065 0.02259 0.01493 -0.0572 0.0224 1.0000
+ 5.000 0.7341 0.02515 0.01785 -0.0561 0.0215 1.0000
+ 5.500 0.7735 0.03029 0.02364 -0.0533 0.0139 1.0000
+ 5.750 0.7759 0.02615 0.02081 -0.0450 0.0333 1.0000
+ 6.000 0.7876 0.03051 0.02556 -0.0424 0.0349 1.0000
+ 6.250 0.7987 0.03441 0.02982 -0.0398 0.0331 1.0000
+ 6.500 0.8071 0.03853 0.03423 -0.0374 0.0316 1.0000
+ 6.750 0.8134 0.04279 0.03873 -0.0353 0.0304 1.0000
+ 7.000 0.8175 0.04716 0.04328 -0.0333 0.0294 1.0000
+ 7.250 0.8197 0.05162 0.04789 -0.0314 0.0286 1.0000
+ 7.500 0.8203 0.05641 0.05277 -0.0299 0.0279 1.0000
+ 7.750 0.8159 0.06249 0.05892 -0.0286 0.0273 1.0000
+ 8.750 0.7514 0.08286 0.07973 -0.0218 0.0266 1.0000
+ 9.000 0.7337 0.08739 0.08434 -0.0222 0.0266 1.0000
diff --git a/Airfoils/Polars/E850_polar_Re_500000.txt b/Airfoils/Polars/E850_polar_Re_500000.txt
new file mode 100644
index 0000000..ce4e2d4
--- /dev/null
+++ b/Airfoils/Polars/E850_polar_Re_500000.txt
@@ -0,0 +1,85 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E850 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.000 -0.4439 0.08243 0.08037 -0.0314 1.0000 0.0124
+ -8.750 -0.4509 0.07837 0.07634 -0.0315 1.0000 0.0125
+ -8.500 -0.4579 0.07485 0.07285 -0.0312 1.0000 0.0128
+ -8.250 -0.4684 0.07064 0.06867 -0.0314 1.0000 0.0127
+ -8.000 -0.4816 0.06628 0.06434 -0.0318 1.0000 0.0127
+ -7.750 -0.5046 0.05978 0.05786 -0.0350 1.0000 0.0122
+ -7.500 -0.5307 0.05718 0.05523 -0.0345 1.0000 0.0118
+ -7.250 -0.5313 0.05034 0.04824 -0.0419 0.9967 0.0119
+ -7.000 -0.5258 0.04384 0.04156 -0.0472 0.9940 0.0122
+ -6.750 -0.5178 0.03902 0.03660 -0.0499 0.9909 0.0127
+ -6.500 -0.5056 0.03468 0.03209 -0.0521 0.9879 0.0134
+ -6.250 -0.4890 0.03041 0.02761 -0.0543 0.9854 0.0143
+ -6.000 -0.4679 0.02629 0.02324 -0.0566 0.9836 0.0155
+ -5.750 -0.4432 0.02226 0.01890 -0.0586 0.9822 0.0172
+ -5.500 -0.4142 0.01903 0.01532 -0.0600 0.9811 0.0199
+ -5.250 -0.3903 0.01998 0.01595 -0.0583 0.9773 0.0220
+ -5.000 -0.3733 0.01315 0.00853 -0.0594 0.9751 0.0238
+ -4.750 -0.3474 0.01059 0.00582 -0.0602 0.9736 0.0260
+ -4.500 -0.3186 0.00871 0.00371 -0.0607 0.9723 0.0284
+ -4.250 -0.3122 0.01860 0.01305 -0.0588 0.9737 0.0195
+ -4.000 -0.2815 0.01581 0.00994 -0.0578 0.9731 0.0096
+ -3.750 -0.2524 0.01387 0.00781 -0.0577 0.9722 0.0080
+ -3.500 -0.2214 0.01259 0.00635 -0.0583 0.9712 0.0070
+ -3.250 -0.1886 0.01182 0.00539 -0.0593 0.9702 0.0065
+ -3.000 -0.1550 0.01144 0.00484 -0.0605 0.9693 0.0065
+ -2.750 -0.1229 0.01125 0.00448 -0.0614 0.9679 0.0076
+ -2.500 -0.1004 0.01109 0.00421 -0.0603 0.9640 0.0093
+ -2.250 -0.0724 0.00975 0.00388 -0.0615 0.9622 0.2791
+ -2.000 -0.0430 0.00906 0.00390 -0.0626 0.9605 0.4811
+ -1.750 -0.0121 0.00886 0.00412 -0.0634 0.9589 0.5970
+ -1.500 0.0213 0.00885 0.00416 -0.0647 0.9576 0.6261
+ -1.250 0.0559 0.00882 0.00416 -0.0663 0.9565 0.6391
+ -1.000 0.0829 0.00878 0.00413 -0.0662 0.9531 0.6489
+ -0.750 0.1117 0.00871 0.00409 -0.0665 0.9495 0.6588
+ -0.500 0.1459 0.00861 0.00402 -0.0680 0.9473 0.6694
+ -0.250 0.1820 0.00850 0.00395 -0.0698 0.9455 0.6807
+ 0.000 0.2202 0.00835 0.00386 -0.0721 0.9441 0.6927
+ 0.250 0.2610 0.00815 0.00373 -0.0749 0.9430 0.7054
+ 0.500 0.2851 0.00802 0.00368 -0.0740 0.9363 0.7187
+ 0.750 0.3268 0.00771 0.00348 -0.0769 0.9339 0.7332
+ 1.000 0.3751 0.00726 0.00315 -0.0810 0.9317 0.7486
+ 1.250 0.4116 0.00685 0.00286 -0.0824 0.9238 0.7656
+ 1.500 0.4507 0.00637 0.00258 -0.0842 0.9124 0.7840
+ 1.750 0.4820 0.00611 0.00246 -0.0846 0.9002 0.8044
+ 2.000 0.5097 0.00589 0.00237 -0.0840 0.8812 0.8277
+ 2.250 0.5350 0.00571 0.00226 -0.0829 0.8435 0.8551
+ 2.500 0.5575 0.00572 0.00217 -0.0810 0.7692 0.8891
+ 2.750 0.5631 0.00660 0.00232 -0.0757 0.5790 0.9482
+ 3.000 0.5738 0.00838 0.00283 -0.0730 0.2924 1.0000
+ 3.250 0.5880 0.00979 0.00335 -0.0709 0.0947 1.0000
+ 3.500 0.6058 0.01117 0.00444 -0.0688 0.0132 1.0000
+ 3.750 0.6279 0.01208 0.00549 -0.0673 0.0113 1.0000
+ 4.000 0.6495 0.01314 0.00665 -0.0658 0.0104 1.0000
+ 4.250 0.6731 0.01382 0.00734 -0.0651 0.0069 1.0000
+ 4.500 0.6940 0.01680 0.01056 -0.0633 0.0057 1.0000
+ 4.750 0.7201 0.01750 0.01137 -0.0628 0.0051 1.0000
+ 5.000 0.7460 0.01954 0.01366 -0.0618 0.0047 1.0000
+ 5.250 0.7698 0.02258 0.01707 -0.0602 0.0047 1.0000
+ 5.500 0.7878 0.02692 0.02191 -0.0575 0.0052 1.0000
+ 5.750 0.8125 0.03169 0.02718 -0.0539 0.0097 1.0000
+ 6.000 0.8276 0.03545 0.03129 -0.0511 0.0097 1.0000
+ 6.250 0.8410 0.03903 0.03518 -0.0484 0.0094 1.0000
+ 6.500 0.8523 0.04276 0.03919 -0.0458 0.0090 1.0000
+ 6.750 0.8617 0.04644 0.04313 -0.0433 0.0085 1.0000
+ 7.000 0.8691 0.05016 0.04707 -0.0409 0.0082 1.0000
+ 7.250 0.8741 0.05394 0.05106 -0.0387 0.0079 1.0000
+ 7.500 0.8766 0.05774 0.05503 -0.0366 0.0076 1.0000
+ 7.750 0.8765 0.06160 0.05906 -0.0345 0.0074 1.0000
+ 8.000 0.8731 0.06556 0.06316 -0.0326 0.0072 1.0000
+ 8.250 0.8670 0.06941 0.06713 -0.0308 0.0071 1.0000
+ 8.500 0.8546 0.07307 0.07089 -0.0285 0.0071 1.0000
+ 8.750 0.8388 0.07703 0.07493 -0.0268 0.0071 1.0000
+ 9.000 0.8225 0.08179 0.07977 -0.0272 0.0072 1.0000
diff --git a/Airfoils/Polars/E851_polar_Re_100000.txt b/Airfoils/Polars/E851_polar_Re_100000.txt
new file mode 100644
index 0000000..a35ef12
--- /dev/null
+++ b/Airfoils/Polars/E851_polar_Re_100000.txt
@@ -0,0 +1,80 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E851 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -8.750 -0.4785 0.09749 0.09277 -0.0410 1.0000 0.0800
+ -8.500 -0.4985 0.09461 0.09001 -0.0433 1.0000 0.0804
+ -8.250 -0.5218 0.09171 0.08724 -0.0446 1.0000 0.0805
+ -8.000 -0.4860 0.08764 0.08312 -0.0377 1.0000 0.0848
+ -7.750 -0.4953 0.08509 0.08066 -0.0364 1.0000 0.0870
+ -7.500 -0.5125 0.08267 0.07835 -0.0351 1.0000 0.0885
+ -7.250 -0.5291 0.07914 0.07490 -0.0367 1.0000 0.0902
+ -7.000 -0.5585 0.07482 0.07045 -0.0426 1.0000 0.0930
+ -6.750 -0.5644 0.07038 0.06602 -0.0424 1.0000 0.0952
+ -6.500 -0.5568 0.06767 0.06340 -0.0394 1.0000 0.0981
+ -6.000 -0.5577 0.06007 0.05560 -0.0409 1.0000 0.1107
+ -5.750 -0.5538 0.05630 0.05158 -0.0425 1.0000 0.1221
+ -5.500 -0.5424 0.05303 0.04831 -0.0415 1.0000 0.1282
+ -5.250 -0.5310 0.04954 0.04465 -0.0420 1.0000 0.1420
+ -5.000 -0.4812 0.03622 0.02939 -0.0453 1.0000 0.0351
+ -4.750 -0.4531 0.03353 0.02578 -0.0443 1.0000 0.0301
+ -4.500 -0.4286 0.03079 0.02260 -0.0440 1.0000 0.0296
+ -4.250 -0.4048 0.02740 0.01899 -0.0446 1.0000 0.0343
+ -4.000 -0.3796 0.02596 0.01705 -0.0439 1.0000 0.0400
+ -3.750 -0.3531 0.02405 0.01481 -0.0429 1.0000 0.0422
+ -3.500 -0.3281 0.02173 0.01255 -0.0423 1.0000 0.0477
+ -3.250 -0.3027 0.02004 0.01098 -0.0418 1.0000 0.0755
+ -3.000 -0.2723 0.01606 0.00928 -0.0433 1.0000 0.4902
+ -2.750 -0.2564 0.01636 0.00975 -0.0399 1.0000 0.6148
+ -2.500 -0.2408 0.01663 0.01002 -0.0367 1.0000 0.6677
+ -2.250 -0.2251 0.01681 0.01016 -0.0338 1.0000 0.7060
+ -2.000 -0.2098 0.01691 0.01023 -0.0308 1.0000 0.7390
+ -1.750 -0.1965 0.01698 0.01029 -0.0273 1.0000 0.7724
+ -1.500 -0.1831 0.01696 0.01026 -0.0241 1.0000 0.8020
+ -1.250 -0.1656 0.01690 0.01003 -0.0222 1.0000 0.8249
+ -1.000 -0.1464 0.01684 0.00990 -0.0208 1.0000 0.8459
+ -0.750 -0.1278 0.01676 0.00978 -0.0194 1.0000 0.8700
+ -0.500 -0.1084 0.01666 0.00968 -0.0181 1.0000 0.9019
+ -0.250 -0.0737 0.01645 0.00948 -0.0205 1.0000 0.9628
+ 0.000 -0.0417 0.01647 0.00934 -0.0235 1.0000 1.0000
+ 0.250 -0.0098 0.01676 0.00941 -0.0261 1.0000 1.0000
+ 0.500 0.0187 0.01711 0.00960 -0.0277 1.0000 1.0000
+ 0.750 0.0450 0.01748 0.00985 -0.0287 1.0000 1.0000
+ 1.000 0.0697 0.01789 0.01016 -0.0293 1.0000 1.0000
+ 1.250 0.0933 0.01832 0.01052 -0.0297 1.0000 1.0000
+ 1.500 0.1282 0.01902 0.01118 -0.0323 0.9954 1.0000
+ 1.750 0.1689 0.01978 0.01192 -0.0359 0.9870 1.0000
+ 2.000 0.2111 0.02062 0.01276 -0.0397 0.9791 1.0000
+ 2.250 0.2513 0.02129 0.01346 -0.0430 0.9701 1.0000
+ 2.500 0.2886 0.02182 0.01404 -0.0457 0.9598 1.0000
+ 2.750 0.3275 0.02236 0.01475 -0.0486 0.9492 1.0000
+ 3.000 0.3686 0.02284 0.01536 -0.0518 0.9379 1.0000
+ 3.250 0.4113 0.02324 0.01590 -0.0550 0.9258 1.0000
+ 3.500 0.4554 0.02347 0.01632 -0.0582 0.9120 1.0000
+ 3.750 0.5060 0.02338 0.01649 -0.0620 0.8955 1.0000
+ 4.000 0.5750 0.02221 0.01580 -0.0674 0.8708 1.0000
+ 4.250 0.6796 0.01851 0.01278 -0.0758 0.8374 1.0000
+ 4.500 0.7379 0.01505 0.00974 -0.0749 0.7585 1.0000
+ 4.750 0.7450 0.01758 0.00888 -0.0676 0.1726 1.0000
+ 5.000 0.7524 0.02028 0.01087 -0.0638 0.0985 1.0000
+ 5.250 0.7694 0.02244 0.01282 -0.0615 0.0644 1.0000
+ 5.500 0.7934 0.02479 0.01511 -0.0605 0.0429 1.0000
+ 5.750 0.8295 0.02900 0.01927 -0.0613 0.0363 1.0000
+ 6.000 0.8602 0.03196 0.02284 -0.0604 0.0349 1.0000
+ 6.250 0.8843 0.03444 0.02582 -0.0588 0.0319 1.0000
+ 6.500 0.9024 0.03664 0.02811 -0.0579 0.0263 1.0000
+ 6.750 0.9181 0.04142 0.03332 -0.0561 0.0253 1.0000
+ 7.000 0.9364 0.04376 0.03613 -0.0537 0.0262 1.0000
+ 7.250 0.9460 0.04873 0.04203 -0.0492 0.0310 1.0000
+ 7.500 0.9527 0.05347 0.04712 -0.0465 0.0340 1.0000
+ 9.500 0.8657 0.09476 0.09067 -0.0265 0.0870 1.0000
+ 9.750 0.8388 0.09881 0.09481 -0.0259 0.0870 1.0000
+ 10.000 0.8129 0.10362 0.09971 -0.0271 0.0869 1.0000
diff --git a/Airfoils/Polars/E851_polar_Re_1000000.txt b/Airfoils/Polars/E851_polar_Re_1000000.txt
new file mode 100644
index 0000000..69a2cdf
--- /dev/null
+++ b/Airfoils/Polars/E851_polar_Re_1000000.txt
@@ -0,0 +1,79 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E851 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.500 -0.4662 0.09767 0.09607 -0.0327 1.0000 0.0056
+ -9.250 -0.4692 0.09407 0.09249 -0.0332 1.0000 0.0059
+ -9.000 -0.4744 0.09059 0.08903 -0.0334 1.0000 0.0059
+ -8.750 -0.4752 0.08623 0.08470 -0.0355 0.9995 0.0059
+ -8.500 -0.4655 0.08056 0.07903 -0.0414 0.9980 0.0063
+ -8.250 -0.4569 0.07398 0.07246 -0.0489 0.9959 0.0062
+ -8.000 -0.4474 0.06513 0.06361 -0.0614 0.9930 0.0062
+ -7.750 -0.3779 0.04143 0.03985 -0.0786 0.9843 0.0068
+ -7.500 -0.3667 0.03514 0.03338 -0.0864 0.9818 0.0071
+ -7.250 -0.3515 0.03215 0.03026 -0.0882 0.9765 0.0076
+ -7.000 -0.3361 0.02772 0.02561 -0.0914 0.9726 0.0077
+ -6.750 -0.3169 0.02308 0.02071 -0.0947 0.9703 0.0078
+ -6.500 -0.2945 0.01892 0.01625 -0.0976 0.9687 0.0078
+ -6.250 -0.2799 0.01580 0.01286 -0.0975 0.9615 0.0078
+ -5.250 -0.2053 0.01174 0.00701 -0.1000 0.9509 0.0042
+ -5.000 -0.1764 0.01050 0.00555 -0.1002 0.9471 0.0033
+ -4.750 -0.1447 0.00978 0.00474 -0.1011 0.9444 0.0031
+ -4.500 -0.1166 0.00905 0.00388 -0.1013 0.9399 0.0029
+ -4.250 -0.0889 0.00828 0.00294 -0.1014 0.9349 0.0030
+ -4.000 -0.0583 0.00776 0.00219 -0.1020 0.9313 0.0036
+ -3.750 -0.0304 0.00737 0.00183 -0.1021 0.9266 0.0261
+ -3.500 -0.0033 0.00692 0.00162 -0.1023 0.9220 0.0887
+ -3.250 0.0245 0.00626 0.00137 -0.1029 0.9181 0.2133
+ -3.000 0.0502 0.00564 0.00116 -0.1031 0.9135 0.3506
+ -2.750 0.0765 0.00521 0.00115 -0.1032 0.9090 0.4816
+ -2.500 0.1055 0.00517 0.00111 -0.1035 0.9055 0.5048
+ -2.250 0.1337 0.00514 0.00108 -0.1037 0.9018 0.5186
+ -2.000 0.1614 0.00512 0.00106 -0.1038 0.8978 0.5315
+ -1.750 0.1896 0.00509 0.00105 -0.1039 0.8941 0.5466
+ -1.500 0.2186 0.00508 0.00102 -0.1042 0.8908 0.5607
+ -1.250 0.2459 0.00506 0.00103 -0.1042 0.8867 0.5729
+ -1.000 0.2737 0.00505 0.00102 -0.1043 0.8825 0.5830
+ -0.750 0.3023 0.00505 0.00101 -0.1045 0.8783 0.5914
+ -0.500 0.3294 0.00503 0.00102 -0.1044 0.8729 0.5996
+ -0.250 0.3571 0.00502 0.00101 -0.1043 0.8672 0.6082
+ 0.000 0.3844 0.00501 0.00103 -0.1043 0.8611 0.6169
+ 0.250 0.4117 0.00500 0.00102 -0.1042 0.8540 0.6260
+ 0.500 0.4388 0.00499 0.00104 -0.1040 0.8465 0.6360
+ 0.750 0.4661 0.00499 0.00105 -0.1039 0.8390 0.6463
+ 1.000 0.4924 0.00497 0.00110 -0.1035 0.8286 0.6572
+ 1.250 0.5186 0.00496 0.00112 -0.1032 0.8163 0.6689
+ 1.500 0.5449 0.00496 0.00115 -0.1028 0.8034 0.6815
+ 1.750 0.5714 0.00496 0.00120 -0.1025 0.7912 0.6952
+ 2.000 0.5974 0.00498 0.00125 -0.1021 0.7758 0.7102
+ 2.250 0.6211 0.00506 0.00129 -0.1011 0.7411 0.7266
+ 2.500 0.6392 0.00541 0.00142 -0.0989 0.6588 0.7441
+ 2.750 0.6472 0.00651 0.00179 -0.0950 0.4785 0.7635
+ 3.000 0.6536 0.00806 0.00234 -0.0913 0.2359 0.7853
+ 3.250 0.6692 0.00900 0.00275 -0.0892 0.1012 0.8111
+ 3.500 0.6908 0.00935 0.00302 -0.0881 0.0592 0.8425
+ 3.750 0.7103 0.00972 0.00332 -0.0864 0.0208 0.8849
+ 4.000 0.7320 0.00999 0.00364 -0.0850 0.0036 1.0000
+ 4.250 0.7559 0.01048 0.00429 -0.0841 0.0021 1.0000
+ 4.500 0.7776 0.01124 0.00521 -0.0827 0.0020 1.0000
+ 4.750 0.7982 0.01213 0.00624 -0.0811 0.0020 1.0000
+ 5.000 0.8194 0.01301 0.00721 -0.0797 0.0020 1.0000
+ 5.250 0.8413 0.01388 0.00822 -0.0784 0.0021 1.0000
+ 5.500 0.8642 0.01471 0.00913 -0.0774 0.0023 1.0000
+ 6.750 0.9391 0.03778 0.03421 -0.0632 0.0056 1.0000
+ 7.000 0.9471 0.04117 0.03788 -0.0601 0.0056 1.0000
+ 7.250 0.9529 0.04465 0.04162 -0.0570 0.0056 1.0000
+ 7.500 0.9565 0.04819 0.04540 -0.0538 0.0056 1.0000
+ 7.750 0.9575 0.05188 0.04932 -0.0506 0.0055 1.0000
+ 8.000 0.9569 0.05549 0.05312 -0.0475 0.0055 1.0000
+ 8.250 0.9542 0.05906 0.05688 -0.0444 0.0055 1.0000
+ 8.500 0.9501 0.06241 0.06037 -0.0414 0.0054 1.0000
+ 8.750 0.9466 0.06487 0.06294 -0.0381 0.0052 1.0000
diff --git a/Airfoils/Polars/E851_polar_Re_200000.txt b/Airfoils/Polars/E851_polar_Re_200000.txt
new file mode 100644
index 0000000..6fc7b54
--- /dev/null
+++ b/Airfoils/Polars/E851_polar_Re_200000.txt
@@ -0,0 +1,86 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E851 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.000 -0.3907 0.08745 0.08426 -0.0382 1.0000 0.0331
+ -8.750 -0.3965 0.08409 0.08095 -0.0377 1.0000 0.0336
+ -8.500 -0.4042 0.08080 0.07771 -0.0371 1.0000 0.0341
+ -8.250 -0.4143 0.07752 0.07449 -0.0363 1.0000 0.0345
+ -8.000 -0.5190 0.08168 0.07841 -0.0453 1.0000 0.0298
+ -7.750 -0.5348 0.07895 0.07565 -0.0446 1.0000 0.0299
+ -7.250 -0.5648 0.06912 0.06588 -0.0430 1.0000 0.0309
+ -7.000 -0.5688 0.06594 0.06272 -0.0417 1.0000 0.0315
+ -6.750 -0.5715 0.06282 0.05957 -0.0409 1.0000 0.0321
+ -6.500 -0.5713 0.05970 0.05638 -0.0405 1.0000 0.0329
+ -6.250 -0.5686 0.05625 0.05282 -0.0406 1.0000 0.0339
+ -6.000 -0.5624 0.05258 0.04898 -0.0409 1.0000 0.0352
+ -5.750 -0.5453 0.04836 0.04446 -0.0430 0.9988 0.0374
+ -5.500 -0.5146 0.04315 0.03839 -0.0478 0.9954 0.0431
+ -5.000 -0.4617 0.03556 0.03049 -0.0516 0.9910 0.0477
+ -4.750 -0.4223 0.02847 0.02235 -0.0502 0.9902 0.0161
+ -4.500 -0.3921 0.02501 0.01837 -0.0510 0.9886 0.0148
+ -4.250 -0.3607 0.02242 0.01535 -0.0515 0.9870 0.0141
+ -4.000 -0.3287 0.02040 0.01294 -0.0519 0.9852 0.0140
+ -3.750 -0.2963 0.01877 0.01113 -0.0526 0.9835 0.0150
+ -3.500 -0.2627 0.01771 0.00992 -0.0538 0.9819 0.0171
+ -3.250 -0.2296 0.01579 0.00819 -0.0554 0.9805 0.0794
+ -3.000 -0.2044 0.01369 0.00791 -0.0566 0.9784 0.4965
+ -2.750 -0.1763 0.01376 0.00796 -0.0567 0.9749 0.5524
+ -2.500 -0.1461 0.01396 0.00815 -0.0572 0.9719 0.5993
+ -2.250 -0.1133 0.01418 0.00833 -0.0583 0.9695 0.6290
+ -2.000 -0.0867 0.01428 0.00838 -0.0582 0.9659 0.6511
+ -1.750 -0.0602 0.01439 0.00843 -0.0579 0.9619 0.6773
+ -1.500 -0.0297 0.01451 0.00857 -0.0585 0.9588 0.6990
+ -1.250 0.0051 0.01463 0.00865 -0.0601 0.9565 0.7135
+ -1.000 0.0298 0.01465 0.00866 -0.0598 0.9520 0.7256
+ -0.750 0.0591 0.01469 0.00870 -0.0604 0.9480 0.7388
+ -0.500 0.0929 0.01475 0.00874 -0.0619 0.9451 0.7527
+ -0.250 0.1278 0.01483 0.00884 -0.0635 0.9426 0.7679
+ 0.000 0.1490 0.01483 0.00889 -0.0625 0.9365 0.7838
+ 0.250 0.1817 0.01484 0.00897 -0.0636 0.9330 0.8016
+ 0.500 0.2182 0.01484 0.00906 -0.0654 0.9305 0.8231
+ 0.750 0.2366 0.01482 0.00916 -0.0636 0.9234 0.8493
+ 1.000 0.2694 0.01471 0.00919 -0.0645 0.9197 0.8854
+ 1.250 0.3193 0.01453 0.00918 -0.0690 0.9178 0.9604
+ 1.500 0.3494 0.01459 0.00923 -0.0702 0.9106 1.0000
+ 1.750 0.3916 0.01453 0.00919 -0.0733 0.9065 1.0000
+ 2.000 0.4296 0.01447 0.00923 -0.0755 0.9009 1.0000
+ 2.250 0.4705 0.01422 0.00906 -0.0779 0.8944 1.0000
+ 2.500 0.5337 0.01339 0.00837 -0.0841 0.8912 1.0000
+ 2.750 0.5920 0.01216 0.00728 -0.0886 0.8791 1.0000
+ 3.000 0.6405 0.01119 0.00653 -0.0913 0.8655 1.0000
+ 3.250 0.6733 0.01069 0.00614 -0.0913 0.8481 1.0000
+ 3.500 0.7052 0.01006 0.00561 -0.0906 0.8193 1.0000
+ 3.750 0.7309 0.00961 0.00516 -0.0887 0.7655 1.0000
+ 4.000 0.7532 0.00972 0.00474 -0.0861 0.6160 1.0000
+ 4.250 0.7489 0.01154 0.00524 -0.0797 0.3779 1.0000
+ 4.500 0.7460 0.01395 0.00628 -0.0747 0.1318 1.0000
+ 4.750 0.7569 0.01579 0.00750 -0.0718 0.0454 1.0000
+ 5.000 0.7728 0.01736 0.00897 -0.0694 0.0176 1.0000
+ 5.250 0.7883 0.01942 0.01111 -0.0670 0.0131 1.0000
+ 5.500 0.8117 0.02108 0.01290 -0.0658 0.0119 1.0000
+ 5.750 0.8392 0.02347 0.01561 -0.0651 0.0111 1.0000
+ 6.000 0.8675 0.02629 0.01874 -0.0645 0.0111 1.0000
+ 6.250 0.8925 0.02958 0.02245 -0.0632 0.0114 1.0000
+ 6.500 0.9119 0.03332 0.02666 -0.0612 0.0120 1.0000
+ 6.750 0.9243 0.03793 0.03174 -0.0585 0.0129 1.0000
+ 7.000 0.9546 0.04634 0.04093 -0.0540 0.0389 1.0000
+ 7.250 0.9669 0.05069 0.04537 -0.0525 0.0379 1.0000
+ 7.500 0.9604 0.05424 0.04956 -0.0479 0.0343 1.0000
+ 7.750 0.9635 0.05759 0.05326 -0.0447 0.0324 1.0000
+ 8.000 0.9634 0.06131 0.05725 -0.0419 0.0312 1.0000
+ 8.250 0.9601 0.06508 0.06124 -0.0393 0.0304 1.0000
+ 8.500 0.9535 0.06879 0.06515 -0.0368 0.0297 1.0000
+ 8.750 0.9429 0.07230 0.06880 -0.0341 0.0292 1.0000
+ 9.000 0.9274 0.07564 0.07226 -0.0313 0.0290 1.0000
+ 9.250 0.9100 0.07935 0.07609 -0.0296 0.0288 1.0000
+ 9.500 0.8912 0.08370 0.08054 -0.0294 0.0288 1.0000
+ 9.750 0.8722 0.08892 0.08584 -0.0310 0.0289 1.0000
diff --git a/Airfoils/Polars/E851_polar_Re_500000.txt b/Airfoils/Polars/E851_polar_Re_500000.txt
new file mode 100644
index 0000000..2de7ddc
--- /dev/null
+++ b/Airfoils/Polars/E851_polar_Re_500000.txt
@@ -0,0 +1,84 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E851 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -8.500 -0.4789 0.08530 0.08318 -0.0354 1.0000 0.0116
+ -8.250 -0.4912 0.08243 0.08036 -0.0343 1.0000 0.0117
+ -8.000 -0.5019 0.07906 0.07704 -0.0346 0.9994 0.0116
+ -7.750 -0.4936 0.07197 0.06996 -0.0443 0.9954 0.0118
+ -7.500 -0.4810 0.06102 0.05889 -0.0624 0.9880 0.0117
+ -7.250 -0.4646 0.05391 0.05157 -0.0716 0.9835 0.0121
+ -7.000 -0.4440 0.04772 0.04512 -0.0782 0.9805 0.0128
+ -6.750 -0.4291 0.04312 0.04028 -0.0804 0.9741 0.0137
+ -6.500 -0.4034 0.03874 0.03559 -0.0834 0.9713 0.0151
+ -6.250 -0.3669 0.03800 0.03457 -0.0848 0.9699 0.0173
+ -6.000 -0.3450 0.03464 0.03088 -0.0858 0.9656 0.0174
+ -5.750 -0.3219 0.03131 0.02722 -0.0866 0.9615 0.0174
+ -5.250 -0.2676 0.01856 0.01333 -0.0883 0.9578 0.0070
+ -5.000 -0.2359 0.01567 0.01003 -0.0891 0.9568 0.0062
+ -4.750 -0.2032 0.01371 0.00785 -0.0901 0.9560 0.0058
+ -4.500 -0.1831 0.01257 0.00655 -0.0885 0.9503 0.0057
+ -4.250 -0.1518 0.01151 0.00532 -0.0894 0.9483 0.0059
+ -4.000 -0.1181 0.01077 0.00439 -0.0907 0.9469 0.0070
+ -3.750 -0.0840 0.00987 0.00362 -0.0922 0.9458 0.0630
+ -3.500 -0.0526 0.00844 0.00314 -0.0944 0.9448 0.3001
+ -3.250 -0.0203 0.00776 0.00306 -0.0960 0.9438 0.4790
+ -3.000 0.0078 0.00768 0.00300 -0.0962 0.9406 0.5092
+ -2.750 0.0362 0.00762 0.00291 -0.0964 0.9372 0.5388
+ -2.500 0.0680 0.00755 0.00282 -0.0973 0.9351 0.5558
+ -2.250 0.1006 0.00746 0.00272 -0.0984 0.9333 0.5713
+ -2.000 0.1337 0.00738 0.00266 -0.0996 0.9318 0.5895
+ -1.750 0.1671 0.00731 0.00260 -0.1008 0.9304 0.6054
+ -1.500 0.1919 0.00731 0.00260 -0.1003 0.9258 0.6164
+ -1.250 0.2214 0.00726 0.00255 -0.1007 0.9227 0.6260
+ -1.000 0.2529 0.00720 0.00249 -0.1016 0.9203 0.6356
+ -0.750 0.2853 0.00711 0.00243 -0.1026 0.9182 0.6446
+ -0.500 0.3128 0.00710 0.00244 -0.1026 0.9142 0.6540
+ -0.250 0.3405 0.00707 0.00245 -0.1026 0.9098 0.6643
+ 0.000 0.3716 0.00699 0.00241 -0.1033 0.9063 0.6750
+ 0.250 0.4006 0.00694 0.00241 -0.1036 0.9019 0.6860
+ 0.500 0.4275 0.00689 0.00241 -0.1034 0.8958 0.6978
+ 0.750 0.4588 0.00679 0.00237 -0.1040 0.8908 0.7107
+ 1.000 0.4838 0.00673 0.00238 -0.1032 0.8822 0.7246
+ 1.250 0.5113 0.00666 0.00241 -0.1030 0.8741 0.7398
+ 1.500 0.5389 0.00656 0.00236 -0.1027 0.8638 0.7566
+ 1.750 0.5639 0.00645 0.00233 -0.1017 0.8499 0.7748
+ 2.250 0.6139 0.00630 0.00237 -0.1000 0.8247 0.8191
+ 2.500 0.6369 0.00623 0.00247 -0.0987 0.8090 0.8469
+ 2.750 0.6568 0.00614 0.00246 -0.0965 0.7794 0.8815
+ 3.000 0.6745 0.00609 0.00242 -0.0938 0.7347 0.9341
+ 3.250 0.6992 0.00647 0.00245 -0.0929 0.6337 1.0000
+ 3.500 0.7000 0.00795 0.00291 -0.0877 0.4276 1.0000
+ 3.750 0.7051 0.00970 0.00354 -0.0839 0.1996 1.0000
+ 4.000 0.7200 0.01089 0.00406 -0.0820 0.0737 1.0000
+ 4.250 0.7407 0.01160 0.00449 -0.0809 0.0255 1.0000
+ 4.500 0.7607 0.01257 0.00548 -0.0791 0.0051 1.0000
+ 4.750 0.7822 0.01335 0.00640 -0.0777 0.0045 1.0000
+ 5.000 0.8021 0.01438 0.00756 -0.0759 0.0042 1.0000
+ 5.250 0.8212 0.01569 0.00902 -0.0740 0.0041 1.0000
+ 5.500 0.8420 0.01741 0.01098 -0.0724 0.0042 1.0000
+ 5.750 0.8665 0.01987 0.01369 -0.0712 0.0045 1.0000
+ 6.000 0.8916 0.02324 0.01740 -0.0701 0.0049 1.0000
+ 6.250 0.9153 0.02585 0.02021 -0.0689 0.0062 1.0000
+ 6.500 0.9177 0.03900 0.03440 -0.0627 0.0136 1.0000
+ 6.750 0.9281 0.04245 0.03815 -0.0600 0.0135 1.0000
+ 7.000 0.9362 0.04598 0.04197 -0.0571 0.0135 1.0000
+ 7.250 0.9422 0.04947 0.04574 -0.0541 0.0135 1.0000
+ 7.500 0.9461 0.05292 0.04945 -0.0512 0.0134 1.0000
+ 7.750 0.9481 0.05627 0.05305 -0.0481 0.0133 1.0000
+ 8.000 0.9625 0.05739 0.05437 -0.0457 0.0125 1.0000
+ 8.250 0.9707 0.06019 0.05737 -0.0432 0.0114 1.0000
+ 8.500 0.9666 0.06400 0.06137 -0.0403 0.0110 1.0000
+ 8.750 0.9565 0.06764 0.06517 -0.0372 0.0107 1.0000
+ 9.000 0.9398 0.07098 0.06863 -0.0337 0.0106 1.0000
+ 9.250 0.9218 0.07463 0.07239 -0.0314 0.0106 1.0000
+ 9.500 0.9018 0.07895 0.07682 -0.0306 0.0107 1.0000
+ 9.750 0.8818 0.08405 0.08200 -0.0317 0.0109 1.0000
diff --git a/Airfoils/Polars/E854_polar_Re_100000.txt b/Airfoils/Polars/E854_polar_Re_100000.txt
new file mode 100644
index 0000000..c246917
--- /dev/null
+++ b/Airfoils/Polars/E854_polar_Re_100000.txt
@@ -0,0 +1,110 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E854 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.000 -0.3733 0.09816 0.09363 -0.0426 1.0000 0.1241
+ -8.750 -0.4009 0.09532 0.09095 -0.0448 1.0000 0.1305
+ -8.500 -0.4470 0.09265 0.08850 -0.0468 1.0000 0.1314
+ -8.250 -0.4036 0.08909 0.08487 -0.0416 1.0000 0.1357
+ -8.000 -0.4105 0.08678 0.08265 -0.0397 1.0000 0.1391
+ -7.750 -0.4374 0.08490 0.08092 -0.0380 1.0000 0.1427
+ -7.250 -0.5402 0.07932 0.07559 -0.0414 1.0000 0.1454
+ -6.750 -0.5644 0.05043 0.04476 -0.0558 0.9954 0.0503
+ -6.500 -0.5337 0.04480 0.03875 -0.0599 0.9891 0.0493
+ -6.250 -0.4965 0.04013 0.03336 -0.0645 0.9833 0.0496
+ -6.000 -0.4631 0.03595 0.02863 -0.0671 0.9761 0.0491
+ -5.750 -0.4207 0.03279 0.02480 -0.0704 0.9710 0.0500
+ -5.500 -0.3889 0.02976 0.02169 -0.0722 0.9635 0.0550
+ -5.250 -0.3458 0.02773 0.01933 -0.0750 0.9586 0.0640
+ -5.000 -0.3132 0.02598 0.01761 -0.0764 0.9499 0.0785
+ -4.750 -0.2742 0.02444 0.01608 -0.0787 0.9436 0.1030
+ -4.500 -0.2397 0.02275 0.01471 -0.0806 0.9358 0.1447
+ -4.250 -0.2038 0.02005 0.01353 -0.0841 0.9301 0.3558
+ -4.000 -0.1694 0.02059 0.01424 -0.0845 0.9208 0.4986
+ -3.750 -0.1389 0.02110 0.01463 -0.0843 0.9112 0.5405
+ -3.500 -0.0993 0.02143 0.01486 -0.0855 0.9053 0.5744
+ -3.250 -0.0723 0.02168 0.01503 -0.0848 0.8951 0.5976
+ -3.000 -0.0304 0.02170 0.01492 -0.0866 0.8906 0.6224
+ -2.750 -0.0058 0.02184 0.01500 -0.0855 0.8804 0.6403
+ -2.500 0.0348 0.02175 0.01485 -0.0867 0.8765 0.6616
+ -2.250 0.0590 0.02180 0.01483 -0.0858 0.8666 0.6776
+ -2.000 0.1032 0.02147 0.01436 -0.0884 0.8628 0.6923
+ -1.750 0.1295 0.02140 0.01421 -0.0882 0.8533 0.7028
+ -1.500 0.1731 0.02100 0.01369 -0.0911 0.8490 0.7140
+ -1.250 0.2008 0.02095 0.01354 -0.0914 0.8401 0.7243
+ -1.000 0.2400 0.02061 0.01313 -0.0932 0.8351 0.7335
+ -0.750 0.2671 0.02059 0.01305 -0.0934 0.8264 0.7432
+ -0.500 0.3055 0.02032 0.01269 -0.0954 0.8209 0.7538
+ -0.250 0.3296 0.02037 0.01273 -0.0948 0.8126 0.7627
+ 0.000 0.3649 0.02016 0.01247 -0.0962 0.8067 0.7731
+ 0.250 0.3901 0.02026 0.01255 -0.0960 0.7986 0.7842
+ 0.500 0.4211 0.02016 0.01243 -0.0965 0.7925 0.7950
+ 0.750 0.4438 0.02029 0.01258 -0.0957 0.7847 0.8064
+ 1.000 0.4727 0.02024 0.01255 -0.0959 0.7783 0.8191
+ 1.250 0.4955 0.02040 0.01273 -0.0951 0.7712 0.8329
+ 1.500 0.5204 0.02045 0.01281 -0.0946 0.7644 0.8483
+ 1.750 0.5451 0.02052 0.01291 -0.0940 0.7582 0.8659
+ 2.000 0.5635 0.02072 0.01320 -0.0924 0.7508 0.8870
+ 2.250 0.5976 0.02047 0.01299 -0.0931 0.7467 0.9123
+ 2.500 0.6222 0.02092 0.01358 -0.0934 0.7376 0.9497
+ 2.750 0.6825 0.02075 0.01345 -0.1002 0.7333 1.0000
+ 3.000 0.7041 0.02151 0.01420 -0.1012 0.7246 1.0000
+ 3.250 0.7436 0.02163 0.01428 -0.1039 0.7195 1.0000
+ 3.500 0.7645 0.02243 0.01509 -0.1039 0.7113 1.0000
+ 3.750 0.7991 0.02262 0.01530 -0.1054 0.7057 1.0000
+ 4.000 0.8194 0.02337 0.01609 -0.1048 0.6978 1.0000
+ 4.250 0.8514 0.02359 0.01634 -0.1057 0.6916 1.0000
+ 4.500 0.8719 0.02428 0.01711 -0.1049 0.6835 1.0000
+ 4.750 0.9049 0.02437 0.01724 -0.1056 0.6768 1.0000
+ 5.000 0.9236 0.02505 0.01802 -0.1045 0.6678 1.0000
+ 5.250 0.9625 0.02479 0.01780 -0.1058 0.6612 1.0000
+ 5.500 0.9805 0.02536 0.01851 -0.1043 0.6506 1.0000
+ 5.750 1.0108 0.02526 0.01848 -0.1041 0.6407 1.0000
+ 6.000 1.0522 0.02447 0.01773 -0.1052 0.6302 1.0000
+ 6.250 1.0817 0.02413 0.01749 -0.1047 0.6174 1.0000
+ 6.500 1.1079 0.02393 0.01738 -0.1037 0.6038 1.0000
+ 6.750 1.1342 0.02370 0.01725 -0.1028 0.5900 1.0000
+ 7.000 1.1616 0.02336 0.01700 -0.1019 0.5753 1.0000
+ 7.250 1.1902 0.02288 0.01659 -0.1010 0.5592 1.0000
+ 7.500 1.2081 0.02272 0.01662 -0.0987 0.5406 1.0000
+ 7.750 1.2281 0.02232 0.01633 -0.0964 0.5192 1.0000
+ 8.000 1.2430 0.02206 0.01618 -0.0934 0.4944 1.0000
+ 8.250 1.2537 0.02187 0.01608 -0.0897 0.4634 1.0000
+ 8.500 1.2586 0.02194 0.01616 -0.0852 0.4224 1.0000
+ 8.750 1.2541 0.02245 0.01644 -0.0793 0.3616 1.0000
+ 9.000 1.2392 0.02382 0.01729 -0.0724 0.2922 1.0000
+ 9.250 1.2218 0.02591 0.01888 -0.0662 0.2372 1.0000
+ 9.500 1.2074 0.02824 0.02084 -0.0611 0.1973 1.0000
+ 9.750 1.1967 0.03064 0.02297 -0.0570 0.1675 1.0000
+ 10.000 1.1903 0.03304 0.02515 -0.0537 0.1447 1.0000
+ 10.250 1.1881 0.03536 0.02732 -0.0510 0.1253 1.0000
+ 10.500 1.1884 0.03770 0.02951 -0.0488 0.1091 1.0000
+ 10.750 1.1940 0.03995 0.03164 -0.0469 0.0944 1.0000
+ 11.000 1.2031 0.04214 0.03375 -0.0453 0.0816 1.0000
+ 11.250 1.2122 0.04424 0.03589 -0.0439 0.0715 1.0000
+ 11.500 1.2238 0.04651 0.03828 -0.0425 0.0622 1.0000
+ 11.750 1.2393 0.04899 0.04079 -0.0415 0.0542 1.0000
+ 12.000 1.2523 0.05135 0.04319 -0.0405 0.0483 1.0000
+ 12.250 1.2673 0.05453 0.04663 -0.0394 0.0433 1.0000
+ 12.500 1.2710 0.05719 0.04951 -0.0381 0.0399 1.0000
+ 12.750 1.2835 0.06035 0.05279 -0.0374 0.0375 1.0000
+ 13.000 1.2965 0.06618 0.05892 -0.0369 0.0357 1.0000
+ 13.250 1.2847 0.06969 0.06280 -0.0354 0.0353 1.0000
+ 13.500 1.2710 0.07367 0.06712 -0.0344 0.0349 1.0000
+ 13.750 1.2552 0.07801 0.07179 -0.0340 0.0345 1.0000
+ 14.000 1.2384 0.08286 0.07694 -0.0341 0.0343 1.0000
+ 14.250 1.2205 0.08810 0.08246 -0.0349 0.0343 1.0000
+ 14.500 1.2009 0.09378 0.08840 -0.0363 0.0343 1.0000
+ 14.750 1.1806 0.09987 0.09473 -0.0383 0.0344 1.0000
+ 15.000 1.1598 0.10646 0.10154 -0.0410 0.0346 1.0000
+ 15.250 1.1401 0.11344 0.10870 -0.0443 0.0350 1.0000
+ 15.500 1.1193 0.12098 0.11641 -0.0483 0.0353 1.0000
+ 15.750 1.1014 0.12881 0.12432 -0.0525 0.0357 1.0000
diff --git a/Airfoils/Polars/E854_polar_Re_1000000.txt b/Airfoils/Polars/E854_polar_Re_1000000.txt
new file mode 100644
index 0000000..a1dc284
--- /dev/null
+++ b/Airfoils/Polars/E854_polar_Re_1000000.txt
@@ -0,0 +1,122 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E854 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -10.500 -0.5365 0.06377 0.06217 -0.0569 1.0000 0.0068
+ -9.750 -0.6095 0.02614 0.02271 -0.0978 0.9832 0.0052
+ -9.500 -0.5948 0.02392 0.02024 -0.0976 0.9765 0.0051
+ -9.250 -0.5680 0.02181 0.01789 -0.0996 0.9738 0.0051
+ -9.000 -0.5379 0.02004 0.01592 -0.1017 0.9720 0.0051
+ -8.750 -0.5187 0.01846 0.01417 -0.1011 0.9654 0.0052
+ -8.500 -0.4910 0.01691 0.01247 -0.1023 0.9615 0.0053
+ -8.250 -0.4595 0.01581 0.01127 -0.1039 0.9591 0.0055
+ -8.000 -0.4397 0.01487 0.01024 -0.1028 0.9498 0.0056
+ -7.750 -0.4087 0.01392 0.00920 -0.1041 0.9458 0.0057
+ -7.500 -0.3815 0.01298 0.00815 -0.1045 0.9370 0.0058
+ -7.250 -0.3430 0.01204 0.00710 -0.1073 0.9326 0.0058
+ -7.000 -0.3032 0.01123 0.00618 -0.1104 0.9248 0.0060
+ -6.750 -0.2552 0.01051 0.00535 -0.1152 0.9175 0.0063
+ -6.500 -0.2133 0.00996 0.00467 -0.1185 0.9031 0.0067
+ -6.250 -0.1797 0.00951 0.00407 -0.1200 0.8835 0.0077
+ -6.000 -0.1515 0.00910 0.00353 -0.1203 0.8626 0.0118
+ -5.750 -0.1259 0.00876 0.00317 -0.1200 0.8429 0.0237
+ -5.500 -0.1004 0.00856 0.00293 -0.1197 0.8245 0.0339
+ -5.250 -0.0750 0.00839 0.00272 -0.1194 0.8080 0.0457
+ -5.000 -0.0495 0.00822 0.00253 -0.1190 0.7924 0.0610
+ -4.750 -0.0242 0.00797 0.00233 -0.1187 0.7777 0.0912
+ -4.500 0.0010 0.00763 0.00212 -0.1185 0.7641 0.1373
+ -4.250 0.0264 0.00725 0.00190 -0.1184 0.7515 0.2010
+ -4.000 0.0515 0.00677 0.00168 -0.1183 0.7393 0.2926
+ -3.750 0.0774 0.00648 0.00156 -0.1182 0.7279 0.3656
+ -3.500 0.1041 0.00640 0.00151 -0.1180 0.7169 0.4012
+ -3.250 0.1314 0.00636 0.00147 -0.1180 0.7065 0.4270
+ -3.000 0.1588 0.00637 0.00144 -0.1179 0.6968 0.4444
+ -2.750 0.1861 0.00639 0.00141 -0.1178 0.6874 0.4560
+ -2.500 0.2138 0.00640 0.00139 -0.1178 0.6781 0.4679
+ -2.250 0.2415 0.00644 0.00137 -0.1177 0.6698 0.4765
+ -2.000 0.2689 0.00646 0.00136 -0.1177 0.6615 0.4851
+ -1.750 0.2968 0.00649 0.00135 -0.1177 0.6537 0.4932
+ -1.250 0.3521 0.00655 0.00137 -0.1176 0.6388 0.5129
+ -1.000 0.3797 0.00662 0.00138 -0.1176 0.6319 0.5206
+ -0.750 0.4076 0.00661 0.00139 -0.1176 0.6254 0.5268
+ -0.500 0.4352 0.00666 0.00140 -0.1176 0.6187 0.5323
+ -0.250 0.4631 0.00669 0.00141 -0.1177 0.6127 0.5374
+ 0.000 0.4909 0.00672 0.00144 -0.1177 0.6067 0.5428
+ 0.250 0.5184 0.00678 0.00146 -0.1177 0.6011 0.5483
+ 0.500 0.5464 0.00679 0.00149 -0.1177 0.5954 0.5537
+ 0.750 0.5739 0.00684 0.00153 -0.1177 0.5899 0.5595
+ 1.000 0.6017 0.00689 0.00157 -0.1177 0.5850 0.5654
+ 1.250 0.6295 0.00691 0.00162 -0.1178 0.5796 0.5716
+ 1.500 0.6568 0.00699 0.00168 -0.1177 0.5745 0.5782
+ 1.750 0.6847 0.00701 0.00174 -0.1178 0.5698 0.5849
+ 2.000 0.7123 0.00705 0.00180 -0.1178 0.5647 0.5924
+ 2.250 0.7394 0.00713 0.00188 -0.1177 0.5595 0.5998
+ 2.500 0.7672 0.00716 0.00195 -0.1177 0.5538 0.6084
+ 2.750 0.7941 0.00721 0.00202 -0.1176 0.5471 0.6172
+ 3.000 0.8213 0.00726 0.00211 -0.1175 0.5404 0.6266
+ 3.250 0.8480 0.00732 0.00219 -0.1174 0.5326 0.6373
+ 3.500 0.8750 0.00737 0.00228 -0.1172 0.5245 0.6488
+ 3.750 0.9010 0.00746 0.00237 -0.1169 0.5155 0.6611
+ 4.000 0.9281 0.00749 0.00248 -0.1168 0.5068 0.6745
+ 4.250 0.9544 0.00757 0.00259 -0.1166 0.4993 0.6897
+ 4.500 0.9808 0.00763 0.00271 -0.1164 0.4903 0.7066
+ 4.750 1.0071 0.00770 0.00285 -0.1162 0.4817 0.7252
+ 5.000 1.0325 0.00779 0.00299 -0.1158 0.4711 0.7457
+ 5.250 1.0579 0.00788 0.00314 -0.1153 0.4597 0.7700
+ 5.500 1.0820 0.00799 0.00331 -0.1147 0.4432 0.7994
+ 5.750 1.1051 0.00811 0.00349 -0.1138 0.4234 0.8366
+ 6.000 1.1237 0.00825 0.00372 -0.1119 0.3991 0.8953
+ 6.250 1.1437 0.00854 0.00396 -0.1104 0.3608 1.0000
+ 6.500 1.1606 0.00924 0.00437 -0.1086 0.3048 1.0000
+ 6.750 1.1738 0.01016 0.00495 -0.1062 0.2399 1.0000
+ 7.000 1.1876 0.01101 0.00551 -0.1039 0.1874 1.0000
+ 7.250 1.2030 0.01172 0.00602 -0.1018 0.1499 1.0000
+ 7.500 1.2179 0.01240 0.00653 -0.0997 0.1171 1.0000
+ 7.750 1.2316 0.01304 0.00702 -0.0973 0.0916 1.0000
+ 8.000 1.2440 0.01367 0.00752 -0.0947 0.0711 1.0000
+ 8.250 1.2578 0.01422 0.00800 -0.0923 0.0565 1.0000
