Updating tutorial scripts

Author: rachealerhard

Regional_Jet_Optimization/Optimize.py

@@ -27,13 +27,17 @@ def main():
 
     ## Base Input Values
     #output = problem.objective()
+    import time
+    t = time.time()
 
-    ## Uncomment to view contours of the design space
-    #variable_sweep(problem)
+    # Uncomment to view contours of the design space
+    variable_sweep(problem)
+    
+    t_elapsed = time.time() - t
 
     ## Uncomment for the first optimization
-    output = scipy_setup.SciPy_Solve(problem,solver='SLSQP')
-    print (output)    
+    #output = scipy_setup.SciPy_Solve(problem,solver='SLSQP')
+    #print (output)    
 
     ## Uncomment these lines when you want to start an optimization problem from a different initial guess
     #inputs                                   = [1.28, 1.38]
@@ -64,10 +68,10 @@ def setup():
     # Inputs
     # -------------------------------------------------------------------
 
-    #   [ tag                            , initial, (lb,ub)             , scaling , units ]
+    #   [ tag                   , initial,        lb,ub)        , scaling , units ]
     problem.inputs = np.array([
-        [ 'wing_area'                    ,  100    , (   90. ,   130.   ) ,   100. , Units.meter**2],
-        [ 'cruise_altitude'              ,  11    , (   9   ,    14.   ) ,   10.  , Units.km],
+        [ 'wing_area'           ,  100   , (   90. ,   130.   ) ,   100.  , Units.meter**2],
+        [ 'cruise_altitude'     ,  11    , (   9   ,    14.   ) ,   10.   , Units.km],
     ])
 
     # -------------------------------------------------------------------
@@ -137,19 +141,18 @@ def variable_sweep(problem):
     objective   = outputs.objective
     constraints = outputs.constraint_val
     plt.figure(0)
-    CS   = plt.contourf(inputs[0,:],inputs[1,:], objective, 20, linewidths=2)
+    CS   = plt.contourf(inputs[0,:],inputs[1,:], objective, 20, linewidths=2,cmap='jet')
     cbar = plt.colorbar(CS)
 
     cbar.ax.set_ylabel('fuel burn (kg)')
-    CS_const = plt.contour(inputs[0,:],inputs[1,:], constraints[0,:,:])
+    CS_const = plt.contour(inputs[0,:],inputs[1,:], constraints[0,:,:],cmap='jet')
     plt.clabel(CS_const, inline=1, fontsize=10)
     cbar = plt.colorbar(CS_const)
     cbar.ax.set_ylabel('fuel margin')
 
     plt.xlabel('Wing Area (m^2)')
     plt.ylabel('Cruise Altitude (km)')
 
-    plt.legend(loc='upper left')  
     plt.show(block=True)    
 
     return

Regional_Jet_Optimization/Procedure.py

@@ -198,12 +198,12 @@ def weight(nexus):
     vehicle=nexus.vehicle_configurations.base
 
     # weight analysis
-    weights = nexus.analyses.base.weights.evaluate()
-    weights = nexus.analyses.cruise.weights.evaluate()
+    weights = nexus.analyses.base.weights.evaluate(method="SUAVE")
+    weights = nexus.analyses.cruise.weights.evaluate(method="SUAVE")
     vehicle.mass_properties.breakdown = weights
-    weights = nexus.analyses.landing.weights.evaluate()
-    weights = nexus.analyses.takeoff.weights.evaluate()
-    weights = nexus.analyses.short_field_takeoff.weights.evaluate()
+    weights = nexus.analyses.landing.weights.evaluate(method="SUAVE")
+    weights = nexus.analyses.takeoff.weights.evaluate(method="SUAVE")
+    weights = nexus.analyses.short_field_takeoff.weights.evaluate(method="SUAVE")
 
     for config in nexus.vehicle_configurations:
         config.mass_properties.zero_fuel_center_of_gravity  = vehicle.mass_properties.zero_fuel_center_of_gravity

tut_mission_B737.py

@@ -36,10 +36,6 @@
 # covered in more detail in a separate tutorial. It does not size the turbofan geometry.
 
 
-# Python Imports
-import numpy as np
-import matplotlib.pyplot as plt  
-
 # ----------------------------------------------------------------------
 #   Main
 # ----------------------------------------------------------------------
@@ -256,7 +252,7 @@ def vehicle_setup():
     wing.areas.reference         = 124.862 * Units['meters**2']  
     wing.twists.root             = 4.0 * Units.degrees
     wing.twists.tip              = 0.0 * Units.degrees
-    wing.origin                  = [[13.61 * Units.meter, 0, -1.27 * Units.meter]]
+    wing.origin                  = [[13.61, 0, -1.27]] * Units.meter
     wing.vertical                = False
     wing.symmetric               = True
     # The high lift flag controls aspects of maximum lift coefficient calculations
@@ -310,12 +306,12 @@ def vehicle_setup():
     wing.tag = 'horizontal_stabilizer'
 
     wing.aspect_ratio            = 6.16     
-    wing.sweeps.quarter_chord    = 40 * Units.deg
+    wing.sweeps.quarter_chord    = 40.0 * Units.deg
     wing.thickness_to_chord      = 0.08
     wing.taper                   = 0.2
     wing.spans.projected         = 14.2 * Units.meter
     wing.chords.root             = 4.7  * Units.meter
-    wing.chords.tip              = .955 * Units.meter
+    wing.chords.tip              = 0.955 * Units.meter
     wing.chords.mean_aerodynamic = 3.0  * Units.meter
     wing.areas.reference         = 32.488   * Units['meters**2']  
     wing.twists.root             = 3.0 * Units.degrees
@@ -413,7 +409,7 @@ def vehicle_setup():
     turbofan.bypass_ratio      = 5.4
     turbofan.engine_length     = 2.71 * Units.meter
     turbofan.nacelle_diameter  = 2.05 * Units.meter
-    turbofan.origin            = [[13.72, 4.86,-1.9],[13.72, -4.86,-1.9]]
+    turbofan.origin            = [[13.72, 4.86,-1.9],[13.72, -4.86,-1.9]] * Units.meter
 
     # Approximate the wetted area
     turbofan.areas.wetted      = 1.1*np.pi*turbofan.nacelle_diameter*turbofan.engine_length
@@ -565,19 +561,20 @@ def vehicle_setup():
     turbofan.append(fan)
 
     # ------------------------------------------------------------------
-    #Component 10 : thrust (to compute the thrust)
+    #  Component 11 - thrust (to compute the thrust)
+    
     thrust = SUAVE.Components.Energy.Processes.Thrust()       
     thrust.tag ='compute_thrust'
 
     # Design thrust is used to determine mass flow at full throttle
     thrust.total_design             = 2*24000. * Units.N #Newtons
- 
-    # Design sizing conditions are also used to determine mass flow
-    altitude      = 35000.0*Units.ft
-    mach_number   = 0.78 
 
     # Add to network
     turbofan.thrust = thrust
+    
+    # Design sizing conditions are also used to determine mass flow
+    altitude      = 35000.0*Units.ft
+    mach_number   = 0.78     
 
     # Determine turbofan behavior at the design condition
     turbofan_sizing(turbofan,mach_number,altitude)   

tut_payload_range.py

@@ -113,7 +113,7 @@ def base_analysis(vehicle):
     #  Aerodynamics Analysis
     aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
     aerodynamics.geometry = vehicle
-    aerodynamics.settings.drag_coefficient_increment = 0.0050
+    aerodynamics.settings.drag_coefficient_increment = 0.000
     analyses.append(aerodynamics)
 
     # ------------------------------------------------------------------