+ 8.500 1.2719 0.01478 0.00850 -0.0900 0.0450 1.0000
+ 8.750 1.2852 0.01536 0.00903 -0.0877 0.0351 1.0000
+ 9.000 1.2989 0.01594 0.00958 -0.0855 0.0285 1.0000
+ 9.250 1.3135 0.01648 0.01013 -0.0834 0.0238 1.0000
+ 9.500 1.3259 0.01712 0.01076 -0.0811 0.0191 1.0000
+ 9.750 1.3394 0.01772 0.01138 -0.0790 0.0163 1.0000
+ 10.000 1.3511 0.01843 0.01209 -0.0767 0.0130 1.0000
+ 10.250 1.3613 0.01922 0.01288 -0.0743 0.0101 1.0000
+ 10.500 1.3719 0.02002 0.01372 -0.0721 0.0080 1.0000
+ 10.750 1.3791 0.02106 0.01477 -0.0695 0.0060 1.0000
+ 11.000 1.3872 0.02209 0.01584 -0.0671 0.0047 1.0000
+ 11.250 1.3929 0.02333 0.01716 -0.0647 0.0036 1.0000
+ 11.500 1.4010 0.02447 0.01837 -0.0627 0.0033 1.0000
+ 11.750 1.4076 0.02578 0.01974 -0.0607 0.0029 1.0000
+ 12.000 1.4074 0.02767 0.02174 -0.0582 0.0025 1.0000
+ 12.250 1.4148 0.02907 0.02321 -0.0566 0.0023 1.0000
+ 12.500 1.4202 0.03070 0.02492 -0.0551 0.0022 1.0000
+ 12.750 1.4254 0.03240 0.02671 -0.0537 0.0022 1.0000
+ 13.000 1.4304 0.03420 0.02858 -0.0524 0.0020 1.0000
+ 13.250 1.4341 0.03619 0.03066 -0.0512 0.0019 1.0000
+ 13.500 1.4372 0.03829 0.03285 -0.0502 0.0018 1.0000
+ 13.750 1.4391 0.04060 0.03525 -0.0493 0.0017 1.0000
+ 14.000 1.4406 0.04305 0.03779 -0.0485 0.0017 1.0000
+ 14.250 1.4415 0.04564 0.04049 -0.0479 0.0017 1.0000
+ 14.500 1.4393 0.04865 0.04360 -0.0474 0.0016 1.0000
+ 14.750 1.4383 0.05166 0.04672 -0.0471 0.0016 1.0000
+ 15.000 1.4328 0.05531 0.05049 -0.0470 0.0015 1.0000
+ 15.250 1.4297 0.05879 0.05408 -0.0471 0.0015 1.0000
+ 15.500 1.4167 0.06373 0.05918 -0.0476 0.0014 1.0000
+ 15.750 1.4160 0.06718 0.06273 -0.0481 0.0015 1.0000
+ 16.000 1.4063 0.07199 0.06769 -0.0490 0.0014 1.0000
+ 16.250 1.3966 0.07700 0.07284 -0.0501 0.0014 1.0000
+ 16.500 1.3903 0.08166 0.07762 -0.0514 0.0014 1.0000
+ 16.750 1.3807 0.08700 0.08310 -0.0531 0.0014 1.0000
+ 17.000 1.3692 0.09280 0.08905 -0.0551 0.0014 1.0000
+ 17.250 1.3594 0.09852 0.09490 -0.0573 0.0014 1.0000
+ 17.500 1.3499 0.10434 0.10086 -0.0597 0.0014 1.0000
diff --git a/Airfoils/Polars/E854_polar_Re_200000.txt b/Airfoils/Polars/E854_polar_Re_200000.txt
new file mode 100644
index 0000000..8c097ba
--- /dev/null
+++ b/Airfoils/Polars/E854_polar_Re_200000.txt
@@ -0,0 +1,123 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E854 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -10.250 -0.3935 0.10398 0.10058 -0.0503 1.0000 0.0499
+ -10.000 -0.4021 0.09911 0.09578 -0.0529 1.0000 0.0501
+ -9.750 -0.4159 0.09339 0.09012 -0.0563 1.0000 0.0502
+ -9.500 -0.4081 0.09084 0.08762 -0.0510 1.0000 0.0517
+ -9.250 -0.4065 0.08853 0.08534 -0.0497 1.0000 0.0526
+ -9.000 -0.4115 0.08547 0.08234 -0.0494 1.0000 0.0532
+ -8.750 -0.4203 0.08223 0.07917 -0.0490 1.0000 0.0541
+ -8.500 -0.4352 0.07876 0.07578 -0.0487 1.0000 0.0547
+ -8.250 -0.4628 0.07469 0.07182 -0.0488 1.0000 0.0547
+ -7.500 -0.4934 0.04219 0.03780 -0.0792 0.9735 0.0311
+ -7.250 -0.4709 0.03517 0.03009 -0.0819 0.9663 0.0245
+ -7.000 -0.4416 0.02897 0.02286 -0.0842 0.9609 0.0219
+ -6.750 -0.4100 0.02621 0.01968 -0.0857 0.9549 0.0219
+ -6.500 -0.3762 0.02394 0.01719 -0.0874 0.9494 0.0226
+ -6.250 -0.3373 0.02213 0.01524 -0.0899 0.9462 0.0247
+ -6.000 -0.3064 0.02108 0.01403 -0.0907 0.9376 0.0280
+ -5.750 -0.2693 0.01917 0.01212 -0.0930 0.9339 0.0337
+ -5.500 -0.2376 0.01768 0.01059 -0.0941 0.9264 0.0434
+ -5.250 -0.1995 0.01640 0.00929 -0.0965 0.9216 0.0646
+ -5.000 -0.1582 0.01504 0.00806 -0.0998 0.9188 0.1001
+ -4.750 -0.1269 0.01366 0.00705 -0.1014 0.9099 0.1746
+ -4.500 -0.0865 0.01215 0.00649 -0.1052 0.9059 0.3966
+ -4.250 -0.0504 0.01206 0.00637 -0.1066 0.8973 0.4588
+ -4.000 -0.0079 0.01201 0.00622 -0.1092 0.8915 0.4945
+ -3.750 0.0273 0.01201 0.00612 -0.1105 0.8815 0.5173
+ -3.500 0.0658 0.01199 0.00597 -0.1123 0.8728 0.5369
+ -3.250 0.1020 0.01196 0.00584 -0.1138 0.8628 0.5523
+ -3.000 0.1338 0.01198 0.00575 -0.1144 0.8509 0.5660
+ -2.750 0.1649 0.01202 0.00569 -0.1148 0.8394 0.5801
+ -2.500 0.1955 0.01209 0.00568 -0.1151 0.8283 0.5944
+ -2.250 0.2262 0.01212 0.00564 -0.1154 0.8181 0.6057
+ -2.000 0.2536 0.01213 0.00558 -0.1152 0.8066 0.6141
+ -1.750 0.2819 0.01215 0.00547 -0.1154 0.7958 0.6229
+ -1.500 0.3103 0.01214 0.00541 -0.1154 0.7862 0.6293
+ -1.250 0.3393 0.01216 0.00531 -0.1157 0.7768 0.6375
+ -1.000 0.3657 0.01216 0.00529 -0.1154 0.7671 0.6436
+ -0.750 0.3947 0.01221 0.00523 -0.1156 0.7589 0.6518
+ -0.500 0.4214 0.01222 0.00523 -0.1154 0.7498 0.6582
+ -0.250 0.4492 0.01228 0.00522 -0.1155 0.7418 0.6665
+ 0.000 0.4768 0.01230 0.00523 -0.1154 0.7342 0.6733
+ 0.250 0.5039 0.01237 0.00527 -0.1154 0.7263 0.6820
+ 0.500 0.5316 0.01241 0.00531 -0.1153 0.7193 0.6896
+ 0.750 0.5582 0.01248 0.00539 -0.1152 0.7119 0.6984
+ 1.000 0.5860 0.01255 0.00544 -0.1152 0.7054 0.7076
+ 1.250 0.6122 0.01262 0.00556 -0.1149 0.6984 0.7168
+ 1.500 0.6396 0.01270 0.00565 -0.1149 0.6919 0.7273
+ 1.750 0.6666 0.01280 0.00578 -0.1148 0.6860 0.7387
+ 2.000 0.6922 0.01287 0.00593 -0.1144 0.6793 0.7502
+ 2.250 0.7206 0.01297 0.00604 -0.1145 0.6741 0.7630
+ 2.500 0.7442 0.01305 0.00626 -0.1137 0.6672 0.7774
+ 2.750 0.7710 0.01313 0.00639 -0.1135 0.6617 0.7939
+ 3.000 0.7953 0.01324 0.00661 -0.1128 0.6558 0.8128
+ 3.250 0.8185 0.01329 0.00680 -0.1118 0.6497 0.8340
+ 3.500 0.8428 0.01334 0.00693 -0.1109 0.6447 0.8615
+ 3.750 0.8613 0.01333 0.00711 -0.1088 0.6379 0.9022
+ 4.000 0.8997 0.01329 0.00714 -0.1109 0.6318 1.0000
+ 4.250 0.9272 0.01349 0.00739 -0.1113 0.6237 1.0000
+ 4.500 0.9582 0.01365 0.00748 -0.1121 0.6168 1.0000
+ 4.750 0.9823 0.01378 0.00768 -0.1115 0.6070 1.0000
+ 5.000 1.0082 0.01388 0.00780 -0.1112 0.5969 1.0000
+ 5.250 1.0351 0.01394 0.00783 -0.1109 0.5866 1.0000
+ 5.500 1.0590 0.01403 0.00797 -0.1101 0.5758 1.0000
+ 5.750 1.0827 0.01415 0.00816 -0.1093 0.5653 1.0000
+ 6.000 1.1072 0.01426 0.00833 -0.1086 0.5548 1.0000
+ 6.250 1.1317 0.01436 0.00845 -0.1079 0.5436 1.0000
+ 6.500 1.1539 0.01444 0.00860 -0.1067 0.5308 1.0000
+ 6.750 1.1750 0.01453 0.00876 -0.1053 0.5162 1.0000
+ 7.000 1.1943 0.01463 0.00894 -0.1036 0.4991 1.0000
+ 7.250 1.2124 0.01474 0.00916 -0.1016 0.4784 1.0000
+ 7.500 1.2289 0.01492 0.00936 -0.0994 0.4534 1.0000
+ 7.750 1.2426 0.01521 0.00963 -0.0967 0.4197 1.0000
+ 8.000 1.2506 0.01576 0.01000 -0.0931 0.3699 1.0000
+ 8.250 1.2483 0.01677 0.01064 -0.0879 0.3036 1.0000
+ 8.500 1.2408 0.01814 0.01163 -0.0821 0.2438 1.0000
+ 8.750 1.2345 0.01968 0.01284 -0.0770 0.1923 1.0000
+ 9.000 1.2303 0.02125 0.01416 -0.0725 0.1527 1.0000
+ 9.250 1.2277 0.02284 0.01557 -0.0686 0.1241 1.0000
+ 9.500 1.2268 0.02444 0.01707 -0.0651 0.1032 1.0000
+ 9.750 1.2260 0.02613 0.01870 -0.0619 0.0870 1.0000
+ 10.000 1.2260 0.02791 0.02048 -0.0591 0.0747 1.0000
+ 10.250 1.2258 0.02982 0.02239 -0.0566 0.0648 1.0000
+ 10.500 1.2236 0.03201 0.02455 -0.0542 0.0568 1.0000
+ 10.750 1.2275 0.03385 0.02644 -0.0525 0.0491 1.0000
+ 11.000 1.2278 0.03611 0.02876 -0.0507 0.0430 1.0000
+ 11.250 1.2294 0.03832 0.03099 -0.0492 0.0377 1.0000
+ 11.500 1.2303 0.04074 0.03350 -0.0477 0.0329 1.0000
+ 11.750 1.2332 0.04301 0.03583 -0.0466 0.0291 1.0000
+ 12.000 1.2315 0.04596 0.03880 -0.0451 0.0259 1.0000
+ 12.250 1.2375 0.04807 0.04105 -0.0443 0.0229 1.0000
+ 12.500 1.2398 0.05057 0.04357 -0.0437 0.0205 1.0000
+ 12.750 1.2444 0.05338 0.04646 -0.0423 0.0189 1.0000
+ 13.000 1.2521 0.05571 0.04901 -0.0414 0.0177 1.0000
+ 13.250 1.2594 0.05823 0.05168 -0.0406 0.0166 1.0000
+ 13.500 1.2650 0.06095 0.05456 -0.0399 0.0158 1.0000
+ 13.750 1.2682 0.06377 0.05749 -0.0396 0.0149 1.0000
+ 14.000 1.2714 0.06687 0.06068 -0.0393 0.0142 1.0000
+ 14.250 1.2706 0.07170 0.06575 -0.0387 0.0135 1.0000
+ 14.500 1.2637 0.07544 0.06974 -0.0392 0.0132 1.0000
+ 14.750 1.2568 0.07962 0.07418 -0.0399 0.0129 1.0000
+ 15.000 1.2483 0.08430 0.07911 -0.0410 0.0126 1.0000
+ 15.250 1.2387 0.08934 0.08439 -0.0424 0.0124 1.0000
+ 15.500 1.2265 0.09494 0.09024 -0.0443 0.0121 1.0000
+ 15.750 1.2126 0.10116 0.09670 -0.0466 0.0121 1.0000
+ 16.000 1.1970 0.10787 0.10365 -0.0496 0.0121 1.0000
+ 16.250 1.1802 0.11512 0.11113 -0.0533 0.0122 1.0000
+ 16.500 1.1626 0.12288 0.11910 -0.0575 0.0123 1.0000
+ 16.750 1.1443 0.13124 0.12766 -0.0625 0.0124 1.0000
+ 17.000 1.1254 0.14027 0.13688 -0.0683 0.0126 1.0000
+ 17.250 1.1061 0.15000 0.14677 -0.0746 0.0128 1.0000
+ 17.500 1.0867 0.16034 0.15722 -0.0814 0.0132 1.0000
+ 17.750 1.0678 0.17112 0.16808 -0.0882 0.0136 1.0000
diff --git a/Airfoils/Polars/E854_polar_Re_500000.txt b/Airfoils/Polars/E854_polar_Re_500000.txt
new file mode 100644
index 0000000..57bac19
--- /dev/null
+++ b/Airfoils/Polars/E854_polar_Re_500000.txt
@@ -0,0 +1,124 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: EPPLER E854 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -12.000 -0.2937 0.11874 0.11662 -0.0412 1.0000 0.0200
+ -11.750 -0.2965 0.11461 0.11252 -0.0426 1.0000 0.0206
+ -8.250 -0.4898 0.02772 0.02335 -0.0943 0.9669 0.0108
+ -8.000 -0.4628 0.02503 0.02043 -0.0958 0.9622 0.0103
+ -7.750 -0.4336 0.02170 0.01674 -0.0979 0.9594 0.0100
+ -7.500 -0.4108 0.01959 0.01438 -0.0977 0.9515 0.0099
+ -7.250 -0.3793 0.01780 0.01240 -0.0991 0.9478 0.0099
+ -7.000 -0.3515 0.01637 0.01081 -0.0995 0.9417 0.0100
+ -6.750 -0.3215 0.01517 0.00949 -0.1003 0.9358 0.0102
+ -6.500 -0.2854 0.01412 0.00833 -0.1024 0.9324 0.0106
+ -6.250 -0.2562 0.01270 0.00681 -0.1035 0.9230 0.0121
+ -6.000 -0.2139 0.01198 0.00600 -0.1068 0.9188 0.0147
+ -5.750 -0.1758 0.01095 0.00494 -0.1095 0.9093 0.0256
+ -5.500 -0.1345 0.01034 0.00437 -0.1127 0.8998 0.0446
+ -5.250 -0.0953 0.00983 0.00391 -0.1155 0.8875 0.0704
+ -5.000 -0.0613 0.00933 0.00351 -0.1173 0.8722 0.1111
+ -4.750 -0.0312 0.00879 0.00311 -0.1182 0.8555 0.1748
+ -4.500 -0.0043 0.00808 0.00274 -0.1188 0.8390 0.2869
+ -4.250 0.0224 0.00771 0.00259 -0.1189 0.8235 0.3799
+ -4.000 0.0498 0.00766 0.00252 -0.1188 0.8087 0.4205
+ -3.750 0.0771 0.00768 0.00246 -0.1187 0.7945 0.4469
+ -3.500 0.1043 0.00772 0.00243 -0.1185 0.7813 0.4655
+ -3.250 0.1315 0.00777 0.00238 -0.1183 0.7689 0.4790
+ -3.000 0.1587 0.00784 0.00235 -0.1181 0.7572 0.4922
+ -2.750 0.1858 0.00789 0.00232 -0.1179 0.7460 0.5025
+ -2.500 0.2129 0.00792 0.00229 -0.1177 0.7350 0.5116
+ -2.250 0.2402 0.00799 0.00226 -0.1176 0.7250 0.5214
+ -2.000 0.2674 0.00805 0.00227 -0.1174 0.7157 0.5323
+ -1.750 0.2944 0.00810 0.00229 -0.1172 0.7062 0.5429
+ -1.500 0.3218 0.00815 0.00228 -0.1172 0.6974 0.5515
+ -1.250 0.3491 0.00819 0.00227 -0.1171 0.6895 0.5577
+ -1.000 0.3767 0.00823 0.00226 -0.1170 0.6813 0.5641
+ -0.750 0.4041 0.00827 0.00226 -0.1169 0.6741 0.5695
+ -0.500 0.4315 0.00829 0.00227 -0.1169 0.6662 0.5756
+ -0.250 0.4592 0.00836 0.00227 -0.1169 0.6598 0.5816
+ 0.000 0.4865 0.00836 0.00230 -0.1168 0.6527 0.5875
+ 0.250 0.5143 0.00845 0.00232 -0.1168 0.6464 0.5941
+ 0.500 0.5415 0.00845 0.00237 -0.1168 0.6399 0.6001
+ 0.750 0.5692 0.00852 0.00241 -0.1168 0.6339 0.6072
+ 1.000 0.5967 0.00857 0.00247 -0.1168 0.6282 0.6138
+ 1.250 0.6241 0.00861 0.00253 -0.1167 0.6221 0.6216
+ 1.500 0.6517 0.00869 0.00260 -0.1167 0.6168 0.6289
+ 1.750 0.6791 0.00873 0.00268 -0.1167 0.6111 0.6378
+ 2.000 0.7063 0.00878 0.00277 -0.1166 0.6057 0.6462
+ 2.250 0.7340 0.00887 0.00287 -0.1167 0.6007 0.6556
+ 2.500 0.7611 0.00890 0.00298 -0.1166 0.5951 0.6663
+ 2.750 0.7882 0.00897 0.00308 -0.1165 0.5899 0.6774
+ 3.000 0.8153 0.00904 0.00321 -0.1164 0.5845 0.6895
+ 3.250 0.8417 0.00907 0.00332 -0.1161 0.5780 0.7028
+ 3.500 0.8681 0.00915 0.00344 -0.1159 0.5715 0.7181
+ 3.750 0.8937 0.00916 0.00355 -0.1155 0.5637 0.7356
+ 4.000 0.9191 0.00922 0.00368 -0.1150 0.5556 0.7552
+ 4.250 0.9434 0.00924 0.00379 -0.1143 0.5464 0.7778
+ 4.500 0.9673 0.00924 0.00392 -0.1134 0.5370 0.8059
+ 5.000 1.0093 0.00918 0.00419 -0.1103 0.5209 0.8984
+ 5.250 1.0369 0.00917 0.00427 -0.1102 0.5118 1.0000
+ 5.500 1.0623 0.00932 0.00441 -0.1098 0.5021 1.0000
+ 5.750 1.0874 0.00944 0.00457 -0.1094 0.4906 1.0000
+ 6.000 1.1117 0.00959 0.00475 -0.1088 0.4773 1.0000
+ 6.250 1.1348 0.00979 0.00492 -0.1080 0.4606 1.0000
+ 6.500 1.1572 0.01001 0.00513 -0.1070 0.4403 1.0000
+ 6.750 1.1778 0.01033 0.00537 -0.1057 0.4131 1.0000
+ 7.000 1.1955 0.01079 0.00569 -0.1039 0.3745 1.0000
+ 7.250 1.2067 0.01162 0.00622 -0.1011 0.3123 1.0000
+ 7.500 1.2133 0.01275 0.00697 -0.0976 0.2452 1.0000
+ 7.750 1.2200 0.01377 0.00769 -0.0941 0.1924 1.0000
+ 8.000 1.2276 0.01469 0.00839 -0.0908 0.1526 1.0000
+ 8.250 1.2361 0.01557 0.00910 -0.0877 0.1222 1.0000
+ 8.500 1.2455 0.01640 0.00981 -0.0848 0.0980 1.0000
+ 8.750 1.2554 0.01721 0.01053 -0.0820 0.0791 1.0000
+ 9.000 1.2653 0.01803 0.01128 -0.0794 0.0641 1.0000
+ 9.250 1.2748 0.01888 0.01210 -0.0768 0.0519 1.0000
+ 9.500 1.2843 0.01975 0.01295 -0.0742 0.0424 1.0000
+ 9.750 1.2923 0.02073 0.01391 -0.0716 0.0348 1.0000
+ 10.000 1.3005 0.02173 0.01490 -0.0691 0.0287 1.0000
+ 10.250 1.3092 0.02274 0.01596 -0.0669 0.0238 1.0000
+ 10.500 1.3121 0.02418 0.01739 -0.0641 0.0192 1.0000
+ 10.750 1.3221 0.02522 0.01848 -0.0624 0.0160 1.0000
+ 11.000 1.3244 0.02687 0.02017 -0.0599 0.0128 1.0000
+ 11.250 1.3314 0.02824 0.02156 -0.0582 0.0103 1.0000
+ 11.500 1.3313 0.03027 0.02369 -0.0560 0.0087 1.0000
+ 11.750 1.3367 0.03194 0.02544 -0.0545 0.0077 1.0000
+ 12.000 1.3395 0.03391 0.02750 -0.0530 0.0069 1.0000
+ 12.250 1.3326 0.03688 0.03058 -0.0511 0.0062 1.0000
+ 12.500 1.3361 0.03901 0.03282 -0.0500 0.0059 1.0000
+ 12.750 1.3386 0.04129 0.03521 -0.0491 0.0057 1.0000
+ 13.000 1.3409 0.04367 0.03770 -0.0483 0.0053 1.0000
+ 13.250 1.3432 0.04613 0.04025 -0.0476 0.0050 1.0000
+ 13.500 1.3441 0.04880 0.04301 -0.0471 0.0047 1.0000
+ 13.750 1.3429 0.05181 0.04611 -0.0468 0.0044 1.0000
+ 14.000 1.3390 0.05524 0.04965 -0.0465 0.0043 1.0000
+ 14.250 1.3322 0.05913 0.05367 -0.0464 0.0042 1.0000
+ 14.500 1.3236 0.06342 0.05810 -0.0464 0.0040 1.0000
+ 14.750 1.3233 0.06676 0.06157 -0.0467 0.0039 1.0000
+ 15.000 1.3222 0.07030 0.06524 -0.0471 0.0039 1.0000
+ 15.250 1.3207 0.07401 0.06909 -0.0476 0.0038 1.0000
+ 15.500 1.3176 0.07806 0.07329 -0.0484 0.0036 1.0000
+ 15.750 1.3139 0.08228 0.07765 -0.0493 0.0036 1.0000
+ 16.000 1.3089 0.08680 0.08235 -0.0505 0.0036 1.0000
+ 16.250 1.3033 0.09160 0.08731 -0.0519 0.0035 1.0000
+ 16.500 1.2968 0.09664 0.09251 -0.0536 0.0035 1.0000
+ 16.750 1.2890 0.10204 0.09808 -0.0557 0.0035 1.0000
+ 17.000 1.2804 0.10775 0.10395 -0.0580 0.0035 1.0000
+ 17.250 1.2710 0.11377 0.11015 -0.0607 0.0034 1.0000
+ 17.500 1.2605 0.12016 0.11671 -0.0638 0.0035 1.0000
+ 17.750 1.2491 0.12691 0.12363 -0.0673 0.0034 1.0000
+ 18.000 1.2375 0.13392 0.13081 -0.0712 0.0035 1.0000
+ 18.250 1.2243 0.14149 0.13856 -0.0756 0.0035 1.0000
+ 18.500 1.2109 0.14940 0.14664 -0.0804 0.0035 1.0000
+ 18.750 1.1967 0.15781 0.15520 -0.0857 0.0036 1.0000
+ 19.000 1.1822 0.16657 0.16413 -0.0913 0.0036 1.0000
+ 19.250 1.1669 0.17597 0.17367 -0.0975 0.0037 1.0000
diff --git a/Airfoils/Polars/NACA_4412_polar_Re_100000.txt b/Airfoils/Polars/NACA_4412_polar_Re_100000.txt
new file mode 100644
index 0000000..faffdd9
--- /dev/null
+++ b/Airfoils/Polars/NACA_4412_polar_Re_100000.txt
@@ -0,0 +1,119 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 4412
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.500 -0.3426 0.10705 0.10189 -0.0379 1.0000 0.1297
+ -9.250 -0.3784 0.10671 0.10171 -0.0397 1.0000 0.1327
+ -9.000 -0.4173 0.10641 0.10160 -0.0391 1.0000 0.1332
+ -8.750 -0.3682 0.09949 0.09457 -0.0361 1.0000 0.1363
+ -8.500 -0.3611 0.09726 0.09235 -0.0338 1.0000 0.1402
+ -8.250 -0.3724 0.09561 0.09079 -0.0319 1.0000 0.1442
+ -8.000 -0.4032 0.09481 0.09013 -0.0296 1.0000 0.1469
+ -7.750 -0.4436 0.09403 0.08951 -0.0287 1.0000 0.1481
+ -7.500 -0.4819 0.09082 0.08639 -0.0339 1.0000 0.1493
+ -7.250 -0.4471 0.08830 0.08388 -0.0235 1.0000 0.1526
+ -7.000 -0.4480 0.08640 0.08202 -0.0209 1.0000 0.1558
+ -6.750 -0.4588 0.08412 0.07980 -0.0206 1.0000 0.1600
+ -6.500 -0.4847 0.07929 0.07495 -0.0287 1.0000 0.1662
+ -6.250 -0.4783 0.07718 0.07290 -0.0243 1.0000 0.1680
+ -6.000 -0.4744 0.07516 0.07091 -0.0219 1.0000 0.1710
+ -5.750 -0.4612 0.06965 0.06522 -0.0315 0.9966 0.1830
+ -5.500 -0.4309 0.06730 0.06289 -0.0316 0.9913 0.1898
+ -5.250 -0.3801 0.03672 0.02954 -0.0608 0.9884 0.1054
+ -5.000 -0.3390 0.03524 0.02804 -0.0644 0.9827 0.1094
+ -4.750 -0.2986 0.03253 0.02478 -0.0679 0.9764 0.1116
+ -4.500 -0.2547 0.03033 0.02199 -0.0716 0.9711 0.1152
+ -4.250 -0.2155 0.02874 0.02008 -0.0743 0.9639 0.1211
+ -4.000 -0.1732 0.02773 0.01892 -0.0774 0.9576 0.1286
+ -3.750 -0.1342 0.02656 0.01755 -0.0797 0.9500 0.1373
+ -3.500 -0.0916 0.02579 0.01667 -0.0827 0.9435 0.1516
+ -3.250 -0.0546 0.02511 0.01596 -0.0846 0.9347 0.1675
+ -3.000 -0.0135 0.02457 0.01552 -0.0873 0.9278 0.1881
+ -2.750 0.0233 0.02414 0.01507 -0.0891 0.9186 0.2119
+ -2.500 0.0621 0.02363 0.01470 -0.0912 0.9110 0.2389
+ -2.250 0.0998 0.02306 0.01425 -0.0931 0.9024 0.2721
+ -2.000 0.1378 0.02248 0.01382 -0.0950 0.8945 0.3153
+ -1.750 0.1759 0.02179 0.01342 -0.0968 0.8865 0.3683
+ -1.500 0.2137 0.02102 0.01305 -0.0985 0.8792 0.4420
+ -1.250 0.2473 0.02001 0.01285 -0.0990 0.8713 0.6046
+ -1.000 0.3011 0.01877 0.01237 -0.1014 0.8681 1.0000
+ -0.750 0.3304 0.01877 0.01214 -0.1016 0.8562 1.0000
+ -0.500 0.3765 0.01838 0.01153 -0.1044 0.8513 1.0000
+ -0.250 0.4028 0.01841 0.01142 -0.1039 0.8387 1.0000
+ 0.000 0.4335 0.01835 0.01123 -0.1041 0.8283 1.0000
+ 0.250 0.4720 0.01803 0.01077 -0.1054 0.8210 1.0000
+ 0.500 0.4988 0.01808 0.01072 -0.1049 0.8090 1.0000
+ 0.750 0.5348 0.01782 0.01035 -0.1058 0.8013 1.0000
+ 1.000 0.5631 0.01782 0.01026 -0.1055 0.7900 1.0000
+ 1.250 0.5902 0.01789 0.01027 -0.1051 0.7784 1.0000
+ 1.500 0.6265 0.01763 0.00990 -0.1059 0.7710 1.0000
+ 1.750 0.6506 0.01784 0.01006 -0.1050 0.7582 1.0000
+ 2.000 0.6774 0.01798 0.01015 -0.1045 0.7469 1.0000
+ 2.250 0.7110 0.01786 0.00993 -0.1050 0.7386 1.0000
+ 2.500 0.7347 0.01814 0.01020 -0.1041 0.7259 1.0000
+ 2.750 0.7611 0.01836 0.01039 -0.1036 0.7150 1.0000
+ 3.000 0.7929 0.01834 0.01028 -0.1038 0.7060 1.0000
+ 3.250 0.8161 0.01870 0.01065 -0.1029 0.6937 1.0000
+ 3.500 0.8423 0.01895 0.01090 -0.1024 0.6829 1.0000
+ 3.750 0.8728 0.01902 0.01091 -0.1024 0.6736 1.0000
+ 4.000 0.8957 0.01941 0.01134 -0.1015 0.6614 1.0000
+ 4.250 0.9217 0.01970 0.01163 -0.1010 0.6508 1.0000
+ 4.500 0.9513 0.01982 0.01170 -0.1009 0.6411 1.0000
+ 4.750 0.9739 0.02025 0.01220 -0.0999 0.6291 1.0000
+ 5.000 0.9995 0.02057 0.01254 -0.0994 0.6185 1.0000
+ 5.250 1.0286 0.02074 0.01268 -0.0992 0.6086 1.0000
+ 5.500 1.0508 0.02119 0.01322 -0.0982 0.5964 1.0000
+ 5.750 1.0755 0.02148 0.01355 -0.0974 0.5844 1.0000
+ 6.000 1.1016 0.02159 0.01364 -0.0966 0.5709 1.0000
+ 6.250 1.1279 0.02168 0.01370 -0.0958 0.5569 1.0000
+ 6.500 1.1534 0.02184 0.01386 -0.0950 0.5432 1.0000
+ 6.750 1.1757 0.02207 0.01413 -0.0937 0.5281 1.0000
+ 7.000 1.1987 0.02216 0.01423 -0.0924 0.5114 1.0000
+ 7.250 1.2212 0.02223 0.01428 -0.0911 0.4939 1.0000
+ 7.500 1.2409 0.02240 0.01449 -0.0893 0.4761 1.0000
+ 7.750 1.2594 0.02262 0.01478 -0.0875 0.4582 1.0000
+ 8.000 1.2769 0.02285 0.01510 -0.0854 0.4393 1.0000
+ 8.250 1.2933 0.02308 0.01539 -0.0832 0.4192 1.0000
+ 8.500 1.3086 0.02333 0.01564 -0.0809 0.3979 1.0000
+ 8.750 1.3197 0.02372 0.01612 -0.0779 0.3730 1.0000
+ 9.000 1.3281 0.02421 0.01659 -0.0746 0.3445 1.0000
+ 9.250 1.3316 0.02495 0.01728 -0.0707 0.3104 1.0000
+ 9.500 1.3276 0.02605 0.01816 -0.0657 0.2733 1.0000
+ 9.750 1.3204 0.02765 0.01952 -0.0608 0.2346 1.0000
+ 10.000 1.3127 0.02965 0.02123 -0.0564 0.2023 1.0000
+ 10.250 1.3077 0.03176 0.02308 -0.0527 0.1788 1.0000
+ 10.500 1.3074 0.03378 0.02497 -0.0497 0.1606 1.0000
+ 10.750 1.3110 0.03571 0.02678 -0.0472 0.1474 1.0000
+ 11.000 1.3189 0.03755 0.02847 -0.0451 0.1369 1.0000
+ 11.250 1.3289 0.03922 0.03015 -0.0434 0.1279 1.0000
+ 11.500 1.3439 0.04092 0.03184 -0.0420 0.1203 1.0000
+ 11.750 1.3595 0.04250 0.03339 -0.0408 0.1136 1.0000
+ 12.000 1.3787 0.04430 0.03522 -0.0399 0.1076 1.0000
+ 12.250 1.3912 0.04604 0.03710 -0.0386 0.1025 1.0000
+ 12.500 1.4202 0.04801 0.03893 -0.0387 0.0966 1.0000
+ 12.750 1.4239 0.05002 0.04125 -0.0367 0.0932 1.0000
+ 13.000 1.4319 0.05205 0.04343 -0.0351 0.0895 1.0000
+ 13.250 1.4604 0.05447 0.04574 -0.0355 0.0845 1.0000
+ 13.500 1.4540 0.05696 0.04860 -0.0329 0.0825 1.0000
+ 13.750 1.4507 0.05963 0.05156 -0.0309 0.0802 1.0000
+ 14.000 1.4512 0.06218 0.05428 -0.0294 0.0775 1.0000
+ 14.250 1.4783 0.06519 0.05713 -0.0297 0.0734 1.0000
+ 14.500 1.4597 0.06850 0.06082 -0.0274 0.0727 1.0000
+ 14.750 1.4404 0.07234 0.06501 -0.0257 0.0720 1.0000
+ 15.000 1.4201 0.07667 0.06967 -0.0247 0.0712 1.0000
+ 15.250 1.3983 0.08150 0.07481 -0.0243 0.0706 1.0000
+ 15.500 1.3742 0.08690 0.08051 -0.0247 0.0701 1.0000
+ 15.750 1.3474 0.09302 0.08692 -0.0259 0.0699 1.0000
+ 16.000 1.3171 0.10006 0.09424 -0.0282 0.0701 1.0000
+ 16.250 1.2836 0.10833 0.10277 -0.0317 0.0706 1.0000
+ 16.500 1.2473 0.11795 0.11263 -0.0367 0.0714 1.0000
+ 16.750 1.2101 0.12886 0.12372 -0.0430 0.0723 1.0000
+ 17.000 1.1753 0.14068 0.13565 -0.0500 0.0732 1.0000
diff --git a/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt b/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt
new file mode 100644
index 0000000..97538d4
--- /dev/null
+++ b/Airfoils/Polars/NACA_4412_polar_Re_1000000.txt
@@ -0,0 +1,148 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 4412
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -15.750 -0.8374 0.08373 0.08141 -0.0585 1.0000 0.0169
+ -15.500 -0.9127 0.06837 0.06591 -0.0687 1.0000 0.0166
+ -15.250 -1.0965 0.03328 0.03022 -0.0993 1.0000 0.0153
+ -15.000 -1.1161 0.03120 0.02803 -0.0956 1.0000 0.0154
+ -14.750 -1.1210 0.02977 0.02651 -0.0926 1.0000 0.0156
+ -14.500 -1.1215 0.02857 0.02523 -0.0896 1.0000 0.0159
+ -14.250 -1.1181 0.02751 0.02407 -0.0870 1.0000 0.0162
+ -14.000 -1.0990 0.02637 0.02282 -0.0871 0.9992 0.0166
+ -13.750 -1.0711 0.02533 0.02165 -0.0885 0.9979 0.0170
+ -13.500 -1.0462 0.02365 0.01985 -0.0903 0.9963 0.0177
+ -13.250 -1.0163 0.02288 0.01905 -0.0918 0.9951 0.0183
+ -13.000 -0.9847 0.02237 0.01850 -0.0933 0.9943 0.0189
+ -12.750 -0.9549 0.02183 0.01790 -0.0943 0.9930 0.0195
+ -12.500 -0.9260 0.02126 0.01724 -0.0952 0.9911 0.0201
+ -12.250 -0.8954 0.02078 0.01666 -0.0963 0.9894 0.0206
+ -12.000 -0.8682 0.01946 0.01525 -0.0976 0.9877 0.0214
+ -11.750 -0.8365 0.01894 0.01471 -0.0990 0.9866 0.0220
+ -11.500 -0.8038 0.01852 0.01424 -0.1004 0.9857 0.0227
+ -11.250 -0.7707 0.01808 0.01375 -0.1019 0.9849 0.0235
+ -11.000 -0.7369 0.01769 0.01328 -0.1035 0.9843 0.0242
+ -10.750 -0.7070 0.01745 0.01297 -0.1041 0.9819 0.0246
+ -10.500 -0.6803 0.01619 0.01161 -0.1049 0.9793 0.0257
+ -10.250 -0.6491 0.01569 0.01109 -0.1060 0.9775 0.0264
+ -10.000 -0.6172 0.01529 0.01065 -0.1071 0.9759 0.0271
+ -9.750 -0.5850 0.01491 0.01022 -0.1082 0.9742 0.0279
+ -9.500 -0.5547 0.01456 0.00981 -0.1089 0.9718 0.0287
+ -9.250 -0.5287 0.01426 0.00944 -0.1085 0.9665 0.0292
+ -9.000 -0.5023 0.01345 0.00855 -0.1085 0.9622 0.0299
+ -8.750 -0.4769 0.01285 0.00791 -0.1082 0.9574 0.0309
+ -8.500 -0.4513 0.01249 0.00752 -0.1078 0.9519 0.0317
+ -8.250 -0.4243 0.01214 0.00713 -0.1076 0.9474 0.0324
+ -8.000 -0.3979 0.01184 0.00678 -0.1073 0.9422 0.0333
+ -7.750 -0.3715 0.01155 0.00644 -0.1070 0.9363 0.0340
+ -7.500 -0.3442 0.01127 0.00609 -0.1068 0.9313 0.0345
+ -7.250 -0.3183 0.01080 0.00556 -0.1064 0.9249 0.0354
+ -7.000 -0.2921 0.01033 0.00505 -0.1061 0.9186 0.0365
+ -6.750 -0.2649 0.01003 0.00471 -0.1059 0.9125 0.0375
+ -6.500 -0.2377 0.00977 0.00441 -0.1057 0.9053 0.0384
+ -6.000 -0.1825 0.00935 0.00389 -0.1054 0.8910 0.0404
+ -5.750 -0.1549 0.00912 0.00360 -0.1052 0.8835 0.0414
+ -5.500 -0.1275 0.00880 0.00325 -0.1051 0.8751 0.0435
+ -5.000 -0.0718 0.00845 0.00283 -0.1049 0.8578 0.0476
+ -4.750 -0.0441 0.00824 0.00259 -0.1047 0.8488 0.0519
+ -4.500 -0.0162 0.00810 0.00243 -0.1046 0.8388 0.0569
+ -4.250 0.0117 0.00793 0.00228 -0.1045 0.8288 0.0655
+ -4.000 0.0394 0.00780 0.00213 -0.1044 0.8184 0.0745
+ -3.750 0.0674 0.00769 0.00201 -0.1044 0.8073 0.0820
+ -3.500 0.0954 0.00761 0.00191 -0.1043 0.7964 0.0890
+ -3.250 0.1232 0.00752 0.00180 -0.1042 0.7851 0.0977
+ -3.000 0.1512 0.00745 0.00171 -0.1041 0.7733 0.1066
+ -2.750 0.1791 0.00737 0.00163 -0.1040 0.7616 0.1182
+ -2.500 0.2069 0.00729 0.00156 -0.1040 0.7497 0.1332
+ -2.250 0.2346 0.00723 0.00150 -0.1039 0.7378 0.1502
+ -2.000 0.2625 0.00715 0.00145 -0.1038 0.7254 0.1697
+ -1.750 0.2903 0.00709 0.00142 -0.1038 0.7132 0.1927
+ -1.500 0.3180 0.00703 0.00141 -0.1037 0.7012 0.2214
+ -1.250 0.3456 0.00701 0.00139 -0.1036 0.6886 0.2466
+ -1.000 0.3734 0.00697 0.00138 -0.1035 0.6754 0.2686
+ -0.750 0.4012 0.00694 0.00137 -0.1035 0.6626 0.2903
+ -0.500 0.4288 0.00691 0.00138 -0.1034 0.6497 0.3203
+ -0.250 0.4562 0.00686 0.00139 -0.1033 0.6365 0.3629
+ 0.000 0.4833 0.00678 0.00141 -0.1032 0.6232 0.4192
+ 0.250 0.5102 0.00658 0.00146 -0.1031 0.6101 0.5177
+ 0.500 0.5366 0.00635 0.00153 -0.1029 0.5975 0.6393
+ 0.750 0.5622 0.00617 0.00160 -0.1024 0.5856 0.7449
+ 1.000 0.5842 0.00594 0.00170 -0.1009 0.5740 0.8717
+ 1.250 0.6163 0.00588 0.00177 -0.1014 0.5622 0.9842
+ 1.500 0.6525 0.00598 0.00181 -0.1033 0.5505 1.0000
+ 1.750 0.6788 0.00611 0.00186 -0.1029 0.5398 1.0000
+ 2.000 0.7055 0.00622 0.00192 -0.1026 0.5294 1.0000
+ 2.250 0.7325 0.00633 0.00199 -0.1024 0.5204 1.0000
+ 2.500 0.7592 0.00646 0.00206 -0.1022 0.5112 1.0000
+ 2.750 0.7865 0.00656 0.00213 -0.1020 0.5029 1.0000
+ 3.250 0.8405 0.00681 0.00231 -0.1016 0.4847 1.0000
+ 3.500 0.8672 0.00696 0.00240 -0.1014 0.4746 1.0000
+ 3.750 0.8941 0.00709 0.00250 -0.1012 0.4646 1.0000
+ 4.000 0.9210 0.00722 0.00260 -0.1010 0.4540 1.0000
+ 4.250 0.9473 0.00739 0.00272 -0.1007 0.4426 1.0000
+ 4.500 0.9734 0.00758 0.00284 -0.1004 0.4273 1.0000
+ 4.750 0.9993 0.00778 0.00297 -0.1001 0.4110 1.0000
+ 5.000 1.0254 0.00797 0.00311 -0.0998 0.3979 1.0000
+ 5.250 1.0518 0.00813 0.00326 -0.0995 0.3861 1.0000
+ 5.500 1.0777 0.00834 0.00342 -0.0992 0.3731 1.0000
+ 5.750 1.1031 0.00857 0.00359 -0.0988 0.3575 1.0000
+ 6.000 1.1280 0.00884 0.00379 -0.0983 0.3398 1.0000
+ 6.250 1.1523 0.00914 0.00401 -0.0978 0.3207 1.0000
+ 6.500 1.1761 0.00948 0.00426 -0.0971 0.2993 1.0000
+ 6.750 1.1988 0.00989 0.00455 -0.0963 0.2737 1.0000
+ 7.000 1.2208 0.01036 0.00488 -0.0954 0.2461 1.0000
+ 7.250 1.2417 0.01089 0.00526 -0.0943 0.2173 1.0000
+ 7.500 1.2614 0.01149 0.00569 -0.0931 0.1865 1.0000
+ 7.750 1.2793 0.01220 0.00621 -0.0915 0.1526 1.0000
+ 8.000 1.2973 0.01288 0.00672 -0.0900 0.1252 1.0000
+ 8.250 1.3164 0.01345 0.00719 -0.0887 0.1065 1.0000
+ 8.500 1.3346 0.01404 0.00769 -0.0872 0.0893 1.0000
+ 8.750 1.3514 0.01469 0.00823 -0.0854 0.0729 1.0000
+ 9.000 1.3676 0.01527 0.00875 -0.0836 0.0622 1.0000
+ 9.250 1.3835 0.01581 0.00926 -0.0817 0.0563 1.0000
+ 9.500 1.4004 0.01631 0.00976 -0.0799 0.0521 1.0000
+ 9.750 1.4171 0.01682 0.01028 -0.0782 0.0491 1.0000
+ 10.000 1.4317 0.01746 0.01091 -0.0762 0.0459 1.0000
+ 10.250 1.4484 0.01797 0.01147 -0.0746 0.0442 1.0000
+ 10.500 1.4653 0.01849 0.01203 -0.0731 0.0427 1.0000
+ 10.750 1.4805 0.01911 0.01267 -0.0714 0.0411 1.0000
+ 11.000 1.4938 0.01986 0.01343 -0.0695 0.0392 1.0000
+ 11.250 1.5061 0.02069 0.01430 -0.0676 0.0376 1.0000
+ 11.500 1.5221 0.02129 0.01495 -0.0662 0.0368 1.0000
+ 11.750 1.5369 0.02199 0.01570 -0.0647 0.0356 1.0000
+ 12.000 1.5500 0.02282 0.01656 -0.0631 0.0343 1.0000
+ 12.250 1.5608 0.02382 0.01758 -0.0614 0.0330 1.0000
+ 12.500 1.5688 0.02506 0.01888 -0.0594 0.0316 1.0000
+ 12.750 1.5831 0.02588 0.01975 -0.0582 0.0308 1.0000
+ 13.000 1.5959 0.02683 0.02075 -0.0569 0.0298 1.0000
+ 13.250 1.6066 0.02796 0.02192 -0.0555 0.0286 1.0000
+ 13.500 1.6141 0.02939 0.02338 -0.0540 0.0273 1.0000
+ 13.750 1.6213 0.03089 0.02494 -0.0526 0.0262 1.0000
+ 14.000 1.6325 0.03209 0.02620 -0.0516 0.0252 1.0000
+ 14.250 1.6414 0.03354 0.02770 -0.0505 0.0241 1.0000
+ 14.500 1.6474 0.03528 0.02947 -0.0493 0.0229 1.0000
+ 14.750 1.6508 0.03731 0.03156 -0.0482 0.0218 1.0000
+ 15.000 1.6585 0.03899 0.03332 -0.0474 0.0209 1.0000
+ 15.250 1.6638 0.04096 0.03533 -0.0466 0.0198 1.0000
+ 15.500 1.6661 0.04330 0.03772 -0.0458 0.0187 1.0000
+ 15.750 1.6666 0.04589 0.04037 -0.0451 0.0179 1.0000
+ 16.000 1.6698 0.04827 0.04284 -0.0447 0.0171 1.0000
+ 16.250 1.6706 0.05099 0.04562 -0.0443 0.0164 1.0000
+ 16.500 1.6692 0.05402 0.04871 -0.0440 0.0157 1.0000
+ 16.750 1.6638 0.05759 0.05235 -0.0439 0.0151 1.0000
+ 17.000 1.6605 0.06101 0.05587 -0.0439 0.0146 1.0000
+ 17.250 1.6584 0.06435 0.05931 -0.0441 0.0142 1.0000
+ 17.500 1.6548 0.06793 0.06298 -0.0444 0.0138 1.0000
+ 17.750 1.6497 0.07175 0.06689 -0.0448 0.0134 1.0000
+ 18.000 1.6430 0.07583 0.07106 -0.0453 0.0131 1.0000
+ 18.250 1.6346 0.08024 0.07555 -0.0461 0.0128 1.0000
+ 18.500 1.6237 0.08507 0.08047 -0.0470 0.0124 1.0000
+ 18.750 1.6097 0.09040 0.08590 -0.0482 0.0121 1.0000
diff --git a/Airfoils/Polars/NACA_4412_polar_Re_200000.txt b/Airfoils/Polars/NACA_4412_polar_Re_200000.txt
new file mode 100644
index 0000000..64d2048
--- /dev/null
+++ b/Airfoils/Polars/NACA_4412_polar_Re_200000.txt
@@ -0,0 +1,115 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 4412
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -8.500 -0.4088 0.08983 0.08647 -0.0344 1.0000 0.0813
+ -8.250 -0.4231 0.08831 0.08501 -0.0315 1.0000 0.0823
+ -8.000 -0.4442 0.08695 0.08373 -0.0282 1.0000 0.0831
+ -7.750 -0.4937 0.05313 0.04929 -0.0671 0.9865 0.0673
+ -7.500 -0.4712 0.04655 0.04243 -0.0720 0.9817 0.0641
+ -7.250 -0.4535 0.03731 0.03244 -0.0777 0.9755 0.0622
+ -7.000 -0.4231 0.03226 0.02667 -0.0816 0.9717 0.0637
+ -6.750 -0.3952 0.02910 0.02295 -0.0831 0.9655 0.0647
+ -6.500 -0.3582 0.02703 0.02031 -0.0857 0.9620 0.0660
+ -6.250 -0.3204 0.02462 0.01772 -0.0887 0.9600 0.0684
+ -6.000 -0.2930 0.02357 0.01656 -0.0890 0.9523 0.0702
+ -5.750 -0.2544 0.02238 0.01517 -0.0913 0.9489 0.0725
+ -5.500 -0.2129 0.02134 0.01390 -0.0940 0.9465 0.0756
+ -5.250 -0.1843 0.02032 0.01270 -0.0942 0.9393 0.0785
+ -5.000 -0.1466 0.01934 0.01174 -0.0963 0.9352 0.0824
+ -4.750 -0.1047 0.01855 0.01086 -0.0990 0.9326 0.0877
+ -4.500 -0.0618 0.01758 0.00989 -0.1020 0.9308 0.0951
+ -4.250 -0.0351 0.01719 0.00943 -0.1016 0.9217 0.1030
+ -4.000 0.0030 0.01638 0.00871 -0.1036 0.9180 0.1150
+ -3.750 0.0426 0.01571 0.00806 -0.1058 0.9151 0.1299
+ -3.500 0.0687 0.01531 0.00771 -0.1053 0.9057 0.1437
+ -3.250 0.1047 0.01477 0.00720 -0.1066 0.9011 0.1620
+ -3.000 0.1342 0.01437 0.00686 -0.1068 0.8934 0.1821
+ -2.750 0.1659 0.01392 0.00653 -0.1073 0.8865 0.2077
+ -2.500 0.1964 0.01352 0.00623 -0.1075 0.8792 0.2377
+ -2.250 0.2255 0.01312 0.00593 -0.1075 0.8706 0.2707
+ -2.000 0.2541 0.01274 0.00566 -0.1074 0.8618 0.3065
+ -1.750 0.2836 0.01233 0.00537 -0.1074 0.8533 0.3489
+ -1.500 0.3098 0.01195 0.00520 -0.1069 0.8424 0.4005
+ -1.250 0.3373 0.01136 0.00499 -0.1066 0.8336 0.5000
+ -1.000 0.3583 0.01070 0.00498 -0.1046 0.8222 0.6832
+ -0.500 0.4431 0.00999 0.00468 -0.1085 0.8021 1.0000
+ -0.250 0.4683 0.01000 0.00456 -0.1077 0.7890 1.0000
+ 0.000 0.4938 0.01005 0.00448 -0.1071 0.7758 1.0000
+ 0.250 0.5198 0.01011 0.00441 -0.1065 0.7629 1.0000
+ 0.500 0.5464 0.01019 0.00435 -0.1060 0.7504 1.0000
+ 0.750 0.5733 0.01027 0.00429 -0.1056 0.7383 1.0000
+ 1.000 0.5988 0.01039 0.00431 -0.1050 0.7246 1.0000
+ 1.250 0.6247 0.01052 0.00435 -0.1045 0.7115 1.0000
+ 1.500 0.6510 0.01066 0.00439 -0.1040 0.6992 1.0000
+ 1.750 0.6778 0.01082 0.00443 -0.1036 0.6876 1.0000
+ 2.000 0.7038 0.01097 0.00451 -0.1032 0.6752 1.0000
+ 2.250 0.7297 0.01115 0.00463 -0.1027 0.6629 1.0000
+ 2.500 0.7561 0.01133 0.00474 -0.1023 0.6517 1.0000
+ 2.750 0.7828 0.01152 0.00483 -0.1020 0.6409 1.0000
+ 3.000 0.8084 0.01171 0.00500 -0.1014 0.6289 1.0000
+ 3.250 0.8345 0.01191 0.00515 -0.1010 0.6180 1.0000
+ 3.500 0.8613 0.01213 0.00528 -0.1007 0.6079 1.0000
+ 3.750 0.8865 0.01233 0.00549 -0.1002 0.5965 1.0000
+ 4.000 0.9127 0.01256 0.00569 -0.0998 0.5863 1.0000
+ 4.250 0.9390 0.01280 0.00587 -0.0995 0.5765 1.0000
+ 4.500 0.9643 0.01303 0.00613 -0.0990 0.5659 1.0000
+ 4.750 0.9905 0.01330 0.00636 -0.0986 0.5563 1.0000
+ 5.000 1.0156 0.01352 0.00657 -0.0980 0.5448 1.0000
+ 5.250 1.0396 0.01373 0.00679 -0.0972 0.5317 1.0000
+ 5.500 1.0639 0.01396 0.00701 -0.0965 0.5191 1.0000
+ 5.750 1.0888 0.01422 0.00724 -0.0959 0.5080 1.0000
+ 6.000 1.1122 0.01444 0.00743 -0.0950 0.4940 1.0000
+ 6.250 1.1342 0.01464 0.00764 -0.0938 0.4781 1.0000
+ 6.500 1.1562 0.01486 0.00788 -0.0927 0.4627 1.0000
+ 6.750 1.1783 0.01510 0.00815 -0.0916 0.4482 1.0000
+ 7.000 1.2002 0.01536 0.00845 -0.0905 0.4333 1.0000
+ 7.250 1.2215 0.01565 0.00876 -0.0893 0.4177 1.0000
+ 7.500 1.2420 0.01596 0.00908 -0.0880 0.4012 1.0000
+ 7.750 1.2616 0.01630 0.00943 -0.0866 0.3835 1.0000
+ 8.000 1.2799 0.01670 0.00982 -0.0849 0.3635 1.0000
+ 8.250 1.2967 0.01714 0.01026 -0.0831 0.3392 1.0000
+ 8.500 1.3110 0.01772 0.01076 -0.0809 0.3110 1.0000
+ 8.750 1.3218 0.01847 0.01138 -0.0781 0.2773 1.0000
+ 9.000 1.3267 0.01944 0.01217 -0.0745 0.2406 1.0000
+ 9.250 1.3283 0.02069 0.01320 -0.0706 0.1997 1.0000
+ 9.500 1.3275 0.02223 0.01449 -0.0667 0.1617 1.0000
+ 9.750 1.3266 0.02391 0.01597 -0.0630 0.1356 1.0000
+ 10.000 1.3275 0.02555 0.01752 -0.0598 0.1187 1.0000
+ 10.250 1.3289 0.02726 0.01916 -0.0569 0.1080 1.0000
+ 10.500 1.3326 0.02887 0.02077 -0.0544 0.1001 1.0000
+ 10.750 1.3366 0.03055 0.02249 -0.0520 0.0942 1.0000
+ 11.000 1.3429 0.03209 0.02408 -0.0501 0.0890 1.0000
+ 11.250 1.3438 0.03415 0.02607 -0.0479 0.0848 1.0000
+ 11.500 1.3530 0.03558 0.02764 -0.0464 0.0812 1.0000
+ 11.750 1.3601 0.03722 0.02934 -0.0450 0.0776 1.0000
+ 12.000 1.3641 0.03917 0.03125 -0.0433 0.0746 1.0000
+ 12.250 1.3713 0.04092 0.03309 -0.0418 0.0719 1.0000
+ 12.500 1.3790 0.04262 0.03491 -0.0407 0.0690 1.0000
+ 12.750 1.3857 0.04441 0.03676 -0.0395 0.0664 1.0000
+ 13.000 1.3923 0.04629 0.03859 -0.0382 0.0639 1.0000
+ 13.250 1.4001 0.04814 0.04054 -0.0369 0.0616 1.0000
+ 13.500 1.4059 0.05012 0.04268 -0.0360 0.0593 1.0000
+ 13.750 1.4115 0.05212 0.04477 -0.0351 0.0571 1.0000
+ 14.000 1.4180 0.05407 0.04670 -0.0342 0.0549 1.0000
+ 14.250 1.4266 0.05607 0.04875 -0.0329 0.0527 1.0000
+ 14.500 1.4286 0.05852 0.05141 -0.0322 0.0509 1.0000
+ 14.750 1.4308 0.06101 0.05404 -0.0316 0.0491 1.0000
+ 15.000 1.4342 0.06336 0.05644 -0.0312 0.0473 1.0000
+ 15.250 1.4467 0.06514 0.05814 -0.0297 0.0450 1.0000
+ 15.500 1.4416 0.06845 0.06172 -0.0297 0.0440 1.0000
+ 15.750 1.4385 0.07175 0.06524 -0.0296 0.0427 1.0000
+ 16.000 1.4362 0.07501 0.06865 -0.0296 0.0414 1.0000
+ 16.250 1.4351 0.07813 0.07188 -0.0298 0.0401 1.0000
+ 16.500 1.4392 0.08059 0.07434 -0.0296 0.0388 1.0000
+ 16.750 1.4401 0.08377 0.07759 -0.0291 0.0376 1.0000
+ 17.000 1.4274 0.08864 0.08274 -0.0304 0.0369 1.0000
+ 17.250 1.4151 0.09369 0.08804 -0.0318 0.0363 1.0000
diff --git a/Airfoils/Polars/NACA_4412_polar_Re_50000.txt b/Airfoils/Polars/NACA_4412_polar_Re_50000.txt
new file mode 100644
index 0000000..3650d36
--- /dev/null
+++ b/Airfoils/Polars/NACA_4412_polar_Re_50000.txt
@@ -0,0 +1,106 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 4412
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -9.250 -0.3484 0.11519 0.10799 -0.0307 1.0000 0.2459
+ -9.000 -0.3263 0.10991 0.10270 -0.0293 1.0000 0.2536
+ -8.750 -0.3430 0.10940 0.10231 -0.0284 1.0000 0.2620
+ -8.500 -0.3340 0.10549 0.09844 -0.0271 1.0000 0.2680
+ -8.250 -0.3380 0.10368 0.09671 -0.0255 1.0000 0.2771
+ -8.000 -0.3862 0.10566 0.09896 -0.0226 1.0000 0.2805
+ -7.750 -0.3424 0.09889 0.09208 -0.0219 1.0000 0.2911
+ -7.500 -0.3860 0.10017 0.09361 -0.0180 1.0000 0.2969
+ -7.250 -0.3576 0.09514 0.08853 -0.0169 1.0000 0.3077
+ -7.000 -0.3988 0.09590 0.08951 -0.0124 1.0000 0.3135
+ -6.750 -0.3812 0.09202 0.08562 -0.0104 1.0000 0.3258
+ -6.500 -0.4296 0.09289 0.08673 -0.0077 1.0000 0.3315
+ -6.250 -0.4111 0.08914 0.08297 -0.0042 1.0000 0.3457
+ -6.000 -0.4133 0.08662 0.08052 -0.0015 1.0000 0.3557
+ -5.750 -0.4339 0.08523 0.07926 0.0005 1.0000 0.3681
+ -5.500 -0.4490 0.08389 0.07801 0.0021 1.0000 0.3833
+ -5.250 -0.4372 0.08080 0.07495 0.0060 1.0000 0.3957
+ -5.000 -0.4394 0.07850 0.07271 0.0087 1.0000 0.4103
+ -4.750 -0.4076 0.05599 0.04887 -0.0396 1.0000 0.1988
+ -4.500 -0.3867 0.05206 0.04465 -0.0418 1.0000 0.1939
+ -4.250 -0.3630 0.04815 0.04039 -0.0444 1.0000 0.1911
+ -4.000 -0.3354 0.04419 0.03592 -0.0475 1.0000 0.1881
+ -3.750 -0.3092 0.04146 0.03274 -0.0493 1.0000 0.1908
+ -3.500 -0.2800 0.03898 0.02952 -0.0515 1.0000 0.1969
+ -3.250 -0.2589 0.03764 0.02820 -0.0515 1.0000 0.2030
+ -3.000 -0.2321 0.03627 0.02629 -0.0526 1.0000 0.2137
+ -2.750 -0.2112 0.03545 0.02553 -0.0525 1.0000 0.2250
+ -2.500 -0.1884 0.03470 0.02464 -0.0528 1.0000 0.2394
+ -2.250 -0.1656 0.03410 0.02391 -0.0531 1.0000 0.2564
+ -2.000 -0.1432 0.03369 0.02338 -0.0533 1.0000 0.2769
+ -1.750 -0.1176 0.03335 0.02302 -0.0542 0.9988 0.3012
+ -1.500 -0.0654 0.03317 0.02283 -0.0594 0.9883 0.3463
+ -1.250 -0.0158 0.03295 0.02274 -0.0639 0.9776 0.4048
+ -1.000 0.0297 0.03258 0.02279 -0.0676 0.9671 0.4910
+ -0.750 0.0671 0.03154 0.02306 -0.0683 0.9582 0.6968
+ -0.500 0.1031 0.03143 0.02286 -0.0702 0.9435 1.0000
+ -0.250 0.1469 0.03237 0.02333 -0.0742 0.9312 1.0000
+ 0.000 0.1804 0.03318 0.02384 -0.0763 0.9178 1.0000
+ 0.250 0.2127 0.03403 0.02444 -0.0782 0.9051 1.0000
+ 0.500 0.2476 0.03490 0.02509 -0.0804 0.8929 1.0000
+ 0.750 0.2900 0.03575 0.02574 -0.0836 0.8821 1.0000
+ 1.000 0.3166 0.03657 0.02643 -0.0844 0.8694 1.0000
+ 1.250 0.3420 0.03746 0.02721 -0.0850 0.8570 1.0000
+ 1.500 0.3718 0.03836 0.02800 -0.0862 0.8458 1.0000
+ 1.750 0.4130 0.03909 0.02862 -0.0888 0.8357 1.0000
+ 2.000 0.4299 0.04014 0.02962 -0.0881 0.8232 1.0000
+ 2.250 0.4529 0.04117 0.03059 -0.0883 0.8118 1.0000
+ 2.500 0.4924 0.04188 0.03125 -0.0905 0.8022 1.0000
+ 2.750 0.5095 0.04305 0.03239 -0.0899 0.7903 1.0000
+ 3.000 0.5279 0.04426 0.03358 -0.0894 0.7788 1.0000
+ 3.250 0.5662 0.04496 0.03427 -0.0912 0.7696 1.0000
+ 3.500 0.5815 0.04627 0.03560 -0.0904 0.7577 1.0000
+ 3.750 0.5962 0.04773 0.03705 -0.0896 0.7465 1.0000
+ 4.000 0.6350 0.04832 0.03768 -0.0911 0.7370 1.0000
+ 4.250 0.6464 0.04992 0.03930 -0.0901 0.7251 1.0000
+ 4.500 0.6592 0.05153 0.04095 -0.0891 0.7135 1.0000
+ 4.750 0.6949 0.05221 0.04167 -0.0901 0.7038 1.0000
+ 5.000 0.7094 0.05375 0.04328 -0.0893 0.6918 1.0000
+ 5.250 0.7169 0.05575 0.04532 -0.0880 0.6799 1.0000
+ 5.500 0.7421 0.05688 0.04651 -0.0879 0.6689 1.0000
+ 5.750 0.7723 0.05768 0.04742 -0.0881 0.6579 1.0000
+ 6.000 0.7737 0.06012 0.04991 -0.0865 0.6450 1.0000
+ 6.250 0.7860 0.06203 0.05189 -0.0857 0.6330 1.0000
+ 6.500 0.8168 0.06274 0.05272 -0.0856 0.6214 1.0000
+ 6.750 0.8415 0.06375 0.05385 -0.0851 0.6091 1.0000
+ 7.000 0.8430 0.06634 0.05651 -0.0836 0.5952 1.0000
+ 7.250 0.8513 0.06850 0.05876 -0.0825 0.5813 1.0000
+ 7.500 0.8671 0.07004 0.06040 -0.0814 0.5669 1.0000
+ 7.750 0.8865 0.07116 0.06166 -0.0802 0.5517 1.0000
+ 8.000 0.9081 0.07193 0.06256 -0.0789 0.5355 1.0000
+ 8.250 0.9309 0.07238 0.06316 -0.0773 0.5186 1.0000
+ 8.500 0.9547 0.07259 0.06354 -0.0755 0.5015 1.0000
+ 8.750 0.9857 0.07176 0.06289 -0.0734 0.4840 1.0000
+ 9.000 1.2725 0.04155 0.03319 -0.0743 0.4435 1.0000
+ 9.250 1.2985 0.04001 0.03163 -0.0714 0.4108 1.0000
+ 9.500 1.3102 0.03953 0.03112 -0.0674 0.3762 1.0000
+ 9.750 1.3184 0.03942 0.03081 -0.0631 0.3384 1.0000
+ 10.000 1.3248 0.03991 0.03091 -0.0590 0.3012 1.0000
+ 10.250 1.3265 0.04130 0.03202 -0.0549 0.2702 1.0000
+ 10.500 1.3283 0.04312 0.03369 -0.0514 0.2452 1.0000
+ 10.750 1.3401 0.04512 0.03548 -0.0492 0.2231 1.0000
+ 11.000 1.3527 0.04742 0.03778 -0.0474 0.2062 1.0000
+ 11.250 1.3686 0.04976 0.04008 -0.0460 0.1920 1.0000
+ 11.500 1.3924 0.05231 0.04258 -0.0457 0.1796 1.0000
+ 11.750 1.3929 0.05510 0.04569 -0.0429 0.1727 1.0000
+ 12.000 1.4092 0.05796 0.04860 -0.0420 0.1644 1.0000
+ 12.250 1.3997 0.06112 0.05211 -0.0387 0.1604 1.0000
+ 12.500 1.4298 0.06425 0.05517 -0.0393 0.1521 1.0000
+ 12.750 1.4065 0.06782 0.05914 -0.0353 0.1508 1.0000
+ 13.000 1.3810 0.07194 0.06362 -0.0321 0.1498 1.0000
+ 13.250 1.3520 0.07677 0.06876 -0.0299 0.1493 1.0000
+ 13.500 1.3181 0.08256 0.07484 -0.0288 0.1496 1.0000
+ 13.750 1.2801 0.08958 0.08211 -0.0292 0.1505 1.0000
+ 14.000 1.2403 0.09792 0.09061 -0.0311 0.1517 1.0000
diff --git a/Airfoils/Polars/NACA_4412_polar_Re_500000.txt b/Airfoils/Polars/NACA_4412_polar_Re_500000.txt
new file mode 100644
index 0000000..567a133
--- /dev/null
+++ b/Airfoils/Polars/NACA_4412_polar_Re_500000.txt
@@ -0,0 +1,131 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 4412
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -12.250 -0.8284 0.04176 0.03861 -0.0832 1.0000 0.0246
+ -12.000 -0.8522 0.03908 0.03574 -0.0799 1.0000 0.0247
+ -11.750 -0.8699 0.03540 0.03185 -0.0776 0.9997 0.0251
+ -11.500 -0.8415 0.03435 0.03082 -0.0796 0.9977 0.0258
+ -11.250 -0.8114 0.03347 0.02989 -0.0818 0.9957 0.0265
+ -11.000 -0.7822 0.03202 0.02830 -0.0843 0.9939 0.0275
+ -10.750 -0.7591 0.02982 0.02579 -0.0862 0.9905 0.0285
+ -10.500 -0.7309 0.02822 0.02384 -0.0882 0.9877 0.0293
+ -10.250 -0.7053 0.02556 0.02100 -0.0903 0.9856 0.0304
+ -10.000 -0.6714 0.02497 0.02039 -0.0924 0.9843 0.0314
+ -9.750 -0.6435 0.02420 0.01952 -0.0932 0.9807 0.0325
+ -9.500 -0.6134 0.02311 0.01823 -0.0945 0.9778 0.0336
+ -9.250 -0.5802 0.02227 0.01716 -0.0962 0.9758 0.0345
+ -9.000 -0.5492 0.02031 0.01504 -0.0982 0.9742 0.0359
+ -8.750 -0.5132 0.01962 0.01432 -0.1004 0.9731 0.0371
+ -8.500 -0.4849 0.01894 0.01355 -0.1008 0.9683 0.0381
+ -8.250 -0.4502 0.01823 0.01273 -0.1025 0.9656 0.0394
+ -8.000 -0.4135 0.01756 0.01192 -0.1045 0.9637 0.0405
+ -7.750 -0.3794 0.01639 0.01059 -0.1062 0.9619 0.0416
+ -7.500 -0.3455 0.01547 0.00965 -0.1078 0.9602 0.0430
+ -7.250 -0.3201 0.01501 0.00917 -0.1073 0.9539 0.0442
+ -7.000 -0.2893 0.01445 0.00854 -0.1080 0.9502 0.0454
+ -6.750 -0.2573 0.01389 0.00791 -0.1088 0.9471 0.0467
+ -6.500 -0.2303 0.01351 0.00746 -0.1085 0.9415 0.0478
+ -6.250 -0.2035 0.01278 0.00668 -0.1083 0.9357 0.0494
+ -6.000 -0.1743 0.01225 0.00613 -0.1085 0.9314 0.0512
+ -5.750 -0.1485 0.01190 0.00577 -0.1080 0.9242 0.0529
+ -5.500 -0.1206 0.01155 0.00536 -0.1078 0.9180 0.0550
+ -5.250 -0.0931 0.01118 0.00493 -0.1076 0.9117 0.0575
+ -5.000 -0.0666 0.01079 0.00457 -0.1072 0.9038 0.0614
+ -4.750 -0.0382 0.01055 0.00427 -0.1071 0.8974 0.0656
+ -4.500 -0.0118 0.01018 0.00393 -0.1067 0.8884 0.0727
+ -4.250 0.0161 0.00990 0.00363 -0.1066 0.8810 0.0812
+ -4.000 0.0434 0.00973 0.00343 -0.1063 0.8716 0.0897
+ -3.750 0.0710 0.00948 0.00321 -0.1061 0.8628 0.1006
+ -3.500 0.0987 0.00929 0.00301 -0.1060 0.8534 0.1110
+ -3.250 0.1262 0.00913 0.00285 -0.1058 0.8432 0.1221
+ -3.000 0.1541 0.00898 0.00269 -0.1056 0.8335 0.1353
+ -2.750 0.1815 0.00882 0.00255 -0.1054 0.8224 0.1520
+ -2.500 0.2090 0.00867 0.00244 -0.1053 0.8113 0.1733
+ -2.250 0.2366 0.00855 0.00235 -0.1051 0.8004 0.1993
+ -2.000 0.2641 0.00844 0.00229 -0.1049 0.7887 0.2300
+ -1.750 0.2915 0.00833 0.00224 -0.1048 0.7765 0.2579
+ -1.500 0.3190 0.00826 0.00218 -0.1046 0.7645 0.2832
+ -1.250 0.3463 0.00818 0.00213 -0.1044 0.7525 0.3117
+ -1.000 0.3734 0.00807 0.00209 -0.1042 0.7397 0.3503
+ -0.750 0.4003 0.00792 0.00208 -0.1040 0.7267 0.4055
+ -0.500 0.4265 0.00766 0.00208 -0.1038 0.7141 0.5024
+ -0.250 0.4513 0.00733 0.00214 -0.1031 0.7014 0.6445
+ 0.000 0.4741 0.00703 0.00220 -0.1018 0.6886 0.7752
+ 0.250 0.4944 0.00681 0.00229 -0.0995 0.6757 0.9086
+ 0.500 0.5429 0.00683 0.00228 -0.1037 0.6616 0.9948
+ 0.750 0.5727 0.00693 0.00228 -0.1041 0.6484 1.0000
+ 1.000 0.5983 0.00705 0.00231 -0.1035 0.6356 1.0000
+ 1.250 0.6239 0.00719 0.00235 -0.1030 0.6230 1.0000
+ 1.500 0.6499 0.00731 0.00240 -0.1026 0.6104 1.0000
+ 1.750 0.6761 0.00744 0.00246 -0.1022 0.5984 1.0000
+ 2.000 0.7023 0.00759 0.00253 -0.1018 0.5871 1.0000
+ 2.250 0.7283 0.00775 0.00260 -0.1014 0.5758 1.0000
+ 2.500 0.7548 0.00788 0.00269 -0.1011 0.5646 1.0000
+ 2.750 0.7811 0.00804 0.00279 -0.1007 0.5544 1.0000
+ 3.000 0.8073 0.00820 0.00290 -0.1004 0.5444 1.0000
+ 3.250 0.8339 0.00834 0.00302 -0.1001 0.5349 1.0000
+ 3.500 0.8600 0.00854 0.00314 -0.0998 0.5260 1.0000
+ 3.750 0.8867 0.00866 0.00327 -0.0995 0.5166 1.0000
+ 4.000 0.9127 0.00884 0.00341 -0.0992 0.5067 1.0000
+ 4.250 0.9385 0.00902 0.00355 -0.0988 0.4957 1.0000
+ 4.500 0.9647 0.00917 0.00370 -0.0984 0.4849 1.0000
+ 4.750 0.9903 0.00937 0.00385 -0.0980 0.4741 1.0000
+ 5.000 1.0155 0.00957 0.00400 -0.0975 0.4618 1.0000
+ 5.250 1.0403 0.00977 0.00416 -0.0970 0.4460 1.0000
+ 5.500 1.0652 0.00997 0.00433 -0.0964 0.4309 1.0000
+ 5.750 1.0906 0.01016 0.00452 -0.0960 0.4187 1.0000
+ 6.000 1.1155 0.01038 0.00472 -0.0955 0.4067 1.0000
+ 6.250 1.1399 0.01063 0.00495 -0.0949 0.3935 1.0000
+ 6.500 1.1639 0.01090 0.00518 -0.0942 0.3784 1.0000
+ 6.750 1.1875 0.01119 0.00543 -0.0935 0.3609 1.0000
+ 7.000 1.2103 0.01152 0.00571 -0.0927 0.3410 1.0000
+ 7.250 1.2318 0.01193 0.00603 -0.0917 0.3191 1.0000
+ 7.500 1.2528 0.01238 0.00639 -0.0906 0.2935 1.0000
+ 7.750 1.2718 0.01294 0.00682 -0.0892 0.2643 1.0000
+ 8.000 1.2891 0.01362 0.00734 -0.0876 0.2320 1.0000
+ 8.250 1.3042 0.01442 0.00794 -0.0856 0.1973 1.0000
+ 8.500 1.3175 0.01531 0.00862 -0.0834 0.1607 1.0000
+ 8.750 1.3282 0.01622 0.00934 -0.0807 0.1301 1.0000
+ 9.000 1.3388 0.01711 0.01011 -0.0781 0.1080 1.0000
+ 9.250 1.3500 0.01798 0.01089 -0.0756 0.0904 1.0000
+ 9.500 1.3615 0.01884 0.01169 -0.0732 0.0782 1.0000
+ 9.750 1.3735 0.01968 0.01251 -0.0709 0.0704 1.0000
+ 10.000 1.3845 0.02060 0.01340 -0.0687 0.0648 1.0000
+ 10.250 1.3969 0.02144 0.01429 -0.0667 0.0610 1.0000
+ 10.500 1.4094 0.02230 0.01519 -0.0648 0.0580 1.0000
+ 10.750 1.4184 0.02340 0.01630 -0.0626 0.0550 1.0000
+ 11.000 1.4270 0.02457 0.01753 -0.0605 0.0528 1.0000
+ 11.250 1.4395 0.02551 0.01855 -0.0589 0.0510 1.0000
+ 11.500 1.4503 0.02659 0.01969 -0.0573 0.0492 1.0000
+ 11.750 1.4584 0.02791 0.02104 -0.0555 0.0473 1.0000
+ 12.000 1.4602 0.02976 0.02294 -0.0533 0.0455 1.0000
+ 12.250 1.4688 0.03116 0.02443 -0.0519 0.0443 1.0000
+ 12.500 1.4791 0.03246 0.02581 -0.0507 0.0429 1.0000
+ 12.750 1.4875 0.03395 0.02738 -0.0495 0.0414 1.0000
+ 13.000 1.4938 0.03568 0.02917 -0.0483 0.0400 1.0000
+ 13.250 1.4941 0.03801 0.03154 -0.0469 0.0387 1.0000
+ 13.500 1.4941 0.04045 0.03407 -0.0457 0.0375 1.0000
+ 13.750 1.5030 0.04212 0.03584 -0.0450 0.0363 1.0000
+ 14.000 1.5091 0.04410 0.03790 -0.0444 0.0349 1.0000
+ 14.250 1.5135 0.04630 0.04017 -0.0438 0.0337 1.0000
+ 14.500 1.5126 0.04913 0.04305 -0.0432 0.0326 1.0000
+ 14.750 1.5064 0.05260 0.04659 -0.0426 0.0315 1.0000
+ 15.000 1.5124 0.05486 0.04898 -0.0424 0.0305 1.0000
+ 15.250 1.5153 0.05748 0.05169 -0.0423 0.0294 1.0000
+ 15.500 1.5171 0.06031 0.05459 -0.0423 0.0283 1.0000
+ 15.750 1.5155 0.06359 0.05793 -0.0425 0.0274 1.0000
+ 16.000 1.5055 0.06789 0.06228 -0.0426 0.0265 1.0000
+ 16.250 1.5080 0.07082 0.06534 -0.0429 0.0257 1.0000
+ 16.500 1.5081 0.07407 0.06871 -0.0433 0.0248 1.0000
+ 16.750 1.5068 0.07756 0.07229 -0.0439 0.0240 1.0000
+ 17.000 1.5043 0.08125 0.07605 -0.0445 0.0233 1.0000
+ 17.250 1.4987 0.08538 0.08024 -0.0453 0.0226 1.0000
diff --git a/Airfoils/Polars/NACA_63_412_polar_Re_100000.txt b/Airfoils/Polars/NACA_63_412_polar_Re_100000.txt
new file mode 100644
index 0000000..a1f3a80
--- /dev/null
+++ b/Airfoils/Polars/NACA_63_412_polar_Re_100000.txt
@@ -0,0 +1,114 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 63-412 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -10.500 -0.4783 0.08951 0.08431 -0.0435 1.0000 0.0331
+ -10.250 -0.4905 0.08189 0.07675 -0.0480 1.0000 0.0327
+ -10.000 -0.5183 0.07077 0.06565 -0.0563 1.0000 0.0317
+ -9.750 -0.5512 0.06300 0.05780 -0.0619 1.0000 0.0311
+ -9.500 -0.5841 0.05816 0.05286 -0.0629 1.0000 0.0306
+ -9.250 -0.6054 0.05429 0.04882 -0.0623 1.0000 0.0305
+ -9.000 -0.6197 0.05077 0.04510 -0.0608 1.0000 0.0305
+ -8.750 -0.6276 0.04780 0.04192 -0.0589 1.0000 0.0306
+ -8.500 -0.6327 0.04510 0.03900 -0.0567 1.0000 0.0309
+ -8.250 -0.6350 0.04262 0.03627 -0.0543 1.0000 0.0312
+ -8.000 -0.6320 0.04016 0.03353 -0.0525 0.9994 0.0317
+ -7.750 -0.6042 0.03712 0.03003 -0.0553 0.9922 0.0335
+ -7.500 -0.5747 0.03401 0.02632 -0.0577 0.9856 0.0354
+ -7.250 -0.5451 0.03138 0.02311 -0.0594 0.9786 0.0365
+ -7.000 -0.5139 0.02893 0.02042 -0.0612 0.9732 0.0378
+ -6.750 -0.4828 0.02752 0.01891 -0.0629 0.9664 0.0402
+ -6.500 -0.4482 0.02608 0.01724 -0.0648 0.9617 0.0430
+ -6.250 -0.4174 0.02465 0.01560 -0.0656 0.9546 0.0449
+ -6.000 -0.3825 0.02324 0.01405 -0.0673 0.9499 0.0474
+ -5.750 -0.3523 0.02220 0.01305 -0.0684 0.9426 0.0511
+ -5.500 -0.3195 0.02124 0.01200 -0.0697 0.9363 0.0551
+ -5.250 -0.2886 0.02028 0.01096 -0.0706 0.9293 0.0590
+ -5.000 -0.2581 0.01951 0.01018 -0.0716 0.9217 0.0661
+ -4.750 -0.2267 0.01870 0.00933 -0.0727 0.9150 0.0745
+ -4.500 -0.1977 0.01799 0.00862 -0.0734 0.9068 0.0891
+ -4.250 -0.1667 0.01711 0.00789 -0.0745 0.9003 0.1256
+ -4.000 -0.1415 0.01578 0.00726 -0.0752 0.8916 0.2573
+ -3.750 -0.1144 0.01487 0.00712 -0.0757 0.8849 0.4407
+ -3.500 -0.0882 0.01474 0.00715 -0.0751 0.8763 0.5184
+ -3.250 -0.0597 0.01473 0.00720 -0.0748 0.8696 0.5748
+ -3.000 -0.0344 0.01484 0.00734 -0.0738 0.8610 0.6176
+ -2.750 -0.0073 0.01497 0.00745 -0.0729 0.8543 0.6488
+ -2.500 0.0176 0.01509 0.00754 -0.0718 0.8457 0.6695
+ -2.250 0.0463 0.01510 0.00744 -0.0716 0.8393 0.6843
+ -2.000 0.0725 0.01510 0.00736 -0.0712 0.8305 0.6966
+ -1.750 0.1018 0.01505 0.00720 -0.0712 0.8244 0.7083
+ -1.500 0.1269 0.01507 0.00718 -0.0705 0.8155 0.7176
+ -1.250 0.1558 0.01502 0.00703 -0.0704 0.8093 0.7277
+ -1.000 0.1813 0.01504 0.00701 -0.0700 0.8003 0.7376
+ -0.750 0.2092 0.01500 0.00692 -0.0696 0.7941 0.7465
+ -0.500 0.2351 0.01503 0.00691 -0.0693 0.7852 0.7568
+ -0.250 0.2622 0.01500 0.00685 -0.0688 0.7788 0.7656
+ 0.000 0.2876 0.01504 0.00689 -0.0683 0.7701 0.7753
+ 0.250 0.3154 0.01503 0.00683 -0.0681 0.7636 0.7855
+ 0.500 0.3399 0.01508 0.00690 -0.0674 0.7551 0.7947
+ 0.750 0.3672 0.01508 0.00687 -0.0671 0.7484 0.8049
+ 1.000 0.3923 0.01514 0.00697 -0.0665 0.7400 0.8154
+ 1.250 0.4185 0.01514 0.00697 -0.0659 0.7333 0.8252
+ 1.500 0.4434 0.01521 0.00708 -0.0653 0.7250 0.8362
+ 1.750 0.4699 0.01523 0.00711 -0.0648 0.7182 0.8476
+ 2.000 0.4939 0.01529 0.00723 -0.0640 0.7099 0.8588
+ 2.250 0.5195 0.01529 0.00727 -0.0633 0.7029 0.8706
+ 2.500 0.5438 0.01535 0.00740 -0.0626 0.6946 0.8834
+ 2.750 0.5695 0.01534 0.00745 -0.0619 0.6874 0.8970
+ 3.000 0.5940 0.01539 0.00759 -0.0612 0.6786 0.9119
+ 3.250 0.6222 0.01536 0.00761 -0.0611 0.6713 0.9278
+ 3.500 0.6518 0.01542 0.00781 -0.0615 0.6616 0.9469
+ 3.750 0.6880 0.01543 0.00790 -0.0632 0.6529 0.9711
+ 4.000 0.7192 0.01553 0.00807 -0.0642 0.6423 1.0000
+ 4.250 0.7467 0.01568 0.00829 -0.0643 0.6301 1.0000
+ 4.500 0.7742 0.01578 0.00844 -0.0643 0.6163 1.0000
+ 4.750 0.8012 0.01585 0.00855 -0.0642 0.5995 1.0000
+ 5.000 0.8279 0.01590 0.00863 -0.0638 0.5803 1.0000
+ 5.250 0.8527 0.01600 0.00877 -0.0632 0.5553 1.0000
+ 5.500 0.8763 0.01611 0.00883 -0.0623 0.5224 1.0000
+ 6.000 0.9189 0.01676 0.00924 -0.0599 0.4276 1.0000
+ 6.250 0.9359 0.01747 0.00961 -0.0583 0.3536 1.0000
+ 6.500 0.9469 0.01877 0.01034 -0.0561 0.2602 1.0000
+ 6.750 0.9566 0.02034 0.01138 -0.0541 0.1783 1.0000
+ 7.000 0.9673 0.02189 0.01254 -0.0523 0.1203 1.0000
+ 7.250 0.9796 0.02327 0.01370 -0.0506 0.0900 1.0000
+ 7.500 0.9931 0.02449 0.01485 -0.0489 0.0736 1.0000
+ 7.750 1.0056 0.02574 0.01605 -0.0472 0.0638 1.0000
+ 8.000 1.0186 0.02687 0.01724 -0.0455 0.0568 1.0000
+ 8.250 1.0281 0.02813 0.01853 -0.0434 0.0522 1.0000
+ 8.500 1.0395 0.02934 0.01984 -0.0416 0.0482 1.0000
+ 8.750 1.0494 0.03070 0.02122 -0.0398 0.0451 1.0000
+ 9.000 1.0595 0.03220 0.02276 -0.0381 0.0426 1.0000
+ 9.250 1.0728 0.03354 0.02423 -0.0368 0.0397 1.0000
+ 9.500 1.0861 0.03501 0.02578 -0.0355 0.0378 1.0000
+ 9.750 1.0998 0.03659 0.02741 -0.0344 0.0363 1.0000
+ 10.000 1.1158 0.03843 0.02925 -0.0335 0.0349 1.0000
+ 10.250 1.1351 0.04023 0.03120 -0.0327 0.0337 1.0000
+ 10.500 1.1522 0.04202 0.03321 -0.0319 0.0322 1.0000
+ 10.750 1.1665 0.04388 0.03527 -0.0310 0.0308 1.0000
+ 11.000 1.1787 0.04583 0.03742 -0.0300 0.0296 1.0000
+ 11.250 1.1908 0.04801 0.03978 -0.0291 0.0288 1.0000
+ 11.500 1.2009 0.05036 0.04230 -0.0282 0.0282 1.0000
+ 11.750 1.2091 0.05293 0.04503 -0.0273 0.0277 1.0000
+ 12.000 1.2152 0.05586 0.04815 -0.0265 0.0273 1.0000
+ 12.250 1.2175 0.05930 0.05178 -0.0256 0.0269 1.0000
+ 12.500 1.2111 0.06277 0.05558 -0.0246 0.0268 1.0000
+ 12.750 1.2019 0.06659 0.05972 -0.0241 0.0266 1.0000
+ 13.000 1.1897 0.07081 0.06425 -0.0240 0.0265 1.0000
+ 13.250 1.1750 0.07547 0.06920 -0.0245 0.0265 1.0000
+ 13.500 1.1579 0.08067 0.07468 -0.0257 0.0264 1.0000
+ 13.750 1.1390 0.08644 0.08071 -0.0276 0.0264 1.0000
+ 14.000 1.1186 0.09282 0.08733 -0.0304 0.0264 1.0000
+ 14.250 1.0970 0.09992 0.09465 -0.0341 0.0264 1.0000
+ 14.500 1.0747 0.10777 0.10271 -0.0387 0.0265 1.0000
+ 14.750 1.0520 0.11657 0.11168 -0.0443 0.0267 1.0000
+ 15.000 1.0291 0.12642 0.12167 -0.0508 0.0269 1.0000
diff --git a/Airfoils/Polars/NACA_63_412_polar_Re_1000000.txt b/Airfoils/Polars/NACA_63_412_polar_Re_1000000.txt
new file mode 100644
index 0000000..18c0a74
--- /dev/null
+++ b/Airfoils/Polars/NACA_63_412_polar_Re_1000000.txt
@@ -0,0 +1,151 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 63-412 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -16.500 -0.9681 0.07832 0.07592 -0.0461 1.0000 0.0061
+ -16.250 -1.0143 0.06526 0.06263 -0.0546 1.0000 0.0060
+ -16.000 -1.0409 0.05685 0.05404 -0.0604 1.0000 0.0059
+ -15.750 -1.0622 0.05000 0.04703 -0.0650 1.0000 0.0060
+ -15.500 -1.0788 0.04436 0.04123 -0.0686 1.0000 0.0061
+ -15.250 -1.0869 0.04026 0.03700 -0.0709 1.0000 0.0061
+ -15.000 -1.0925 0.03675 0.03336 -0.0726 1.0000 0.0062
+ -14.750 -1.0945 0.03389 0.03039 -0.0736 1.0000 0.0062
+ -14.500 -1.0920 0.03170 0.02811 -0.0741 1.0000 0.0063
+ -14.250 -1.0919 0.02948 0.02578 -0.0741 1.0000 0.0064
+ -14.000 -1.0864 0.02795 0.02416 -0.0737 1.0000 0.0066
+ -13.750 -1.0846 0.02637 0.02249 -0.0727 1.0000 0.0066
+ -13.500 -1.0834 0.02518 0.02122 -0.0707 1.0000 0.0067
+ -13.250 -1.0741 0.02419 0.02015 -0.0694 1.0000 0.0069
+ -13.000 -1.0625 0.02321 0.01909 -0.0682 1.0000 0.0070
+ -12.750 -1.0492 0.02232 0.01811 -0.0670 1.0000 0.0072
+ -12.500 -1.0339 0.02154 0.01726 -0.0659 1.0000 0.0074
+ -12.250 -1.0185 0.02073 0.01637 -0.0647 1.0000 0.0075
+ -12.000 -1.0004 0.02005 0.01561 -0.0639 0.9998 0.0078
+ -11.500 -0.9441 0.01844 0.01381 -0.0662 0.9889 0.0082
+ -11.250 -0.9140 0.01757 0.01287 -0.0678 0.9836 0.0085
+ -11.000 -0.8823 0.01693 0.01219 -0.0694 0.9772 0.0089
+ -10.750 -0.8496 0.01630 0.01150 -0.0712 0.9695 0.0092
+ -10.500 -0.8160 0.01574 0.01086 -0.0731 0.9607 0.0096
+ -10.250 -0.7840 0.01518 0.01022 -0.0746 0.9485 0.0100
+ -9.750 -0.7302 0.01431 0.00914 -0.0752 0.9190 0.0106
+ -9.500 -0.7061 0.01382 0.00856 -0.0749 0.9060 0.0110
+ -9.250 -0.6815 0.01341 0.00809 -0.0746 0.8944 0.0115
+ -9.000 -0.6564 0.01311 0.00772 -0.0744 0.8832 0.0120
+ -8.500 -0.6057 0.01251 0.00696 -0.0739 0.8590 0.0130
+ -8.250 -0.5800 0.01220 0.00657 -0.0737 0.8473 0.0135
+ -8.000 -0.5541 0.01192 0.00619 -0.0736 0.8375 0.0138
+ -7.750 -0.5282 0.01153 0.00574 -0.0735 0.8284 0.0143
+ -7.500 -0.5018 0.01122 0.00539 -0.0735 0.8205 0.0151
+ -7.250 -0.4750 0.01097 0.00510 -0.0734 0.8123 0.0158
+ -7.000 -0.4480 0.01073 0.00481 -0.0734 0.8047 0.0165
+ -6.750 -0.4209 0.01050 0.00451 -0.0735 0.7966 0.0171
+ -6.500 -0.3936 0.01028 0.00423 -0.0735 0.7893 0.0176
+ -6.250 -0.3664 0.01000 0.00390 -0.0735 0.7820 0.0186
+ -6.000 -0.3389 0.00977 0.00365 -0.0736 0.7752 0.0199
+ -5.750 -0.3112 0.00957 0.00341 -0.0737 0.7677 0.0209
+ -5.500 -0.2835 0.00940 0.00319 -0.0738 0.7607 0.0219
+ -5.250 -0.2555 0.00923 0.00298 -0.0739 0.7536 0.0227
+ -5.000 -0.2278 0.00903 0.00275 -0.0740 0.7471 0.0251
+ -4.750 -0.1996 0.00886 0.00257 -0.0741 0.7402 0.0272
+ -4.500 -0.1715 0.00874 0.00239 -0.0742 0.7330 0.0288
+ -4.250 -0.1433 0.00856 0.00222 -0.0744 0.7263 0.0324
+ -4.000 -0.1151 0.00844 0.00207 -0.0746 0.7191 0.0360
+ -3.750 -0.0868 0.00828 0.00192 -0.0747 0.7125 0.0429
+ -3.500 -0.0585 0.00813 0.00178 -0.0749 0.7051 0.0536
+ -3.000 -0.0019 0.00773 0.00151 -0.0754 0.6912 0.1021
+ -2.750 0.0262 0.00746 0.00138 -0.0756 0.6844 0.1495
+ -2.500 0.0545 0.00712 0.00124 -0.0760 0.6773 0.2160
+ -2.250 0.0826 0.00673 0.00110 -0.0764 0.6703 0.3012
+ -2.000 0.1111 0.00641 0.00101 -0.0768 0.6635 0.3768
+ -1.750 0.1395 0.00623 0.00096 -0.0771 0.6563 0.4309
+ -1.500 0.1681 0.00605 0.00094 -0.0774 0.6495 0.4880
+ -1.250 0.1967 0.00599 0.00093 -0.0776 0.6423 0.5168
+ -1.000 0.2254 0.00594 0.00093 -0.0778 0.6355 0.5432
+ -0.750 0.2541 0.00592 0.00094 -0.0780 0.6282 0.5616
+ -0.500 0.2828 0.00594 0.00094 -0.0781 0.6216 0.5722
+ -0.250 0.3115 0.00595 0.00095 -0.0783 0.6140 0.5829
+ 0.000 0.3401 0.00597 0.00096 -0.0785 0.6070 0.5917
+ 0.250 0.3688 0.00599 0.00098 -0.0786 0.5991 0.6000
+ 0.500 0.3973 0.00602 0.00101 -0.0788 0.5916 0.6077
+ 0.750 0.4259 0.00605 0.00104 -0.0789 0.5830 0.6156
+ 1.000 0.4544 0.00609 0.00108 -0.0790 0.5749 0.6236
+ 1.250 0.4829 0.00614 0.00112 -0.0792 0.5660 0.6320
+ 1.500 0.5113 0.00618 0.00118 -0.0793 0.5572 0.6407
+ 1.750 0.5395 0.00625 0.00124 -0.0794 0.5466 0.6500
+ 2.250 0.5954 0.00643 0.00139 -0.0795 0.5199 0.6683
+ 2.500 0.6232 0.00655 0.00147 -0.0795 0.5047 0.6774
+ 2.750 0.6508 0.00668 0.00157 -0.0795 0.4863 0.6855
+ 3.000 0.6778 0.00687 0.00168 -0.0794 0.4592 0.6934
+ 3.250 0.7032 0.00722 0.00185 -0.0791 0.4094 0.7008
+ 3.500 0.7280 0.00765 0.00206 -0.0787 0.3560 0.7081
+ 3.750 0.7529 0.00807 0.00231 -0.0783 0.3092 0.7149
+ 4.000 0.7768 0.00861 0.00261 -0.0779 0.2532 0.7222
+ 4.250 0.8001 0.00920 0.00295 -0.0773 0.1951 0.7288
+ 4.500 0.8233 0.00979 0.00331 -0.0768 0.1426 0.7362
+ 4.750 0.8470 0.01032 0.00366 -0.0763 0.1009 0.7433
+ 5.000 0.8715 0.01074 0.00398 -0.0758 0.0740 0.7509
+ 5.250 0.8962 0.01112 0.00428 -0.0755 0.0549 0.7583
+ 5.500 0.9212 0.01146 0.00459 -0.0751 0.0424 0.7665
+ 5.750 0.9463 0.01177 0.00488 -0.0748 0.0341 0.7743
+ 6.000 0.9715 0.01206 0.00518 -0.0744 0.0295 0.7830
+ 6.250 0.9966 0.01234 0.00547 -0.0741 0.0261 0.7914
+ 6.500 1.0216 0.01262 0.00578 -0.0737 0.0234 0.8007
+ 6.750 1.0461 0.01294 0.00612 -0.0733 0.0208 0.8100
+ 7.000 1.0707 0.01322 0.00644 -0.0728 0.0190 0.8197
+ 7.250 1.0945 0.01356 0.00680 -0.0723 0.0169 0.8306
+ 7.500 1.1181 0.01388 0.00718 -0.0717 0.0155 0.8416
+ 7.750 1.1414 0.01421 0.00756 -0.0711 0.0144 0.8533
+ 8.000 1.1640 0.01458 0.00796 -0.0703 0.0134 0.8659
+ 8.250 1.1854 0.01500 0.00844 -0.0694 0.0125 0.8797
+ 8.500 1.2068 0.01535 0.00886 -0.0684 0.0121 0.8951
+ 8.750 1.2267 0.01569 0.00930 -0.0671 0.0117 0.9152
+ 9.000 1.2460 0.01599 0.00972 -0.0657 0.0114 1.0000
+ 9.250 1.2672 0.01647 0.01022 -0.0649 0.0110 1.0000
+ 9.500 1.2877 0.01696 0.01074 -0.0640 0.0106 1.0000
+ 9.750 1.3072 0.01750 0.01130 -0.0629 0.0102 1.0000
+ 10.000 1.3250 0.01811 0.01195 -0.0616 0.0099 1.0000
+ 10.250 1.3394 0.01879 0.01268 -0.0597 0.0096 1.0000
+ 10.500 1.3520 0.01947 0.01341 -0.0575 0.0094 1.0000
+ 10.750 1.3654 0.02011 0.01410 -0.0555 0.0093 1.0000
+ 11.000 1.3783 0.02079 0.01484 -0.0536 0.0092 1.0000
+ 11.250 1.3903 0.02156 0.01568 -0.0517 0.0091 1.0000
+ 11.500 1.4018 0.02239 0.01658 -0.0498 0.0090 1.0000
+ 11.750 1.4128 0.02329 0.01755 -0.0481 0.0089 1.0000
+ 12.000 1.4237 0.02426 0.01858 -0.0464 0.0087 1.0000
+ 12.250 1.4339 0.02530 0.01969 -0.0449 0.0086 1.0000
+ 12.500 1.4446 0.02636 0.02080 -0.0435 0.0084 1.0000
+ 12.750 1.4534 0.02761 0.02212 -0.0420 0.0083 1.0000
+ 13.000 1.4622 0.02891 0.02350 -0.0407 0.0082 1.0000
+ 13.250 1.4705 0.03030 0.02497 -0.0395 0.0081 1.0000
+ 13.500 1.4780 0.03181 0.02655 -0.0384 0.0079 1.0000
+ 13.750 1.4848 0.03344 0.02826 -0.0374 0.0078 1.0000
+ 14.000 1.4907 0.03521 0.03010 -0.0365 0.0077 1.0000
+ 14.250 1.4960 0.03710 0.03207 -0.0357 0.0076 1.0000
+ 14.500 1.5001 0.03918 0.03423 -0.0350 0.0075 1.0000
+ 14.750 1.5032 0.04142 0.03655 -0.0345 0.0074 1.0000
+ 15.000 1.5033 0.04407 0.03930 -0.0341 0.0073 1.0000
+ 15.250 1.5014 0.04705 0.04238 -0.0338 0.0072 1.0000
+ 15.500 1.4958 0.05058 0.04603 -0.0339 0.0070 1.0000
+ 15.750 1.4985 0.05326 0.04881 -0.0341 0.0070 1.0000
+ 16.000 1.5000 0.05619 0.05185 -0.0344 0.0069 1.0000
+ 16.250 1.4988 0.05957 0.05534 -0.0350 0.0069 1.0000
+ 16.500 1.4969 0.06318 0.05906 -0.0359 0.0068 1.0000
+ 16.750 1.4949 0.06694 0.06293 -0.0369 0.0068 1.0000
+ 17.000 1.4909 0.07110 0.06721 -0.0382 0.0067 1.0000
+ 17.250 1.4857 0.07560 0.07184 -0.0398 0.0067 1.0000
+ 17.500 1.4802 0.08031 0.07666 -0.0416 0.0066 1.0000
+ 17.750 1.4734 0.08538 0.08186 -0.0437 0.0066 1.0000
+ 18.000 1.4642 0.09101 0.08762 -0.0462 0.0065 1.0000
+ 18.250 1.4542 0.09698 0.09371 -0.0491 0.0065 1.0000
+ 18.500 1.4433 0.10330 0.10017 -0.0523 0.0065 1.0000
+ 18.750 1.4301 0.11024 0.10725 -0.0559 0.0064 1.0000
+ 19.000 1.4164 0.11746 0.11460 -0.0599 0.0064 1.0000
+ 19.250 1.4014 0.12518 0.12246 -0.0644 0.0064 1.0000
diff --git a/Airfoils/Polars/NACA_63_412_polar_Re_200000.txt b/Airfoils/Polars/NACA_63_412_polar_Re_200000.txt
new file mode 100644
index 0000000..2357deb
--- /dev/null
+++ b/Airfoils/Polars/NACA_63_412_polar_Re_200000.txt
@@ -0,0 +1,123 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 63-412 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -11.250 -0.5922 0.06654 0.06268 -0.0567 1.0000 0.0176
+ -11.000 -0.6185 0.05869 0.05470 -0.0629 1.0000 0.0176
+ -10.750 -0.6387 0.05337 0.04925 -0.0664 1.0000 0.0175
+ -10.500 -0.6626 0.04883 0.04454 -0.0678 1.0000 0.0175
+ -10.250 -0.6859 0.04557 0.04111 -0.0665 1.0000 0.0175
+ -10.000 -0.7050 0.04177 0.03699 -0.0647 1.0000 0.0176
+ -9.750 -0.7158 0.03843 0.03329 -0.0625 1.0000 0.0178
+ -9.500 -0.7160 0.03625 0.03096 -0.0604 1.0000 0.0180
+ -9.250 -0.7064 0.03457 0.02915 -0.0595 0.9983 0.0184
+ -9.000 -0.6795 0.03247 0.02681 -0.0618 0.9912 0.0190
+ -8.750 -0.6503 0.03076 0.02489 -0.0642 0.9852 0.0201
+ -8.500 -0.6232 0.02850 0.02223 -0.0659 0.9777 0.0213
+ -8.250 -0.5942 0.02623 0.01947 -0.0675 0.9714 0.0224
+ -8.000 -0.5650 0.02427 0.01728 -0.0689 0.9649 0.0232
+ -7.750 -0.5349 0.02305 0.01598 -0.0703 0.9581 0.0241
+ -7.500 -0.5035 0.02210 0.01492 -0.0718 0.9514 0.0255
+ -7.250 -0.4725 0.02108 0.01370 -0.0729 0.9445 0.0272
+ -7.000 -0.4418 0.02000 0.01244 -0.0739 0.9370 0.0284
+ -6.750 -0.4122 0.01882 0.01117 -0.0747 0.9295 0.0295
+ -6.500 -0.3830 0.01800 0.01034 -0.0755 0.9209 0.0313
+ -6.250 -0.3545 0.01734 0.00960 -0.0760 0.9118 0.0334
+ -6.000 -0.3259 0.01664 0.00880 -0.0764 0.9033 0.0353
+ -5.750 -0.2995 0.01603 0.00809 -0.0763 0.8936 0.0369
+ -5.500 -0.2731 0.01528 0.00731 -0.0765 0.8852 0.0397
+ -5.250 -0.2468 0.01483 0.00680 -0.0765 0.8759 0.0433
+ -5.000 -0.2200 0.01439 0.00626 -0.0764 0.8677 0.0467
+ -4.750 -0.1937 0.01387 0.00571 -0.0764 0.8592 0.0521
+ -4.500 -0.1670 0.01347 0.00526 -0.0764 0.8511 0.0603
+ -4.250 -0.1402 0.01303 0.00483 -0.0764 0.8430 0.0753
+ -4.000 -0.1138 0.01252 0.00446 -0.0765 0.8350 0.1122
+ -3.750 -0.0878 0.01177 0.00405 -0.0768 0.8273 0.2014
+ -3.500 -0.0625 0.01089 0.00375 -0.0771 0.8195 0.3463
+ -3.250 -0.0360 0.01049 0.00365 -0.0771 0.8119 0.4466
+ -3.000 -0.0089 0.01033 0.00360 -0.0769 0.8044 0.5004
+ -2.750 0.0183 0.01025 0.00359 -0.0767 0.7970 0.5456
+ -2.500 0.0455 0.01022 0.00360 -0.0764 0.7897 0.5836
+ -2.250 0.0725 0.01020 0.00360 -0.0760 0.7820 0.6103
+ -2.000 0.1000 0.01022 0.00359 -0.0757 0.7749 0.6315
+ -1.750 0.1275 0.01022 0.00357 -0.0755 0.7672 0.6441
+ -1.500 0.1553 0.01023 0.00353 -0.0754 0.7602 0.6549
+ -1.250 0.1832 0.01024 0.00349 -0.0753 0.7523 0.6653
+ -1.000 0.2112 0.01025 0.00346 -0.0752 0.7455 0.6748
+ -0.750 0.2389 0.01026 0.00345 -0.0751 0.7375 0.6836
+ -0.500 0.2672 0.01029 0.00341 -0.0751 0.7308 0.6934
+ -0.250 0.2946 0.01030 0.00343 -0.0750 0.7227 0.7012
+ 0.000 0.3227 0.01034 0.00341 -0.0749 0.7161 0.7103
+ 0.250 0.3501 0.01036 0.00346 -0.0748 0.7080 0.7184
+ 0.500 0.3780 0.01040 0.00347 -0.0747 0.7013 0.7272
+ 0.750 0.4054 0.01043 0.00354 -0.0746 0.6932 0.7357
+ 1.000 0.4330 0.01048 0.00357 -0.0744 0.6864 0.7441
+ 1.250 0.4605 0.01053 0.00366 -0.0743 0.6783 0.7531
+ 1.500 0.4877 0.01058 0.00371 -0.0741 0.6714 0.7613
+ 1.750 0.5152 0.01064 0.00382 -0.0740 0.6632 0.7707
+ 2.000 0.5420 0.01070 0.00390 -0.0736 0.6560 0.7789
+ 2.250 0.5693 0.01077 0.00402 -0.0735 0.6475 0.7882
+ 2.500 0.5959 0.01084 0.00412 -0.0731 0.6400 0.7969
+ 2.750 0.6226 0.01091 0.00426 -0.0728 0.6309 0.8060
+ 3.000 0.6491 0.01100 0.00437 -0.0725 0.6226 0.8156
+ 3.250 0.6749 0.01107 0.00452 -0.0720 0.6125 0.8246
+ 3.500 0.7010 0.01115 0.00465 -0.0715 0.6008 0.8347
+ 3.750 0.7261 0.01123 0.00477 -0.0708 0.5873 0.8448
+ 4.000 0.7506 0.01131 0.00489 -0.0701 0.5709 0.8550
+ 4.250 0.7749 0.01141 0.00500 -0.0692 0.5526 0.8663
+ 4.750 0.8202 0.01171 0.00523 -0.0669 0.4968 0.8916
+ 5.000 0.8410 0.01193 0.00537 -0.0655 0.4590 0.9069
+ 5.250 0.8612 0.01220 0.00555 -0.0640 0.4164 0.9258
+ 5.500 0.8827 0.01269 0.00582 -0.0631 0.3565 0.9584
+ 5.750 0.9022 0.01365 0.00637 -0.0624 0.2750 1.0000
+ 6.000 0.9187 0.01482 0.00710 -0.0613 0.1962 1.0000
+ 6.250 0.9358 0.01594 0.00787 -0.0603 0.1347 1.0000
+ 6.500 0.9536 0.01697 0.00863 -0.0593 0.0915 1.0000
+ 6.750 0.9724 0.01788 0.00939 -0.0583 0.0661 1.0000
+ 7.000 0.9919 0.01869 0.01013 -0.0574 0.0526 1.0000
+ 7.250 1.0110 0.01948 0.01090 -0.0564 0.0446 1.0000
+ 7.500 1.0306 0.02020 0.01168 -0.0555 0.0394 1.0000
+ 7.750 1.0476 0.02108 0.01255 -0.0543 0.0353 1.0000
+ 8.000 1.0644 0.02193 0.01349 -0.0529 0.0328 1.0000
+ 8.250 1.0800 0.02277 0.01441 -0.0514 0.0306 1.0000
+ 8.500 1.0930 0.02368 0.01537 -0.0496 0.0288 1.0000
+ 8.750 1.1019 0.02488 0.01659 -0.0473 0.0272 1.0000
+ 9.000 1.1143 0.02590 0.01770 -0.0456 0.0260 1.0000
+ 9.250 1.1268 0.02695 0.01885 -0.0440 0.0246 1.0000
+ 9.500 1.1387 0.02808 0.02006 -0.0424 0.0235 1.0000
+ 9.750 1.1498 0.02932 0.02136 -0.0409 0.0227 1.0000
+ 10.000 1.1604 0.03066 0.02276 -0.0395 0.0220 1.0000
+ 10.250 1.1699 0.03218 0.02433 -0.0381 0.0214 1.0000
+ 10.500 1.1798 0.03390 0.02609 -0.0367 0.0208 1.0000
+ 10.750 1.1923 0.03538 0.02772 -0.0356 0.0204 1.0000
+ 11.000 1.2047 0.03697 0.02945 -0.0345 0.0199 1.0000
+ 11.250 1.2168 0.03866 0.03129 -0.0334 0.0195 1.0000
+ 11.500 1.2281 0.04043 0.03321 -0.0324 0.0190 1.0000
+ 11.750 1.2381 0.04228 0.03521 -0.0315 0.0185 1.0000
+ 12.000 1.2465 0.04415 0.03723 -0.0306 0.0180 1.0000
+ 12.250 1.2536 0.04614 0.03935 -0.0298 0.0176 1.0000
+ 12.500 1.2594 0.04821 0.04155 -0.0292 0.0171 1.0000
+ 12.750 1.2644 0.05049 0.04395 -0.0286 0.0168 1.0000
+ 13.000 1.2684 0.05302 0.04661 -0.0281 0.0166 1.0000
+ 13.250 1.2708 0.05589 0.04961 -0.0277 0.0164 1.0000
+ 13.500 1.2700 0.05924 0.05315 -0.0274 0.0162 1.0000
+ 13.750 1.2651 0.06291 0.05708 -0.0274 0.0160 1.0000
+ 14.000 1.2577 0.06701 0.06145 -0.0279 0.0159 1.0000
+ 14.250 1.2476 0.07160 0.06630 -0.0287 0.0158 1.0000
+ 14.500 1.2353 0.07669 0.07165 -0.0301 0.0158 1.0000
+ 14.750 1.2206 0.08236 0.07757 -0.0321 0.0157 1.0000
+ 15.000 1.2039 0.08874 0.08420 -0.0349 0.0156 1.0000
+ 15.250 1.1855 0.09582 0.09152 -0.0384 0.0156 1.0000
+ 15.500 1.1654 0.10373 0.09966 -0.0428 0.0156 1.0000
+ 15.750 1.1439 0.11253 0.10868 -0.0481 0.0156 1.0000
+ 16.000 1.1203 0.12262 0.11898 -0.0546 0.0157 1.0000
+ 16.250 1.0940 0.13428 0.13082 -0.0624 0.0158 1.0000
+ 16.500 1.0626 0.14877 0.14548 -0.0720 0.0160 1.0000
diff --git a/Airfoils/Polars/NACA_63_412_polar_Re_50000.txt b/Airfoils/Polars/NACA_63_412_polar_Re_50000.txt
new file mode 100644
index 0000000..7a1e655
--- /dev/null
+++ b/Airfoils/Polars/NACA_63_412_polar_Re_50000.txt
@@ -0,0 +1,109 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 63-412 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -10.750 -0.4759 0.10169 0.09432 -0.0425 1.0000 0.0575
+ -10.500 -0.4749 0.09720 0.08987 -0.0440 1.0000 0.0571
+ -10.250 -0.4768 0.09223 0.08495 -0.0462 1.0000 0.0567
+ -10.000 -0.4815 0.08675 0.07952 -0.0491 1.0000 0.0561
+ -9.750 -0.4911 0.08100 0.07383 -0.0525 1.0000 0.0556
+ -9.500 -0.5061 0.07551 0.06838 -0.0559 1.0000 0.0551
+ -9.250 -0.5259 0.07084 0.06374 -0.0581 1.0000 0.0546
+ -9.000 -0.5460 0.06689 0.05978 -0.0588 1.0000 0.0542
+ -8.750 -0.5620 0.06310 0.05590 -0.0587 1.0000 0.0539
+ -8.500 -0.5747 0.05950 0.05216 -0.0580 1.0000 0.0538
+ -8.250 -0.5831 0.05611 0.04857 -0.0568 1.0000 0.0538
+ -8.000 -0.5881 0.05292 0.04513 -0.0554 1.0000 0.0543
+ -7.750 -0.5903 0.04991 0.04178 -0.0538 1.0000 0.0552
+ -7.500 -0.5891 0.04703 0.03848 -0.0521 1.0000 0.0564
+ -7.250 -0.5841 0.04432 0.03525 -0.0505 1.0000 0.0575
+ -7.000 -0.5745 0.04174 0.03244 -0.0491 1.0000 0.0586
+ -6.750 -0.5624 0.03952 0.03001 -0.0478 1.0000 0.0596
+ -6.500 -0.5485 0.03754 0.02781 -0.0465 1.0000 0.0610
+ -6.250 -0.5334 0.03591 0.02598 -0.0453 1.0000 0.0636
+ -6.000 -0.5169 0.03434 0.02410 -0.0441 1.0000 0.0670
+ -5.750 -0.4988 0.03284 0.02222 -0.0428 1.0000 0.0695
+ -5.500 -0.4673 0.03111 0.02040 -0.0439 0.9948 0.0726
+ -5.250 -0.4338 0.02988 0.01905 -0.0454 0.9881 0.0790
+ -5.000 -0.4009 0.02870 0.01773 -0.0465 0.9815 0.0855
+ -4.750 -0.3671 0.02764 0.01657 -0.0480 0.9749 0.0939
+ -4.500 -0.3361 0.02661 0.01552 -0.0493 0.9669 0.1064
+ -4.250 -0.3029 0.02552 0.01439 -0.0511 0.9598 0.1257
+ -3.750 -0.2443 0.02196 0.01275 -0.0550 0.9451 0.4112
+ -3.500 -0.2174 0.02206 0.01317 -0.0540 0.9366 0.5528
+ -3.250 -0.1933 0.02254 0.01373 -0.0520 0.9280 0.6249
+ -3.000 -0.1707 0.02326 0.01446 -0.0490 0.9199 0.6833
+ -2.750 -0.1506 0.02372 0.01484 -0.0460 0.9111 0.7197
+ -2.500 -0.1219 0.02392 0.01488 -0.0452 0.9040 0.7447
+ -2.250 -0.0966 0.02395 0.01473 -0.0445 0.8955 0.7618
+ -2.000 -0.0643 0.02393 0.01452 -0.0452 0.8889 0.7773
+ -1.750 -0.0386 0.02390 0.01434 -0.0449 0.8802 0.7916
+ -1.500 -0.0070 0.02383 0.01411 -0.0456 0.8734 0.8054
+ -1.250 0.0176 0.02381 0.01397 -0.0452 0.8647 0.8192
+ -1.000 0.0489 0.02374 0.01377 -0.0458 0.8580 0.8326
+ -0.750 0.0728 0.02373 0.01368 -0.0453 0.8492 0.8460
+ -0.500 0.1040 0.02366 0.01352 -0.0459 0.8425 0.8594
+ -0.250 0.1285 0.02366 0.01346 -0.0455 0.8338 0.8738
+ 0.000 0.1608 0.02358 0.01332 -0.0464 0.8271 0.8882
+ 0.250 0.1881 0.02360 0.01331 -0.0466 0.8187 0.9038
+ 0.500 0.2238 0.02355 0.01322 -0.0482 0.8120 0.9190
+ 0.750 0.2591 0.02359 0.01325 -0.0501 0.8044 0.9352
+ 1.000 0.3013 0.02360 0.01325 -0.0532 0.7974 0.9508
+ 1.250 0.3476 0.02362 0.01328 -0.0572 0.7911 0.9662
+ 1.500 0.3902 0.02373 0.01342 -0.0608 0.7828 0.9869
+ 1.750 0.4180 0.02388 0.01356 -0.0616 0.7749 1.0000
+ 2.000 0.4383 0.02413 0.01381 -0.0611 0.7659 1.0000
+ 2.250 0.4599 0.02448 0.01416 -0.0609 0.7568 1.0000
+ 2.500 0.4894 0.02467 0.01437 -0.0616 0.7495 1.0000
+ 2.750 0.5119 0.02512 0.01485 -0.0616 0.7398 1.0000
+ 3.000 0.5443 0.02527 0.01505 -0.0625 0.7332 1.0000
+ 3.250 0.5662 0.02580 0.01564 -0.0623 0.7228 1.0000
+ 3.500 0.6008 0.02586 0.01576 -0.0632 0.7167 1.0000
+ 3.750 0.6216 0.02645 0.01646 -0.0628 0.7054 1.0000
+ 4.000 0.6482 0.02682 0.01693 -0.0629 0.6962 1.0000
+ 4.250 0.6784 0.02699 0.01721 -0.0631 0.6877 1.0000
+ 4.500 0.7013 0.02748 0.01785 -0.0627 0.6762 1.0000
+ 4.750 0.7294 0.02769 0.01820 -0.0626 0.6661 1.0000
+ 5.000 0.7613 0.02762 0.01828 -0.0626 0.6560 1.0000
+ 5.250 0.7860 0.02782 0.01865 -0.0619 0.6423 1.0000
+ 5.500 0.8128 0.02777 0.01881 -0.0611 0.6275 1.0000
+ 5.750 0.8424 0.02738 0.01859 -0.0602 0.6106 1.0000
+ 6.000 0.8654 0.02716 0.01854 -0.0585 0.5884 1.0000
+ 6.250 0.8910 0.02650 0.01802 -0.0565 0.5606 1.0000
+ 6.500 0.9104 0.02608 0.01767 -0.0539 0.5234 1.0000
+ 6.750 0.9270 0.02590 0.01748 -0.0511 0.4754 1.0000
+ 7.000 0.9401 0.02612 0.01757 -0.0483 0.4122 1.0000
+ 7.250 0.9485 0.02693 0.01791 -0.0451 0.3202 1.0000
+ 7.500 0.9472 0.02883 0.01909 -0.0417 0.2324 1.0000
+ 7.750 0.9432 0.03108 0.02083 -0.0386 0.1761 1.0000
+ 8.000 0.9425 0.03336 0.02280 -0.0362 0.1413 1.0000
+ 8.250 0.9449 0.03557 0.02483 -0.0343 0.1206 1.0000
+ 8.500 0.9496 0.03769 0.02684 -0.0326 0.1057 1.0000
+ 8.750 0.9578 0.03965 0.02879 -0.0312 0.0943 1.0000
+ 9.000 0.9689 0.04152 0.03072 -0.0299 0.0862 1.0000
+ 9.250 0.9817 0.04332 0.03250 -0.0288 0.0788 1.0000
+ 9.500 1.0031 0.04488 0.03423 -0.0278 0.0722 1.0000
+ 9.750 1.0260 0.04648 0.03585 -0.0271 0.0667 1.0000
+ 10.000 1.0540 0.04828 0.03785 -0.0266 0.0615 1.0000
+ 10.250 1.0829 0.05037 0.04020 -0.0264 0.0578 1.0000
+ 10.500 1.1069 0.05267 0.04255 -0.0263 0.0549 1.0000
+ 10.750 1.1236 0.05550 0.04568 -0.0257 0.0525 1.0000
+ 11.000 1.1309 0.05848 0.04908 -0.0245 0.0507 1.0000
+ 11.250 1.1352 0.06171 0.05266 -0.0235 0.0494 1.0000
+ 11.500 1.1352 0.06522 0.05650 -0.0225 0.0487 1.0000
+ 11.750 1.1305 0.06897 0.06057 -0.0216 0.0482 1.0000
+ 12.000 1.1213 0.07308 0.06499 -0.0212 0.0478 1.0000
+ 12.250 1.1079 0.07762 0.06981 -0.0213 0.0476 1.0000
+ 12.500 1.0904 0.08277 0.07525 -0.0222 0.0476 1.0000
+ 12.750 1.0688 0.08870 0.08144 -0.0240 0.0477 1.0000
+ 13.000 1.0430 0.09570 0.08869 -0.0272 0.0480 1.0000
+ 13.250 1.0136 0.10409 0.09729 -0.0319 0.0485 1.0000
+ 13.500 0.9815 0.11425 0.10762 -0.0384 0.0493 1.0000
diff --git a/Airfoils/Polars/NACA_63_412_polar_Re_500000.txt b/Airfoils/Polars/NACA_63_412_polar_Re_500000.txt
new file mode 100644
index 0000000..47f6101
--- /dev/null
+++ b/Airfoils/Polars/NACA_63_412_polar_Re_500000.txt
@@ -0,0 +1,141 @@
+
+ XFOIL Version 6.96
+
+ Calculated polar for: NACA 63-412 AIRFOIL
+
+ 1 1 Reynolds number fixed Mach number fixed
+
+ xtrf = 1.000 (top) 1.000 (bottom)
+ Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
+
+ alpha CL CD CDp CM Top_Xtr Bot_Xtr
+ ------ -------- --------- --------- -------- -------- --------
+ -13.500 -0.8163 0.05641 0.05349 -0.0612 1.0000 0.0093
+ -13.250 -0.8408 0.04909 0.04595 -0.0664 1.0000 0.0093
+ -13.000 -0.8587 0.04361 0.04028 -0.0698 1.0000 0.0093
+ -12.750 -0.8743 0.03910 0.03556 -0.0717 1.0000 0.0094
+ -12.500 -0.8787 0.03634 0.03264 -0.0723 1.0000 0.0095
+ -12.250 -0.8858 0.03366 0.02979 -0.0720 1.0000 0.0097
+ -12.000 -0.8929 0.03159 0.02756 -0.0704 1.0000 0.0098
+ -11.750 -0.8958 0.02978 0.02557 -0.0684 1.0000 0.0100
+ -11.500 -0.8925 0.02801 0.02360 -0.0667 1.0000 0.0102
+ -11.250 -0.8851 0.02651 0.02191 -0.0651 1.0000 0.0104
+ -11.000 -0.8750 0.02522 0.02044 -0.0635 1.0000 0.0106
+ -10.750 -0.8602 0.02381 0.01890 -0.0629 0.9986 0.0109
+ -10.500 -0.8323 0.02274 0.01772 -0.0644 0.9920 0.0112
+ -10.250 -0.8024 0.02190 0.01680 -0.0661 0.9863 0.0116
+ -10.000 -0.7721 0.02119 0.01601 -0.0676 0.9802 0.0122
+ -9.750 -0.7407 0.02030 0.01500 -0.0694 0.9742 0.0128
+ -9.500 -0.7108 0.01935 0.01389 -0.0707 0.9661 0.0133
+ -9.250 -0.6795 0.01848 0.01284 -0.0722 0.9582 0.0138
+ -9.000 -0.6490 0.01743 0.01170 -0.0738 0.9492 0.0144
+ -8.750 -0.6192 0.01681 0.01102 -0.0748 0.9386 0.0150
+ -8.500 -0.5904 0.01625 0.01038 -0.0756 0.9280 0.0156
+ -8.250 -0.5631 0.01576 0.00979 -0.0759 0.9168 0.0163
+ -8.000 -0.5374 0.01528 0.00919 -0.0758 0.9050 0.0171
+ -7.750 -0.5121 0.01485 0.00864 -0.0756 0.8931 0.0177
+ -7.500 -0.4883 0.01417 0.00788 -0.0752 0.8813 0.0184
+ -7.250 -0.4632 0.01373 0.00736 -0.0750 0.8706 0.0192
+ -7.000 -0.4376 0.01337 0.00693 -0.0748 0.8602 0.0201
+ -6.750 -0.4116 0.01303 0.00652 -0.0747 0.8514 0.0211
+ -6.500 -0.3856 0.01269 0.00609 -0.0745 0.8424 0.0221
+ -6.250 -0.3593 0.01232 0.00563 -0.0744 0.8343 0.0229
+ -6.000 -0.3334 0.01187 0.00513 -0.0743 0.8261 0.0242
+ -5.750 -0.3066 0.01157 0.00479 -0.0743 0.8180 0.0257
+ -5.500 -0.2796 0.01132 0.00446 -0.0743 0.8101 0.0274
+ -5.250 -0.2523 0.01108 0.00415 -0.0743 0.8026 0.0288
+ -5.000 -0.2253 0.01074 0.00377 -0.0743 0.7952 0.0313
+ -4.750 -0.1977 0.01050 0.00350 -0.0744 0.7878 0.0341
+ -4.500 -0.1701 0.01032 0.00325 -0.0744 0.7803 0.0371
+ -4.250 -0.1425 0.01007 0.00300 -0.0745 0.7732 0.0432
+ -4.000 -0.1147 0.00986 0.00278 -0.0746 0.7661 0.0511
+ -3.750 -0.0869 0.00964 0.00257 -0.0747 0.7592 0.0654
+ -3.250 -0.0316 0.00895 0.00215 -0.0752 0.7448 0.1510
+ -3.000 -0.0041 0.00844 0.00193 -0.0756 0.7375 0.2369
+ -2.750 0.0233 0.00790 0.00173 -0.0761 0.7307 0.3439
+ -2.500 0.0514 0.00757 0.00164 -0.0764 0.7232 0.4223
+ -2.250 0.0795 0.00744 0.00157 -0.0765 0.7164 0.4656
+ -2.000 0.1078 0.00729 0.00156 -0.0767 0.7090 0.5146
+ -1.750 0.1359 0.00723 0.00155 -0.0767 0.7020 0.5494
+ -1.500 0.1644 0.00719 0.00155 -0.0769 0.6948 0.5735
+ -1.250 0.1927 0.00717 0.00153 -0.0769 0.6876 0.5910
+ -1.000 0.2213 0.00717 0.00152 -0.0770 0.6806 0.6023
+ -0.750 0.2497 0.00717 0.00151 -0.0771 0.6732 0.6127
+ -0.500 0.2781 0.00718 0.00151 -0.0772 0.6663 0.6220
+ -0.250 0.3066 0.00720 0.00152 -0.0773 0.6589 0.6313
+ 0.000 0.3350 0.00722 0.00154 -0.0774 0.6519 0.6398
+ 0.250 0.3634 0.00725 0.00157 -0.0775 0.6444 0.6489
+ 0.500 0.3917 0.00729 0.00160 -0.0775 0.6374 0.6579
+ 0.750 0.4201 0.00732 0.00165 -0.0776 0.6296 0.6664
+ 1.000 0.4484 0.00738 0.00169 -0.0777 0.6224 0.6755
+ 1.250 0.4766 0.00742 0.00175 -0.0778 0.6143 0.6837
+ 1.500 0.5048 0.00748 0.00181 -0.0778 0.6065 0.6923
+ 1.750 0.5329 0.00753 0.00188 -0.0778 0.5979 0.7000
+ 2.000 0.5611 0.00760 0.00195 -0.0779 0.5896 0.7083
+ 2.250 0.5888 0.00767 0.00204 -0.0779 0.5802 0.7158
+ 2.500 0.6167 0.00775 0.00212 -0.0779 0.5689 0.7240
+ 2.750 0.6441 0.00784 0.00222 -0.0778 0.5555 0.7312
+ 3.000 0.6714 0.00796 0.00233 -0.0777 0.5404 0.7393
+ 3.250 0.6981 0.00809 0.00245 -0.0775 0.5221 0.7468
+ 3.500 0.7246 0.00827 0.00257 -0.0772 0.5013 0.7550
+ 3.750 0.7499 0.00851 0.00273 -0.0768 0.4658 0.7627
+ 4.000 0.7744 0.00887 0.00292 -0.0763 0.4192 0.7711
+ 4.250 0.7989 0.00922 0.00316 -0.0758 0.3816 0.7791
+ 4.500 0.8219 0.00974 0.00347 -0.0752 0.3270 0.7877
+ 4.750 0.8428 0.01047 0.00391 -0.0743 0.2584 0.7964
+ 5.000 0.8639 0.01120 0.00437 -0.0734 0.1974 0.8053
+ 5.250 0.8843 0.01198 0.00487 -0.0724 0.1377 0.8149
+ 5.500 0.9050 0.01266 0.00536 -0.0715 0.0933 0.8244
+ 5.750 0.9268 0.01324 0.00581 -0.0707 0.0653 0.8347
+ 6.000 0.9495 0.01367 0.00622 -0.0700 0.0504 0.8455
+ 6.250 0.9720 0.01407 0.00662 -0.0692 0.0411 0.8567
+ 6.500 0.9942 0.01448 0.00705 -0.0683 0.0345 0.8688
+ 6.750 1.0164 0.01483 0.00744 -0.0674 0.0307 0.8821
+ 7.000 1.0368 0.01525 0.00789 -0.0662 0.0268 0.8975
+ 7.250 1.0568 0.01555 0.00828 -0.0649 0.0246 0.9178
+ 7.500 1.0781 0.01590 0.00871 -0.0638 0.0222 0.9809
+ 7.750 1.1004 0.01643 0.00926 -0.0632 0.0204 1.0000
+ 8.000 1.1226 0.01692 0.00979 -0.0626 0.0190 1.0000
+ 8.250 1.1438 0.01746 0.01036 -0.0618 0.0178 1.0000
+ 8.500 1.1635 0.01808 0.01100 -0.0608 0.0167 1.0000
+ 8.750 1.1818 0.01877 0.01173 -0.0596 0.0159 1.0000
+ 9.000 1.2002 0.01939 0.01242 -0.0584 0.0154 1.0000
+ 9.250 1.2164 0.02005 0.01313 -0.0568 0.0149 1.0000
+ 9.500 1.2308 0.02073 0.01387 -0.0550 0.0144 1.0000
+ 9.750 1.2442 0.02147 0.01466 -0.0531 0.0140 1.0000
+ 10.000 1.2570 0.02226 0.01551 -0.0512 0.0136 1.0000
+ 10.250 1.2689 0.02312 0.01642 -0.0494 0.0132 1.0000
+ 10.500 1.2790 0.02415 0.01750 -0.0475 0.0128 1.0000
+ 10.750 1.2858 0.02545 0.01886 -0.0454 0.0124 1.0000
+ 11.000 1.2965 0.02653 0.02003 -0.0439 0.0123 1.0000
+ 11.250 1.3063 0.02774 0.02133 -0.0424 0.0121 1.0000
+ 11.500 1.3153 0.02905 0.02273 -0.0409 0.0119 1.0000
+ 11.750 1.3236 0.03048 0.02427 -0.0396 0.0117 1.0000
+ 12.000 1.3317 0.03199 0.02587 -0.0383 0.0115 1.0000
+ 12.250 1.3390 0.03361 0.02759 -0.0372 0.0114 1.0000
+ 12.500 1.3462 0.03530 0.02938 -0.0362 0.0112 1.0000
+ 12.750 1.3526 0.03711 0.03129 -0.0352 0.0110 1.0000
+ 13.000 1.3591 0.03897 0.03324 -0.0345 0.0108 1.0000
+ 13.250 1.3651 0.04091 0.03528 -0.0338 0.0106 1.0000
+ 13.500 1.3709 0.04292 0.03737 -0.0333 0.0104 1.0000
+ 13.750 1.3753 0.04514 0.03969 -0.0329 0.0102 1.0000
+ 14.000 1.3791 0.04747 0.04213 -0.0327 0.0101 1.0000
+ 14.250 1.3811 0.05009 0.04484 -0.0326 0.0099 1.0000
+ 14.500 1.3816 0.05298 0.04783 -0.0326 0.0098 1.0000
+ 14.750 1.3804 0.05617 0.05114 -0.0328 0.0097 1.0000
+ 15.000 1.3768 0.05980 0.05489 -0.0331 0.0095 1.0000
+ 15.250 1.3725 0.06367 0.05891 -0.0338 0.0094 1.0000
+ 15.500 1.3704 0.06740 0.06278 -0.0347 0.0094 1.0000
+ 15.750 1.3669 0.07146 0.06699 -0.0358 0.0093 1.0000
+ 16.000 1.3626 0.07582 0.07151 -0.0373 0.0092 1.0000
+ 16.250 1.3568 0.08057 0.07643 -0.0391 0.0092 1.0000
+ 16.500 1.3492 0.08575 0.08177 -0.0412 0.0091 1.0000
+ 16.750 1.3406 0.09132 0.08751 -0.0437 0.0090 1.0000
+ 17.000 1.3309 0.09731 0.09367 -0.0466 0.0090 1.0000
+ 17.250 1.3192 0.10390 0.10044 -0.0500 0.0089 1.0000
+ 17.500 1.3064 0.11094 0.10764 -0.0538 0.0088 1.0000
+ 17.750 1.2922 0.11855 0.11543 -0.0582 0.0088 1.0000
+ 18.000 1.2762 0.12686 0.12391 -0.0632 0.0088 1.0000
+ 18.250 1.2584 0.13587 0.13309 -0.0688 0.0088 1.0000
+ 18.500 1.2377 0.14599 0.14340 -0.0752 0.0087 1.0000
+ 18.750 1.2128 0.15775 0.15534 -0.0828 0.0088 1.0000
diff --git a/Airfoils/supersonic_tail.txt b/Airfoils/supersonic_tail.txt
new file mode 100644
index 0000000..227379a
--- /dev/null
+++ b/Airfoils/supersonic_tail.txt
@@ -0,0 +1,156 @@
+NACA 65-204 no camber, 151 panels
+ 76. 76.
+
+ 0.00000 -0.00022
+ 0.00047 0.00078
+ 0.00213 0.00194
+ 0.00592 0.00341
+ 0.01468 0.00508
+ 0.02821 0.00670
+ 0.04141 0.00802
+ 0.05535 0.00923
+ 0.06944 0.01030
+ 0.08363 0.01128
+ 0.09782 0.01217
+ 0.11209 0.01299
+ 0.12637 0.01373
+ 0.14066 0.01442
+ 0.15493 0.01505
+ 0.16920 0.01564
+ 0.18347 0.01618
+ 0.19775 0.01668
+ 0.21204 0.01714
+ 0.22633 0.01757
+ 0.24062 0.01795
+ 0.25491 0.01830
+ 0.26920 0.01861
+ 0.28349 0.01889
+ 0.29778 0.01914
+ 0.31208 0.01936
+ 0.32638 0.01954
+ 0.34067 0.01969
+ 0.35497 0.01982
+ 0.36927 0.01991
+ 0.38356 0.01997
+ 0.39786 0.01999
+ 0.41215 0.01998
+ 0.42644 0.01993
+ 0.44074 0.01985
+ 0.45503 0.01973
+ 0.46932 0.01957
+ 0.48361 0.01937
+ 0.49790 0.01913
+ 0.51219 0.01885
+ 0.52648 0.01853
+ 0.54077 0.01817
+ 0.55506 0.01778
+ 0.56936 0.01736
+ 0.58365 0.01692
+ 0.59794 0.01645
+ 0.61223 0.01595
+ 0.62653 0.01544
+ 0.64082 0.01490
+ 0.65511 0.01435
+ 0.66940 0.01377
+ 0.68370 0.01318
+ 0.69799 0.01258
+ 0.71228 0.01195
+ 0.72658 0.01132
+ 0.74087 0.01067
+ 0.75517 0.01002
+ 0.76945 0.00936
+ 0.78373 0.00868
+ 0.79801 0.00801
+ 0.81230 0.00733
+ 0.82655 0.00665
+ 0.84080 0.00597
+ 0.85507 0.00530
+ 0.86925 0.00465
+ 0.88334 0.00400
+ 0.89747 0.00337
+ 0.91160 0.00275
+ 0.92534 0.00217
+ 0.93891 0.00163
+ 0.95284 0.00114
+ 0.96581 0.00074
+ 0.97701 0.00045
+ 0.98662 0.00023
+ 0.99499 0.00007
+ 1.00000 -0.00002
+
+ 0.00000 -0.00022
+ 0.00058 -0.00119
+ 0.00226 -0.00227
+ 0.00595 -0.00363
+ 0.01433 -0.00523
+ 0.02789 -0.00684
+ 0.04133 -0.00817
+ 0.05529 -0.00940
+ 0.06945 -0.01046
+ 0.08367 -0.01143
+ 0.09787 -0.01232
+ 0.11213 -0.01313
+ 0.12642 -0.01387
+ 0.14071 -0.01454
+ 0.15499 -0.01517
+ 0.16927 -0.01575
+ 0.18354 -0.01629
+ 0.19782 -0.01678
+ 0.21211 -0.01724
+ 0.22640 -0.01765
+ 0.24069 -0.01803
+ 0.25498 -0.01837
+ 0.26927 -0.01868
+ 0.28356 -0.01895
+ 0.29786 -0.01920
+ 0.31215 -0.01941
+ 0.32645 -0.01958
+ 0.34074 -0.01973
+ 0.35504 -0.01985
+ 0.36934 -0.01994
+ 0.38363 -0.01999
+ 0.39793 -0.02001
+ 0.41222 -0.02000
+ 0.42652 -0.01995
+ 0.44081 -0.01986
+ 0.45510 -0.01974
+ 0.46940 -0.01958
+ 0.48369 -0.01938
+ 0.49798 -0.01914
+ 0.51227 -0.01886
+ 0.52655 -0.01854
+ 0.54085 -0.01818
+ 0.55514 -0.01780
+ 0.56943 -0.01738
+ 0.58372 -0.01694
+ 0.59802 -0.01647
+ 0.61231 -0.01598
+ 0.62660 -0.01546
+ 0.64089 -0.01493
+ 0.65519 -0.01437
+ 0.66948 -0.01380
+ 0.68377 -0.01321
+ 0.69806 -0.01260
+ 0.71235 -0.01198
+ 0.72665 -0.01135
+ 0.74095 -0.01070
+ 0.75524 -0.01005
+ 0.76953 -0.00939
+ 0.78381 -0.00872
+ 0.79809 -0.00804
+ 0.81237 -0.00736
+ 0.82663 -0.00668
+ 0.84087 -0.00601
+ 0.85514 -0.00534
+ 0.86932 -0.00468
+ 0.88341 -0.00403
+ 0.89754 -0.00340
+ 0.91167 -0.00278
+ 0.92539 -0.00220
+ 0.93897 -0.00167
+ 0.95290 -0.00117
+ 0.96585 -0.00078
+ 0.97702 -0.00048
+ 0.98663 -0.00027
+ 0.99499 -0.00010
+ 1.00000 -0.00002
\ No newline at end of file
diff --git a/tut_Aerodynamic_Polars_X57Mod3.py b/tut_Aerodynamic_Polars_X57Mod3.py
new file mode 100644
index 0000000..be4a8f1
--- /dev/null
+++ b/tut_Aerodynamic_Polars_X57Mod3.py
@@ -0,0 +1,752 @@
+# tut_Aerodynamic_Polars_X57.py
+#
+# Created: Dec 2021, M. Clarke
+
+""" setup file for the Whisper Drone Vehicle Polar Analysis
+"""
+
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+# SUAVE Imports
+import SUAVE
+#assert SUAVE.__version__=='2.5.0', 'These tutorials only work with the SUAVE 2.5.0 release'
+
+from SUAVE.Core import Units, Data
+from SUAVE.Plots.Performance.Mission_Plots import *
+from SUAVE.Plots.Geometry import *
+from SUAVE.Methods.Aerodynamics.Common.Fidelity_Zero.Lift import VLM
+from SUAVE.Components.Energy.Networks.Lift_Cruise import Lift_Cruise
+from SUAVE.Methods.Propulsion import propeller_design
+from SUAVE.Methods.Power.Battery.Sizing import initialize_from_mass
+from SUAVE.Methods.Propulsion.electric_motor_sizing import size_optimal_motor
+from SUAVE.Methods.Geometry.Two_Dimensional.Planform import segment_properties
+from copy import deepcopy
+import matplotlib.cm as cm
+
+# ----------------------------------------------------------------------
+# Generic Function Call
+# ----------------------------------------------------------------------
+def main():
+ # call X57 - Mod 3 vehicle setup function
+ vehicle = vehicle_setup()
+
+ # define wing and control surface
+ wing_tag = 'main_wing'
+ control_surface_tag = 'flap' # can leave empty ie. control_surface_tag = ''
+ control_surface_deflection_angles = np.linspace(-10,20,7)*Units.degrees # can be zero i.e. control_surface_deflection_angles = np.array([0])
+ airspeed = 150.* Units['mph']
+ altitude = 2500.0*Units.feet
+ alpha_range = np.linspace(-2,10,13)*Units.degrees
+
+ # call polar analysis function
+ setup_vehicle_polar_analyses(vehicle,wing_tag,control_surface_tag,control_surface_deflection_angles,airspeed,altitude ,alpha_range)
+
+ return
+
+# -----------------------------------------
+# Setup for Aircraft Polars
+# -----------------------------------------
+def setup_vehicle_polar_analyses(vehicle,wing_tag,control_surface_tag,deflection_angles,V,Alt,alpha_range):
+
+ MAC = vehicle.wings['main_wing'].chords.mean_aerodynamic
+ S_ref = vehicle.reference_area
+ num_def = len(deflection_angles)
+
+ #------------------------------------------------------------------------
+ # setup figures
+ #------------------------------------------------------------------------
+
+ plt.rcParams['axes.linewidth'] = 2.
+ plt.rcParams["font.family"] = "Times New Roman"
+ parameters = {'axes.labelsize': 24,
+ 'legend.fontsize': 20,
+ 'xtick.labelsize': 24,
+ 'ytick.labelsize': 24,
+ 'axes.titlesize': 28}
+ plt.rcParams.update(parameters)
+
+ header_text = ' : $S_{ref}$ = ' + str(round(S_ref,2)) + ' MAC =' + str(round(MAC,2))
+
+ fig1 = plt.figure('C_L_vs_AoA')
+ fig1.set_size_inches(8,6)
+ axes1 = fig1.add_subplot(1,1,1)
+ axes1.set_title('$C_L$ vs AoA' + header_text )
+ axes1.set_ylabel('Coefficient of Lift')
+ axes1.set_xlabel('Angle of Attack (degrees)')
+
+ fig2 = plt.figure('C_Di_vs_AoA')
+ fig2.set_size_inches(8,6)
+ axes2 = fig2.add_subplot(1,1,1)
+ axes2.set_title('$C_{Di}$ vs. AoA'+ header_text)
+ axes2.set_ylabel('Coefficient of Drag')
+ axes2.set_xlabel('Angle of Attack (degrees)')
+
+ fig3 = plt.figure('C_Di_vs_C_L2')
+ fig3.set_size_inches(8,6)
+ axes3 = fig3.add_subplot(1,1,1)
+ axes3.set_title('$C_{Di}$ vs. $C_L^{2}$'+ header_text)
+ axes3.set_ylabel('Coefficient of Drag')
+ axes3.set_xlabel('Lineraized Coefficient of Lift ($CL^2$)')
+
+ fig4 = plt.figure('C_M_vs_AoA')
+ fig4.set_size_inches(8,6)
+ axes4 = fig4.add_subplot(1,1,1)
+ axes4.set_title('$C_M$ vs. AoA'+ header_text)
+ axes4.set_ylabel('Coefficient of Moment')
+ axes4.set_xlabel('Angle of Attack (degrees)')
+
+ linecolor_1 = cm.jet(np.linspace(0, 1,num_def))
+ linestyle_1 = ['-']*num_def
+ marker_1 = ['o']*num_def
+
+ #------------------------------------------------------------------------
+ # setup flight conditions
+ #------------------------------------------------------------------------
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude=Alt)
+ a = atmo_data.speed_of_sound[0]
+ Mach = V/a
+ AoA_range = np.atleast_2d(alpha_range).T
+
+ # check if control surface is defined
+ CS_flag = False
+ for wing in vehicle.wings:
+ if 'control_surfaces' in wing:
+ if control_surface_tag in wing.control_surfaces:
+ CS_flag = True
+
+ for i in range (num_def):
+ # change control surface deflection
+ if CS_flag:
+ vehicle.wings[wing_tag].control_surfaces[control_surface_tag].deflection = deflection_angles[i]
+
+ # compute polar
+ results = compute_polars(vehicle,AoA_range,Mach,Alt)
+ if CS_flag:
+ line_label = wing_tag + ',' + control_surface_tag + ' ' +\
+ str(round( deflection_angles[i]/Units.degrees,3)) + '$\degree$ defl.'
+ else:
+ line_label = ''
+
+ # plot
+ plot_polars(axes1,axes2,axes3,axes4,AoA_range,Mach,results,linestyle_1[i], linecolor_1[i],marker_1[i],line_label)
+
+ # append legend
+ axes1.legend(loc='upper left', prop={'size': 14})
+ axes2.legend(loc='upper left', prop={'size': 14})
+ axes3.legend(loc='upper left', prop={'size': 14})
+ axes4.legend(loc='upper right', prop={'size': 14})
+
+ # format figure
+ fig1.tight_layout()
+ fig2.tight_layout()
+ fig3.tight_layout()
+ fig4.tight_layout()
+ return
+
+
+# -----------------------------------------
+# Compute Aircraft Polars
+# -----------------------------------------
+def compute_polars(vehicle,AoA_range,Mach,Alt):
+
+ MAC = vehicle.wings['main_wing'].chords.mean_aerodynamic
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude=Alt)
+ P = atmo_data.pressure[0]
+ T = atmo_data.temperature[0]
+ rho = atmo_data.density[0]
+ a = atmo_data.speed_of_sound[0]
+ mu = atmo_data.dynamic_viscosity[0]
+ V = a*Mach
+ re = (V*rho*MAC)/mu
+
+ n_aoa = len(AoA_range)
+ vortices = 4
+
+ state = SUAVE.Analyses.Mission.Segments.Conditions.State()
+ state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
+ state.conditions.freestream.mach_number = Mach * np.ones_like(AoA_range)
+ state.conditions.freestream.density = rho * np.ones_like(AoA_range)
+ state.conditions.freestream.dynamic_viscosity = mu * np.ones_like(AoA_range)
+ state.conditions.freestream.temperature = T * np.ones_like(AoA_range)
+ state.conditions.freestream.pressure = P * np.ones_like(AoA_range)
+ state.conditions.freestream.reynolds_number = re * np.ones_like(AoA_range)
+ state.conditions.freestream.velocity = V * np.ones_like(AoA_range)
+ state.conditions.aerodynamics.angle_of_attack = AoA_range
+
+ # -----------------------------------------------------------------
+ # VLM No Surrogate (Inviscid)
+ # -----------------------------------------------------------------
+ settings = Data()
+ settings.use_surrogate = False
+ settings.number_spanwise_vortices = vortices **2
+ settings.number_chordwise_vortices = vortices
+ settings.propeller_wake_model = False
+ settings.initial_timestep_offset = 0
+ settings.wake_development_time = 0.05
+ settings.use_bemt_wake_model = False
+ settings.number_of_wake_timesteps = 30
+ settings.leading_edge_suction_multiplier = 1.0
+ settings.spanwise_cosine_spacing = True
+ settings.model_fuselage = False
+ settings.model_nacelle = False
+ settings.wing_spanwise_vortices = None
+ settings.wing_chordwise_vortices = None
+ settings.fuselage_spanwise_vortices = None
+ settings.discretize_control_surfaces = True
+ settings.fuselage_chordwise_vortices = None
+ settings.floating_point_precision = np.float32
+ settings.use_VORLAX_matrix_calculation = False
+ settings.use_surrogate = True
+ results = VLM(state.conditions,settings,vehicle)
+
+
+ # pack results
+ Aero_Results = Data()
+ Aero_Results.CL_Inv = results.CL
+ Aero_Results.CDi_Inv = results.CDi
+ Aero_Results.CM_Inv = results.CM
+
+
+ # plot aircraft
+ plot_vehicle_vlm_panelization(vehicle, elevation_angle = 30,azimuthal_angle = 135, axis_limits = 6 ,plot_control_points = False,save_filename = 'X47_M3')
+
+ return Aero_Results
+
+
+# -----------------------------------------
+# Plot Polars Aircraft Polars
+# -----------------------------------------
+def plot_polars(axes1,axes2,axes3,axes4,AoA_range,Mach,results,linestyle_1,
+ linecolor_1,marker_1,line_label):
+
+ CL_Inv = results.CL_Inv
+ CDi_Inv = results.CDi_Inv
+ CM_Inv = results.CM_Inv
+
+ axes1.plot(AoA_range/Units.degrees,CL_Inv,linestyle = linestyle_1,linewidth = 2, markersize = 10, color = linecolor_1, marker = marker_1,label = line_label)
+
+ axes2.plot(AoA_range/Units.degrees,CDi_Inv,linestyle = linestyle_1,linewidth = 2, markersize = 10, color = linecolor_1, marker = marker_1,label = line_label)
+
+ axes3.plot(CL_Inv**2,CDi_Inv,linestyle = linestyle_1,linewidth = 2, markersize = 10, color = linecolor_1, marker = marker_1,label = line_label)
+
+ axes4.plot(AoA_range/Units.degrees,CM_Inv,linestyle = linestyle_1,linewidth = 2, markersize = 10, color = linecolor_1, marker = marker_1,label = line_label)
+
+ return
+
+def vehicle_setup():
+
+ # ------------------------------------------------------------------
+ # Initialize the Vehicle
+ # ------------------------------------------------------------------
+
+ vehicle = SUAVE.Vehicle()
+ vehicle.tag = 'X57_Modification_3'
+
+ # ------------------------------------------------------------------
+ # Vehicle-level Properties
+ # ------------------------------------------------------------------
+
+ # mass properties
+ vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
+ vehicle.mass_properties.takeoff = 2550. * Units.pounds
+ vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
+ vehicle.mass_properties.cargo = 0.
+ vehicle.mass_properties.center_of_gravity = [[ 3.35, 0. , 0.34 ]]
+
+ # envelope properties
+ vehicle.envelope.ultimate_load = 5.7
+ vehicle.envelope.limit_load = 3.8
+
+ # basic parameters
+ vehicle.reference_area = 66.66 *Units.feet**2
+ vehicle.passengers = 4
+
+ # ------------------------------------------------------------------
+ # Main Wing
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Main_Wing()
+ wing.tag = 'main_wing'
+
+ wing.sweeps.quarter_chord = 0.0 * Units.deg
+ wing.sweeps.leading_edge = 0.0 * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 66.66 *Units.feet**2
+ wing.spans.projected = 31.633 * Units.feet
+ wing.chords.root = 0.7 * Units.meter
+ wing.chords.tip = 0.6 * Units.meter
+ wing.chords.mean_aerodynamic = 0.649224 # 2.13 * Units.feet
+ wing.taper = wing.chords.tip / wing.chords.root
+ wing.aspect_ratio = wing.spans.projected ** 2. / wing.areas.reference
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[3.05, 0., 0.784]]
+ wing.aerodynamic_center = [0.558, 0., 0.784]
+ wing.vertical = False
+ wing.symmetric = True
+ wing.high_lift = True
+ airfoil = SUAVE.Components.Airfoils.Airfoil()
+ #airfoil.coordinate_file = 'NACA_63_412.txt'
+ wing.append_airfoil(airfoil)
+ wing.dynamic_pressure_ratio = 1.0
+
+ # Wing Segments
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'root'
+ segment.percent_span_location = 0.0
+ segment.twist = 3. * Units.degrees
+ segment.root_chord_percent = 1.
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ #segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'tip'
+ segment.percent_span_location = 1.
+ segment.twist = 3. * Units.degrees
+ segment.root_chord_percent = wing.taper
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ #segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+
+ flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
+ flap.tag = 'flap'
+ flap.span_fraction_start = 0.15
+ flap.span_fraction_end = 0.8
+ flap.deflection = 20.0 * Units.degrees
+ flap.chord_fraction = 0.20
+ wing.append_control_surface(flap)
+
+ segment_properties(wing)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Horizontal Stabilizer
+ # ------------------------------------------------------------------
+ wing = SUAVE.Components.Wings.Wing()
+ wing.tag = 'horizontal_stabilizer'
+ wing.sweeps.quarter_chord = 0.0 * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 2.540
+ wing.spans.projected = 3.3 * Units.meter
+ wing.sweeps.quarter_chord = 0 * Units.deg
+ wing.chords.root = 0.769 * Units.meter
+ wing.chords.tip = 0.769 * Units.meter
+ wing.chords.mean_aerodynamic = 0.769 * Units.meter
+ wing.taper = 1.
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 1.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[7.7, 0., 0.25]]
+ wing.aerodynamic_center = [7.8, 0., 0.25]
+ wing.vertical = False
+ wing.winglet_fraction = 0.0
+ wing.symmetric = True
+ wing.high_lift = False
+ 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.sweeps.quarter_chord = 25. * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 2.258 * Units['meters**2']
+ wing.spans.projected = 1.854 * Units.meter
+ wing.chords.root = 1.6764 * Units.meter
+ wing.chords.tip = 0.6858 * Units.meter
+ wing.chords.mean_aerodynamic = 1.21 * Units.meter
+ wing.taper = wing.chords.tip/wing.chords.root
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[6.75 ,0, 0.]]
+ wing.aerodynamic_center = [0.508 ,0,0]
+ wing.vertical = True
+ wing.symmetric = False
+ wing.t_tail = False
+ wing.winglet_fraction = 0.0
+ wing.dynamic_pressure_ratio = 1.0
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+ # ------------------------------------------------------------------
+ # Fuselage
+ # ------------------------------------------------------------------
+ fuselage = SUAVE.Components.Fuselages.Fuselage()
+ fuselage.tag = 'fuselage'
+ fuselage.seats_abreast = 2.
+ fuselage.fineness.nose = 1.6
+ fuselage.fineness.tail = 2.
+ fuselage.lengths.nose = 60. * Units.inches
+ fuselage.lengths.tail = 161. * Units.inches
+ fuselage.lengths.cabin = 105. * Units.inches
+ fuselage.lengths.total = 332.2 * Units.inches
+ fuselage.lengths.fore_space = 0.
+ fuselage.lengths.aft_space = 0.
+ fuselage.width = 42. * Units.inches
+ fuselage.heights.maximum = 62. * Units.inches
+ fuselage.heights.at_quarter_length = 62. * Units.inches
+ fuselage.heights.at_three_quarters_length = 62. * Units.inches
+ fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
+ fuselage.areas.side_projected = 8000. * Units.inches ** 2.
+ fuselage.areas.wetted = 30000. * Units.inches ** 2.
+ fuselage.areas.front_projected = 42. * 62. * Units.inches ** 2.
+ fuselage.effective_diameter = 50. * Units.inches
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_0'
+ segment.percent_x_location = 0
+ segment.percent_z_location = 0
+ segment.height = 0.01
+ segment.width = 0.01
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_1'
+ segment.percent_x_location = 0.007279116466
+ segment.percent_z_location = 0.002502014453
+ segment.height = 0.1669064748
+ segment.width = 0.2780205877
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_2'
+ segment.percent_x_location = 0.01941097724
+ segment.percent_z_location = 0.001216095397
+ segment.height = 0.3129496403
+ segment.width = 0.4365777215
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_3'
+ segment.percent_x_location = 0.06308567604
+ segment.percent_z_location = 0.007395489231
+ segment.height = 0.5841726619
+ segment.width = 0.6735119903
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_4'
+ segment.percent_x_location = 0.1653761217
+ segment.percent_z_location = 0.02891281352
+ segment.height = 1.064028777
+ segment.width = 1.067200529
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_5'
+ segment.percent_x_location = 0.2426372155
+ segment.percent_z_location = 0.04214148761
+ segment.height = 1.293766653
+ segment.width = 1.183058255
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_6'
+ segment.percent_x_location = 0.2960174029
+ segment.percent_z_location = 0.04705241831
+ segment.height = 1.377026712
+ segment.width = 1.181540054
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_7'
+ segment.percent_x_location = 0.3809404284
+ segment.percent_z_location = 0.05313580461
+ segment.height = 1.439568345
+ segment.width = 1.178218989
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_8'
+ segment.percent_x_location = 0.5046854083
+ segment.percent_z_location = 0.04655492473
+ segment.height = 1.29352518
+ segment.width = 1.054390707
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_9'
+ segment.percent_x_location = 0.6454149933
+ segment.percent_z_location = 0.03741966266
+ segment.height = 0.8971223022
+ segment.width = 0.8501926505
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_10'
+ segment.percent_x_location = 0.985107095
+ segment.percent_z_location = 0.04540283436
+ segment.height = 0.2920863309
+ segment.width = 0.2012565415
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_11'
+ segment.percent_x_location = 1
+ segment.percent_z_location = 0.04787575562
+ segment.height = 0.1251798561
+ segment.width = 0.1206021048
+ fuselage.Segments.append(segment)
+
+ # add to vehicle
+ vehicle.append_component(fuselage)
+
+ #---------------------------------------------------------------------------------------------
+ # DEFINE NETWORK
+ #---------------------------------------------------------------------------------------------
+ # Component 1
+ net = Lift_Cruise()
+ net.number_of_propeller_engines = 2
+ net.propeller_thrust_angle = 0. * Units.degrees
+ net.propeller_nacelle_diameter = 1.166 * Units.feet
+ net.propeller_engine_length = 3 * Units.feet
+
+ net.number_of_rotor_engines = 12
+ net.rotor_thrust_angle = 0. * Units.degrees
+ net.rotor_nacelle_diameter = 0.5 * Units.feet
+ net.rotor_engine_length = 1 * Units.feet
+
+ net.voltage = 400.
+
+ # Component 2 Electronic Speed Controller --------------------------------------------------------
+ rotor_esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
+ rotor_esc.efficiency = 0.95
+ net.rotor_esc = rotor_esc
+
+ propeller_esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
+ propeller_esc.efficiency = 0.95
+ net.propeller_esc = propeller_esc
+
+ # Component 3 the Propeller -------------------------------------------------------------
+ # Cruise Propeller
+ prop_cr = SUAVE.Components.Energy.Converters.Propeller()
+ prop_cr.tag = 'cruise_propeller'
+ prop_cr.number_of_blades = 3
+ prop_cr.freestream_velocity = 173.984 * Units['mph']
+ prop_cr.tip_radius = 1.523/2
+ prop_cr.hub_radius = 0.1
+ prop_cr.design_Cl = 0.75
+ prop_cr.design_tip_mach = 0.6
+ prop_cr.angular_velocity = 2250 * Units.rpm # prop_cr.design_tip_mach*speed_of_sound/prop_cr.tip_radius
+ prop_cr.design_altitude = 8000. * Units.feet
+ prop_cr.design_power = 48100 # 115.
+ prop_cr_origins = [[2.5, 4.97584, 1.01],[2.5, -4.97584, 1.01]]
+ prop_cr_rotations = [-1,1]
+ prop_cr = propeller_design(prop_cr)
+ prop_cr.symmetry = True
+
+
+ # Appending rotors with different origins
+ for ii in range(net.number_of_propeller_engines):
+ cruise_prop = deepcopy(prop_cr)
+ cruise_prop.tag = 'cruise_prop_' + str(ii+1)
+ cruise_prop.rotation = prop_cr_rotations[ii]
+ cruise_prop.origin = [prop_cr_origins[ii]]
+ net.propellers.append(cruise_prop)
+
+ # Design Highlift Propeller
+ prop_hl = SUAVE.Components.Energy.Converters.Rotor()
+ prop_hl.tag = 'high_lift_propeller'
+ prop_hl.number_of_blades = 5
+ prop_hl.freestream_velocity = 63.379 * Units['mph']
+ prop_hl.tip_radius = 0.58/2
+ prop_hl.hub_radius = 0.1
+ prop_hl.design_Cl = 0.75
+ prop_hl.design_tip_mach = 0.6
+ prop_hl.angular_velocity = 2250 * Units.rpm
+ prop_hl.design_altitude = 100. * Units.feet
+ prop_hl.design_power = 1400
+ prop_pitch = 0.6
+ prop_hl_origins = [[2.5, (1.05 + prop_pitch*0), 1.01], [2.5, (1.05 + prop_pitch*1), 1.01],[2.5, (1.05 + prop_pitch*2), 1.01],
+ [2.5, (1.05 + prop_pitch*3), 1.01],[2.5, (1.05 + prop_pitch*4), 1.01],[2.5, (1.05 + prop_pitch*5), 1.01],
+ [2.5,-(1.05 + prop_pitch*0) ,1.01], [2.5,-(1.05 + prop_pitch*1), 1.01],[2.5,-(1.05 + prop_pitch*2), 1.01],
+ [2.5,-(1.05 + prop_pitch*3), 1.01],[2.5 ,-(1.05 + prop_pitch*4), 1.01],[2.5,-(1.05 + prop_pitch*5), 1.01]]
+ prop_hl_rotations = [-1,-1,-1,-1,-1,-1,1,1,1,1,1,1]
+ prop_hl = propeller_design(prop_hl)
+ prop_hl.symmetry = True
+
+ for ii in range(net.number_of_rotor_engines):
+ prop_high_lift = deepcopy(prop_hl)
+ prop_high_lift.tag = 'high_lift_propeller_' + str(ii+1)
+ prop_high_lift.rotation = prop_hl_rotations[ii]
+ prop_high_lift.origin = [prop_hl_origins[ii]]
+ net.lift_rotors.append(prop_high_lift)
+
+ # Component 4 the Battery --------------------------------------------------------------------
+ bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
+ bat.mass_properties.mass = 300. * Units.kg
+ bat.specific_energy = 200. * Units.Wh/Units.kg
+ bat.resistance = 0.006
+ bat.max_voltage = 400.
+
+ initialize_from_mass(bat,bat.mass_properties.mass)
+ net.battery = bat
+ net.voltage = bat.max_voltage
+
+ # Component 5 the Motor --------------------------------------------------------------------
+ # Cruise Propeller motor
+ motor_cr = SUAVE.Components.Energy.Converters.Motor()
+ motor_cr.efficiency = 0.95
+ motor_cr.gearbox_efficiency = 1.
+ motor_cr.nominal_voltage = bat.max_voltage*0.75
+ motor_cr.propeller_radius = prop_cr.tip_radius
+ motor_cr.no_load_current = 0.1
+ motor_cr.origin = prop_cr.origin
+ motor_cr = size_optimal_motor(motor_cr,prop_cr)
+ net.propeller_motor = motor_cr
+
+ # High Lift Propeller motor
+ motor_hl = SUAVE.Components.Energy.Converters.Motor()
+ motor_hl.efficiency = 0.9
+ motor_hl.gearbox_efficiency = 1.
+ motor_hl.nominal_voltage = bat.max_voltage
+ motor_hl.propeller_radius = prop_hl.tip_radius
+ motor_hl.no_load_current = 0.1
+ motor_hl.origin = prop_hl.origin
+ motor_hl = size_optimal_motor(motor_hl,prop_hl)
+ net.rotor_motor = motor_hl
+
+ # ------------------------------------------------------------------
+ # Nacelles
+ # ------------------------------------------------------------------
+ nacelle = SUAVE.Components.Nacelles.Nacelle()
+ nacelle.tag = 'rotor_nacelle'
+ nacelle.length = 0.5
+ nacelle.diameter = 0.2
+ nacelle.orientation_euler_angles = [0,0,0.]
+ nacelle.flow_through = False
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_1'
+ nac_segment.percent_x_location = 0.0
+ nac_segment.height = 0.0
+ nac_segment.width = 0.0
+ nacelle.append_segment(nac_segment)
+
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_2'
+ nac_segment.percent_x_location = 0.1
+ nac_segment.height = 0.25
+ nac_segment.width = 0.25
+ nacelle.append_segment(nac_segment)
+
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_3'
+ nac_segment.percent_x_location = 0.35
+ nac_segment.height = 0.3
+ nac_segment.width = 0.3
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_4'
+ nac_segment.percent_x_location = 0.5
+ nac_segment.height = 0.4
+ nac_segment.width = 0.4
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_5'
+ nac_segment.percent_x_location = 0.85
+ nac_segment.height = 0.3
+ nac_segment.width = 0.3
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_6'
+ nac_segment.percent_x_location = 0.9
+ nac_segment.height = 0.25
+ nac_segment.width = 0.25
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_6'
+ nac_segment.percent_x_location = 1.0
+ nac_segment.height = 0.0
+ nac_segment.width = 0.0
+ nacelle.append_segment(nac_segment)
+
+ lift_rotor_nacelle_origins = prop_hl_origins
+
+ for ii in range(net.number_of_rotor_engines):
+ rotor_nacelle = deepcopy(nacelle)
+ rotor_nacelle.tag = 'rotor_nacelle_' + str(ii+1)
+ rotor_nacelle.origin = [lift_rotor_nacelle_origins[ii]]
+ vehicle.append_component(rotor_nacelle)
+
+
+ # Update for cruise propeller
+ nacelle.tag = 'cruise_prop_nacelle'
+ nacelle.length = 1.0
+ nacelle.diameter = 0.6
+
+ propeller_nacelle_origins = [[2.5, 4.97584, 1.01],[2.5, -4.97584, 1.01]]
+
+ for ii in range(net.number_of_propeller_engines):
+ propeller_nacelle = deepcopy(nacelle)
+ propeller_nacelle.tag = 'propeller_nacelle_' + str(ii+1)
+ propeller_nacelle.origin = [propeller_nacelle_origins[ii]]
+ vehicle.append_component(propeller_nacelle)
+
+ # Component 6 the Payload -----------------------------------------------------------------
+ payload = SUAVE.Components.Energy.Peripherals.Payload()
+ payload.power_draw = 10. #Watts
+ payload.mass_properties.mass = 1.0 * Units.kg
+ net.payload = payload
+
+ # Component 7 the Avionics-----------------------------------------------------------------
+ avionics = SUAVE.Components.Energy.Peripherals.Avionics()
+ avionics.power_draw = 20. #Watts
+ net.avionics = avionics
+
+ # Component 9 Miscellaneous Systems
+ sys = SUAVE.Components.Systems.System()
+ sys.mass_properties.mass = 5 # kg
+
+
+ # add the solar network to the vehicle
+ vehicle.append_component(net)
+
+ return vehicle
+
+if __name__ == '__main__':
+ main()
+ plt.show()
+
+
+
+
+
+
+
\ No newline at end of file
diff --git a/BWB_CFD/BWB.py b/tut_CFD_BWB/BWB.py
similarity index 100%
rename from BWB_CFD/BWB.py
rename to tut_CFD_BWB/BWB.py
diff --git a/BWB_CFD/base_data_10.txt b/tut_CFD_BWB/base_data_10.txt
similarity index 100%
rename from BWB_CFD/base_data_10.txt
rename to tut_CFD_BWB/base_data_10.txt
diff --git a/BWB_CFD/base_data_1500.txt b/tut_CFD_BWB/base_data_1500.txt
similarity index 100%
rename from BWB_CFD/base_data_1500.txt
rename to tut_CFD_BWB/base_data_1500.txt
diff --git a/tut_Control_Surface_Sizing_Navion/Missions.py b/tut_Control_Surface_Sizing_Navion/Missions.py
new file mode 100644
index 0000000..bba0813
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/Missions.py
@@ -0,0 +1,83 @@
+# Missions.py
+
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+
+import SUAVE
+from SUAVE.Core import Units
+import numpy as np
+
+# ----------------------------------------------------------------------
+# Define the Mission
+# ----------------------------------------------------------------------
+def stick_fixed_stability_setup(analyses,vehicle):
+ missions = SUAVE.Analyses.Mission.Mission.Container()
+ max_speed_multiplier = 1.0 # this multiplier is used to compute V_max from V_nominal
+ missions.stick_fixed_cruise = base_mission_setup(vehicle,max_speed_multiplier)
+
+ return missions
+
+def elevator_sizing_setup(analyses,vehicle):
+ missions = SUAVE.Analyses.Mission.Mission.Container()
+ max_speed_multiplier = 1.4 # this multiplier is used to compute V_max from V_nominal
+ missions.elevator_sizing = base_mission_setup(vehicle,max_speed_multiplier)
+
+ return missions
+
+def aileron_rudder_sizing_setup(analyses,vehicle):
+ missions = SUAVE.Analyses.Mission.Mission.Container()
+ max_speed_multiplier = 1.0
+ missions.aileron_sizing = base_mission_setup(vehicle,max_speed_multiplier)
+ max_speed_multiplier = 1.4 # this multiplier is used to compute V_max from V_nominal
+ missions.turn_criteria = base_mission_setup(vehicle,max_speed_multiplier)
+
+ return missions
+
+def flap_sizing_setup(analyses,vehicle):
+ missions = SUAVE.Analyses.Mission.Mission.Container()
+ max_speed_multiplier = 1.0
+ missions.flap_sizing = base_mission_setup(vehicle,max_speed_multiplier)
+ return missions
+
+
+# ------------------------------------------------------------------
+# Initialize the Mission
+# ------------------------------------------------------------------
+
+def base_mission_setup(vehicle,max_speed_multiplier):
+ '''
+ This sets up the nominal cruise of the aircraft
+ '''
+
+ mission = SUAVE.Analyses.Mission.Sequential_Segments()
+ mission.tag = 'mission'
+
+ # airport
+ airport = SUAVE.Attributes.Airports.Airport()
+ airport.altitude = 0. * Units.ft
+ airport.delta_isa = 0.0
+ airport.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
+
+ mission.airport = airport
+
+ # unpack Segments module
+ Segments = SUAVE.Analyses.Mission.Segments
+
+ # base segment
+ base_segment = Segments.Segment()
+ base_segment.process.initialize.initialize_battery = SUAVE.Methods.Missions.Segments.Common.Energy.initialize_battery
+ base_segment.process.iterate.conditions.planet_position = SUAVE.Methods.skip
+ base_segment.state.numerics.number_control_points = 4
+
+ # Cruise Segment: constant Speed, constant altitude
+ segment = Segments.Cruise.Constant_Speed_Constant_Altitude(base_segment)
+ segment.tag = "cruise"
+ segment.battery_energy = vehicle.networks.battery_propeller.battery.max_energy * 0.89
+ segment.altitude = 8012 * Units.feet
+ segment.air_speed = 120.91 * Units['mph'] * max_speed_multiplier
+ segment.distance = 20. * Units.nautical_mile
+ segment = vehicle.networks.battery_propeller.add_unknowns_and_residuals_to_segment(segment)
+ mission.append_segment(segment)
+
+ return mission
diff --git a/tut_Control_Surface_Sizing_Navion/Optimize.py b/tut_Control_Surface_Sizing_Navion/Optimize.py
new file mode 100644
index 0000000..5f82746
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/Optimize.py
@@ -0,0 +1,474 @@
+# Optimize.py
+
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+import SUAVE
+from SUAVE.Core import Units, Data
+import numpy as np
+import Vehicles
+import Missions
+import Procedure
+import SUAVE.Optimization.Package_Setups.scipy_setup as scipy_setup
+from SUAVE.Optimization import Nexus
+import time
+# ----------------------------------------------------------------------
+# Run the whole thing
+# ----------------------------------------------------------------------
+def main():
+ '''
+ STICK FIXED (STATIC STABILITY AND DRAG) OTIMIZATION
+ '''
+ ti = time.time()
+ solver_name = 'SLSQP'
+ planform_optimization_problem = stick_fixed_stability_and_drag_optimization_setup()
+ output = scipy_setup.SciPy_Solve(planform_optimization_problem,solver=solver_name, sense_step = 1E-3, tolerance = 1E-3)
+ print (output)
+ tf = time.time()
+ elapsed_time = round((tf-ti)/60,2)
+ print('Stick Fixed Stability and Drag Otimization Simulation Time: ' + str(elapsed_time))
+
+ '''
+ ELEVATOR SIZING
+ '''
+ # define vehicle for elevator sizing
+ optimized_vehicle_v1 = planform_optimization_problem.vehicle_configurations.stick_fixed_cruise
+ optimized_vehicle_v1.maximum_elevator_deflection = 30*Units.degrees
+ optimized_vehicle_v1.maxiumum_load_factor = 3.0
+ optimized_vehicle_v1.minimum_load_factor = -1
+
+ ti = time.time()
+ solver_name = 'SLSQP'
+ elevator_sizing_optimization_problem = elevator_sizing_optimization_setup(optimized_vehicle_v1)
+ output = scipy_setup.SciPy_Solve(elevator_sizing_optimization_problem,solver=solver_name, sense_step = 1E-3, tolerance = 1E-3)
+ print (output)
+ tf = time.time()
+ elapsed_time = round((tf-ti)/60,2)
+ print('Elevator Sizing Simulation Time: ' + str(elapsed_time))
+
+ '''
+ AILERON AND RUDDER SIZING
+ '''
+ # define vehicle for aileron and rudder sizing
+ optimized_vehicle_v2 = elevator_sizing_optimization_problem.vehicle_configurations.elevator_sizing
+ optimized_vehicle_v2.rudder_flag = True
+ optimized_vehicle_v2.maximum_aileron_rudder_deflection = 30*Units.degrees
+ optimized_vehicle_v2.crosswind_velocity = 20 * Units.knots
+
+ ti = time.time()
+ solver_name = 'SLSQP'
+ aileron_rudder_sizing_optimization_problem = aileron_rudder_sizing_optimization_setup(optimized_vehicle_v2)
+ output = scipy_setup.SciPy_Solve(aileron_rudder_sizing_optimization_problem,solver=solver_name, sense_step = 1E-3, tolerance = 1E-3)
+ print (output)
+ tf = time.time()
+ elapsed_time = round((tf-ti)/60,2)
+ print('Aileron and Rudder Sizing Simulation Time: ' + str(elapsed_time))
+
+ '''
+ FLAP SIZING
+ '''
+ # define vehicle for flap sizing
+ optimized_vehicle_v3 = aileron_rudder_sizing_optimization_problem.vehicle_configurations.aileron_rudder_sizing
+ optimized_vehicle_v3.maximum_flap_deflection = 40*Units.degrees
+
+ ti = time.time()
+ solver_name = 'SLSQP'
+ flap_sizing_optimization_problem = flap_sizing_optimization_setup(optimized_vehicle_v3)
+ output = scipy_setup.SciPy_Solve(flap_sizing_optimization_problem,solver=solver_name, sense_step = 1E-3, tolerance = 1E-3)
+ print (output)
+ tf = time.time()
+ elapsed_time = round((tf-ti)/60,2)
+ print('Flap Sizing Simulation Time: ' + str(elapsed_time))
+
+ '''
+ PRINT VEHICLE CONTROL SURFACES
+ '''
+ optimized_vehicle_v4 = flap_sizing_optimization_problem.vehicle_configurations.flap_sizing
+ print_vehicle_control_surface_geoemtry(optimized_vehicle_v4)
+
+ return
+
+def stick_fixed_stability_and_drag_optimization_setup():
+ nexus = Nexus()
+ problem = Data()
+ nexus.optimization_problem = problem
+
+ # -------------------------------------------------------------------
+ # Inputs
+ # -------------------------------------------------------------------
+
+ # [ tag , initial, (lb , ub) , scaling , units ]
+ problem.inputs = np.array([
+ #[ 'mw_taper' , 0.5 , 0.4 , 1.0 , 1.0 , 1*Units.less],
+ #[ 'mw_area' , 15.39, 14.0 , 16.0 , 100. , 1*Units.meter**2],
+ #[ 'mw_AR' , 11.0 , 9.0 , 14.0 , 100 , 1*Units.less],
+ [ 'mw_root_twist' , 3.0 , -5.0 , 5.0 , 10. , 1*Units.degree],
+ [ 'mw_tip_twist' , 0.0 , -5.0 , 5.0 , 10. , 1*Units.degree],
+ #[ 'mw_dihedral' , 0.0 , 0.0 , 5.0 , 10. , 1*Units.degree],
+ #[ 'hs_AR' , 4.287, 4.0 , 4.2 , 10. , 1*Units.less],
+ #[ 'hs_area' , 2.54 , 2.30 , 3.0 , 10. , 1*Units.meter**2],
+ #[ 'hs_root_twist' , 0.0 , -5.0 , 5.0 , 10. , 1*Units.degree],
+ #[ 'hs_tip_twist' , 0.0 , -5.0 , 5.0 , 10. , 1*Units.degree],
+ [ 'c_g_x' , 3.1 , 2.0 , 4.0 , 10 , 1*Units.less],
+
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Objective
+ # -------------------------------------------------------------------
+
+ # [ tag, scaling, units ]
+ problem.objective = np.array([
+ [ 'CD' , 1.0 , 1*Units.less]
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Constraints
+ # -------------------------------------------------------------------
+
+ # [ tag, sense, edge, scaling, units ]
+ problem.constraints = np.array([
+ [ 'CM_residual' , '<' , 1E-2 , 1E-2 , 1*Units.less], # close to zero 2 works
+ [ 'static_margin' , '>' , 0.1 , 0.1 , 1*Units.less],
+ [ 'CM_alpha' , '<' , 0.0 , 1.0 , 1*Units.less],
+ #[ 'phugoid_damping_ratio' , '>' , 0.04 , 1.0 , 1*Units.less],
+ #[ 'short_period_frequency', '>' , 1.34 , 1.0 , 1*Units.less],
+ #[ 'dutch_roll_frequency' , '>' , 1.0 , 1.0 , 1*Units.less],
+ #[ 'spiral_doubling_time' , '>' , 4.0 , 1.0 , 1*Units.less],
+ #[ 'spiral_criteria' , '>' , 1.0 , 1.0 , 1*Units.less],
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Aliases
+ # -------------------------------------------------------------------
+
+ # [ 'alias' , ['data.path1.name','data.path2.name'] ]
+ problem.aliases = [
+ [ 'CD' , 'summary.CD' ],
+ [ 'CM_residual' , 'summary.CM_residual' ],
+ [ 'CM_alpha' , 'summary.CM_alpha' ],
+ [ 'static_margin' , 'summary.static_margin' ],
+ #[ 'phugoid_damping_ratio' , 'summary.phugoid_damping_ratio' ],
+ #[ 'short_period_frequency' , 'summary.short_period_frequency' ],
+ #[ 'dutch_roll_frequency' , 'summary.dutch_roll_frequency' ],
+ #[ 'spiral_doubling_time' , 'summary.spiral_doubling_time' ],
+ #[ 'spiral_criteria' , 'summary.spiral_criteria' ],
+ #[ 'mw_area' , 'vehicle_configurations.*.wings.main_wing.aspect_ratio'],
+ #[ 'mw_taper' , 'vehicle_configurations.*.wings.main_wing.taper'],
+ #[ 'mw_AR' , 'vehicle_configurations.*.wings.main_wing.aspect_ratio'],
+ [ 'mw_root_twist' , 'vehicle_configurations.*.wings.main_wing.twists.root' ],
+ [ 'mw_tip_twist' , 'vehicle_configurations.*.wings.main_wing.twists.tip' ],
+ #[ 'mw_dihedral' , 'vehicle_configurations.*.wings.main_wing.dihedral' ],
+ #[ 'hs_AR' , 'vehicle_configurations.*.wings.horizontal_stabilizer.aspect_ratio'],
+ #[ 'hs_area' , 'vehicle_configurations.*.wings.horizontal_stabilizer.aspect_ratio'],
+ #[ 'hs_taper' , 'vehicle_configurations.*.wings.horizontal_stabilizer.taper'],
+ #[ 'hs_root_twist' , 'vehicle_configurations.*.wings.horizontal_stabilizer.twists.root' ],
+ #[ 'hs_tip_twist' , 'vehicle_configurations.*.wings.horizontal_stabilizer.twists.tip' ],
+ #[ 'hs_dihedral' , 'vehicle_configurations.*.wings.horizontal_stabilizer.dihedral' ],
+ [ 'c_g_x' , 'vehicle_configurations.*.mass_properties.center_of_gravity[0][0]' ],
+ ]
+
+ # -------------------------------------------------------------------
+ # Vehicles
+ # -------------------------------------------------------------------
+ nexus.vehicle_configurations = Vehicles.stick_fixed_stability_setup()
+
+ # -------------------------------------------------------------------
+ # Analyses
+ # -------------------------------------------------------------------
+ nexus.analyses = None
+
+ # -------------------------------------------------------------------
+ # Missions
+ # -------------------------------------------------------------------
+ nexus.missions = Missions.stick_fixed_stability_setup(nexus.analyses,nexus.vehicle_configurations.stick_fixed_cruise)
+
+ # -------------------------------------------------------------------
+ # Procedure
+ # -------------------------------------------------------------------
+ nexus.procedure = Procedure.stick_fixed_stability_and_drag_procedure()
+
+ # -------------------------------------------------------------------
+ # Summary
+ # -------------------------------------------------------------------
+ nexus.summary = Data()
+ return nexus
+
+def elevator_sizing_optimization_setup(vehicle):
+
+ nexus = Nexus()
+ problem = Data()
+ nexus.optimization_problem = problem
+
+ # -------------------------------------------------------------------
+ # Inputs
+ # -------------------------------------------------------------------
+
+ # [ tag , initial, (lb , ub) , scaling , units ]
+ problem.inputs = np.array([
+ [ 'hs_elevator_chord_fraction' , 0.2 , 0.1 , 0.3 , 1.0 , 1*Units.less],
+ [ 'hs_elevator_span_frac_start', 0.25 , 0.05 , 0.45 , 1.0 , 1*Units.less],
+ [ 'hs_elevator_span_frac_end' , 0.75 , 0.55 , 0.9 , 1.0 , 1*Units.less],
+
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Objective
+ # -------------------------------------------------------------------
+
+ # [ tag, scaling, units ]
+ problem.objective = np.array([
+ [ 'elevator_surface_area', 1. , 1*Units.kg],
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Constraints
+ # -------------------------------------------------------------------
+
+ # [ tag, sense, edge, scaling, units ]
+ problem.constraints = np.array([
+ [ 'elevator_push_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ [ 'elevator_pull_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Aliases
+ # -------------------------------------------------------------------
+
+ # [ 'alias' , ['data.path1.name','data.path2.name'] ]
+ problem.aliases = [
+ [ 'elevator_surface_area' , 'summary.elevator_surface_area' ],
+ [ 'elevator_push_deflection_residual' , 'summary.elevator_push_deflection_residual' ],
+ [ 'elevator_pull_deflection_residual' , 'summary.elevator_pull_deflection_residual' ],
+ [ 'hs_elevator_chord_fraction' , 'vehicle_configurations.*.wings.horizontal_stabilizer.control_surfaces.elevator.chord_fraction'],
+ [ 'hs_elevator_span_frac_start' , 'vehicle_configurations.*.wings.horizontal_stabilizer.control_surfaces.elevator.span_fraction_start'],
+ [ 'hs_elevator_span_frac_end' , 'vehicle_configurations.*.wings.horizontal_stabilizer.control_surfaces.elevator.span_fraction_end'],
+ ]
+
+ # -------------------------------------------------------------------
+ # Vehicles
+ # -------------------------------------------------------------------
+ nexus.vehicle_configurations = Vehicles.elevator_sizing_setup(vehicle)
+
+ # -------------------------------------------------------------------
+ # Analyses
+ # -------------------------------------------------------------------
+ nexus.analyses = None
+
+ # -------------------------------------------------------------------
+ # Missions
+ # -------------------------------------------------------------------
+ nexus.missions = Missions.elevator_sizing_setup(nexus.analyses,nexus.vehicle_configurations.elevator_sizing)
+
+ # -------------------------------------------------------------------
+ # Procedure
+ # -------------------------------------------------------------------
+ nexus.procedure = Procedure.elevator_sizing_setup()
+
+ # -------------------------------------------------------------------
+ # Summary
+ # -------------------------------------------------------------------
+ nexus.summary = Data()
+ return nexus
+
+
+
+
+def aileron_rudder_sizing_optimization_setup(vehicle):
+
+ nexus = Nexus()
+ problem = Data()
+ nexus.optimization_problem = problem
+
+ # -------------------------------------------------------------------
+ # Inputs
+ # -------------------------------------------------------------------
+
+ # [ tag , initial, (lb , ub) , scaling , units ]
+ if vehicle.rudder_flag:
+ problem.inputs = np.array([
+ [ 'mw_aileron_chord_fraction' , 0.2 , 0.15 , 0.3 , 1.0 , 1*Units.less],
+ [ 'mw_aileron_span_frac_start' , 0.75 , 0.55 , 0.8 , 1.0 , 1*Units.less],
+ [ 'mw_aileron_span_frac_end' , 0.9 , 0.85 , 0.95 , 1.0 , 1*Units.less],
+ [ 'vs_rudder_chord_fraction' , 0.2 , 0.15 , 0.3 , 1.0 , 1*Units.less],
+ [ 'vs_rudder_span_frac_start' , 0.25 , 0.05 , 0.35 , 1.0 , 1*Units.less],
+ [ 'vs_rudder_span_frac_end' , 0.75 , 0.5 , 0.95 , 1.0 , 1*Units.less] ],dtype=object)
+ else:
+ problem.inputs = np.array([
+ [ 'mw_aileron_chord_fraction' , 0.2 , 0.15 , 0.3 , 1.0 , 1*Units.less],
+ [ 'mw_aileron_span_frac_start' , 0.75 , 0.55 , 0.8 , 1.0 , 1*Units.less],
+ [ 'mw_aileron_span_frac_end' , 0.9 , 0.85 , 0.95 , 1.0 , 1*Units.less],
+
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Objective
+ # -------------------------------------------------------------------
+
+ # [ tag, scaling, units ]
+ problem.objective = np.array([
+ [ 'aileron_rudder_surface_area', 1. , 1*Units.kg],
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Constraints
+ # -------------------------------------------------------------------
+
+ # [ tag, sense, edge, scaling, units ]
+ if vehicle.rudder_flag:
+ problem.constraints = np.array([
+ [ 'aileron_roll_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ [ 'rudder_roll_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ [ 'aileron_crosswind_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ [ 'rudder_crosswind_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ ],dtype=object)
+ else:
+ problem.constraints = np.array([
+ [ 'aileron_roll_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ [ 'aileron_crosswind_deflection_residual' , '>' , 0 , 1.0 , 1*Units.less],
+ ],dtype=object)
+
+
+ # -------------------------------------------------------------------
+ # Aliases
+ # -------------------------------------------------------------------
+
+ # [ 'alias' , ['data.path1.name','data.path2.name'] ]
+ if vehicle.rudder_flag:
+ problem.aliases = [
+ [ 'aileron_rudder_surface_area' , 'summary.aileron_rudder_surface_area' ],
+ [ 'aileron_roll_deflection_residual' , 'summary.aileron_roll_deflection_residual' ],
+ [ 'rudder_roll_deflection_residual' , 'summary.rudder_roll_deflection_residual' ],
+ [ 'aileron_crosswind_deflection_residual' , 'summary.aileron_crosswind_deflection_residual' ],
+ [ 'rudder_crosswind_deflection_residual' , 'summary.rudder_crosswind_deflection_residual' ],
+ [ 'mw_aileron_chord_fraction' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.aileron.chord_fraction'],
+ [ 'mw_aileron_span_frac_start' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.aileron.span_fraction_start'],
+ [ 'mw_aileron_span_frac_end' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.aileron.span_fraction_end'],
+ [ 'vs_rudder_chord_fraction' , 'vehicle_configurations.*.wings.vertical_stabilizer.control_surfaces.rudder.chord_fraction'],
+ [ 'vs_rudder_span_frac_start' , 'vehicle_configurations.*.wings.vertical_stabilizer.control_surfaces.rudder.span_fraction_start'],
+ [ 'vs_rudder_span_frac_end' , 'vehicle_configurations.*.wings.vertical_stabilizer.control_surfaces.rudder.span_fraction_end']]
+ else:
+ problem.aliases = [
+ [ 'aileron_rudder_surface_area' , 'summary.aileron_rudder_surface_area' ],
+ [ 'aileron_roll_deflection_residual' , 'summary.aileron_roll_deflection_residual' ],
+ [ 'aileron_crosswind_deflection_residual' , 'summary.aileron_crosswind_deflection_residual' ],
+ [ 'mw_aileron_chord_fraction' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.aileron.chord_fraction'],
+ [ 'mw_aileron_span_frac_start' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.aileron.span_fraction_start'],
+ [ 'mw_aileron_span_frac_end' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.aileron.span_fraction_end']]
+
+ # -------------------------------------------------------------------
+ # Vehicles
+ # -------------------------------------------------------------------
+ nexus.vehicle_configurations = Vehicles.aileron_rudder_sizing_setup(vehicle)
+
+ # -------------------------------------------------------------------
+ # Analyses
+ # -------------------------------------------------------------------
+ nexus.analyses = None
+
+ # -------------------------------------------------------------------
+ # Missions
+ # -------------------------------------------------------------------
+ nexus.missions = Missions.aileron_rudder_sizing_setup(nexus.analyses,nexus.vehicle_configurations.aileron_rudder_sizing)
+
+ # -------------------------------------------------------------------
+ # Procedure
+ # -------------------------------------------------------------------
+ nexus.procedure = Procedure.aileron_rudder_sizing_setup()
+
+ # -------------------------------------------------------------------
+ # Summary
+ # -------------------------------------------------------------------
+ nexus.summary = Data()
+ return nexus
+
+
+def flap_sizing_optimization_setup(optimized_vehicle):
+
+ nexus = Nexus()
+ problem = Data()
+ nexus.optimization_problem = problem
+
+ # -------------------------------------------------------------------
+ # Inputs
+ # -------------------------------------------------------------------
+
+ # [ tag , initial, (lb , ub) , scaling , units ]
+ problem.inputs = np.array([
+ [ 'mw_flap_chord_fraction' , 0.2 , 0.15 , 0.4 , 1.0 , 1*Units.less],
+ [ 'mw_flap_span_frac_start' , 0.2 , 0.05 , 0.25 , 1.0 , 1*Units.less],
+ [ 'mw_flap_span_frac_end' , 0.4 , 0.3 , 0.5 , 1.0 , 1*Units.less],
+
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Objective
+ # -------------------------------------------------------------------
+
+ # [ tag, scaling, units ]
+ problem.objective = np.array([
+ [ 'flap_surface_area', 1. , 1*Units.kg],
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Constraints
+ # -------------------------------------------------------------------
+
+ # [ tag, sense, edge, scaling, units ]
+ problem.constraints = np.array([
+ [ 'flap_criteria' , '>' , 0. , 1.0 , 1*Units.less],
+ ],dtype=object)
+
+ # -------------------------------------------------------------------
+ # Aliases
+ # -------------------------------------------------------------------
+
+ # [ 'alias' , ['data.path1.name','data.path2.name'] ]
+ problem.aliases = [
+ [ 'flap_surface_area' , 'summary.flap_surface_area' ],
+ [ 'flap_criteria' , 'summary.flap_criteria' ],
+ [ 'mw_flap_chord_fraction' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.flap.chord_fraction'],
+ [ 'mw_flap_span_frac_start' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.flap.span_fraction_start'],
+ [ 'mw_flap_span_frac_end' , 'vehicle_configurations.*.wings.main_wing.control_surfaces.flap.span_fraction_end'],
+ ]
+
+ # -------------------------------------------------------------------
+ # Vehicles
+ # -------------------------------------------------------------------
+ nexus.vehicle_configurations = Vehicles.flap_sizing_setup(optimized_vehicle)
+
+ # -------------------------------------------------------------------
+ # Analyses
+ # -------------------------------------------------------------------
+ nexus.analyses = None
+
+ # -------------------------------------------------------------------
+ # Missions
+ # -------------------------------------------------------------------
+ nexus.missions = Missions.flap_sizing_setup(nexus.analyses,nexus.vehicle_configurations.flap_sizing)
+
+ # -------------------------------------------------------------------
+ # Procedure
+ # -------------------------------------------------------------------
+ nexus.procedure = Procedure.flap_sizing_setup()
+
+ # -------------------------------------------------------------------
+ # Summary
+ # -------------------------------------------------------------------
+ nexus.summary = Data()
+ return nexus
+
+def print_vehicle_control_surface_geoemtry(vehicle):
+ for wing in vehicle.wings:
+ if 'control_surfaces' in wing:
+ for CS in wing.control_surfaces:
+ print('Wing : ' + wing.tag)
+ print('Control Surface : ' + CS.tag)
+ print('Span Fraction Start : ' + str(CS.span_fraction_start))
+ print('Span Fraction End : ' + str(CS.span_fraction_end))
+ print('Chord Fraction : ' + str(CS.chord_fraction))
+ print("\n\n")
+
+ return
+if __name__ == '__main__':
+ main()
diff --git a/tut_Control_Surface_Sizing_Navion/Procedure.py b/tut_Control_Surface_Sizing_Navion/Procedure.py
new file mode 100644
index 0000000..a4af0c0
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/Procedure.py
@@ -0,0 +1,431 @@
+# Procedure.py
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+import SUAVE
+from SUAVE.Core import Units, Data
+import numpy as np
+from SUAVE.Analyses.Process import Process
+from SUAVE.Methods.Weights.Correlations.UAV import empty
+from SUAVE.Methods.Weights.Buildups.eVTOL.empty import empty
+
+from SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics import Aerodynamics
+# Routines
+import Missions
+
+# ----------------------------------------------------------------------
+# Setup
+# ----------------------------------------------------------------------
+
+def stick_fixed_stability_and_drag_procedure():
+ procedure = Process()
+ procedure.modify_vehicle = modify_stick_fixed_vehicle
+ procedure.post_process = longitudinal_static_stability_and_drag_post_process
+
+ return procedure
+
+def elevator_sizing_setup():
+ procedure = Process()
+ procedure.post_process = elevator_sizing_post_process
+ return procedure
+
+def aileron_rudder_sizing_setup():
+ procedure = Process()
+ procedure.post_process = aileron_rudder_sizing_post_process
+ return procedure
+
+def flap_sizing_setup():
+ procedure = Process()
+ procedure.post_process = flap_sizing_post_process
+ return procedure
+
+# ----------------------------------------------------------------------
+# Modify Vehicle
+# ----------------------------------------------------------------------
+
+def modify_stick_fixed_vehicle(nexus):
+ '''
+ This function takes the updated design variables and modifies the aircraft
+ '''
+ # Pull out the vehicles
+ vehicle = nexus.vehicle_configurations.stick_fixed_cruise
+
+ # Update Wing
+ for wing in vehicle.wings:
+ update_chords(wing)
+
+ # ----------------------------------------------------------------------
+ # Update Vehicle Mass
+ # ----------------------------------------------------------------------
+ #weight_breakdown = empty(vehicle)
+ #vehicle.mass_properties.max_takeoff = weight_breakdown.total
+ #vehicle.mass_properties.takeoff = weight_breakdown.total
+
+ # Update Mission
+ nexus.missions = Missions.stick_fixed_stability_setup(nexus.analyses,vehicle)
+
+ # diff the new data
+ vehicle.store_diff()
+
+ return nexus
+
+def update_chords(wing):
+ '''
+ Updates the wing planform each iteration
+ '''
+ Sref = wing.areas.reference # fixed
+ span = wing.spans.projected # optimization input
+ taper = wing.taper # optimization input
+ croot = 2*Sref/((taper+1)*span) # set by Sref and current design point
+ ctip = taper * croot # set by Sref and current design point
+ wing.chords.root = croot
+ wing.chords.tip = ctip
+
+ # Wing Segments
+ if 'Segments' in wing:
+ for seg in wing.Segments:
+ seg.twist = (wing.twists.tip-wing.twists.root)*seg.percent_span_location + wing.twists.root
+
+ return wing
+
+def longitudinal_static_stability_and_drag_post_process(nexus):
+ '''
+ This function analyses and post processes the aircraft at cruise conditions.
+ The objective of is to minimize the drag of a trimmed aircraft
+ '''
+ summary = nexus.summary
+ vehicle = nexus.vehicle_configurations.stick_fixed_cruise
+ g = 9.81
+ L = g*vehicle.mass_properties.max_takeoff
+ S = vehicle.reference_area
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude = \
+ nexus.missions['stick_fixed_cruise'].segments['cruise'].altitude )
+
+
+ run_conditions = Aerodynamics()
+ run_conditions.freestream.density = atmo_data.density[0,0]
+ run_conditions.freestream.gravity = g
+ run_conditions.freestream.speed_of_sound = atmo_data.speed_of_sound[0,0]
+ run_conditions.freestream.velocity = nexus.missions['stick_fixed_cruise'].segments['cruise'].air_speed
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+ run_conditions.aerodynamics.angle_of_attack = np.array([0.0])
+ run_conditions.aerodynamics.lift_coefficient = L/(S*(0.5*run_conditions.freestream.density*(run_conditions.freestream.velocity**2)))
+ run_conditions.aerodynamics.roll_rate_coefficient = 0.0
+ run_conditions.aerodynamics.pitch_rate_coefficient = 0.0
+
+ stability_stick_fixed = SUAVE.Analyses.Stability.AVL()
+ stability_stick_fixed.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ stability_stick_fixed.geometry = nexus.vehicle_configurations.stick_fixed_cruise
+ results_stick_fixed = stability_stick_fixed.evaluate_conditions(run_conditions, trim_aircraft = True )
+
+
+ summary.CD = results_stick_fixed.aerodynamics.drag_breakdown.induced.total[0,0]
+ summary.CM_residual = abs(results_stick_fixed.aerodynamics.pitch_moment_coefficient[0,0])
+ summary.spiral_criteria = results_stick_fixed.stability.static.spiral_criteria[0,0]
+ NP = results_stick_fixed.stability.static.neutral_point[0,0]
+ cg = vehicle.mass_properties.center_of_gravity[0][0]
+ MAC = vehicle.wings.main_wing.chords.mean_aerodynamic
+ summary.static_margin = (NP - cg)/MAC
+ summary.CM_alpha = results_stick_fixed.stability.static.Cm_alpha[0,0]
+
+ if np.count_nonzero(vehicle.mass_properties.moments_of_inertia.tensor) > 0:
+ summary.phugoid_damping_ratio = results_stick_fixed.dynamic_stability.LongModes.phugoidDamp[0,0]
+ summary.short_period_frequency = results_stick_fixed.dynamic_stability.LongModes.shortPeriodFreqHz[0,0]
+ summary.dutch_roll_frequency = results_stick_fixed.dynamic_stability.LatModes.dutchRollFreqHz[0,0]
+ summary.spiral_doubling_time = results_stick_fixed.dynamic_stability.LatModes.spiralTimeDoubleHalf[0,0]
+ print("Drag Coefficient : " + str(summary.CD))
+ print("Moment Coefficient : " + str(summary.CM_residual))
+ print("Static Margin : " + str(summary.static_margin))
+ print("CM alpla : " + str(summary.CM_alpha))
+ print("Phugoid Damping Ratio : " + str(summary.phugoid_damping_ratio))
+ print("Short Period Frequency: " + str(summary.short_period_frequency))
+ print("Dutch Roll Frequency : " + str(summary.dutch_roll_frequency))
+ print("Spiral Doubling Time : " + str(summary.spiral_doubling_time))
+ print("Spiral Criteria : " + str(summary.spiral_criteria))
+ print("\n\n")
+
+ else:
+ summary.phugoid_damping_ratio = 0
+ summary.short_period_frequency = 0
+ summary.dutch_roll_frequency = 0
+ summary.spiral_doubling_time = 0
+ summary.spiral_criteria = 0
+ print("Drag Coefficient : " + str(summary.CD))
+ print("Moment Coefficient : " + str(summary.CM_residual))
+ print("Static Margin : " + str(summary.static_margin))
+ print("CM alpla : " + str(summary.CM_alpha))
+ print("Spiral Criteria : " + str(summary.spiral_criteria))
+ print("\n\n")
+
+
+ vehicle.trim_cl = run_conditions.aerodynamics.lift_coefficient
+ vehicle.trim_airspeed = run_conditions.freestream.velocity
+
+ return nexus
+
+def elevator_sizing_post_process(nexus):
+ '''
+ This function analyses and post processes the aircraft at the flight conditions required to size
+ the elevator. These conditions are:
+ 1) Stick pull maneuver with a load factor of 3.0
+ 2) Stick push maneuver with a load factor of -1
+ '''
+ summary = nexus.summary
+ trim_aircraft = True
+ g = 9.81
+ vehicle = nexus.vehicle_configurations.elevator_sizing
+ m = vehicle.mass_properties.max_takeoff
+ S = vehicle.reference_area
+ V_trim = vehicle.trim_airspeed
+ max_defl = vehicle.maximum_elevator_deflection
+ V_max = nexus.missions['elevator_sizing'].segments['cruise'].air_speed
+
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude = nexus.missions['elevator_sizing'].segments['cruise'].altitude )
+ run_conditions = Aerodynamics()
+ run_conditions.freestream.density = atmo_data.density[0,0]
+ run_conditions.freestream.gravity = g
+ run_conditions.freestream.speed_of_sound = atmo_data.speed_of_sound[0,0]
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+ run_conditions.aerodynamics.angle_of_attack = np.array([0.0])
+ run_conditions.aerodynamics.roll_rate_coefficient = 0.0
+ run_conditions.aerodynamics.pitch_rate_coefficient = 0.0
+
+ q = 0.5*(V_trim**2)*atmo_data.density[0,0]
+ CL_pull_man = vehicle.maxiumum_load_factor*m*g/(S*q)
+ CL_push_man = vehicle.minimum_load_factor*m*g/(S*q)
+
+ stability_pull_maneuver = SUAVE.Analyses.Stability.AVL()
+ stability_pull_maneuver.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ run_conditions.aerodynamics.lift_coefficient = CL_pull_man
+ run_conditions.freestream.velocity = V_max
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ stability_pull_maneuver.settings.number_spanwise_vortices = 40
+ stability_pull_maneuver.geometry = vehicle
+ results_pull_maneuver = stability_pull_maneuver.evaluate_conditions(run_conditions, trim_aircraft )
+ AoA_pull = results_pull_maneuver.aerodynamics.AoA[0,0]
+ elevator_pull_deflection = results_pull_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.elevator.deflection
+
+ stability_push_maneuver = SUAVE.Analyses.Stability.AVL()
+ stability_push_maneuver.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ run_conditions.aerodynamics.lift_coefficient = CL_push_man
+ run_conditions.freestream.velocity = V_trim
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ stability_pull_maneuver.settings.number_spanwise_vortices = 40
+ stability_push_maneuver.geometry = vehicle
+ results_push_maneuver = stability_push_maneuver.evaluate_conditions(run_conditions, trim_aircraft )
+ AoA_push = results_push_maneuver.aerodynamics.AoA[0,0]
+ elevator_push_deflection = results_push_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.elevator.deflection
+
+ summary.elevator_pull_deflection_residual = (max_defl/Units.degrees - abs(elevator_pull_deflection))*Units.degrees
+ summary.elevator_push_deflection_residual = (max_defl/Units.degrees - abs(elevator_push_deflection))*Units.degrees
+
+ # compute control surface area
+ control_surfaces = ['elevator']
+ total_control_surface_area = compute_control_surface_areas(control_surfaces,vehicle)
+ summary.elevator_surface_area = total_control_surface_area
+
+
+ print("Elevator Area : " + str(summary.elevator_surface_area))
+ print("Aircraft CL Pull : " + str(CL_pull_man))
+ print("Aircraft AoA Pull : " + str(AoA_pull))
+ print("Elevator Pull Defl.: " + str(elevator_pull_deflection))
+ print("Aircraft CL Push : " + str(CL_push_man))
+ print("Aircraft AoA Push : " + str(AoA_push))
+ print("Elevator Push Defl.: " + str(elevator_push_deflection))
+ print("\n\n")
+
+ return nexus
+
+
+
+def aileron_rudder_sizing_post_process(nexus):
+ '''
+ This function analyses and post processes the aircraft at the flight conditions required to size
+ the aileron and rudder. These conditions are:
+ 1) A controlled roll at a rate of 0.07
+ 2) Trimmed flight in a 20 knot crosswind
+ '''
+ summary = nexus.summary
+ trim_aircraft = True
+ g = 9.81
+ vehicle = nexus.vehicle_configurations.aileron_rudder_sizing
+ CL_trim = vehicle.trim_cl
+ max_defl = vehicle.maximum_aileron_rudder_deflection
+ V_crosswind = vehicle.crosswind_velocity
+
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude = nexus.missions['aileron_sizing'].segments['cruise'].altitude )
+ run_conditions = Aerodynamics()
+ run_conditions.freestream.density = atmo_data.density[0,0]
+ run_conditions.freestream.gravity = g
+ run_conditions.freestream.speed_of_sound = atmo_data.speed_of_sound[0,0]
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+ run_conditions.aerodynamics.angle_of_attack = np.array([0.0])
+
+
+ stability_roll_maneuver = SUAVE.Analyses.Stability.AVL()
+ stability_roll_maneuver.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ stability_roll_maneuver.settings.number_spanwise_vortices = 40
+ run_conditions.aerodynamics.lift_coefficient = CL_trim
+ stability_roll_maneuver.geometry = vehicle
+ run_conditions.freestream.velocity = nexus.missions['aileron_sizing'].segments['cruise'].air_speed
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ run_conditions.aerodynamics.roll_rate_coefficient = 0.07
+ run_conditions.aerodynamics.pitch_rate_coefficient = 0.0
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+ results_roll_maneuver = stability_roll_maneuver.evaluate_conditions(run_conditions, trim_aircraft )
+ aileron_roll_deflection = results_roll_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.aileron.deflection
+
+ summary.aileron_roll_deflection_residual = (max_defl/Units.degrees - abs(aileron_roll_deflection))*Units.degrees
+ if vehicle.rudder_flag:
+ rudder_roll_deflection = results_roll_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.rudder.deflection
+ summary.rudder_roll_deflection_residual = (max_defl/Units.degrees - abs(rudder_roll_deflection))*Units.degrees
+ else:
+ rudder_roll_deflection = 0
+ summary.rudder_roll_deflection_residual = 0
+
+ stability_cross_wind_maneuver = SUAVE.Analyses.Stability.AVL()
+ stability_cross_wind_maneuver.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ run_conditions.aerodynamics.lift_coefficient = CL_trim
+ stability_cross_wind_maneuver.geometry = vehicle
+ run_conditions.freestream.velocity = nexus.missions['aileron_sizing'].segments['cruise'].air_speed
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ run_conditions.aerodynamics.roll_rate_coefficient = 0.0
+ run_conditions.aerodynamics.pitch_rate_coefficient = 0.0
+ run_conditions.aerodynamics.side_slip_angle = np.tan(V_crosswind/nexus.missions['aileron_sizing'].segments['cruise'].air_speed) # beta
+ results_cross_wind_maneuver = stability_cross_wind_maneuver.evaluate_conditions(run_conditions, trim_aircraft )
+ aileron_cross_wind_deflection = results_cross_wind_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.aileron.deflection
+
+ # criteria
+ summary.aileron_crosswind_deflection_residual = (max_defl/Units.degrees - abs(aileron_cross_wind_deflection))*Units.degrees
+
+ if vehicle.rudder_flag:
+ rudder_cross_wind_deflection = results_cross_wind_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.rudder.deflection
+ summary.rudder_crosswind_deflection_residual = (max_defl/Units.degrees - abs(rudder_cross_wind_deflection))*Units.degrees
+ else:
+ rudder_cross_wind_deflection = 0
+ summary.rudder_crosswind_deflection_residual = 0
+
+ # compute control surface area
+ control_surfaces = ['aileron','rudder']
+ total_control_surface_area = compute_control_surface_areas(control_surfaces,vehicle)
+ summary.aileron_rudder_surface_area = total_control_surface_area
+
+ print("Total Rudder Aileron Surface Area : " + str(summary.aileron_rudder_surface_area))
+ print("Aileron Roll Defl : " + str(aileron_roll_deflection))
+ print("Rudder Roll Defl : " + str(rudder_roll_deflection))
+ print("Aileron Crosswind Defl : " + str(aileron_cross_wind_deflection))
+ print("Rudder Crosswind Defl : " + str(rudder_cross_wind_deflection ))
+ print("\n\n")
+
+ return nexus
+
+
+def flap_sizing_post_process(nexus):
+ '''
+ This function analyses and post processes the aircraft at the flight conditions required to size
+ the flap. These conditions are:
+ 1) A comparison of clean and deployed flap at 12 deg. angle of attack
+ '''
+ summary = nexus.summary
+ trim_aircraft = False
+ g = 9.81
+ vehicle = nexus.vehicle_configurations.flap_sizing
+ max_defl = vehicle.maximum_flap_deflection
+ V_max = nexus.missions['flap_sizing'].segments['cruise'].air_speed
+
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude = nexus.missions['flap_sizing'].segments['cruise'].altitude )
+ run_conditions = Aerodynamics()
+ run_conditions.freestream.density = atmo_data.density[0,0]
+ run_conditions.freestream.gravity = g
+ run_conditions.freestream.speed_of_sound = atmo_data.speed_of_sound[0,0]
+ run_conditions.freestream.velocity = V_max
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+ run_conditions.aerodynamics.lift_coefficient = None
+ run_conditions.aerodynamics.angle_of_attack = np.array([12.0])*Units.degrees
+ run_conditions.aerodynamics.roll_rate_coefficient = 0.0
+ run_conditions.aerodynamics.pitch_rate_coefficient = 0.0
+
+ stability_no_flap = SUAVE.Analyses.Stability.AVL()
+ stability_no_flap.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ stability_no_flap.settings.number_spanwise_vortices = 40
+ vehicle.wings.main_wing.control_surfaces.flap.deflection = 0.0
+ stability_no_flap.geometry = vehicle
+ results_no_flap = stability_no_flap.evaluate_conditions(run_conditions, trim_aircraft)
+ CL_12_deg_no_flap = results_no_flap.aerodynamics.lift_coefficient[0,0]
+
+
+ stability_flap = SUAVE.Analyses.Stability.AVL()
+ stability_flap.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ stability_flap.settings.number_spanwise_vortices = 40
+ vehicle.wings.main_wing.control_surfaces.flap.deflection = max_defl
+ stability_flap.geometry = vehicle
+ results_flap = stability_flap.evaluate_conditions(run_conditions, trim_aircraft)
+ CL_12_deg_flap = results_flap.aerodynamics.lift_coefficient[0,0]
+
+ # critera
+ flap_criteria = (CL_12_deg_flap-CL_12_deg_no_flap) - 0.95*(CL_12_deg_flap-CL_12_deg_no_flap)
+ # compute control surface area
+ control_surfaces = ['flap']
+ total_control_surface_area = compute_control_surface_areas(control_surfaces,vehicle)
+ summary.flap_surface_area = total_control_surface_area
+ summary.flap_criteria = flap_criteria
+
+ print("Flap Area : " + str(summary.flap_surface_area))
+ print("Flap Criteria : " + str(flap_criteria)) # https://aviation.stackexchange.com/questions/48715/how-is-the-area-of-flaps-determined
+ print("\n\n")
+
+ return nexus
+
+
+def compute_control_surface_areas(control_surfaces,vehicle):
+ '''
+ This function computes the control suface area used in the objectives of the
+ control surface sizing scripts
+ '''
+ total_control_surface_area = 0
+ for cs_idx in range(len(control_surfaces)):
+ for wing in vehicle.wings:
+ if getattr(wing,'control_surfaces',False):
+ for CS in wing.control_surfaces:
+ if CS.tag == control_surfaces[cs_idx]:
+ if wing.Segments:
+ num_segs = len(wing.Segments)
+ for seg_id in range(num_segs-1):
+ if (CS.span_fraction_start >= wing.Segments[seg_id].percent_span_location) \
+ and (CS.span_fraction_end <= wing.Segments[seg_id+1].percent_span_location):
+ root_chord = wing.Segments[seg_id].root_chord_percent*wing.chords.root
+ tip_chord = wing.Segments[seg_id+1].root_chord_percent*wing.chords.root
+ span = (wing.Segments[seg_id+1].percent_span_location-wing.Segments[seg_id].percent_span_location)*wing.spans.projected
+ rel_start_percent_span = CS.span_fraction_start - wing.Segments[seg_id].percent_span_location
+ rel_end_percent_span = CS.span_fraction_end - wing.Segments[seg_id].percent_span_location
+ chord_fraction = CS.chord_fraction
+ area = conpute_control_surface_area(root_chord,tip_chord,span,rel_start_percent_span,rel_end_percent_span,chord_fraction)
+ total_control_surface_area += area
+ else:
+ root_chord = wing.chords.root
+ tip_chord = wing.chords.tip
+ span = wing.spans.projected
+ rel_start_percent_span = CS.span_fraction_start
+ rel_end_percent_span = CS.span_fraction_end
+ chord_fraction = CS.chord_fraction
+ area = conpute_control_surface_area(root_chord,tip_chord,span,rel_start_percent_span,rel_end_percent_span,chord_fraction)
+ total_control_surface_area += area
+
+ return total_control_surface_area
+
+def conpute_control_surface_area(root_chord,tip_chord,span,rel_start_percent_span,rel_end_percent_span,chord_fraction):
+ '''
+ This is a simple function that computes the area of a single control surface
+ '''
+ cs_start_chord = (root_chord + ((tip_chord-root_chord)/span)*(rel_start_percent_span*span))*chord_fraction
+ cs_end_chord = (root_chord + ((tip_chord-root_chord)/span)*(rel_end_percent_span*span))*chord_fraction
+ cs_span = (rel_end_percent_span-rel_start_percent_span)*span
+ cs_area = 0.5*(cs_start_chord+cs_end_chord)*cs_span
+ return cs_area
+
diff --git a/tut_Control_Surface_Sizing_Navion/Vehicles.py b/tut_Control_Surface_Sizing_Navion/Vehicles.py
new file mode 100644
index 0000000..f2d1dd7
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/Vehicles.py
@@ -0,0 +1,621 @@
+# Vehicle.py
+
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+
+import SUAVE
+from SUAVE.Core import Units , Data
+import numpy as np
+
+from copy import deepcopy
+from SUAVE.Plots.Performance.Mission_Plots import *
+from SUAVE.Plots.Geometry import *
+from SUAVE.Components.Energy.Networks.Battery_Propeller import Battery_Propeller
+from SUAVE.Methods.Propulsion import propeller_design
+from SUAVE.Methods.Propulsion.electric_motor_sizing import size_optimal_motor
+from SUAVE.Methods.Power.Battery.Sizing import initialize_from_mass
+from SUAVE.Methods.Geometry.Two_Dimensional.Planform import segment_properties
+
+# ----------------------------------------------------------------------
+# Define the Vehicle
+# ----------------------------------------------------------------------
+
+def stick_fixed_stability_setup():
+ vehicle = vehicle_setup()
+ configs = stick_fixed_stability_configs_setup(vehicle)
+ return configs
+
+
+def elevator_sizing_setup(vehicle):
+ hs_wing = vehicle.wings.horizontal_stabilizer
+ elevator = SUAVE.Components.Wings.Control_Surfaces.Elevator()
+ elevator.tag = 'elevator'
+ elevator.span_fraction_start = 0.1
+ elevator.span_fraction_end = 0.9
+ elevator.deflection = 0.0 * Units.deg
+ elevator.chord_fraction = 0.3
+ hs_wing.append_control_surface(elevator)
+ configs = elevator_sizing_configs_setup(vehicle)
+ return configs
+
+def aileron_rudder_sizing_setup(vehicle):
+ mw_wing = vehicle.wings.main_wing
+ aileron = SUAVE.Components.Wings.Control_Surfaces.Aileron()
+ aileron.tag = 'aileron'
+ aileron.span_fraction_start = 0.7
+ aileron.span_fraction_end = 0.9
+ aileron.deflection = 0.0 * Units.degrees
+ aileron.chord_fraction = 0.2
+ mw_wing.append_control_surface(aileron)
+
+ if vehicle.rudder_flag:
+ vs_wing = vehicle.wings.vertical_stabilizer
+ rudder = SUAVE.Components.Wings.Control_Surfaces.Rudder()
+ rudder.tag = 'rudder'
+ rudder.span_fraction_start = 0.2
+ rudder.span_fraction_end = 0.8
+ rudder.deflection = 0.0 * Units.deg
+ rudder.chord_fraction = 0.2
+ vs_wing.append_control_surface(rudder)
+
+ configs = aileron_rudder_sizing_configs_setup(vehicle)
+ return configs
+
+def flap_sizing_setup(vehicle):
+ mw_wing = vehicle.wings.main_wing
+ flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
+ flap.tag = 'flap'
+ flap.span_fraction_start = 0.2
+ flap.span_fraction_end = 0.5
+ flap.deflection = 0.0 * Units.degrees
+ flap.chord_fraction = 0.20
+ mw_wing.append_control_surface(flap)
+ configs = flap_sizing_configs_setup(vehicle)
+ return configs
+
+# ----------------------------------------------------------------------
+# Define the Configurations
+# ---------------------------------------------------------------------
+
+def stick_fixed_stability_configs_setup(vehicle):
+ configs = SUAVE.Components.Configs.Config.Container()
+ base_config = SUAVE.Components.Configs.Config(vehicle)
+ config = SUAVE.Components.Configs.Config(base_config)
+ config.tag = 'stick_fixed_cruise'
+ configs.append(config)
+ return configs
+
+def elevator_sizing_configs_setup(vehicle):
+ configs = SUAVE.Components.Configs.Config.Container()
+ base_config = SUAVE.Components.Configs.Config(vehicle)
+ config = SUAVE.Components.Configs.Config(base_config)
+ config.tag = 'elevator_sizing'
+ configs.append(config)
+ return configs
+
+def aileron_rudder_sizing_configs_setup(vehicle):
+ configs = SUAVE.Components.Configs.Config.Container()
+ base_config = SUAVE.Components.Configs.Config(vehicle)
+ config = SUAVE.Components.Configs.Config(base_config)
+ config.tag = 'aileron_rudder_sizing'
+ configs.append(config)
+ return configs
+
+def flap_sizing_configs_setup(vehicle):
+ configs = SUAVE.Components.Configs.Config.Container()
+ base_config = SUAVE.Components.Configs.Config(vehicle)
+ config = SUAVE.Components.Configs.Config(base_config)
+ config.tag = 'flap_sizing'
+ configs.append(config)
+ return configs
+
+
+# ----------------------------------------------------------------------
+# Define Vehicle
+# ---------------------------------------------------------------------
+def vehicle_setup():
+
+ '''
+ This function defines the base vehicle including
+ 1) center of gravity (either hard coded or use suave's built in function)
+ 2) mass moment of interita (optional)
+
+ Key Notes:
+ 1) The wing that is intended to be the main must be given the tag "main wing". This wing will be used to append
+ a flap and an aileron
+
+ 2) If present, the wing that is intended to be the horizontal stabilizer must be given the tag "horizontal_stabilizer"
+ This wing will be used to append an elevator
+
+ 3) If present, The wing that is intended to be the vertical stabilizer must be given the tag "vertical_stabilizer"
+ This wing will be used to append a rudder (optional)
+
+
+ '''
+ # ------------------------------------------------------------------
+ # Initialize the Vehicle
+ # ------------------------------------------------------------------
+ vehicle = SUAVE.Vehicle()
+ vehicle.tag = 'X57_Mod2'
+
+
+ # ------------------------------------------------------------------
+ # Vehicle-level Properties
+ # ------------------------------------------------------------------
+
+ # mass properties
+ vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
+ vehicle.mass_properties.takeoff = 2550. * Units.pounds
+ vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
+ vehicle.mass_properties.moments_of_inertia.tensor = np.array([[164627.7,0.0,0.0],[0.0,471262.4,0.0],[0.0,0.0,554518.7]]) # Navion's
+ vehicle.envelope.ultimate_load = 5.7
+ vehicle.envelope.limit_load = 3.8
+ vehicle.reference_area = 14.76
+ vehicle.passengers = 4
+ vehicle.systems.control = "fully powered"
+ vehicle.systems.accessories = "commuter"
+
+ cruise_speed = 135.*Units['mph']
+ altitude = 2500. * Units.ft
+ atmo = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ freestream = atmo.compute_values (0.)
+ freestream0 = atmo.compute_values (altitude)
+ mach_number = (cruise_speed/freestream.speed_of_sound)[0][0]
+ vehicle.design_dynamic_pressure = ( .5 *freestream0.density*(cruise_speed*cruise_speed))[0][0]
+ vehicle.design_mach_number = mach_number
+
+ # ------------------------------------------------------------------
+ # Main Wing
+ # ------------------------------------------------------------------
+ wing = SUAVE.Components.Wings.Main_Wing()
+ wing.tag = 'main_wing'
+ wing.sweeps.quarter_chord = 0.0 * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 14.76
+ wing.spans.projected = 11.4
+ wing.chords.root = 1.46
+ wing.chords.tip = 0.92
+ wing.chords.mean_aerodynamic = 1.19
+ wing.taper = wing.chords.root/wing.chords.tip
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 3.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[2.93, 0., 1.01]]
+ wing.aerodynamic_center = [3., 0., 1.01]
+ wing.vertical = False
+ wing.symmetric = True
+ wing.high_lift = True
+ wing.winglet_fraction = 0.0
+ wing.dynamic_pressure_ratio = 1.0
+ airfoil = SUAVE.Components.Airfoils.Airfoil()
+ airfoil.coordinate_file = 'Airfoils/NACA_63_412.txt'
+
+ cg_x = wing.origin[0][0] + 0.25*wing.chords.mean_aerodynamic
+ cg_z = wing.origin[0][2] - 0.2*wing.chords.mean_aerodynamic
+ vehicle.mass_properties.center_of_gravity = [[cg_x, 0. , cg_z ]] # SOURCE: Design and aerodynamic analysis of a twin-engine commuter aircraft
+
+ # Wing Segments
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'inboard'
+ segment.percent_span_location = 0.0
+ segment.twist = 3. * Units.degrees
+ segment.root_chord_percent = 1.
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'outboard'
+ segment.percent_span_location = 0.5438
+ segment.twist = 2.* Units.degrees
+ segment.root_chord_percent = 1.
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ # Wing Segments
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'winglet'
+ segment.percent_span_location = 0.98
+ segment.twist = 1. * Units.degrees
+ segment.root_chord_percent = 0.630
+ segment.dihedral_outboard = 75. * Units.degrees
+ segment.sweeps.quarter_chord = 15. * Units.degrees
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'tip'
+ segment.percent_span_location = 1.
+ segment.twist = 0. * Units.degrees
+ segment.root_chord_percent = 0.12
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ # Fill out more segment properties automatically
+ wing = segment_properties(wing)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Horizontal Stabilizer
+ # ------------------------------------------------------------------
+ wing = SUAVE.Components.Wings.Wing()
+ wing.tag = 'horizontal_stabilizer'
+ wing.sweeps.quarter_chord = 0.0 * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 2.540
+ wing.spans.projected = 3.3 * Units.meter
+ wing.sweeps.quarter_chord = 0 * Units.deg
+ wing.chords.root = 0.769 * Units.meter
+ wing.chords.tip = 0.769 * Units.meter
+ wing.chords.mean_aerodynamic = 0.769 * Units.meter
+ wing.taper = 1.
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[7.7, 0., 0.25]]
+ wing.aerodynamic_center = [7.8, 0., 0.25]
+ wing.vertical = False
+ wing.winglet_fraction = 0.0
+ wing.symmetric = True
+ wing.high_lift = False
+ 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.sweeps.quarter_chord = 25. * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 2.258 * Units['meters**2']
+ wing.spans.projected = 1.854 * Units.meter
+ wing.chords.root = 1.6764 * Units.meter
+ wing.chords.tip = 0.6858 * Units.meter
+ wing.chords.mean_aerodynamic = 1.21 * Units.meter
+ wing.taper = wing.chords.tip/wing.chords.root
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[6.75 ,0, 0.0]]
+ wing.aerodynamic_center = [0.508 ,0,0]
+ wing.vertical = True
+ wing.symmetric = False
+ wing.t_tail = False
+ wing.winglet_fraction = 0.0
+ wing.dynamic_pressure_ratio = 1.0
+
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Fuselage
+ # ------------------------------------------------------------------
+ fuselage = SUAVE.Components.Fuselages.Fuselage()
+ fuselage.tag = 'fuselage'
+ fuselage.seats_abreast = 2.
+ fuselage.fineness.nose = 1.6
+ fuselage.fineness.tail = 2.
+ fuselage.lengths.nose = 60. * Units.inches
+ fuselage.lengths.tail = 161. * Units.inches
+ fuselage.lengths.cabin = 105. * Units.inches
+ fuselage.lengths.total = 332.2* Units.inches
+ fuselage.lengths.fore_space = 0.
+ fuselage.lengths.aft_space = 0.
+ fuselage.width = 42. * Units.inches
+ fuselage.heights.maximum = 62. * Units.inches
+ fuselage.heights.at_quarter_length = 62. * Units.inches
+ fuselage.heights.at_three_quarters_length = 62. * Units.inches
+ fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
+ fuselage.areas.side_projected = 8000. * Units.inches**2.
+ fuselage.areas.wetted = 30000. * Units.inches**2.
+ fuselage.areas.front_projected = 42.* 62. * Units.inches**2.
+ fuselage.effective_diameter = 50. * Units.inches
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_0'
+ segment.percent_x_location = 0
+ segment.percent_z_location = 0
+ segment.height = 0.01
+ segment.width = 0.01
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_1'
+ segment.percent_x_location = 0.007279116466
+ segment.percent_z_location = 0.002502014453
+ segment.height = 0.1669064748
+ segment.width = 0.2780205877
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_2'
+ segment.percent_x_location = 0.01941097724
+ segment.percent_z_location = 0.001216095397
+ segment.height = 0.3129496403
+ segment.width = 0.4365777215
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_3'
+ segment.percent_x_location = 0.06308567604
+ segment.percent_z_location = 0.007395489231
+ segment.height = 0.5841726619
+ segment.width = 0.6735119903
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_4'
+ segment.percent_x_location = 0.1653761217
+ segment.percent_z_location = 0.02891281352
+ segment.height = 1.064028777
+ segment.width = 1.067200529
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_5'
+ segment.percent_x_location = 0.2426372155
+ segment.percent_z_location = 0.04214148761
+ segment.height = 1.293766653
+ segment.width = 1.183058255
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_6'
+ segment.percent_x_location = 0.2960174029
+ segment.percent_z_location = 0.04705241831
+ segment.height = 1.377026712
+ segment.width = 1.181540054
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_7'
+ segment.percent_x_location = 0.3809404284
+ segment.percent_z_location = 0.05313580461
+ segment.height = 1.439568345
+ segment.width = 1.178218989
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_8'
+ segment.percent_x_location = 0.5046854083
+ segment.percent_z_location = 0.04655492473
+ segment.height = 1.29352518
+ segment.width = 1.054390707
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_9'
+ segment.percent_x_location = 0.6454149933
+ segment.percent_z_location = 0.03741966266
+ segment.height = 0.8971223022
+ segment.width = 0.8501926505
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_10'
+ segment.percent_x_location = 0.985107095
+ segment.percent_z_location = 0.04540283436
+ segment.height = 0.2920863309
+ segment.width = 0.2012565415
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_11'
+ segment.percent_x_location = 1
+ segment.percent_z_location = 0.04787575562
+ segment.height = 0.1251798561
+ segment.width = 0.1206021048
+ fuselage.Segments.append(segment)
+
+ # add to vehicle
+ vehicle.append_component(fuselage)
+
+ # ------------------------------------------------------------------
+ # Nacelles
+ # ------------------------------------------------------------------
+ nacelle = SUAVE.Components.Nacelles.Nacelle()
+ nacelle.tag = 'nacelle_1'
+ nacelle.length = 2
+ nacelle.diameter = 42 * Units.inches
+ nacelle.areas.wetted = 0.01*(2*np.pi*0.01/2)
+ nacelle.origin = [[2.5,2.5,1.0]]
+ nacelle.flow_through = False
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_1'
+ nac_segment.percent_x_location = 0.0
+ nac_segment.height = 0.0
+ nac_segment.width = 0.0
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_2'
+ nac_segment.percent_x_location = 0.1
+ nac_segment.height = 0.5
+ nac_segment.width = 0.65
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_3'
+ nac_segment.percent_x_location = 0.3
+ nac_segment.height = 0.52
+ nac_segment.width = 0.7
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_4'
+ nac_segment.percent_x_location = 0.5
+ nac_segment.height = 0.5
+ nac_segment.width = 0.65
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_5'
+ nac_segment.percent_x_location = 0.7
+ nac_segment.height = 0.4
+ nac_segment.width = 0.6
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_6'
+ nac_segment.percent_x_location = 0.9
+ nac_segment.height = 0.3
+ nac_segment.width = 0.5
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_7'
+ nac_segment.percent_x_location = 1.0
+ nac_segment.height = 0.0
+ nac_segment.width = 0.0
+ nacelle.append_segment(nac_segment)
+
+ vehicle.append_component(nacelle)
+
+ nacelle_2 = deepcopy(nacelle)
+ nacelle_2.tag = 'nacelle_2'
+ nacelle_2.origin = [[2.5,-2.5,1.0]]
+ vehicle.append_component(nacelle_2)
+
+ #---------------------------------------------------------------------------------------------
+ # DEFINE PROPELLER
+ #---------------------------------------------------------------------------------------------
+ # build network
+ net = Battery_Propeller()
+ net.number_of_propeller_engines = 2.
+ net.identical_propellers = True
+
+ # Component 1 the ESC
+ esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
+ esc.efficiency = 0.95 # Gundlach for brushless motors
+ net.esc = esc
+
+ # Component 2 the Propeller
+ prop = SUAVE.Components.Energy.Converters.Propeller()
+ prop.tag = 'propeller_1'
+ prop.number_of_blades = 2.0
+ prop.freestream_velocity = 135.*Units['mph']
+ prop.angular_velocity = 1300. * Units.rpm
+ prop.tip_radius = 76./2. * Units.inches
+ prop.hub_radius = 8. * Units.inches
+ prop.design_Cl = 0.8
+ prop.design_altitude = 12000. * Units.feet
+ prop.design_altitude = 12000. * Units.feet
+ prop.design_thrust = 1200.
+ prop.origin = [[2.,2.5,0.784]]
+ prop.rotation = -1
+ prop.symmetry = True
+ prop.variable_pitch = True
+ prop.airfoil_geometry = ['Airfoils/NACA_4412.txt']
+ prop.airfoil_polars = [['Airfoils/Polars/NACA_4412_polar_Re_50000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_100000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_200000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_500000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_1000000.txt' ]]
+
+ prop.airfoil_polar_stations = [0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0]
+ prop = propeller_design(prop)
+
+ prop_left = deepcopy(prop)
+ prop_left.tag = 'propeller_2'
+ prop_left.origin = [[2.,-2.5,0.784]]
+ prop_left.rotation = 1
+
+ net.propellers.append(prop)
+ net.propellers.append(prop_left)
+
+
+ # Component 3 the Battery
+ bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiFePO4_18650()
+
+ bat.mass_properties.mass = 500. * Units.kg
+ bat.max_voltage = 500.
+ initialize_from_mass(bat)
+
+ # Assume a battery pack module shape. This step is optional but
+ # required for thermal analysis of the pack
+ number_of_modules = 10
+ bat.module_config.total = int(np.ceil(bat.pack_config.total/number_of_modules))
+ bat.module_config.normal_count = int(np.ceil(bat.module_config.total/bat.pack_config.series))
+ bat.module_config.parallel_count = int(np.ceil(bat.module_config.total/bat.pack_config.parallel))
+ net.battery = bat
+
+ net.battery = bat
+ net.voltage = bat.max_voltage
+
+ # Component 4 Miscellaneous Systems
+ sys = SUAVE.Components.Systems.System()
+ sys.mass_properties.mass = 5 # kg
+
+ # Component 5 the Motor
+ motor = SUAVE.Components.Energy.Converters.Motor()
+ motor.efficiency = 0.95
+ motor.gearbox_efficiency = 1.
+ motor.origin = [[2., 2.5, 0.784]]
+ motor.nominal_voltage = bat.max_voltage *3/4
+ motor.propeller_radius = prop.tip_radius
+ motor.no_load_current = 4.0
+ motor = size_optimal_motor(motor,prop)
+ motor.mass_properties.mass = 10. * Units.kg
+
+ # append right motor
+ net.propeller_motors.append(motor)
+
+ # append left motor
+ motor_left = deepcopy(motor)
+ motor_left.origin = [[2., -2.5, 0.784]]
+ net.propeller_motors.append(motor_left)
+
+ # Component 6 the Payload
+ payload = SUAVE.Components.Energy.Peripherals.Payload()
+ payload.power_draw = 10. # Watts
+ payload.mass_properties.mass = 1.0 * Units.kg
+ net.payload = payload
+
+ # Component 7 the Avionics
+ avionics = SUAVE.Components.Energy.Peripherals.Avionics()
+ avionics.power_draw = 20. # Watts
+ net.avionics = avionics
+
+ # add the solar network to the vehicle
+ vehicle.append_component(net)
+
+ # ------------------------------------------------------------------
+ # Vehicle Definition Complete
+ # ------------------------------------------------------------------
+
+ return vehicle
+
+
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/Whisper_Drone.avl b/tut_Control_Surface_Sizing_Navion/avl_files/Whisper_Drone.avl
new file mode 100644
index 0000000..a80f595
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/Whisper_Drone.avl
@@ -0,0 +1,301 @@
+base
+
+#Mach
+ 0.0
+
+#Iysym IZsym Zsym
+ 0 0 0.0
+
+#Sref Cref Bref <meters>
+0.929030400000002 0.3875547628787364 2.4291211518269122
+
+#Xref Yref Zref <meters>
+0.2 0.0 0.0
+
+
+
+#---------------------------------------------------------
+SURFACE
+propulsor_pylon
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0 20 -1.1
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1219 0.0 -0.0732 0.1334 0.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1787 0.0 0.0253 0.1143 0.0
+
+
+#---------------------------------------------------------
+SURFACE
+012m_htailnosubsurfaces
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0 20 1.0
+
+YDUPLICATE
+0.0
+
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.6477 1.2192 0.0945 0.3048 0.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.8028 1.8288 0.4464 0.1143 0.0
+
+
+#---------------------------------------------------------
+SURFACE
+012m_vtailnosubsurfaces_1
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0 20 -1.1
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.6477 -1.2192 0.0792 0.3048 0.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.7454 -1.2192 0.2012 0.1524 0.0
+
+
+#---------------------------------------------------------
+SURFACE
+012m_vtailnosubsurfaces_2
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0 20 -1.1
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.6477 1.2192 0.0792 0.3048 0.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.7454 1.2192 0.2012 0.1524 0.0
+
+
+#---------------------------------------------------------
+SURFACE
+main_wing
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0 20 1.0
+
+YDUPLICATE
+0.0
+
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.277 0.0 0.0 0.4589 0.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.0628 1.2146 0.1707 0.306 0.0
+
+
+#---------------------------------------------------------
+SURFACE
+Booms_1_horizontal
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 1.1811 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 1.1887 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 1.1963 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 1.204 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 1.2116 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.0914 1.2192 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 1.2268 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 1.2344 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 1.2421 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 1.2497 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 1.2573 0.093 nan 0.0 1 0
+
+
+#---------------------------------------------------------
+SURFACE
+Booms_1_vertical
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 1.2192 0.0549 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 1.2192 0.0625 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 1.2192 0.0701 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 1.2192 0.0777 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 1.2192 0.0853 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.0914 1.2192 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 1.2192 0.1006 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 1.2192 0.1082 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 1.2192 0.1158 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 1.2192 0.1234 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 1.2192 0.1311 nan 0.0 1 0
+
+
+#---------------------------------------------------------
+SURFACE
+Booms_2_horizontal
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 -1.2573 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 -1.2497 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 -1.2421 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 -1.2344 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 -1.2268 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.0914 -1.2192 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 -1.2116 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 -1.204 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 -1.1963 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 -1.1887 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 -1.1811 0.093 nan 0.0 1 0
+
+
+#---------------------------------------------------------
+SURFACE
+Booms_2_vertical
+#Nchordwise Cspace Nspanwise Sspace
+10 1.0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 -1.2192 0.0549 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 -1.2192 0.0625 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 -1.2192 0.0701 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 -1.2192 0.0777 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 -1.2192 0.0853 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.0914 -1.2192 0.093 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+-0.039 -1.2192 0.1006 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0252 -1.2192 0.1082 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.0967 -1.2192 0.1158 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.1761 -1.2192 0.1234 nan 0.0 1 0
+
+SECTION
+#Xle Yle Zle Chord Ainc Nspanwise Sspace
+0.2734 -1.2192 0.1311 nan 0.0 1 0
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/Whisper_Drone.mass b/tut_Control_Surface_Sizing_Navion/avl_files/Whisper_Drone.mass
new file mode 100644
index 0000000..6089ef7
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/Whisper_Drone.mass
@@ -0,0 +1,32 @@
+
+#-------------------------------------------------
+# Whisper_Drone
+#
+# Dimensional unit and parameter data.
+# Mass & Inertia breakdown.
+#-------------------------------------------------
+
+# Names and scalings for units to be used for trim and eigenmode calculations.
+# The Lunit and Munit values scale the mass, xyz, and inertia table data below.
+# Lunit value will also scale all lengths and areas in the AVL input file.
+Lunit = 1.0 m
+Munit = 1.0 kg
+Tunit = 1.0 s
+
+#-------------------------
+# Gravity and density to be used as default values in trim setup.
+# Must be in the units given above.
+g = 9.81
+rho = 1.2250000002007604
+
+#-------------------------
+# Mass & Inertia breakdown.
+# x y z is location of item's own CG.
+# Ixx... are item's inertias about item's own CG.
+#
+# x,y,z system here must be exactly the same one used in the AVL input file
+# (same orientation, same origin location, same length units)
+#
+# mass x y z Ixx Iyy Izz Component Name
+#
+ 24.947580350000003 0.2 0.0 0.0 0.0 0.0 0.0 ! Whisper_Drone
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/batch_01.run b/tut_Control_Surface_Sizing_Navion/avl_files/batch_01.run
new file mode 100644
index 0000000..76f4b2f
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/batch_01.run
@@ -0,0 +1,252 @@
+
+
+ ---------------------------------------------
+ Run case 1: case_01_01
+
+ alpha -> alpha = -2.0
+ beta -> beta = 0.0
+ pb/2V -> pb/2V = 0.00000
+ qc/2V -> qc/2V = 0.00000
+ rb/2V -> rb/2V = 0.00000
+
+ alpha = 0.00000 deg
+ beta = 0.00000 deg
+ pb/2V = 0.00000
+ qc/2V = 0.00000
+ rb/2V = 0.00000
+ CL = 0.00000
+ CDo = 0.0
+ bank = 0.00000 deg
+ elevation = 0.00000 deg
+ heading = 0.00000 deg
+ Mach = 0.05
+ velocity = 0.0 m/s
+ density = 1.225 kg/m^3
+ grav.acc. = 9.81 m/s^2
+ turn_rad. = 0.00000 m
+ load_fac. = 0.00000
+ X_cg = 0.2 m
+ Y_cg = 0.0 m
+ Z_cg = 0.0 m
+ mass = 0.0 kg
+ Ixx = 0.0 kg-m^2
+ Iyy = 0.0 kg-m^2
+ Izz = 0.0 kg-m^2
+ Ixy = 0.0 kg-m^2
+ Iyz = 0.0 kg-m^2
+ Izx = 0.0 kg-m^2
+ visc CL_a = 0.00000
+ visc CL_u = 0.00000
+ visc CM_a = 0.00000
+ visc CM_u = 0.00000
+
+
+
+ ---------------------------------------------
+ Run case 2: case_01_02
+
+ alpha -> alpha = 0.0
+ beta -> beta = 0.0
+ pb/2V -> pb/2V = 0.00000
+ qc/2V -> qc/2V = 0.00000
+ rb/2V -> rb/2V = 0.00000
+
+ alpha = 0.00000 deg
+ beta = 0.00000 deg
+ pb/2V = 0.00000
+ qc/2V = 0.00000
+ rb/2V = 0.00000
+ CL = 0.00000
+ CDo = 0.0
+ bank = 0.00000 deg
+ elevation = 0.00000 deg
+ heading = 0.00000 deg
+ Mach = 0.05
+ velocity = 0.0 m/s
+ density = 1.225 kg/m^3
+ grav.acc. = 9.81 m/s^2
+ turn_rad. = 0.00000 m
+ load_fac. = 0.00000
+ X_cg = 0.2 m
+ Y_cg = 0.0 m
+ Z_cg = 0.0 m
+ mass = 0.0 kg
+ Ixx = 0.0 kg-m^2
+ Iyy = 0.0 kg-m^2
+ Izz = 0.0 kg-m^2
+ Ixy = 0.0 kg-m^2
+ Iyz = 0.0 kg-m^2
+ Izx = 0.0 kg-m^2
+ visc CL_a = 0.00000
+ visc CL_u = 0.00000
+ visc CM_a = 0.00000
+ visc CM_u = 0.00000
+
+
+
+ ---------------------------------------------
+ Run case 3: case_01_03
+
+ alpha -> alpha = 2.0
+ beta -> beta = 0.0
+ pb/2V -> pb/2V = 0.00000
+ qc/2V -> qc/2V = 0.00000
+ rb/2V -> rb/2V = 0.00000
+
+ alpha = 0.00000 deg
+ beta = 0.00000 deg
+ pb/2V = 0.00000
+ qc/2V = 0.00000
+ rb/2V = 0.00000
+ CL = 0.00000
+ CDo = 0.0
+ bank = 0.00000 deg
+ elevation = 0.00000 deg
+ heading = 0.00000 deg
+ Mach = 0.05
+ velocity = 0.0 m/s
+ density = 1.225 kg/m^3
+ grav.acc. = 9.81 m/s^2
+ turn_rad. = 0.00000 m
+ load_fac. = 0.00000
+ X_cg = 0.2 m
+ Y_cg = 0.0 m
+ Z_cg = 0.0 m
+ mass = 0.0 kg
+ Ixx = 0.0 kg-m^2
+ Iyy = 0.0 kg-m^2
+ Izz = 0.0 kg-m^2
+ Ixy = 0.0 kg-m^2
+ Iyz = 0.0 kg-m^2
+ Izx = 0.0 kg-m^2
+ visc CL_a = 0.00000
+ visc CL_u = 0.00000
+ visc CM_a = 0.00000
+ visc CM_u = 0.00000
+
+
+
+ ---------------------------------------------
+ Run case 4: case_01_04
+
+ alpha -> alpha = 5.0
+ beta -> beta = 0.0
+ pb/2V -> pb/2V = 0.00000
+ qc/2V -> qc/2V = 0.00000
+ rb/2V -> rb/2V = 0.00000
+
+ alpha = 0.00000 deg
+ beta = 0.00000 deg
+ pb/2V = 0.00000
+ qc/2V = 0.00000
+ rb/2V = 0.00000
+ CL = 0.00000
+ CDo = 0.0
+ bank = 0.00000 deg
+ elevation = 0.00000 deg
+ heading = 0.00000 deg
+ Mach = 0.05
+ velocity = 0.0 m/s
+ density = 1.225 kg/m^3
+ grav.acc. = 9.81 m/s^2
+ turn_rad. = 0.00000 m
+ load_fac. = 0.00000
+ X_cg = 0.2 m
+ Y_cg = 0.0 m
+ Z_cg = 0.0 m
+ mass = 0.0 kg
+ Ixx = 0.0 kg-m^2
+ Iyy = 0.0 kg-m^2
+ Izz = 0.0 kg-m^2
+ Ixy = 0.0 kg-m^2
+ Iyz = 0.0 kg-m^2
+ Izx = 0.0 kg-m^2
+ visc CL_a = 0.00000
+ visc CL_u = 0.00000
+ visc CM_a = 0.00000
+ visc CM_u = 0.00000
+
+
+
+ ---------------------------------------------
+ Run case 5: case_01_05
+
+ alpha -> alpha = 7.0
+ beta -> beta = 0.0
+ pb/2V -> pb/2V = 0.00000
+ qc/2V -> qc/2V = 0.00000
+ rb/2V -> rb/2V = 0.00000
+
+ alpha = 0.00000 deg
+ beta = 0.00000 deg
+ pb/2V = 0.00000
+ qc/2V = 0.00000
+ rb/2V = 0.00000
+ CL = 0.00000
+ CDo = 0.0
+ bank = 0.00000 deg
+ elevation = 0.00000 deg
+ heading = 0.00000 deg
+ Mach = 0.05
+ velocity = 0.0 m/s
+ density = 1.225 kg/m^3
+ grav.acc. = 9.81 m/s^2
+ turn_rad. = 0.00000 m
+ load_fac. = 0.00000
+ X_cg = 0.2 m
+ Y_cg = 0.0 m
+ Z_cg = 0.0 m
+ mass = 0.0 kg
+ Ixx = 0.0 kg-m^2
+ Iyy = 0.0 kg-m^2
+ Izz = 0.0 kg-m^2
+ Ixy = 0.0 kg-m^2
+ Iyz = 0.0 kg-m^2
+ Izx = 0.0 kg-m^2
+ visc CL_a = 0.00000
+ visc CL_u = 0.00000
+ visc CM_a = 0.00000
+ visc CM_u = 0.00000
+
+
+
+ ---------------------------------------------
+ Run case 6: case_01_06
+
+ alpha -> alpha = 10.0
+ beta -> beta = 0.0
+ pb/2V -> pb/2V = 0.00000
+ qc/2V -> qc/2V = 0.00000
+ rb/2V -> rb/2V = 0.00000
+
+ alpha = 0.00000 deg
+ beta = 0.00000 deg
+ pb/2V = 0.00000
+ qc/2V = 0.00000
+ rb/2V = 0.00000
+ CL = 0.00000
+ CDo = 0.0
+ bank = 0.00000 deg
+ elevation = 0.00000 deg
+ heading = 0.00000 deg
+ Mach = 0.05
+ velocity = 0.0 m/s
+ density = 1.225 kg/m^3
+ grav.acc. = 9.81 m/s^2
+ turn_rad. = 0.00000 m
+ load_fac. = 0.00000
+ X_cg = 0.2 m
+ Y_cg = 0.0 m
+ Z_cg = 0.0 m
+ mass = 0.0 kg
+ Ixx = 0.0 kg-m^2
+ Iyy = 0.0 kg-m^2
+ Izz = 0.0 kg-m^2
+ Ixy = 0.0 kg-m^2
+ Iyz = 0.0 kg-m^2
+ Izx = 0.0 kg-m^2
+ visc CL_a = 0.00000
+ visc CL_u = 0.00000
+ visc CM_a = 0.00000
+ visc CM_u = 0.00000
+
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_01.txt b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_01.txt
new file mode 100644
index 0000000..b5d6aea
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_01.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_01
+
+ Alpha = -2.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Geometry-axis derivatives...
+
+ axial vel. u sideslip vel. v normal vel. w
+ ---------------- ---------------- ----------------
+ x force CX | CXu = NaN CXv = NaN CXw = NaN
+ y force CY | CYu = NaN CYv = NaN CYw = NaN
+ z force CZ | CZu = NaN CZv = NaN CZw = NaN
+ x mom. Cl | Clu = NaN Clv = NaN Clw = NaN
+ y mom. Cm | Cmu = NaN Cmv = NaN Cmw = NaN
+ z mom. Cn | Cnu = NaN Cnv = NaN Cnw = NaN
+
+ roll rate p pitch rate q yaw rate r
+ ---------------- ---------------- ----------------
+ x force CX | CXp = NaN CXq = NaN CXr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ z force CZ | CZp = NaN CZq = NaN CZr = NaN
+ x mom. Cl | Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z mom. Cn | Cnp = NaN Cnq = NaN Cnr = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_02.txt b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_02.txt
new file mode 100644
index 0000000..b213cbe
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_02.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_02
+
+ Alpha = 0.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Geometry-axis derivatives...
+
+ axial vel. u sideslip vel. v normal vel. w
+ ---------------- ---------------- ----------------
+ x force CX | CXu = NaN CXv = NaN CXw = NaN
+ y force CY | CYu = NaN CYv = NaN CYw = NaN
+ z force CZ | CZu = NaN CZv = NaN CZw = NaN
+ x mom. Cl | Clu = NaN Clv = NaN Clw = NaN
+ y mom. Cm | Cmu = NaN Cmv = NaN Cmw = NaN
+ z mom. Cn | Cnu = NaN Cnv = NaN Cnw = NaN
+
+ roll rate p pitch rate q yaw rate r
+ ---------------- ---------------- ----------------
+ x force CX | CXp = NaN CXq = NaN CXr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ z force CZ | CZp = NaN CZq = NaN CZr = NaN
+ x mom. Cl | Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z mom. Cn | Cnp = NaN Cnq = NaN Cnr = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_03.txt b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_03.txt
new file mode 100644
index 0000000..b9cc531
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_03.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_03
+
+ Alpha = 2.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Geometry-axis derivatives...
+
+ axial vel. u sideslip vel. v normal vel. w
+ ---------------- ---------------- ----------------
+ x force CX | CXu = NaN CXv = NaN CXw = NaN
+ y force CY | CYu = NaN CYv = NaN CYw = NaN
+ z force CZ | CZu = NaN CZv = NaN CZw = NaN
+ x mom. Cl | Clu = NaN Clv = NaN Clw = NaN
+ y mom. Cm | Cmu = NaN Cmv = NaN Cmw = NaN
+ z mom. Cn | Cnu = NaN Cnv = NaN Cnw = NaN
+
+ roll rate p pitch rate q yaw rate r
+ ---------------- ---------------- ----------------
+ x force CX | CXp = NaN CXq = NaN CXr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ z force CZ | CZp = NaN CZq = NaN CZr = NaN
+ x mom. Cl | Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z mom. Cn | Cnp = NaN Cnq = NaN Cnr = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_04.txt b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_04.txt
new file mode 100644
index 0000000..460ab3b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_04.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_04
+
+ Alpha = 5.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Geometry-axis derivatives...
+
+ axial vel. u sideslip vel. v normal vel. w
+ ---------------- ---------------- ----------------
+ x force CX | CXu = NaN CXv = NaN CXw = NaN
+ y force CY | CYu = NaN CYv = NaN CYw = NaN
+ z force CZ | CZu = NaN CZv = NaN CZw = NaN
+ x mom. Cl | Clu = NaN Clv = NaN Clw = NaN
+ y mom. Cm | Cmu = NaN Cmv = NaN Cmw = NaN
+ z mom. Cn | Cnu = NaN Cnv = NaN Cnw = NaN
+
+ roll rate p pitch rate q yaw rate r
+ ---------------- ---------------- ----------------
+ x force CX | CXp = NaN CXq = NaN CXr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ z force CZ | CZp = NaN CZq = NaN CZr = NaN
+ x mom. Cl | Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z mom. Cn | Cnp = NaN Cnq = NaN Cnr = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_05.txt b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_05.txt
new file mode 100644
index 0000000..a6d5e8d
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_05.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_05
+
+ Alpha = 7.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Geometry-axis derivatives...
+
+ axial vel. u sideslip vel. v normal vel. w
+ ---------------- ---------------- ----------------
+ x force CX | CXu = NaN CXv = NaN CXw = NaN
+ y force CY | CYu = NaN CYv = NaN CYw = NaN
+ z force CZ | CZu = NaN CZv = NaN CZw = NaN
+ x mom. Cl | Clu = NaN Clv = NaN Clw = NaN
+ y mom. Cm | Cmu = NaN Cmv = NaN Cmw = NaN
+ z mom. Cn | Cnu = NaN Cnv = NaN Cnw = NaN
+
+ roll rate p pitch rate q yaw rate r
+ ---------------- ---------------- ----------------
+ x force CX | CXp = NaN CXq = NaN CXr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ z force CZ | CZp = NaN CZq = NaN CZr = NaN
+ x mom. Cl | Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z mom. Cn | Cnp = NaN Cnq = NaN Cnr = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_06.txt b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_06.txt
new file mode 100644
index 0000000..b08e217
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/body_axis_derivatives_case_01_06.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_06
+
+ Alpha = 10.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Geometry-axis derivatives...
+
+ axial vel. u sideslip vel. v normal vel. w
+ ---------------- ---------------- ----------------
+ x force CX | CXu = NaN CXv = NaN CXw = NaN
+ y force CY | CYu = NaN CYv = NaN CYw = NaN
+ z force CZ | CZu = NaN CZv = NaN CZw = NaN
+ x mom. Cl | Clu = NaN Clv = NaN Clw = NaN
+ y mom. Cm | Cmu = NaN Cmv = NaN Cmw = NaN
+ z mom. Cn | Cnu = NaN Cnv = NaN Cnw = NaN
+
+ roll rate p pitch rate q yaw rate r
+ ---------------- ---------------- ----------------
+ x force CX | CXp = NaN CXq = NaN CXr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ z force CZ | CZp = NaN CZq = NaN CZr = NaN
+ x mom. Cl | Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z mom. Cn | Cnp = NaN Cnq = NaN Cnr = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/commands_01.deck b/tut_Control_Surface_Sizing_Navion/avl_files/commands_01.deck
new file mode 100644
index 0000000..768cf09
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/commands_01.deck
@@ -0,0 +1,70 @@
+MASS Whisper_Drone.mass
+mset
+0
+PLOP
+G
+
+CASE batch_01.run
+OPER
+1
+x
+st
+stability_axis_derivatives_case_01_01.txt
+fn
+surface_forces_case_01_01.txt
+fs
+strip_forces_case_01_01.txt
+sb
+body_axis_derivatives_case_01_01.txt
+2
+x
+st
+stability_axis_derivatives_case_01_02.txt
+fn
+surface_forces_case_01_02.txt
+fs
+strip_forces_case_01_02.txt
+sb
+body_axis_derivatives_case_01_02.txt
+3
+x
+st
+stability_axis_derivatives_case_01_03.txt
+fn
+surface_forces_case_01_03.txt
+fs
+strip_forces_case_01_03.txt
+sb
+body_axis_derivatives_case_01_03.txt
+4
+x
+st
+stability_axis_derivatives_case_01_04.txt
+fn
+surface_forces_case_01_04.txt
+fs
+strip_forces_case_01_04.txt
+sb
+body_axis_derivatives_case_01_04.txt
+5
+x
+st
+stability_axis_derivatives_case_01_05.txt
+fn
+surface_forces_case_01_05.txt
+fs
+strip_forces_case_01_05.txt
+sb
+body_axis_derivatives_case_01_05.txt
+6
+x
+st
+stability_axis_derivatives_case_01_06.txt
+fn
+surface_forces_case_01_06.txt
+fs
+strip_forces_case_01_06.txt
+sb
+body_axis_derivatives_case_01_06.txt
+
+QUIT
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_01.txt b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_01.txt
new file mode 100644
index 0000000..28a5b15
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_01.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_01
+
+ Alpha = -2.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Stability-axis derivatives...
+
+ alpha beta
+ ---------------- ----------------
+ z' force CL | CLa = NaN CLb = NaN
+ y force CY | CYa = NaN CYb = NaN
+ x' mom. Cl'| Cla = NaN Clb = NaN
+ y mom. Cm | Cma = NaN Cmb = NaN
+ z' mom. Cn'| Cna = NaN Cnb = NaN
+
+ roll rate p' pitch rate q' yaw rate r'
+ ---------------- ---------------- ----------------
+ z' force CL | CLp = NaN CLq = NaN CLr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ x' mom. Cl'| Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z' mom. Cn'| Cnp = NaN Cnq = NaN Cnr = NaN
+
+ Neutral point Xnp = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_02.txt b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_02.txt
new file mode 100644
index 0000000..5d8263d
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_02.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_02
+
+ Alpha = 0.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Stability-axis derivatives...
+
+ alpha beta
+ ---------------- ----------------
+ z' force CL | CLa = NaN CLb = NaN
+ y force CY | CYa = NaN CYb = NaN
+ x' mom. Cl'| Cla = NaN Clb = NaN
+ y mom. Cm | Cma = NaN Cmb = NaN
+ z' mom. Cn'| Cna = NaN Cnb = NaN
+
+ roll rate p' pitch rate q' yaw rate r'
+ ---------------- ---------------- ----------------
+ z' force CL | CLp = NaN CLq = NaN CLr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ x' mom. Cl'| Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z' mom. Cn'| Cnp = NaN Cnq = NaN Cnr = NaN
+
+ Neutral point Xnp = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_03.txt b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_03.txt
new file mode 100644
index 0000000..e79c0d7
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_03.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_03
+
+ Alpha = 2.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Stability-axis derivatives...
+
+ alpha beta
+ ---------------- ----------------
+ z' force CL | CLa = NaN CLb = NaN
+ y force CY | CYa = NaN CYb = NaN
+ x' mom. Cl'| Cla = NaN Clb = NaN
+ y mom. Cm | Cma = NaN Cmb = NaN
+ z' mom. Cn'| Cna = NaN Cnb = NaN
+
+ roll rate p' pitch rate q' yaw rate r'
+ ---------------- ---------------- ----------------
+ z' force CL | CLp = NaN CLq = NaN CLr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ x' mom. Cl'| Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z' mom. Cn'| Cnp = NaN Cnq = NaN Cnr = NaN
+
+ Neutral point Xnp = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_04.txt b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_04.txt
new file mode 100644
index 0000000..7790e0b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_04.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_04
+
+ Alpha = 5.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Stability-axis derivatives...
+
+ alpha beta
+ ---------------- ----------------
+ z' force CL | CLa = NaN CLb = NaN
+ y force CY | CYa = NaN CYb = NaN
+ x' mom. Cl'| Cla = NaN Clb = NaN
+ y mom. Cm | Cma = NaN Cmb = NaN
+ z' mom. Cn'| Cna = NaN Cnb = NaN
+
+ roll rate p' pitch rate q' yaw rate r'
+ ---------------- ---------------- ----------------
+ z' force CL | CLp = NaN CLq = NaN CLr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ x' mom. Cl'| Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z' mom. Cn'| Cnp = NaN Cnq = NaN Cnr = NaN
+
+ Neutral point Xnp = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_05.txt b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_05.txt
new file mode 100644
index 0000000..e577e5f
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_05.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_05
+
+ Alpha = 7.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Stability-axis derivatives...
+
+ alpha beta
+ ---------------- ----------------
+ z' force CL | CLa = NaN CLb = NaN
+ y force CY | CYa = NaN CYb = NaN
+ x' mom. Cl'| Cla = NaN Clb = NaN
+ y mom. Cm | Cma = NaN Cmb = NaN
+ z' mom. Cn'| Cna = NaN Cnb = NaN
+
+ roll rate p' pitch rate q' yaw rate r'
+ ---------------- ---------------- ----------------
+ z' force CL | CLp = NaN CLq = NaN CLr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ x' mom. Cl'| Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z' mom. Cn'| Cnp = NaN Cnq = NaN Cnr = NaN
+
+ Neutral point Xnp = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_06.txt b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_06.txt
new file mode 100644
index 0000000..71e7bac
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/stability_axis_derivatives_case_01_06.txt
@@ -0,0 +1,51 @@
+ ---------------------------------------------------------------
+ Vortex Lattice Output -- Total Forces
+
+ Configuration: base
+ # Surfaces = 11
+ # Strips = 180
+ # Vortices =1800
+
+ Sref = 0.92903 Cref = 0.38755 Bref = 2.4291
+ Xref = 0.20000 Yref = 0.0000 Zref = 0.0000
+
+ Standard axis orientation, X fwd, Z down
+
+ Run case: case_01_06
+
+ Alpha = 10.00000 pb/2V = 0.00000 p'b/2V = 0.00000
+ Beta = 0.00000 qc/2V = 0.00000
+ Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000
+
+ CXtot = NaN Cltot = NaN Cl'tot = NaN
+ CYtot = NaN Cmtot = NaN
+ CZtot = NaN Cntot = NaN Cn'tot = NaN
+
+ CLtot = NaN
+ CDtot = NaN
+ CDvis = NaN CDind = NaN
+ CLff = NaN CDff = NaN | Trefftz
+ CYff = NaN e = NaN | Plane
+
+
+ ---------------------------------------------------------------
+
+ Stability-axis derivatives...
+
+ alpha beta
+ ---------------- ----------------
+ z' force CL | CLa = NaN CLb = NaN
+ y force CY | CYa = NaN CYb = NaN
+ x' mom. Cl'| Cla = NaN Clb = NaN
+ y mom. Cm | Cma = NaN Cmb = NaN
+ z' mom. Cn'| Cna = NaN Cnb = NaN
+
+ roll rate p' pitch rate q' yaw rate r'
+ ---------------- ---------------- ----------------
+ z' force CL | CLp = NaN CLq = NaN CLr = NaN
+ y force CY | CYp = NaN CYq = NaN CYr = NaN
+ x' mom. Cl'| Clp = NaN Clq = NaN Clr = NaN
+ y mom. Cm | Cmp = NaN Cmq = NaN Cmr = NaN
+ z' mom. Cn'| Cnp = NaN Cnq = NaN Cnr = NaN
+
+ Neutral point Xnp = NaN
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_01.txt b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_01.txt
new file mode 100644
index 0000000..888303b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_01.txt
@@ -0,0 +1,351 @@
+ ---------------------------------------------------------------
+ Surface and Strip Forces by surface
+
+ Forces referred to Sref, Cref, Bref about Xref, Yref, Zref
+ Standard axis orientation, X fwd, Z down
+
+ Surface # 1 propulsor_pylon
+ # Chordwise = 10 # Spanwise = 20 First strip = 1
+ Surface area = 0.012199 Ave. chord = 0.123850
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 1 0.0000 0.1333 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 2 0.0000 0.1329 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 3 0.0000 0.1324 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 4 0.0000 0.1316 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 5 0.0000 0.1307 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 6 0.0000 0.1296 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 7 0.0000 0.1284 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 8 0.0000 0.1270 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 9 0.0000 0.1256 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 10 0.0000 0.1242 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 11 0.0000 0.1227 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 12 0.0000 0.1213 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 13 0.0000 0.1199 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 14 0.0000 0.1186 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 15 0.0000 0.1175 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 16 0.0000 0.1165 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 17 0.0000 0.1156 0.0004 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 18 0.0000 0.1150 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 19 0.0000 0.1146 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 20 0.0000 0.1143 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 2 012m_htailnosubsurfaces
+ # Chordwise = 10 # Spanwise = 20 First strip = 21
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 21 1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 22 1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 23 1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 24 1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 25 1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 26 1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 27 1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 28 1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 29 1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 30 1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 31 1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 32 1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 33 1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 34 1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 35 1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 36 1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 37 1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 38 1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 39 1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 40 1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 3 012m_htailnosubsurfaces (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip = 41
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 41 -1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 42 -1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 43 -1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 44 -1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 45 -1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 46 -1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 47 -1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 48 -1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 49 -1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 50 -1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 51 -1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 52 -1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 53 -1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 54 -1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 55 -1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 56 -1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 57 -1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 58 -1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 59 -1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 60 -1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 4 012m_vtailnosubsurfaces_1
+ # Chordwise = 10 # Spanwise = 20 First strip = 61
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 61 -1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 62 -1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 63 -1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 64 -1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 65 -1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 66 -1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 67 -1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 68 -1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 69 -1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 70 -1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 71 -1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 72 -1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 73 -1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 74 -1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 75 -1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 76 -1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 77 -1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 78 -1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 79 -1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 80 -1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 5 012m_vtailnosubsurfaces_2
+ # Chordwise = 10 # Spanwise = 20 First strip = 81
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 81 1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 82 1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 83 1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 84 1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 85 1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 86 1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 87 1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 88 1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 89 1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 90 1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 91 1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 92 1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 93 1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 94 1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 95 1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 96 1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 97 1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 98 1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 99 1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 100 1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 6 main_wing
+ # Chordwise = 10 # Spanwise = 20 First strip =101
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 101 0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 102 0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 103 0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 104 0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 105 0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 106 0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 107 0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 108 0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 109 0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 110 0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 111 0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 112 0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 113 0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 114 0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 115 1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 116 1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 117 1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 118 1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 119 1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 120 1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 7 main_wing (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip =121
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 121 -0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 122 -0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 123 -0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 124 -0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 125 -0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 126 -0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 127 -0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 128 -0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 129 -0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 130 -0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 131 -0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 132 -0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 133 -0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 134 -0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 135 -1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 136 -1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 137 -1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 138 -1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 139 -1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 140 -1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 8 Booms_1_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =141
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 141 1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 142 1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 143 1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 144 1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 145 1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 146 1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 147 1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 148 1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 149 1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 150 1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 9 Booms_1_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =151
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 151 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 152 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 153 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 154 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 155 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 156 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 157 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 158 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 159 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 160 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #10 Booms_2_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =161
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 161 -1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 162 -1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 163 -1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 164 -1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 165 -1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 166 -1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 167 -1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 168 -1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 169 -1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 170 -1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #11 Booms_2_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =171
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 171 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 172 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 173 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 174 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 175 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 176 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 177 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 178 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 179 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 180 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_02.txt b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_02.txt
new file mode 100644
index 0000000..888303b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_02.txt
@@ -0,0 +1,351 @@
+ ---------------------------------------------------------------
+ Surface and Strip Forces by surface
+
+ Forces referred to Sref, Cref, Bref about Xref, Yref, Zref
+ Standard axis orientation, X fwd, Z down
+
+ Surface # 1 propulsor_pylon
+ # Chordwise = 10 # Spanwise = 20 First strip = 1
+ Surface area = 0.012199 Ave. chord = 0.123850
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 1 0.0000 0.1333 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 2 0.0000 0.1329 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 3 0.0000 0.1324 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 4 0.0000 0.1316 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 5 0.0000 0.1307 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 6 0.0000 0.1296 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 7 0.0000 0.1284 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 8 0.0000 0.1270 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 9 0.0000 0.1256 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 10 0.0000 0.1242 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 11 0.0000 0.1227 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 12 0.0000 0.1213 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 13 0.0000 0.1199 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 14 0.0000 0.1186 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 15 0.0000 0.1175 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 16 0.0000 0.1165 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 17 0.0000 0.1156 0.0004 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 18 0.0000 0.1150 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 19 0.0000 0.1146 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 20 0.0000 0.1143 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 2 012m_htailnosubsurfaces
+ # Chordwise = 10 # Spanwise = 20 First strip = 21
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 21 1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 22 1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 23 1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 24 1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 25 1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 26 1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 27 1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 28 1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 29 1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 30 1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 31 1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 32 1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 33 1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 34 1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 35 1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 36 1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 37 1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 38 1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 39 1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 40 1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 3 012m_htailnosubsurfaces (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip = 41
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 41 -1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 42 -1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 43 -1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 44 -1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 45 -1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 46 -1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 47 -1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 48 -1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 49 -1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 50 -1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 51 -1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 52 -1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 53 -1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 54 -1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 55 -1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 56 -1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 57 -1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 58 -1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 59 -1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 60 -1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 4 012m_vtailnosubsurfaces_1
+ # Chordwise = 10 # Spanwise = 20 First strip = 61
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 61 -1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 62 -1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 63 -1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 64 -1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 65 -1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 66 -1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 67 -1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 68 -1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 69 -1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 70 -1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 71 -1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 72 -1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 73 -1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 74 -1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 75 -1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 76 -1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 77 -1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 78 -1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 79 -1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 80 -1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 5 012m_vtailnosubsurfaces_2
+ # Chordwise = 10 # Spanwise = 20 First strip = 81
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 81 1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 82 1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 83 1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 84 1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 85 1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 86 1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 87 1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 88 1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 89 1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 90 1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 91 1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 92 1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 93 1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 94 1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 95 1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 96 1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 97 1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 98 1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 99 1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 100 1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 6 main_wing
+ # Chordwise = 10 # Spanwise = 20 First strip =101
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 101 0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 102 0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 103 0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 104 0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 105 0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 106 0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 107 0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 108 0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 109 0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 110 0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 111 0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 112 0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 113 0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 114 0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 115 1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 116 1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 117 1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 118 1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 119 1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 120 1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 7 main_wing (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip =121
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 121 -0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 122 -0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 123 -0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 124 -0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 125 -0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 126 -0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 127 -0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 128 -0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 129 -0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 130 -0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 131 -0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 132 -0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 133 -0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 134 -0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 135 -1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 136 -1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 137 -1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 138 -1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 139 -1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 140 -1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 8 Booms_1_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =141
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 141 1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 142 1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 143 1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 144 1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 145 1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 146 1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 147 1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 148 1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 149 1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 150 1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 9 Booms_1_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =151
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 151 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 152 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 153 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 154 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 155 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 156 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 157 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 158 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 159 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 160 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #10 Booms_2_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =161
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 161 -1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 162 -1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 163 -1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 164 -1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 165 -1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 166 -1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 167 -1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 168 -1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 169 -1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 170 -1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #11 Booms_2_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =171
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 171 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 172 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 173 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 174 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 175 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 176 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 177 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 178 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 179 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 180 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_03.txt b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_03.txt
new file mode 100644
index 0000000..888303b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_03.txt
@@ -0,0 +1,351 @@
+ ---------------------------------------------------------------
+ Surface and Strip Forces by surface
+
+ Forces referred to Sref, Cref, Bref about Xref, Yref, Zref
+ Standard axis orientation, X fwd, Z down
+
+ Surface # 1 propulsor_pylon
+ # Chordwise = 10 # Spanwise = 20 First strip = 1
+ Surface area = 0.012199 Ave. chord = 0.123850
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 1 0.0000 0.1333 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 2 0.0000 0.1329 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 3 0.0000 0.1324 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 4 0.0000 0.1316 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 5 0.0000 0.1307 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 6 0.0000 0.1296 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 7 0.0000 0.1284 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 8 0.0000 0.1270 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 9 0.0000 0.1256 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 10 0.0000 0.1242 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 11 0.0000 0.1227 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 12 0.0000 0.1213 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 13 0.0000 0.1199 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 14 0.0000 0.1186 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 15 0.0000 0.1175 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 16 0.0000 0.1165 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 17 0.0000 0.1156 0.0004 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 18 0.0000 0.1150 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 19 0.0000 0.1146 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 20 0.0000 0.1143 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 2 012m_htailnosubsurfaces
+ # Chordwise = 10 # Spanwise = 20 First strip = 21
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 21 1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 22 1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 23 1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 24 1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 25 1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 26 1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 27 1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 28 1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 29 1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 30 1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 31 1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 32 1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 33 1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 34 1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 35 1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 36 1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 37 1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 38 1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 39 1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 40 1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 3 012m_htailnosubsurfaces (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip = 41
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 41 -1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 42 -1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 43 -1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 44 -1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 45 -1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 46 -1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 47 -1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 48 -1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 49 -1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 50 -1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 51 -1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 52 -1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 53 -1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 54 -1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 55 -1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 56 -1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 57 -1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 58 -1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 59 -1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 60 -1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 4 012m_vtailnosubsurfaces_1
+ # Chordwise = 10 # Spanwise = 20 First strip = 61
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 61 -1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 62 -1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 63 -1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 64 -1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 65 -1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 66 -1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 67 -1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 68 -1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 69 -1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 70 -1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 71 -1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 72 -1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 73 -1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 74 -1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 75 -1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 76 -1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 77 -1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 78 -1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 79 -1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 80 -1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 5 012m_vtailnosubsurfaces_2
+ # Chordwise = 10 # Spanwise = 20 First strip = 81
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 81 1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 82 1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 83 1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 84 1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 85 1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 86 1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 87 1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 88 1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 89 1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 90 1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 91 1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 92 1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 93 1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 94 1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 95 1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 96 1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 97 1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 98 1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 99 1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 100 1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 6 main_wing
+ # Chordwise = 10 # Spanwise = 20 First strip =101
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 101 0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 102 0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 103 0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 104 0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 105 0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 106 0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 107 0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 108 0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 109 0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 110 0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 111 0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 112 0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 113 0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 114 0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 115 1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 116 1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 117 1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 118 1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 119 1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 120 1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 7 main_wing (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip =121
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 121 -0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 122 -0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 123 -0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 124 -0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 125 -0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 126 -0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 127 -0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 128 -0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 129 -0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 130 -0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 131 -0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 132 -0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 133 -0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 134 -0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 135 -1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 136 -1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 137 -1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 138 -1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 139 -1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 140 -1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 8 Booms_1_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =141
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 141 1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 142 1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 143 1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 144 1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 145 1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 146 1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 147 1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 148 1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 149 1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 150 1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 9 Booms_1_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =151
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 151 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 152 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 153 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 154 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 155 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 156 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 157 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 158 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 159 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 160 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #10 Booms_2_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =161
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 161 -1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 162 -1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 163 -1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 164 -1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 165 -1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 166 -1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 167 -1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 168 -1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 169 -1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 170 -1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #11 Booms_2_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =171
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 171 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 172 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 173 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 174 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 175 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 176 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 177 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 178 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 179 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 180 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_04.txt b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_04.txt
new file mode 100644
index 0000000..888303b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_04.txt
@@ -0,0 +1,351 @@
+ ---------------------------------------------------------------
+ Surface and Strip Forces by surface
+
+ Forces referred to Sref, Cref, Bref about Xref, Yref, Zref
+ Standard axis orientation, X fwd, Z down
+
+ Surface # 1 propulsor_pylon
+ # Chordwise = 10 # Spanwise = 20 First strip = 1
+ Surface area = 0.012199 Ave. chord = 0.123850
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 1 0.0000 0.1333 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 2 0.0000 0.1329 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 3 0.0000 0.1324 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 4 0.0000 0.1316 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 5 0.0000 0.1307 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 6 0.0000 0.1296 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 7 0.0000 0.1284 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 8 0.0000 0.1270 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 9 0.0000 0.1256 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 10 0.0000 0.1242 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 11 0.0000 0.1227 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 12 0.0000 0.1213 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 13 0.0000 0.1199 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 14 0.0000 0.1186 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 15 0.0000 0.1175 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 16 0.0000 0.1165 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 17 0.0000 0.1156 0.0004 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 18 0.0000 0.1150 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 19 0.0000 0.1146 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 20 0.0000 0.1143 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 2 012m_htailnosubsurfaces
+ # Chordwise = 10 # Spanwise = 20 First strip = 21
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 21 1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 22 1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 23 1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 24 1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 25 1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 26 1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 27 1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 28 1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 29 1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 30 1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 31 1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 32 1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 33 1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 34 1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 35 1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 36 1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 37 1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 38 1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 39 1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 40 1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 3 012m_htailnosubsurfaces (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip = 41
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 41 -1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 42 -1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 43 -1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 44 -1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 45 -1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 46 -1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 47 -1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 48 -1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 49 -1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 50 -1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 51 -1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 52 -1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 53 -1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 54 -1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 55 -1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 56 -1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 57 -1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 58 -1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 59 -1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 60 -1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 4 012m_vtailnosubsurfaces_1
+ # Chordwise = 10 # Spanwise = 20 First strip = 61
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 61 -1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 62 -1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 63 -1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 64 -1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 65 -1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 66 -1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 67 -1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 68 -1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 69 -1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 70 -1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 71 -1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 72 -1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 73 -1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 74 -1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 75 -1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 76 -1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 77 -1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 78 -1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 79 -1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 80 -1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 5 012m_vtailnosubsurfaces_2
+ # Chordwise = 10 # Spanwise = 20 First strip = 81
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 81 1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 82 1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 83 1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 84 1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 85 1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 86 1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 87 1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 88 1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 89 1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 90 1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 91 1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 92 1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 93 1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 94 1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 95 1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 96 1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 97 1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 98 1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 99 1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 100 1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 6 main_wing
+ # Chordwise = 10 # Spanwise = 20 First strip =101
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 101 0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 102 0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 103 0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 104 0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 105 0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 106 0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 107 0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 108 0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 109 0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 110 0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 111 0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 112 0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 113 0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 114 0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 115 1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 116 1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 117 1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 118 1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 119 1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 120 1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 7 main_wing (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip =121
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 121 -0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 122 -0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 123 -0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 124 -0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 125 -0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 126 -0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 127 -0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 128 -0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 129 -0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 130 -0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 131 -0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 132 -0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 133 -0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 134 -0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 135 -1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 136 -1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 137 -1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 138 -1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 139 -1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 140 -1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 8 Booms_1_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =141
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 141 1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 142 1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 143 1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 144 1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 145 1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 146 1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 147 1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 148 1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 149 1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 150 1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 9 Booms_1_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =151
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 151 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 152 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 153 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 154 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 155 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 156 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 157 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 158 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 159 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 160 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #10 Booms_2_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =161
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 161 -1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 162 -1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 163 -1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 164 -1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 165 -1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 166 -1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 167 -1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 168 -1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 169 -1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 170 -1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #11 Booms_2_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =171
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 171 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 172 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 173 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 174 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 175 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 176 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 177 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 178 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 179 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 180 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_05.txt b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_05.txt
new file mode 100644
index 0000000..888303b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_05.txt
@@ -0,0 +1,351 @@
+ ---------------------------------------------------------------
+ Surface and Strip Forces by surface
+
+ Forces referred to Sref, Cref, Bref about Xref, Yref, Zref
+ Standard axis orientation, X fwd, Z down
+
+ Surface # 1 propulsor_pylon
+ # Chordwise = 10 # Spanwise = 20 First strip = 1
+ Surface area = 0.012199 Ave. chord = 0.123850
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 1 0.0000 0.1333 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 2 0.0000 0.1329 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 3 0.0000 0.1324 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 4 0.0000 0.1316 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 5 0.0000 0.1307 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 6 0.0000 0.1296 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 7 0.0000 0.1284 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 8 0.0000 0.1270 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 9 0.0000 0.1256 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 10 0.0000 0.1242 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 11 0.0000 0.1227 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 12 0.0000 0.1213 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 13 0.0000 0.1199 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 14 0.0000 0.1186 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 15 0.0000 0.1175 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 16 0.0000 0.1165 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 17 0.0000 0.1156 0.0004 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 18 0.0000 0.1150 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 19 0.0000 0.1146 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 20 0.0000 0.1143 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 2 012m_htailnosubsurfaces
+ # Chordwise = 10 # Spanwise = 20 First strip = 21
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 21 1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 22 1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 23 1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 24 1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 25 1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 26 1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 27 1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 28 1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 29 1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 30 1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 31 1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 32 1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 33 1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 34 1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 35 1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 36 1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 37 1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 38 1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 39 1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 40 1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 3 012m_htailnosubsurfaces (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip = 41
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 41 -1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 42 -1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 43 -1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 44 -1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 45 -1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 46 -1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 47 -1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 48 -1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 49 -1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 50 -1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 51 -1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 52 -1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 53 -1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 54 -1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 55 -1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 56 -1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 57 -1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 58 -1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 59 -1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 60 -1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 4 012m_vtailnosubsurfaces_1
+ # Chordwise = 10 # Spanwise = 20 First strip = 61
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 61 -1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 62 -1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 63 -1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 64 -1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 65 -1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 66 -1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 67 -1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 68 -1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 69 -1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 70 -1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 71 -1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 72 -1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 73 -1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 74 -1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 75 -1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 76 -1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 77 -1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 78 -1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 79 -1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 80 -1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 5 012m_vtailnosubsurfaces_2
+ # Chordwise = 10 # Spanwise = 20 First strip = 81
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 81 1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 82 1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 83 1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 84 1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 85 1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 86 1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 87 1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 88 1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 89 1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 90 1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 91 1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 92 1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 93 1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 94 1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 95 1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 96 1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 97 1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 98 1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 99 1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 100 1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 6 main_wing
+ # Chordwise = 10 # Spanwise = 20 First strip =101
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 101 0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 102 0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 103 0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 104 0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 105 0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 106 0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 107 0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 108 0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 109 0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 110 0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 111 0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 112 0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 113 0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 114 0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 115 1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 116 1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 117 1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 118 1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 119 1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 120 1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 7 main_wing (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip =121
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 121 -0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 122 -0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 123 -0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 124 -0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 125 -0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 126 -0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 127 -0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 128 -0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 129 -0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 130 -0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 131 -0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 132 -0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 133 -0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 134 -0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 135 -1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 136 -1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 137 -1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 138 -1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 139 -1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 140 -1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 8 Booms_1_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =141
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 141 1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 142 1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 143 1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 144 1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 145 1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 146 1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 147 1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 148 1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 149 1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 150 1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 9 Booms_1_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =151
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 151 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 152 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 153 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 154 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 155 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 156 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 157 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 158 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 159 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 160 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #10 Booms_2_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =161
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 161 -1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 162 -1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 163 -1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 164 -1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 165 -1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 166 -1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 167 -1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 168 -1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 169 -1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 170 -1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #11 Booms_2_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =171
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 171 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 172 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 173 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 174 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 175 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 176 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 177 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 178 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 179 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 180 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_06.txt b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_06.txt
new file mode 100644
index 0000000..888303b
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/strip_forces_case_01_06.txt
@@ -0,0 +1,351 @@
+ ---------------------------------------------------------------
+ Surface and Strip Forces by surface
+
+ Forces referred to Sref, Cref, Bref about Xref, Yref, Zref
+ Standard axis orientation, X fwd, Z down
+
+ Surface # 1 propulsor_pylon
+ # Chordwise = 10 # Spanwise = 20 First strip = 1
+ Surface area = 0.012199 Ave. chord = 0.123850
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 1 0.0000 0.1333 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 2 0.0000 0.1329 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 3 0.0000 0.1324 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 4 0.0000 0.1316 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 5 0.0000 0.1307 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 6 0.0000 0.1296 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 7 0.0000 0.1284 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 8 0.0000 0.1270 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 9 0.0000 0.1256 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 10 0.0000 0.1242 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 11 0.0000 0.1227 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 12 0.0000 0.1213 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 13 0.0000 0.1199 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 14 0.0000 0.1186 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 15 0.0000 0.1175 0.0007 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 16 0.0000 0.1165 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 17 0.0000 0.1156 0.0004 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 18 0.0000 0.1150 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 19 0.0000 0.1146 0.0002 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 20 0.0000 0.1143 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 2 012m_htailnosubsurfaces
+ # Chordwise = 10 # Spanwise = 20 First strip = 21
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 21 1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 22 1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 23 1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 24 1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 25 1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 26 1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 27 1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 28 1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 29 1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 30 1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 31 1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 32 1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 33 1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 34 1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 35 1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 36 1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 37 1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 38 1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 39 1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 40 1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 3 012m_htailnosubsurfaces (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip = 41
+ Surface area = 0.147498 Ave. chord = 0.209550
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 41 -1.2201 0.3045 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 42 -1.2276 0.3022 0.0039 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 43 -1.2424 0.2975 0.0063 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 44 -1.2641 0.2908 0.0084 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 45 -1.2922 0.2820 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 46 -1.3260 0.2714 0.0114 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 47 -1.3647 0.2593 0.0122 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 48 -1.4074 0.2460 0.0126 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 49 -1.4528 0.2318 0.0124 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 50 -1.5001 0.2170 0.0119 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 51 -1.5479 0.2021 0.0111 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 52 -1.5952 0.1873 0.0101 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 53 -1.6406 0.1731 0.0088 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 54 -1.6833 0.1598 0.0075 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 55 -1.7220 0.1477 0.0062 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 56 -1.7558 0.1371 0.0049 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 57 -1.7839 0.1283 0.0037 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 58 -1.8056 0.1216 0.0026 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 59 -1.8204 0.1169 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 60 -1.8279 0.1146 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 4 012m_vtailnosubsurfaces_1
+ # Chordwise = 10 # Spanwise = 20 First strip = 61
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 61 -1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 62 -1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 63 -1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 64 -1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 65 -1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 66 -1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 67 -1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 68 -1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 69 -1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 70 -1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 71 -1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 72 -1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 73 -1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 74 -1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 75 -1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 76 -1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 77 -1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 78 -1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 79 -1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 80 -1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 5 012m_vtailnosubsurfaces_2
+ # Chordwise = 10 # Spanwise = 20 First strip = 81
+ Surface area = 0.027889 Ave. chord = 0.228599
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 81 1.2192 0.3040 0.0005 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 82 1.2192 0.3011 0.0009 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 83 1.2192 0.2966 0.0013 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 84 1.2192 0.2906 0.0016 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 85 1.2192 0.2831 0.0018 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 86 1.2192 0.2744 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 87 1.2192 0.2646 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 88 1.2192 0.2540 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 89 1.2192 0.2428 0.0022 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 90 1.2192 0.2313 0.0021 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 91 1.2192 0.2196 0.0020 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 92 1.2192 0.2082 0.0019 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 93 1.2192 0.1973 0.0017 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 94 1.2192 0.1871 0.0015 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 95 1.2192 0.1778 0.0012 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 96 1.2192 0.1698 0.0010 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 97 1.2192 0.1631 0.0008 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 98 1.2192 0.1579 0.0006 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 99 1.2192 0.1544 0.0003 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 100 1.2192 0.1526 0.0001 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 6 main_wing
+ # Chordwise = 10 # Spanwise = 20 First strip =101
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 101 0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 102 0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 103 0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 104 0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 105 0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 106 0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 107 0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 108 0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 109 0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 110 0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 111 0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 112 0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 113 0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 114 0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 115 1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 116 1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 117 1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 118 1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 119 1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 120 1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 7 main_wing (YDUP)
+ # Chordwise = 10 # Spanwise = 20 First strip =121
+ Surface area = 0.469089 Ave. chord = 0.382450
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = 0.00000
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 121 -0.0019 0.4587 0.0035 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 122 -0.0168 0.4568 0.0103 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 123 -0.0462 0.4531 0.0167 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 124 -0.0895 0.4476 0.0225 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 125 -0.1455 0.4406 0.0275 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 126 -0.2129 0.4321 0.0316 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 127 -0.2900 0.4224 0.0347 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 128 -0.3749 0.4117 0.0366 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 129 -0.4655 0.4003 0.0375 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 130 -0.5597 0.3884 0.0373 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 131 -0.6549 0.3765 0.0361 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 132 -0.7491 0.3646 0.0341 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 133 -0.8397 0.3532 0.0314 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 134 -0.9246 0.3425 0.0281 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 135 -1.0017 0.3328 0.0244 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 136 -1.0691 0.3243 0.0203 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 137 -1.1251 0.3173 0.0160 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 138 -1.1684 0.3118 0.0115 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 139 -1.1978 0.3081 0.0069 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 140 -1.2127 0.3062 0.0023 NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 8 Booms_1_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =141
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 141 1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 142 1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 143 1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 144 1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 145 1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 146 1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 147 1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 148 1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 149 1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 150 1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface # 9 Booms_1_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =151
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 151 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 152 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 153 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 154 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 155 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 156 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 157 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 158 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 159 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 160 1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #10 Booms_2_horizontal
+ # Chordwise = 10 # Spanwise = 10 First strip =161
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 161 -1.2535 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 162 -1.2459 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 163 -1.2383 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 164 -1.2306 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 165 -1.2230 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 166 -1.2154 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 167 -1.2078 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 168 -1.2002 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 169 -1.1925 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 170 -1.1849 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+
+ Surface #11 Booms_2_vertical
+ # Chordwise = 10 # Spanwise = 10 First strip =171
+ Surface area = NaN Ave. chord = NaN
+ CLsurf = NaN Clsurf = NaN
+ CYsurf = NaN Cmsurf = NaN
+ CDsurf = NaN Cnsurf = NaN
+ CDisurf = NaN CDvsurf = NaN
+
+ Forces referred to Ssurf, Cave about hinge axis thru LE
+ CLsurf = NaN CDsurf = NaN
+ Deflect =
+
+ Strip Forces referred to Strip Area, Chord
+ j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
+ 171 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 172 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 173 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 174 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 175 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 176 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 177 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 178 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 179 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ 180 -1.2192 NaN NaN NaN NaN NaN NaN NaN 0.0000 NaN NaN
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_01.txt b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_01.txt
new file mode 100644
index 0000000..807c3c2
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_01.txt
@@ -0,0 +1,35 @@
+ ---------------------------------------------------------------
+ Surface Forces (referred to Sref,Cref,Bref about Xref,Yref,Zref)
+ Standard axis orientation, X fwd, Z down
+
+ Sref = 0.9290 Cref = 0.3876 Bref = 2.4291
+ Xref = 0.2000 Yref = 0.0000 Zref = 0.0000
+
+ n Area CL CD Cm CY Cn Cl CDi CDv
+ 1 0.012 NaN NaN NaN NaN NaN NaN NaN 0.0000 propulsor_pylon
+ 2 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces
+ 3 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing
+ 7 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_vertical
+10 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_horizontal
+11 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_vertical
+
+ Surface Forces (referred to Ssurf, Cave about root LE on hinge axis)
+
+ n Ssurf Cave cl cd cdv cm_LE
+ 1 0.012 0.124 NaN NaN 0.0000 0.0000 propulsor_pylon
+ 2 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces
+ 3 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing
+ 7 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN 0.0000 Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN 0.0000 Booms_1_vertical
+ 10 NaN NaN NaN NaN NaN 0.0000 Booms_2_horizontal
+ 11 NaN NaN NaN NaN NaN 0.0000 Booms_2_vertical
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_02.txt b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_02.txt
new file mode 100644
index 0000000..807c3c2
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_02.txt
@@ -0,0 +1,35 @@
+ ---------------------------------------------------------------
+ Surface Forces (referred to Sref,Cref,Bref about Xref,Yref,Zref)
+ Standard axis orientation, X fwd, Z down
+
+ Sref = 0.9290 Cref = 0.3876 Bref = 2.4291
+ Xref = 0.2000 Yref = 0.0000 Zref = 0.0000
+
+ n Area CL CD Cm CY Cn Cl CDi CDv
+ 1 0.012 NaN NaN NaN NaN NaN NaN NaN 0.0000 propulsor_pylon
+ 2 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces
+ 3 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing
+ 7 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_vertical
+10 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_horizontal
+11 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_vertical
+
+ Surface Forces (referred to Ssurf, Cave about root LE on hinge axis)
+
+ n Ssurf Cave cl cd cdv cm_LE
+ 1 0.012 0.124 NaN NaN 0.0000 0.0000 propulsor_pylon
+ 2 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces
+ 3 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing
+ 7 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN 0.0000 Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN 0.0000 Booms_1_vertical
+ 10 NaN NaN NaN NaN NaN 0.0000 Booms_2_horizontal
+ 11 NaN NaN NaN NaN NaN 0.0000 Booms_2_vertical
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_03.txt b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_03.txt
new file mode 100644
index 0000000..807c3c2
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_03.txt
@@ -0,0 +1,35 @@
+ ---------------------------------------------------------------
+ Surface Forces (referred to Sref,Cref,Bref about Xref,Yref,Zref)
+ Standard axis orientation, X fwd, Z down
+
+ Sref = 0.9290 Cref = 0.3876 Bref = 2.4291
+ Xref = 0.2000 Yref = 0.0000 Zref = 0.0000
+
+ n Area CL CD Cm CY Cn Cl CDi CDv
+ 1 0.012 NaN NaN NaN NaN NaN NaN NaN 0.0000 propulsor_pylon
+ 2 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces
+ 3 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing
+ 7 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_vertical
+10 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_horizontal
+11 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_vertical
+
+ Surface Forces (referred to Ssurf, Cave about root LE on hinge axis)
+
+ n Ssurf Cave cl cd cdv cm_LE
+ 1 0.012 0.124 NaN NaN 0.0000 0.0000 propulsor_pylon
+ 2 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces
+ 3 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing
+ 7 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN 0.0000 Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN 0.0000 Booms_1_vertical
+ 10 NaN NaN NaN NaN NaN 0.0000 Booms_2_horizontal
+ 11 NaN NaN NaN NaN NaN 0.0000 Booms_2_vertical
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_04.txt b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_04.txt
new file mode 100644
index 0000000..807c3c2
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_04.txt
@@ -0,0 +1,35 @@
+ ---------------------------------------------------------------
+ Surface Forces (referred to Sref,Cref,Bref about Xref,Yref,Zref)
+ Standard axis orientation, X fwd, Z down
+
+ Sref = 0.9290 Cref = 0.3876 Bref = 2.4291
+ Xref = 0.2000 Yref = 0.0000 Zref = 0.0000
+
+ n Area CL CD Cm CY Cn Cl CDi CDv
+ 1 0.012 NaN NaN NaN NaN NaN NaN NaN 0.0000 propulsor_pylon
+ 2 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces
+ 3 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing
+ 7 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_vertical
+10 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_horizontal
+11 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_vertical
+
+ Surface Forces (referred to Ssurf, Cave about root LE on hinge axis)
+
+ n Ssurf Cave cl cd cdv cm_LE
+ 1 0.012 0.124 NaN NaN 0.0000 0.0000 propulsor_pylon
+ 2 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces
+ 3 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing
+ 7 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN 0.0000 Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN 0.0000 Booms_1_vertical
+ 10 NaN NaN NaN NaN NaN 0.0000 Booms_2_horizontal
+ 11 NaN NaN NaN NaN NaN 0.0000 Booms_2_vertical
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_05.txt b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_05.txt
new file mode 100644
index 0000000..807c3c2
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_05.txt
@@ -0,0 +1,35 @@
+ ---------------------------------------------------------------
+ Surface Forces (referred to Sref,Cref,Bref about Xref,Yref,Zref)
+ Standard axis orientation, X fwd, Z down
+
+ Sref = 0.9290 Cref = 0.3876 Bref = 2.4291
+ Xref = 0.2000 Yref = 0.0000 Zref = 0.0000
+
+ n Area CL CD Cm CY Cn Cl CDi CDv
+ 1 0.012 NaN NaN NaN NaN NaN NaN NaN 0.0000 propulsor_pylon
+ 2 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces
+ 3 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing
+ 7 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_vertical
+10 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_horizontal
+11 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_vertical
+
+ Surface Forces (referred to Ssurf, Cave about root LE on hinge axis)
+
+ n Ssurf Cave cl cd cdv cm_LE
+ 1 0.012 0.124 NaN NaN 0.0000 0.0000 propulsor_pylon
+ 2 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces
+ 3 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing
+ 7 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN 0.0000 Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN 0.0000 Booms_1_vertical
+ 10 NaN NaN NaN NaN NaN 0.0000 Booms_2_horizontal
+ 11 NaN NaN NaN NaN NaN 0.0000 Booms_2_vertical
+ ---------------------------------------------------------------
diff --git a/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_06.txt b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_06.txt
new file mode 100644
index 0000000..807c3c2
--- /dev/null
+++ b/tut_Control_Surface_Sizing_Navion/avl_files/surface_forces_case_01_06.txt
@@ -0,0 +1,35 @@
+ ---------------------------------------------------------------
+ Surface Forces (referred to Sref,Cref,Bref about Xref,Yref,Zref)
+ Standard axis orientation, X fwd, Z down
+
+ Sref = 0.9290 Cref = 0.3876 Bref = 2.4291
+ Xref = 0.2000 Yref = 0.0000 Zref = 0.0000
+
+ n Area CL CD Cm CY Cn Cl CDi CDv
+ 1 0.012 NaN NaN NaN NaN NaN NaN NaN 0.0000 propulsor_pylon
+ 2 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces
+ 3 0.147 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 NaN NaN NaN NaN NaN NaN NaN 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing
+ 7 0.469 NaN NaN NaN NaN NaN NaN NaN 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_1_vertical
+10 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_horizontal
+11 NaN NaN NaN NaN NaN NaN NaN NaN NaN Booms_2_vertical
+
+ Surface Forces (referred to Ssurf, Cave about root LE on hinge axis)
+
+ n Ssurf Cave cl cd cdv cm_LE
+ 1 0.012 0.124 NaN NaN 0.0000 0.0000 propulsor_pylon
+ 2 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces
+ 3 0.147 0.210 NaN NaN 0.0000 0.0000 012m_htailnosubsurfaces (YDUP)
+ 4 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_1
+ 5 0.028 0.229 NaN NaN 0.0000 0.0000 012m_vtailnosubsurfaces_2
+ 6 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing
+ 7 0.469 0.382 NaN NaN 0.0000 0.0000 main_wing (YDUP)
+ 8 NaN NaN NaN NaN NaN 0.0000 Booms_1_horizontal
+ 9 NaN NaN NaN NaN NaN 0.0000 Booms_1_vertical
+ 10 NaN NaN NaN NaN NaN 0.0000 Booms_2_horizontal
+ 11 NaN NaN NaN NaN NaN 0.0000 Booms_2_vertical
+ ---------------------------------------------------------------
diff --git a/B737_AVL_Tutorial/tut_mission_B737_AVL.py b/tut_Mission_AVL_Tutorial_B737/tut_mission_B737_AVL.py
similarity index 100%
rename from B737_AVL_Tutorial/tut_mission_B737_AVL.py
rename to tut_Mission_AVL_Tutorial_B737/tut_mission_B737_AVL.py
diff --git a/tut_concorde.py b/tut_Mission_Concorde.py
similarity index 100%
rename from tut_concorde.py
rename to tut_Mission_Concorde.py
diff --git a/tut_solar_uav.py b/tut_Mission_Solar_UAV.py
similarity index 100%
rename from tut_solar_uav.py
rename to tut_Mission_Solar_UAV.py
diff --git a/tut_payload_range.py b/tut_Payload_Range_E190.py
similarity index 100%
rename from tut_payload_range.py
rename to tut_Payload_Range_E190.py
diff --git a/Regional_Jet_Optimization/Analyses.py b/tut_Regional_Jet_Optimization/Analyses.py
similarity index 98%
rename from Regional_Jet_Optimization/Analyses.py
rename to tut_Regional_Jet_Optimization/Analyses.py
index d07c455..9f0f70c 100644
--- a/Regional_Jet_Optimization/Analyses.py
+++ b/tut_Regional_Jet_Optimization/Analyses.py
@@ -73,7 +73,7 @@ def base(vehicle):
# ------------------------------------------------------------------
# Energy
energy= SUAVE.Analyses.Energy.Energy()
- energy.network = vehicle.propulsors
+ energy.network = vehicle.networks
analyses.append(energy)
# ------------------------------------------------------------------
diff --git a/Regional_Jet_Optimization/Missions.py b/tut_Regional_Jet_Optimization/Missions.py
similarity index 100%
rename from Regional_Jet_Optimization/Missions.py
rename to tut_Regional_Jet_Optimization/Missions.py
diff --git a/Regional_Jet_Optimization/Optimize.py b/tut_Regional_Jet_Optimization/Optimize.py
similarity index 100%
rename from Regional_Jet_Optimization/Optimize.py
rename to tut_Regional_Jet_Optimization/Optimize.py
diff --git a/Regional_Jet_Optimization/Plot_Mission.py b/tut_Regional_Jet_Optimization/Plot_Mission.py
similarity index 100%
rename from Regional_Jet_Optimization/Plot_Mission.py
rename to tut_Regional_Jet_Optimization/Plot_Mission.py
diff --git a/Regional_Jet_Optimization/Procedure.py b/tut_Regional_Jet_Optimization/Procedure.py
similarity index 100%
rename from Regional_Jet_Optimization/Procedure.py
rename to tut_Regional_Jet_Optimization/Procedure.py
diff --git a/Regional_Jet_Optimization/Vehicles.py b/tut_Regional_Jet_Optimization/Vehicles.py
similarity index 100%
rename from Regional_Jet_Optimization/Vehicles.py
rename to tut_Regional_Jet_Optimization/Vehicles.py
diff --git a/tut_Single_Point_Analysis_X57Mod2.py b/tut_Single_Point_Analysis_X57Mod2.py
new file mode 100644
index 0000000..8588e27
--- /dev/null
+++ b/tut_Single_Point_Analysis_X57Mod2.py
@@ -0,0 +1,632 @@
+# tut_Single_Point_Analsis_X57_Mod2.py
+#
+# Created: Oct 2021, M. Clarke
+
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+# SUAVE Imports
+import SUAVE
+from SUAVE.Core import Data, Units
+from SUAVE.Plots.Performance.Mission_Plots import *
+from SUAVE.Plots.Geometry import *
+from SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics import Aerodynamics
+from SUAVE.Components.Energy.Networks.Battery_Propeller import Battery_Propeller
+from SUAVE.Methods.Propulsion import propeller_design
+from SUAVE.Methods.Power.Battery.Sizing import initialize_from_mass
+from SUAVE.Methods.Propulsion.electric_motor_sizing import size_optimal_motor
+from SUAVE.Methods.Geometry.Two_Dimensional.Planform import segment_properties
+
+
+# Python Imports
+import numpy as np
+from copy import deepcopy
+
+
+def main():
+
+ vehicle = vehicle_setup()
+
+ # Get properties of atmosphere at specified altitude
+ atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ atmo_data = atmosphere.compute_values(altitude = 8012 * Units.feet)
+
+ # Define run conditions
+ run_conditions = Aerodynamics()
+ run_conditions.freestream.density = atmo_data.density[0,0]
+ run_conditions.freestream.gravity = 9.81
+ run_conditions.freestream.speed_of_sound = atmo_data.speed_of_sound[0,0]
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+ run_conditions.aerodynamics.angle_of_attack = np.array([0.0])
+ run_conditions.aerodynamics.lift_coefficient = 0.547
+ run_conditions.freestream.velocity = 120.91 * Units['mph']
+ run_conditions.freestream.mach_number = run_conditions.freestream.velocity/run_conditions.freestream.speed_of_sound
+ run_conditions.aerodynamics.roll_rate_coefficient = 0.07
+ run_conditions.aerodynamics.pitch_rate_coefficient = 0.0
+ run_conditions.aerodynamics.side_slip_angle = 0.0
+
+ # Call AVL Stability Analysis
+ stability_roll_maneuver = SUAVE.Analyses.Stability.AVL()
+ stability_roll_maneuver.settings.filenames.avl_bin_name = '/Users/matthewclarke/Documents/AVL/avl3.35' # change to path of AVL
+ stability_roll_maneuver.settings.number_spanwise_vortices = 40
+ stability_roll_maneuver.geometry = vehicle
+ stability_roll_maneuver.geometry._base = Data()
+ stability_roll_maneuver.geometry._base.tag = vehicle.tag
+ results_roll_maneuver = stability_roll_maneuver.evaluate_conditions(run_conditions, trim_aircraft = True)
+
+ # Extract data
+ CL = results_roll_maneuver.aerodynamics.lift_coefficient[0,0]
+ AoA = results_roll_maneuver.aerodynamics.angle_of_attack[0,0]
+ CD = results_roll_maneuver.aerodynamics.drag_breakdown.induced.total[0,0]
+ CM = results_roll_maneuver.aerodynamics.pitch_moment_coefficient[0,0]
+ spiral_criteria = results_roll_maneuver.stability.static.spiral_criteria[0,0]
+ NP = results_roll_maneuver.stability.static.neutral_point[0,0]
+ cg = vehicle.mass_properties.center_of_gravity[0][0]
+ MAC = vehicle.wings.main_wing.chords.mean_aerodynamic
+ static_margin = (NP - cg)/MAC
+ CM_alpha = results_roll_maneuver.stability.static.Cm_alpha[0,0]
+ phugoid_damping_ratio = results_roll_maneuver.dynamic_stability.LongModes.phugoidDamp[0,0]
+ short_period_frequency = results_roll_maneuver.dynamic_stability.LongModes.shortPeriodFreqHz[0,0]
+ dutch_roll_frequency = results_roll_maneuver.dynamic_stability.LatModes.dutchRollFreqHz[0,0]
+ spiral_doubling_time = results_roll_maneuver.dynamic_stability.LatModes.spiralTimeDoubleHalf[0,0]
+ aileron_roll_deflection = results_roll_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.aileron.deflection
+ rudder_roll_deflection = results_roll_maneuver.stability.static.control_surfaces_cases['case_0001_0001'].control_surfaces.rudder.deflection
+
+ print("\n\n")
+ print("************** RESULTS ************** ")
+ print("Angle of Attack : " + str(AoA))
+ print("Lift Coefficient : " + str(CL))
+ print("Drag Coefficient : " + str(CD))
+ print("Moment Coefficient : " + str(CM))
+ print("Static Margin : " + str(static_margin))
+ print("CM alpla : " + str(CM_alpha))
+ print("Phugoid Damping Ratio : " + str(phugoid_damping_ratio))
+ print("Short Period Frequency : " + str(short_period_frequency))
+ print("Dutch Roll Frequency : " + str(dutch_roll_frequency))
+ print("Spiral Doubling Time : " + str(spiral_doubling_time))
+ print("Spiral Criteria : " + str(spiral_criteria))
+ print("Aileron Roll Defl : " + str(aileron_roll_deflection))
+ print("Rudder Roll Defl : " + str(rudder_roll_deflection))
+
+ return
+
+
+
+def vehicle_setup():
+ # ----------------------------------------------------------------------
+ # Define Vehicle
+ # ---------------------------------------------------------------------
+
+ vehicle = SUAVE.Vehicle()
+ vehicle.tag = 'X57_Mod2'
+
+ # ------------------------------------------------------------------
+ # Vehicle-level Properties
+ # ------------------------------------------------------------------
+
+ # mass properties
+ vehicle.mass_properties.max_takeoff = 2550. * Units.pounds
+ vehicle.mass_properties.takeoff = 2550. * Units.pounds
+ vehicle.mass_properties.max_zero_fuel = 2550. * Units.pounds
+ vehicle.mass_properties.moments_of_inertia.tensor = np.array([[164627.7,0.0,0.0],[0.0,471262.4,0.0],[0.0,0.0,554518.7]]) # Navion
+ vehicle.envelope.ultimate_load = 5.7
+ vehicle.envelope.limit_load = 3.8
+ vehicle.reference_area = 14.76
+ vehicle.passengers = 4
+ vehicle.systems.control = "fully powered"
+ vehicle.systems.accessories = "commuter"
+
+ cruise_speed = 135.*Units['mph']
+ altitude = 2500. * Units.ft
+ atmo = SUAVE.Analyses.Atmospheric.US_Standard_1976()
+ freestream = atmo.compute_values (0.)
+ freestream0 = atmo.compute_values (altitude)
+ mach_number = (cruise_speed/freestream.speed_of_sound)[0][0]
+ vehicle.design_dynamic_pressure = ( .5 *freestream0.density*(cruise_speed*cruise_speed))[0][0]
+ vehicle.design_mach_number = mach_number
+
+ # ------------------------------------------------------------------
+ # Main Wing
+ # ------------------------------------------------------------------
+ wing = SUAVE.Components.Wings.Main_Wing()
+ wing.tag = 'main_wing'
+ wing.sweeps.quarter_chord = 0.0 * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 14.76
+ wing.spans.projected = 11.4
+ wing.chords.root = 1.46
+ wing.chords.tip = 0.92
+ wing.chords.mean_aerodynamic = 1.19
+ wing.taper = wing.chords.root/wing.chords.tip
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 3.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[2.93, 0., 1.01]]
+ wing.aerodynamic_center = [3., 0., 1.01]
+ wing.vertical = False
+ wing.symmetric = True
+ wing.high_lift = True
+ wing.winglet_fraction = 0.0
+ wing.dynamic_pressure_ratio = 1.0
+ airfoil = SUAVE.Components.Airfoils.Airfoil()
+ airfoil.coordinate_file = 'Airfoils/NACA_63_412.txt'
+
+ cg_x = wing.origin[0][0] + 0.25*wing.chords.mean_aerodynamic
+ cg_z = wing.origin[0][2] - 0.2*wing.chords.mean_aerodynamic
+ vehicle.mass_properties.center_of_gravity = [[cg_x, 0. , cg_z ]] # SOURCE: Design and aerodynamic analysis of a twin-engine commuter aircraft
+
+ # Wing Segments
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'inboard'
+ segment.percent_span_location = 0.0
+ segment.twist = 3. * Units.degrees
+ segment.root_chord_percent = 1.
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'outboard'
+ segment.percent_span_location = 0.5438
+ segment.twist = 2.* Units.degrees
+ segment.root_chord_percent = 1.
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ # Wing Segments
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'winglet'
+ segment.percent_span_location = 0.98
+ segment.twist = 1. * Units.degrees
+ segment.root_chord_percent = 0.630
+ segment.dihedral_outboard = 75. * Units.degrees
+ segment.sweeps.quarter_chord = 15. * Units.degrees
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'tip'
+ segment.percent_span_location = 1.
+ segment.twist = 0. * Units.degrees
+ segment.root_chord_percent = 0.12
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = 0.12
+ segment.append_airfoil(airfoil)
+ wing.append_segment(segment)
+
+
+ aileron = SUAVE.Components.Wings.Control_Surfaces.Aileron()
+ aileron.tag = 'aileron'
+ aileron.span_fraction_start = 0.7
+ aileron.span_fraction_end = 0.9
+ aileron.deflection = 0.0 * Units.degrees
+ aileron.chord_fraction = 0.2
+ wing.append_control_surface(aileron)
+
+ flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
+ flap.tag = 'flap'
+ flap.span_fraction_start = 0.2
+ flap.span_fraction_end = 0.5
+ flap.deflection = 0.0 * Units.degrees
+ flap.chord_fraction = 0.20
+ wing.append_control_surface(flap)
+
+ # Fill out more segment properties automatically
+ wing = segment_properties(wing)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Horizontal Stabilizer
+ # ------------------------------------------------------------------
+ wing = SUAVE.Components.Wings.Wing()
+ wing.tag = 'horizontal_stabilizer'
+ wing.sweeps.quarter_chord = 0.0 * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 2.540
+ wing.spans.projected = 3.3 * Units.meter
+ wing.sweeps.quarter_chord = 0 * Units.deg
+ wing.chords.root = 0.769 * Units.meter
+ wing.chords.tip = 0.769 * Units.meter
+ wing.chords.mean_aerodynamic = 0.769 * Units.meter
+ wing.taper = 1.
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[7.7, 0., 0.25]]
+ wing.aerodynamic_center = [7.8, 0., 0.25]
+ wing.vertical = False
+ wing.winglet_fraction = 0.0
+ wing.symmetric = True
+ wing.high_lift = False
+ wing.dynamic_pressure_ratio = 0.9
+
+ elevator = SUAVE.Components.Wings.Control_Surfaces.Elevator()
+ elevator.tag = 'elevator'
+ elevator.span_fraction_start = 0.1
+ elevator.span_fraction_end = 0.9
+ elevator.deflection = 0.0 * Units.deg
+ elevator.chord_fraction = 0.3
+ wing.append_control_surface(elevator)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Vertical Stabilizer
+ # ------------------------------------------------------------------
+ wing = SUAVE.Components.Wings.Wing()
+ wing.tag = 'vertical_stabilizer'
+ wing.sweeps.quarter_chord = 25. * Units.deg
+ wing.thickness_to_chord = 0.12
+ wing.areas.reference = 2.258 * Units['meters**2']
+ wing.spans.projected = 1.854 * Units.meter
+ wing.chords.root = 1.6764 * Units.meter
+ wing.chords.tip = 0.6858 * Units.meter
+ wing.chords.mean_aerodynamic = 1.21 * Units.meter
+ wing.taper = wing.chords.tip/wing.chords.root
+ wing.aspect_ratio = wing.spans.projected**2. / wing.areas.reference
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+ wing.origin = [[6.75 ,0, 0.0]]
+ wing.aerodynamic_center = [0.508 ,0,0]
+ wing.vertical = True
+ wing.symmetric = False
+ wing.t_tail = False
+ wing.winglet_fraction = 0.0
+ wing.dynamic_pressure_ratio = 1.0
+
+ rudder = SUAVE.Components.Wings.Control_Surfaces.Rudder()
+ rudder.tag = 'rudder'
+ rudder.span_fraction_start = 0.2
+ rudder.span_fraction_end = 0.8
+ rudder.deflection = 0.0 * Units.deg
+ rudder.chord_fraction = 0.2
+ wing.append_control_surface(rudder)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Fuselage
+ # ------------------------------------------------------------------
+ fuselage = SUAVE.Components.Fuselages.Fuselage()
+ fuselage.tag = 'fuselage'
+ fuselage.seats_abreast = 2.
+ fuselage.fineness.nose = 1.6
+ fuselage.fineness.tail = 2.
+ fuselage.lengths.nose = 60. * Units.inches
+ fuselage.lengths.tail = 161. * Units.inches
+ fuselage.lengths.cabin = 105. * Units.inches
+ fuselage.lengths.total = 332.2* Units.inches
+ fuselage.lengths.fore_space = 0.
+ fuselage.lengths.aft_space = 0.
+ fuselage.width = 42. * Units.inches
+ fuselage.heights.maximum = 62. * Units.inches
+ fuselage.heights.at_quarter_length = 62. * Units.inches
+ fuselage.heights.at_three_quarters_length = 62. * Units.inches
+ fuselage.heights.at_wing_root_quarter_chord = 23. * Units.inches
+ fuselage.areas.side_projected = 8000. * Units.inches**2.
+ fuselage.areas.wetted = 30000. * Units.inches**2.
+ fuselage.areas.front_projected = 42.* 62. * Units.inches**2.
+ fuselage.effective_diameter = 50. * Units.inches
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_0'
+ segment.percent_x_location = 0
+ segment.percent_z_location = 0
+ segment.height = 0.01
+ segment.width = 0.01
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_1'
+ segment.percent_x_location = 0.007279116466
+ segment.percent_z_location = 0.002502014453
+ segment.height = 0.1669064748
+ segment.width = 0.2780205877
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_2'
+ segment.percent_x_location = 0.01941097724
+ segment.percent_z_location = 0.001216095397
+ segment.height = 0.3129496403
+ segment.width = 0.4365777215
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_3'
+ segment.percent_x_location = 0.06308567604
+ segment.percent_z_location = 0.007395489231
+ segment.height = 0.5841726619
+ segment.width = 0.6735119903
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_4'
+ segment.percent_x_location = 0.1653761217
+ segment.percent_z_location = 0.02891281352
+ segment.height = 1.064028777
+ segment.width = 1.067200529
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_5'
+ segment.percent_x_location = 0.2426372155
+ segment.percent_z_location = 0.04214148761
+ segment.height = 1.293766653
+ segment.width = 1.183058255
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_6'
+ segment.percent_x_location = 0.2960174029
+ segment.percent_z_location = 0.04705241831
+ segment.height = 1.377026712
+ segment.width = 1.181540054
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_7'
+ segment.percent_x_location = 0.3809404284
+ segment.percent_z_location = 0.05313580461
+ segment.height = 1.439568345
+ segment.width = 1.178218989
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_8'
+ segment.percent_x_location = 0.5046854083
+ segment.percent_z_location = 0.04655492473
+ segment.height = 1.29352518
+ segment.width = 1.054390707
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_9'
+ segment.percent_x_location = 0.6454149933
+ segment.percent_z_location = 0.03741966266
+ segment.height = 0.8971223022
+ segment.width = 0.8501926505
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_10'
+ segment.percent_x_location = 0.985107095
+ segment.percent_z_location = 0.04540283436
+ segment.height = 0.2920863309
+ segment.width = 0.2012565415
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_11'
+ segment.percent_x_location = 1
+ segment.percent_z_location = 0.04787575562
+ segment.height = 0.1251798561
+ segment.width = 0.1206021048
+ fuselage.Segments.append(segment)
+
+ # add to vehicle
+ vehicle.append_component(fuselage)
+
+ # ------------------------------------------------------------------
+ # Nacelles
+ # ------------------------------------------------------------------
+ nacelle = SUAVE.Components.Nacelles.Nacelle()
+ nacelle.tag = 'nacelle_1'
+ nacelle.length = 2
+ nacelle.diameter = 42 * Units.inches
+ nacelle.areas.wetted = 0.01*(2*np.pi*0.01/2)
+ nacelle.origin = [[2.5,2.5,1.0]]
+ nacelle.flow_through = False
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_1'
+ nac_segment.percent_x_location = 0.0
+ nac_segment.height = 0.0
+ nac_segment.width = 0.0
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_2'
+ nac_segment.percent_x_location = 0.1
+ nac_segment.height = 0.5
+ nac_segment.width = 0.65
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_3'
+ nac_segment.percent_x_location = 0.3
+ nac_segment.height = 0.52
+ nac_segment.width = 0.7
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_4'
+ nac_segment.percent_x_location = 0.5
+ nac_segment.height = 0.5
+ nac_segment.width = 0.65
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_5'
+ nac_segment.percent_x_location = 0.7
+ nac_segment.height = 0.4
+ nac_segment.width = 0.6
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_6'
+ nac_segment.percent_x_location = 0.9
+ nac_segment.height = 0.3
+ nac_segment.width = 0.5
+ nacelle.append_segment(nac_segment)
+
+ nac_segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ nac_segment.tag = 'segment_7'
+ nac_segment.percent_x_location = 1.0
+ nac_segment.height = 0.0
+ nac_segment.width = 0.0
+ nacelle.append_segment(nac_segment)
+
+ vehicle.append_component(nacelle)
+
+ nacelle_2 = deepcopy(nacelle)
+ nacelle_2.tag = 'nacelle_2'
+ nacelle_2.origin = [[2.5,-2.5,1.0]]
+ vehicle.append_component(nacelle_2)
+
+ #---------------------------------------------------------------------------------------------
+ # DEFINE PROPELLER
+ #---------------------------------------------------------------------------------------------
+ # build network
+ net = Battery_Propeller()
+ net.number_of_propeller_engines = 2.
+ net.identical_propellers = True
+
+ # Component 1 the ESC
+ esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
+ esc.efficiency = 0.95 # Gundlach for brushless motors
+ net.esc = esc
+
+ # Component 2 the Propeller
+ prop = SUAVE.Components.Energy.Converters.Propeller()
+ prop.tag = 'propeller_1'
+ prop.number_of_blades = 2.0
+ prop.freestream_velocity = 135.*Units['mph']
+ prop.angular_velocity = 1300. * Units.rpm
+ prop.tip_radius = 76./2. * Units.inches
+ prop.hub_radius = 8. * Units.inches
+ prop.design_Cl = 0.8
+ prop.design_altitude = 12000. * Units.feet
+ prop.design_altitude = 12000. * Units.feet
+ prop.design_thrust = 1200.
+ prop.origin = [[2.,2.5,0.784]]
+ prop.rotation = -1
+ prop.symmetry = True
+ prop.variable_pitch = True
+ prop.airfoil_geometry = ['Airfoils/NACA_4412.txt']
+ prop.airfoil_polars = [['Airfoils/Polars/NACA_4412_polar_Re_50000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_100000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_200000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_500000.txt' ,
+ 'Airfoils/Polars/NACA_4412_polar_Re_1000000.txt' ]]
+
+ prop.airfoil_polar_stations = [0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0,0]
+ prop = propeller_design(prop)
+
+ prop_left = deepcopy(prop)
+ prop_left.tag = 'propeller_2'
+ prop_left.origin = [[2.,-2.5,0.784]]
+ prop_left.rotation = 1
+
+ net.propellers.append(prop)
+ net.propellers.append(prop_left)
+
+
+ # Component 3 the Battery
+ bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion_LiFePO4_18650()
+
+ bat.mass_properties.mass = 500. * Units.kg
+ bat.max_voltage = 500.
+ initialize_from_mass(bat)
+
+ # Assume a battery pack module shape. This step is optional but
+ # required for thermal analysis of the pack
+ number_of_modules = 10
+ bat.module_config.total = int(np.ceil(bat.pack_config.total/number_of_modules))
+ bat.module_config.normal_count = int(np.ceil(bat.module_config.total/bat.pack_config.series))
+ bat.module_config.parallel_count = int(np.ceil(bat.module_config.total/bat.pack_config.parallel))
+ net.battery = bat
+
+ net.battery = bat
+ net.voltage = bat.max_voltage
+
+ # Component 4 Miscellaneous Systems
+ sys = SUAVE.Components.Systems.System()
+ sys.mass_properties.mass = 5 # kg
+
+ # Component 5 the Motor
+ motor = SUAVE.Components.Energy.Converters.Motor()
+ motor.efficiency = 0.95
+ motor.gearbox_efficiency = 1.
+ motor.origin = [[2., 2.5, 0.784]]
+ motor.nominal_voltage = bat.max_voltage *3/4
+ motor.propeller_radius = prop.tip_radius
+ motor.no_load_current = 4.0
+ motor = size_optimal_motor(motor,prop)
+ motor.mass_properties.mass = 10. * Units.kg
+
+ # append right motor
+ net.propeller_motors.append(motor)
+
+ # append left motor
+ motor_left = deepcopy(motor)
+ motor_left.origin = [[2., -2.5, 0.784]]
+ net.propeller_motors.append(motor_left)
+
+ # Component 6 the Payload
+ payload = SUAVE.Components.Energy.Peripherals.Payload()
+ payload.power_draw = 10. # Watts
+ payload.mass_properties.mass = 1.0 * Units.kg
+ net.payload = payload
+
+ # Component 7 the Avionics
+ avionics = SUAVE.Components.Energy.Peripherals.Avionics()
+ avionics.power_draw = 20. # Watts
+ net.avionics = avionics
+
+ # add the solar network to the vehicle
+ vehicle.append_component(net)
+
+ # ------------------------------------------------------------------
+ # Vehicle Definition Complete
+ # ------------------------------------------------------------------
+
+ 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)
+
+ # done!
+ return configs
+
+
+if __name__ == '__main__':
+ main()
\ No newline at end of file
diff --git a/Solar_UAV_Optimization/Analyses.py b/tut_Solar_UAV_Optimization/Analyses.py
similarity index 100%
rename from Solar_UAV_Optimization/Analyses.py
rename to tut_Solar_UAV_Optimization/Analyses.py
diff --git a/Solar_UAV_Optimization/Missions.py b/tut_Solar_UAV_Optimization/Missions.py
similarity index 100%
rename from Solar_UAV_Optimization/Missions.py
rename to tut_Solar_UAV_Optimization/Missions.py
diff --git a/Solar_UAV_Optimization/Optimize.py b/tut_Solar_UAV_Optimization/Optimize.py
similarity index 100%
rename from Solar_UAV_Optimization/Optimize.py
rename to tut_Solar_UAV_Optimization/Optimize.py
diff --git a/Solar_UAV_Optimization/Plot_Mission.py b/tut_Solar_UAV_Optimization/Plot_Mission.py
similarity index 100%
rename from Solar_UAV_Optimization/Plot_Mission.py
rename to tut_Solar_UAV_Optimization/Plot_Mission.py
diff --git a/Solar_UAV_Optimization/Procedure.py b/tut_Solar_UAV_Optimization/Procedure.py
similarity index 100%
rename from Solar_UAV_Optimization/Procedure.py
rename to tut_Solar_UAV_Optimization/Procedure.py
diff --git a/Solar_UAV_Optimization/Vehicles.py b/tut_Solar_UAV_Optimization/Vehicles.py
similarity index 100%
rename from Solar_UAV_Optimization/Vehicles.py
rename to tut_Solar_UAV_Optimization/Vehicles.py
diff --git a/tut_VSP_Import_Export_B737.py b/tut_VSP_Import_Export_B737.py
new file mode 100644
index 0000000..ac6c810
--- /dev/null
+++ b/tut_VSP_Import_Export_B737.py
@@ -0,0 +1,1034 @@
+# tut_VSP_Import_Export_B737.py
+#
+# Created: Aug 2014, SUAVE Team
+# Modified: Jun 2015, SUAVE Team
+
+""" setup file for a mission with a 737
+"""
+
+# ----------------------------------------------------------------------
+# Imports
+# ----------------------------------------------------------------------
+
+import SUAVE
+from SUAVE.Core import Units , Data, Container
+from SUAVE.Plots.Performance.Mission_Plots import *
+from SUAVE.Plots.Geometry import *
+from SUAVE.Methods.Propulsion.turbofan_sizing import turbofan_sizing
+from SUAVE.Methods.Propulsion.turbofan_sizing import turbofan_sizing
+from SUAVE.Methods.Geometry.Two_Dimensional.Planform import segment_properties
+from copy import deepcopy
+
+import vsp
+from SUAVE.Input_Output.OpenVSP.vsp_write import write
+
+import numpy as np
+import pylab as plt
+# ----------------------------------------------------------------------
+# Main
+# ----------------------------------------------------------------------
+
+def main():
+ vehicle = vsp_export_vehicle_setup()
+ write(vehicle, "B737-800")
+
+ vehicle = vsp_import_vehicle_setup()
+ return
+
+# ----------------------------------------------------------------------
+# Define the Vehicle
+# ----------------------------------------------------------------------
+
+def vsp_export_vehicle_setup():
+
+ # ------------------------------------------------------------------
+ # Initialize the Vehicle
+ # ------------------------------------------------------------------
+
+ vehicle = SUAVE.Vehicle()
+ vehicle.tag = 'Boeing_737800'
+
+ # ------------------------------------------------------------------
+ # Vehicle-level Properties
+ # ------------------------------------------------------------------
+
+ # mass properties
+ vehicle.mass_properties.max_takeoff = 79015.8 # kg
+ vehicle.mass_properties.takeoff = 79015.8 # kg
+ vehicle.mass_properties.operating_empty = 62746.4 # kg
+ vehicle.mass_properties.takeoff = 79015.8 # kg
+ vehicle.mass_properties.max_zero_fuel = 62732.0 # kg
+ vehicle.mass_properties.cargo = 10000. * Units.kilogram
+ vehicle.mass_properties.center_of_gravity = [[ 15.30987849, 0. , -0.48023939]]
+ vehicle.mass_properties.moments_of_inertia.tensor = [[3173074.17, 0 , 28752.77565],[0 , 3019041.443, 0],[0, 0, 5730017.433]] # estimated, not correct
+ vehicle.design_mach_number = 0.78
+ vehicle.design_range = 3582 * Units.miles
+ vehicle.design_cruise_alt = 35000.0 * Units.ft
+
+ # envelope properties
+ vehicle.envelope.ultimate_load = 3.75
+ vehicle.envelope.limit_load = 1.5
+
+ # basic parameters
+ vehicle.reference_area = 124.862
+ vehicle.passengers = 170
+ vehicle.systems.control = "fully powered"
+ vehicle.systems.accessories = "medium range"
+
+ # ------------------------------------------------------------------
+ # Main Wing
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Main_Wing()
+ wing.tag = 'main_wing'
+
+ wing.aspect_ratio = 10.18
+ wing.sweeps.quarter_chord = 25 * Units.deg
+ wing.thickness_to_chord = 0.1
+ wing.taper = 0.1
+
+ wing.spans.projected = 34.32
+
+ 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
+ wing.areas.wetted = 225.08
+
+ wing.twists.root = 4.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+
+ wing.origin = [[13.61,0,-0.93]]
+ wing.aerodynamic_center = [0,0,0]
+
+ wing.vertical = False
+ wing.symmetric = True
+ wing.high_lift = True
+
+ wing.dynamic_pressure_ratio = 1.0
+
+
+ # Wing Segments
+ root_airfoil = SUAVE.Components.Airfoils.Airfoil()
+ root_airfoil.coordinate_file = 'Airfoils/B737a.txt'
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'Root'
+ segment.percent_span_location = 0.0
+ segment.twist = 4. * Units.deg
+ segment.root_chord_percent = 1.
+ segment.thickness_to_chord = 0.1
+ segment.dihedral_outboard = 2.5 * Units.degrees
+ segment.sweeps.quarter_chord = 28.225 * Units.degrees
+ segment.thickness_to_chord = .1
+ segment.append_airfoil(root_airfoil)
+ wing.append_segment(segment)
+
+ yehudi_airfoil = SUAVE.Components.Airfoils.Airfoil()
+ yehudi_airfoil.coordinate_file = 'Airfoils/B737b.txt'
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'Yehudi'
+ segment.percent_span_location = 0.324
+ segment.twist = 0.047193 * Units.deg
+ segment.root_chord_percent = 0.5
+ segment.thickness_to_chord = 0.1
+ segment.dihedral_outboard = 5.5 * Units.degrees
+ segment.sweeps.quarter_chord = 25. * Units.degrees
+ segment.thickness_to_chord = .1
+ segment.append_airfoil(yehudi_airfoil)
+ wing.append_segment(segment)
+
+ mid_airfoil = SUAVE.Components.Airfoils.Airfoil()
+ mid_airfoil.coordinate_file = 'Airfoils/B737c.txt'
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'Section_2'
+ segment.percent_span_location = 0.963
+ segment.twist = 0.00258 * Units.deg
+ segment.root_chord_percent = 0.220
+ segment.thickness_to_chord = 0.1
+ segment.dihedral_outboard = 5.5 * Units.degrees
+ segment.sweeps.quarter_chord = 56.75 * Units.degrees
+ segment.thickness_to_chord = .1
+ segment.append_airfoil(mid_airfoil)
+ wing.append_segment(segment)
+
+ tip_airfoil = SUAVE.Components.Airfoils.Airfoil()
+ tip_airfoil.coordinate_file = 'Airfoils/B737d.txt'
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'Tip'
+ segment.percent_span_location = 1.
+ segment.twist = 0. * Units.degrees
+ segment.root_chord_percent = 0.10077
+ segment.thickness_to_chord = 0.1
+ segment.dihedral_outboard = 0.
+ segment.sweeps.quarter_chord = 0.
+ segment.thickness_to_chord = .1
+ segment.append_airfoil(tip_airfoil)
+ wing.append_segment(segment)
+
+ # Fill out more segment properties automatically
+ wing = segment_properties(wing)
+
+ # control surfaces -------------------------------------------
+ slat = SUAVE.Components.Wings.Control_Surfaces.Slat()
+ slat.tag = 'slat'
+ slat.span_fraction_start = 0.2
+ slat.span_fraction_end = 0.963
+ slat.deflection = 0.0 * Units.degrees
+ slat.chord_fraction = 0.075
+ wing.append_control_surface(slat)
+
+ flap = SUAVE.Components.Wings.Control_Surfaces.Flap()
+ flap.tag = 'flap'
+ flap.span_fraction_start = 0.2
+ flap.span_fraction_end = 0.7
+ flap.deflection = 0.0 * Units.degrees
+ flap.configuration_type = 'double_slotted'
+ flap.chord_fraction = 0.30
+ wing.append_control_surface(flap)
+
+ aileron = SUAVE.Components.Wings.Control_Surfaces.Aileron()
+ aileron.tag = 'aileron'
+ aileron.span_fraction_start = 0.7
+ aileron.span_fraction_end = 0.963
+ aileron.deflection = 0.0 * Units.degrees
+ aileron.chord_fraction = 0.16
+ wing.append_control_surface(aileron)
+
+
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Horizontal Stabilizer
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Horizontal_Tail()
+ wing.tag = 'horizontal_stabilizer'
+
+ wing.aspect_ratio = 4.99
+ wing.sweeps.quarter_chord = 28.2250 * Units.deg
+ wing.thickness_to_chord = 0.08
+ wing.taper = 0.3333
+
+ wing.spans.projected = 14.4
+
+ wing.chords.root = 4.2731
+ wing.chords.tip = 1.4243
+ wing.chords.mean_aerodynamic = 8.0
+
+ wing.areas.reference = 41.49
+ wing.areas.exposed = 59.354 # Exposed area of the horizontal tail
+ wing.areas.wetted = 71.81 # Wetted area of the horizontal tail
+ wing.twists.root = 3.0 * Units.degrees
+ wing.twists.tip = 3.0 * Units.degrees
+
+ wing.origin = [[33.02,0,1.466]]
+ wing.aerodynamic_center = [0,0,0]
+
+ wing.vertical = False
+ wing.symmetric = True
+
+ wing.dynamic_pressure_ratio = 0.9
+
+
+ # Wing Segments
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'root_segment'
+ segment.percent_span_location = 0.0
+ segment.twist = 0. * Units.deg
+ segment.root_chord_percent = 1.0
+ segment.dihedral_outboard = 8.63 * Units.degrees
+ segment.sweeps.quarter_chord = 28.2250 * Units.degrees
+ segment.thickness_to_chord = .1
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'tip_segment'
+ segment.percent_span_location = 1.
+ segment.twist = 0. * Units.deg
+ segment.root_chord_percent = 0.3333
+ segment.dihedral_outboard = 0 * Units.degrees
+ segment.sweeps.quarter_chord = 0 * Units.degrees
+ segment.thickness_to_chord = .1
+ wing.append_segment(segment)
+
+ # Fill out more segment properties automatically
+ wing = segment_properties(wing)
+
+ # control surfaces -------------------------------------------
+ elevator = SUAVE.Components.Wings.Control_Surfaces.Elevator()
+ elevator.tag = 'elevator'
+ elevator.span_fraction_start = 0.09
+ elevator.span_fraction_end = 0.92
+ elevator.deflection = 0.0 * Units.deg
+ elevator.chord_fraction = 0.3
+ wing.append_control_surface(elevator)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Vertical Stabilizer
+ # ------------------------------------------------------------------
+
+ wing = SUAVE.Components.Wings.Vertical_Tail()
+ wing.tag = 'vertical_stabilizer'
+
+ wing.aspect_ratio = 1.98865
+ wing.sweeps.quarter_chord = 31.2 * Units.deg
+ wing.thickness_to_chord = 0.08
+ wing.taper = 0.1183
+
+ wing.spans.projected = 8.33
+ wing.total_length = wing.spans.projected
+
+ wing.chords.root = 10.1
+ wing.chords.tip = 1.20
+ wing.chords.mean_aerodynamic = 4.0
+
+ wing.areas.reference = 34.89
+ wing.areas.wetted = 57.25
+
+ wing.twists.root = 0.0 * Units.degrees
+ wing.twists.tip = 0.0 * Units.degrees
+
+ wing.origin = [[26.944,0,1.54]]
+ wing.aerodynamic_center = [0,0,0]
+
+ wing.vertical = True
+ wing.symmetric = False
+ wing.t_tail = False
+
+ wing.dynamic_pressure_ratio = 1.0
+
+
+ # Wing Segments
+ 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 = 0 * Units.degrees
+ segment.sweeps.quarter_chord = 61.485 * Units.degrees
+ segment.thickness_to_chord = .1
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'segment_1'
+ segment.percent_span_location = 0.2962
+ segment.twist = 0. * Units.deg
+ segment.root_chord_percent = 0.45
+ segment.dihedral_outboard = 0. * Units.degrees
+ segment.sweeps.quarter_chord = 31.2 * Units.degrees
+ segment.thickness_to_chord = .1
+ wing.append_segment(segment)
+
+ segment = SUAVE.Components.Wings.Segment()
+ segment.tag = 'segment_2'
+ segment.percent_span_location = 1.0
+ segment.twist = 0. * Units.deg
+ segment.root_chord_percent = 0.1183
+ segment.dihedral_outboard = 0.0 * Units.degrees
+ segment.sweeps.quarter_chord = 0.0
+ segment.thickness_to_chord = .1
+ wing.append_segment(segment)
+
+ # Fill out more segment properties automatically
+ wing = segment_properties(wing)
+
+ # add to vehicle
+ vehicle.append_component(wing)
+
+
+ # ------------------------------------------------------------------
+ # Fuselage
+ # ------------------------------------------------------------------
+
+ fuselage = SUAVE.Components.Fuselages.Fuselage()
+ fuselage.tag = 'fuselage'
+
+ fuselage.number_coach_seats = vehicle.passengers
+ fuselage.seats_abreast = 6
+ fuselage.seat_pitch = 31. * Units.inches
+ fuselage.fineness.nose = 1.6
+ fuselage.fineness.tail = 2.
+
+ fuselage.lengths.nose = 6.4
+ fuselage.lengths.tail = 8.0
+ fuselage.lengths.cabin = 28.85
+ fuselage.lengths.total = 38.02
+ fuselage.lengths.fore_space = 6.
+ fuselage.lengths.aft_space = 5.
+
+ fuselage.width = 3.74
+
+ fuselage.heights.maximum = 3.74
+ fuselage.heights.at_quarter_length = 3.74
+ fuselage.heights.at_three_quarters_length = 3.65
+ fuselage.heights.at_wing_root_quarter_chord = 3.74
+
+ fuselage.areas.side_projected = 142.1948
+ fuselage.areas.wetted = 385.51
+ fuselage.areas.front_projected = 12.57
+
+ fuselage.effective_diameter = 3.74
+
+ fuselage.differential_pressure = 5.0e4 * Units.pascal # Maximum differential pressure
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_0'
+ segment.percent_x_location = 0.0000
+ segment.percent_z_location = -0.00144
+ segment.height = 0.0100
+ segment.width = 0.0100
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_1'
+ segment.percent_x_location = 0.00576
+ segment.percent_z_location = -0.00144
+ segment.height = 0.7500
+ segment.width = 0.6500
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_2'
+ segment.percent_x_location = 0.02017
+ segment.percent_z_location = 0.00000
+ segment.height = 1.52783
+ segment.width = 1.20043
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_3'
+ segment.percent_x_location = 0.03170
+ segment.percent_z_location = 0.00000
+ segment.height = 1.96435
+ segment.width = 1.52783
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_4'
+ segment.percent_x_location = 0.04899
+ segment.percent_z_location = 0.00431
+ segment.height = 2.72826
+ segment.width = 1.96435
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_5'
+ segment.percent_x_location = 0.07781
+ segment.percent_z_location = 0.00861
+ segment.height = 3.49217
+ segment.width = 2.61913
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_6'
+ segment.percent_x_location = 0.10375
+ segment.percent_z_location = 0.01005
+ segment.height = 3.70130
+ segment.width = 3.05565
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_7'
+ segment.percent_x_location = 0.16427
+ segment.percent_z_location = 0.01148
+ segment.height = 3.92870
+ segment.width = 3.71043
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_8'
+ segment.percent_x_location = 0.22478
+ segment.percent_z_location = 0.01148
+ segment.height = 3.92870
+ segment.width = 3.92870
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_9'
+ segment.percent_x_location = 0.69164
+ segment.percent_z_location = 0.01292
+ segment.height = 3.81957
+ segment.width = 3.81957
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_10'
+ segment.percent_x_location = 0.71758
+ segment.percent_z_location = 0.01292
+ segment.height = 3.81957
+ segment.width = 3.81957
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_11'
+ segment.percent_x_location = 0.78098
+ segment.percent_z_location = 0.01722
+ segment.height = 3.49217
+ segment.width = 3.71043
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_12'
+ segment.percent_x_location = 0.85303
+ segment.percent_z_location = 0.02296
+ segment.height = 3.05565
+ segment.width = 3.16478
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_13'
+ segment.percent_x_location = 0.91931
+ segment.percent_z_location = 0.03157
+ segment.height = 2.40087
+ segment.width = 1.96435
+ fuselage.Segments.append(segment)
+
+ # Segment
+ segment = SUAVE.Components.Lofted_Body_Segment.Segment()
+ segment.tag = 'segment_14'
+ segment.percent_x_location = 1.00
+ segment.percent_z_location = 0.04593
+ segment.height = 1.09130
+ segment.width = 0.21826
+ fuselage.Segments.append(segment)
+
+ # add to vehicle
+ vehicle.append_component(fuselage)
+
+ # ------------------------------------------------------------------
+ # Nacelles
+ # ------------------------------------------------------------------
+ nacelle = SUAVE.Components.Nacelles.Nacelle()
+ nacelle.tag = 'nacelle_1'
+ nacelle.length = 2.71
+ nacelle.inlet_diameter = 1.90
+ nacelle.diameter = 2.05
+ nacelle.areas.wetted = 1.1*np.pi*nacelle.diameter*nacelle.length
+ nacelle.origin = [[13.72, -4.86,-1.9]]
+ nacelle.flow_through = True
+ nacelle_airfoil = SUAVE.Components.Airfoils.Airfoil()
+ nacelle_airfoil.naca_4_series_airfoil = '2410'
+ nacelle.append_airfoil(nacelle_airfoil)
+
+ nacelle_2 = deepcopy(nacelle)
+ nacelle_2.tag = 'nacelle_2'
+ nacelle_2.origin = [[13.72, 4.86,-1.9]]
+
+ vehicle.append_component(nacelle)
+ vehicle.append_component(nacelle_2)
+
+
+ # ------------------------------------------------------------------
+ # Turbofan Network
+ # ------------------------------------------------------------------
+
+ #instantiate the gas turbine network
+ turbofan = SUAVE.Components.Energy.Networks.Turbofan()
+ turbofan.tag = 'turbofan'
+
+ # setup
+ turbofan.number_of_engines = 2.0
+ turbofan.bypass_ratio = 5.4
+ turbofan.engine_length = 2.71
+
+ # This origin is overwritten by compute_component_centers_of_gravity(base,compute_propulsor_origin=True)
+ turbofan.origin = [[13.72, 4.86,-1.9],[13.72, -4.86,-1.9]]
+
+ # working fluid
+ turbofan.working_fluid = SUAVE.Attributes.Gases.Air()
+
+
+ # ------------------------------------------------------------------
+ # Component 1 - Ram
+
+ # to convert freestream static to stagnation quantities
+
+ # instantiate
+ ram = SUAVE.Components.Energy.Converters.Ram()
+ ram.tag = 'ram'
+
+ # add to the network
+ turbofan.append(ram)
+
+
+ # ------------------------------------------------------------------
+ # Component 2 - Inlet Nozzle
+
+ # instantiate
+ inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
+ inlet_nozzle.tag = 'inlet_nozzle'
+
+ # setup
+ inlet_nozzle.polytropic_efficiency = 0.98
+ inlet_nozzle.pressure_ratio = 0.98
+
+ # add to network
+ turbofan.append(inlet_nozzle)
+
+
+ # ------------------------------------------------------------------
+ # Component 3 - Low Pressure Compressor
+
+ # instantiate
+ compressor = SUAVE.Components.Energy.Converters.Compressor()
+ compressor.tag = 'low_pressure_compressor'
+
+ # setup
+ compressor.polytropic_efficiency = 0.91
+ compressor.pressure_ratio = 1.14
+
+ # add to network
+ turbofan.append(compressor)
+
+
+ # ------------------------------------------------------------------
+ # Component 4 - High Pressure Compressor
+
+ # instantiate
+ compressor = SUAVE.Components.Energy.Converters.Compressor()
+ compressor.tag = 'high_pressure_compressor'
+
+ # setup
+ compressor.polytropic_efficiency = 0.91
+ compressor.pressure_ratio = 13.415
+
+ # add to network
+ turbofan.append(compressor)
+
+
+ # ------------------------------------------------------------------
+ # Component 5 - Low Pressure Turbine
+
+ # instantiate
+ turbine = SUAVE.Components.Energy.Converters.Turbine()
+ turbine.tag='low_pressure_turbine'
+
+ # setup
+ turbine.mechanical_efficiency = 0.99
+ turbine.polytropic_efficiency = 0.93
+
+ # add to network
+ turbofan.append(turbine)
+
+
+ # ------------------------------------------------------------------
+ # Component 6 - High Pressure Turbine
+
+ # instantiate
+ turbine = SUAVE.Components.Energy.Converters.Turbine()
+ turbine.tag='high_pressure_turbine'
+
+ # setup
+ turbine.mechanical_efficiency = 0.99
+ turbine.polytropic_efficiency = 0.93
+
+ # add to network
+ turbofan.append(turbine)
+
+
+ # ------------------------------------------------------------------
+ # Component 7 - Combustor
+
+ # instantiate
+ combustor = SUAVE.Components.Energy.Converters.Combustor()
+ combustor.tag = 'combustor'
+
+ # setup
+ combustor.efficiency = 0.99
+ combustor.alphac = 1.0
+ combustor.turbine_inlet_temperature = 1450
+ combustor.pressure_ratio = 0.95
+ combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
+
+ # add to network
+ turbofan.append(combustor)
+
+
+ # ------------------------------------------------------------------
+ # Component 8 - Core Nozzle
+
+ # instantiate
+ nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
+ nozzle.tag = 'core_nozzle'
+
+ # setup
+ nozzle.polytropic_efficiency = 0.95
+ nozzle.pressure_ratio = 0.99
+
+ # add to network
+ turbofan.append(nozzle)
+
+
+ # ------------------------------------------------------------------
+ # Component 9 - Fan Nozzle
+
+ # instantiate
+ nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
+ nozzle.tag = 'fan_nozzle'
+
+ # setup
+ nozzle.polytropic_efficiency = 0.95
+ nozzle.pressure_ratio = 0.99
+
+ # add to network
+ turbofan.append(nozzle)
+
+
+ # ------------------------------------------------------------------
+ # Component 10 - Fan
+
+ # instantiate
+ fan = SUAVE.Components.Energy.Converters.Fan()
+ fan.tag = 'fan'
+
+ # setup
+ fan.polytropic_efficiency = 0.93
+ fan.pressure_ratio = 1.7
+
+ # add to network
+ turbofan.append(fan)
+
+
+ # ------------------------------------------------------------------
+ #Component 10 : thrust (to compute the thrust)
+ thrust = SUAVE.Components.Energy.Processes.Thrust()
+ thrust.tag ='compute_thrust'
+
+ #total design thrust (includes all the engines)
+ thrust.total_design = 2*24000. * Units.N #Newtons
+
+ #design sizing conditions
+ altitude = 35000.0*Units.ft
+ mach_number = 0.78
+ isa_deviation = 0.
+
+ #Engine setup for noise module
+
+
+ # add to network
+ turbofan.thrust = thrust
+
+ turbofan.core_nozzle_diameter = 0.92
+ turbofan.fan_nozzle_diameter = 1.659
+ turbofan.engine_height = 0.5 #Engine centerline heigh above the ground plane
+ turbofan.exa = 1 #distance from fan face to fan exit/ fan diameter)
+ turbofan.plug_diameter = 0.1 #dimater of the engine plug
+ turbofan.geometry_xe = 1. # Geometry information for the installation effects function
+ turbofan.geometry_ye = 1. # Geometry information for the installation effects function
+ turbofan.geometry_Ce = 2. # Geometry information for the installation effects function
+
+
+
+
+
+ #size the turbofan
+ turbofan_sizing(turbofan,mach_number,altitude)
+
+ # add gas turbine network turbofan to the vehicle
+ vehicle.append_component(turbofan)
+
+ # ------------------------------------------------------------------
+ # Fuel
+ # ------------------------------------------------------------------
+ fuel = SUAVE.Components.Physical_Component()
+ vehicle.fuel = fuel
+ fuel.mass_properties.mass = vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_fuel
+ fuel.origin = vehicle.wings.main_wing.mass_properties.center_of_gravity
+ fuel.mass_properties.center_of_gravity= vehicle.wings.main_wing.aerodynamic_center
+
+ # ------------------------------------------------------------------
+ # Landing Gear
+ # ------------------------------------------------------------------
+ landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
+ landing_gear.tag = "main_landing_gear"
+ landing_gear.main_tire_diameter = 1.12000 * Units.m
+ landing_gear.nose_tire_diameter = 0.6858 * Units.m
+ landing_gear.main_strut_length = 1.8 * Units.m
+ landing_gear.nose_strut_length = 1.3 * Units.m
+ landing_gear.main_units = 1 #number of nose landing gear
+ landing_gear.nose_units = 1 #number of nose landing gear
+ landing_gear.main_wheels = 2 #number of wheels on the main landing gear
+ landing_gear.nose_wheels = 2 #number of wheels on the nose landing gear
+ vehicle.landing_gear = landing_gear
+
+ # ------------------------------------------------------------------
+ # Vehicle Definition Complete
+ # ------------------------------------------------------------------
+
+ return vehicle
+
+
+def vsp_import_vehicle_setup():
+
+
+
+
+ # ------------------------------------------------------------------
+ # Turbofan Network
+ # ------------------------------------------------------------------
+
+ #instantiate the gas turbine network
+ turbofan = SUAVE.Components.Energy.Networks.Turbofan()
+ turbofan.tag = 'turbofan'
+
+ # setup
+ turbofan.number_of_engines = 2.0
+ turbofan.bypass_ratio = 5.4
+ turbofan.engine_length = 2.71
+
+ # This origin is overwritten by compute_component_centers_of_gravity(base,compute_propulsor_origin=True)
+ turbofan.origin = [[13.72, 4.86,-1.9],[13.72, -4.86,-1.9]]
+
+ # working fluid
+ turbofan.working_fluid = SUAVE.Attributes.Gases.Air()
+
+
+ # ------------------------------------------------------------------
+ # Component 1 - Ram
+
+ # to convert freestream static to stagnation quantities
+
+ # instantiate
+ ram = SUAVE.Components.Energy.Converters.Ram()
+ ram.tag = 'ram'
+
+ # add to the network
+ turbofan.append(ram)
+
+
+ # ------------------------------------------------------------------
+ # Component 2 - Inlet Nozzle
+
+ # instantiate
+ inlet_nozzle = SUAVE.Components.Energy.Converters.Compression_Nozzle()
+ inlet_nozzle.tag = 'inlet_nozzle'
+
+ # setup
+ inlet_nozzle.polytropic_efficiency = 0.98
+ inlet_nozzle.pressure_ratio = 0.98
+
+ # add to network
+ turbofan.append(inlet_nozzle)
+
+
+ # ------------------------------------------------------------------
+ # Component 3 - Low Pressure Compressor
+
+ # instantiate
+ compressor = SUAVE.Components.Energy.Converters.Compressor()
+ compressor.tag = 'low_pressure_compressor'
+
+ # setup
+ compressor.polytropic_efficiency = 0.91
+ compressor.pressure_ratio = 1.14
+
+ # add to network
+ turbofan.append(compressor)
+
+
+ # ------------------------------------------------------------------
+ # Component 4 - High Pressure Compressor
+
+ # instantiate
+ compressor = SUAVE.Components.Energy.Converters.Compressor()
+ compressor.tag = 'high_pressure_compressor'
+
+ # setup
+ compressor.polytropic_efficiency = 0.91
+ compressor.pressure_ratio = 13.415
+
+ # add to network
+ turbofan.append(compressor)
+
+
+ # ------------------------------------------------------------------
+ # Component 5 - Low Pressure Turbine
+
+ # instantiate
+ turbine = SUAVE.Components.Energy.Converters.Turbine()
+ turbine.tag='low_pressure_turbine'
+
+ # setup
+ turbine.mechanical_efficiency = 0.99
+ turbine.polytropic_efficiency = 0.93
+
+ # add to network
+ turbofan.append(turbine)
+
+
+ # ------------------------------------------------------------------
+ # Component 6 - High Pressure Turbine
+
+ # instantiate
+ turbine = SUAVE.Components.Energy.Converters.Turbine()
+ turbine.tag='high_pressure_turbine'
+
+ # setup
+ turbine.mechanical_efficiency = 0.99
+ turbine.polytropic_efficiency = 0.93
+
+ # add to network
+ turbofan.append(turbine)
+
+
+ # ------------------------------------------------------------------
+ # Component 7 - Combustor
+
+ # instantiate
+ combustor = SUAVE.Components.Energy.Converters.Combustor()
+ combustor.tag = 'combustor'
+
+ # setup
+ combustor.efficiency = 0.99
+ combustor.alphac = 1.0
+ combustor.turbine_inlet_temperature = 1450
+ combustor.pressure_ratio = 0.95
+ combustor.fuel_data = SUAVE.Attributes.Propellants.Jet_A()
+
+ # add to network
+ turbofan.append(combustor)
+
+
+ # ------------------------------------------------------------------
+ # Component 8 - Core Nozzle
+
+ # instantiate
+ nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
+ nozzle.tag = 'core_nozzle'
+
+ # setup
+ nozzle.polytropic_efficiency = 0.95
+ nozzle.pressure_ratio = 0.99
+
+ # add to network
+ turbofan.append(nozzle)
+
+
+ # ------------------------------------------------------------------
+ # Component 9 - Fan Nozzle
+
+ # instantiate
+ nozzle = SUAVE.Components.Energy.Converters.Expansion_Nozzle()
+ nozzle.tag = 'fan_nozzle'
+
+ # setup
+ nozzle.polytropic_efficiency = 0.95
+ nozzle.pressure_ratio = 0.99
+
+ # add to network
+ turbofan.append(nozzle)
+
+
+ # ------------------------------------------------------------------
+ # Component 10 - Fan
+
+ # instantiate
+ fan = SUAVE.Components.Energy.Converters.Fan()
+ fan.tag = 'fan'
+
+ # setup
+ fan.polytropic_efficiency = 0.93
+ fan.pressure_ratio = 1.7
+
+ # add to network
+ turbofan.append(fan)
+
+
+ # ------------------------------------------------------------------
+ #Component 10 : thrust (to compute the thrust)
+ thrust = SUAVE.Components.Energy.Processes.Thrust()
+ thrust.tag ='compute_thrust'
+
+ #total design thrust (includes all the engines)
+ thrust.total_design = 2*24000. * Units.N #Newtons
+
+ #design sizing conditions
+ altitude = 35000.0*Units.ft
+ mach_number = 0.78
+ isa_deviation = 0.
+
+ #Engine setup for noise module
+
+
+ # add to network
+ turbofan.thrust = thrust
+
+ turbofan.core_nozzle_diameter = 0.92
+ turbofan.fan_nozzle_diameter = 1.659
+ turbofan.engine_height = 0.5 #Engine centerline heigh above the ground plane
+ turbofan.exa = 1 #distance from fan face to fan exit/ fan diameter)
+ turbofan.plug_diameter = 0.1 #dimater of the engine plug
+ turbofan.geometry_xe = 1. # Geometry information for the installation effects function
+ turbofan.geometry_ye = 1. # Geometry information for the installation effects function
+ turbofan.geometry_Ce = 2. # Geometry information for the installation effects function
+
+
+
+
+
+ #size the turbofan
+ turbofan_sizing(turbofan,mach_number,altitude)
+
+ # add gas turbine network turbofan to the vehicle
+ vehicle.append_component(turbofan)
+
+ # ------------------------------------------------------------------
+ # Fuel
+ # ------------------------------------------------------------------
+ fuel = SUAVE.Components.Physical_Component()
+ vehicle.fuel = fuel
+ fuel.mass_properties.mass = vehicle.mass_properties.max_takeoff-vehicle.mass_properties.max_fuel
+ fuel.origin = vehicle.wings.main_wing.mass_properties.center_of_gravity
+ fuel.mass_properties.center_of_gravity= vehicle.wings.main_wing.aerodynamic_center
+
+ # ------------------------------------------------------------------
+ # Landing Gear
+ # ------------------------------------------------------------------
+ landing_gear = SUAVE.Components.Landing_Gear.Landing_Gear()
+ landing_gear.tag = "main_landing_gear"
+ landing_gear.main_tire_diameter = 1.12000 * Units.m
+ landing_gear.nose_tire_diameter = 0.6858 * Units.m
+ landing_gear.main_strut_length = 1.8 * Units.m
+ landing_gear.nose_strut_length = 1.3 * Units.m
+ landing_gear.main_units = 1 #number of nose landing gear
+ landing_gear.nose_units = 1 #number of nose landing gear
+ landing_gear.main_wheels = 2 #number of wheels on the main landing gear
+ landing_gear.nose_wheels = 2 #number of wheels on the nose landing gear
+ vehicle.landing_gear = landing_gear
+
+ # ------------------------------------------------------------------
+ # Vehicle Definition Complete
+ # ------------------------------------------------------------------
+
+ return vehicle
+
+
+if __name__ == '__main__':
+ main()
+ plt.show()
+
+
+
+