Improve init!(body_aero) performance (#139)

* Rename panel distribution

* Runs with less allocs

* Tests pass

* Low allocations on updating panels

* Remove unimplemented keyword arguments

* Non-allocating init pos

* Reduce proj area allocs

* Non-allocating proj area

* Add reinit allocation tests

* Fix test

* Fix test

* Really fix tests

* Fix docs

Author: 1-Bart-1

README.md

@@ -104,7 +104,7 @@ alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
 
 # Step 2: Create wing geometry with linear panel distribution
-wing = Wing(n_panels, spanwise_panel_distribution=LINEAR)
+wing = Wing(n_panels, spanwise_distribution=LINEAR)
 
 # Add wing sections - defining only tip sections with inviscid airfoil model
 add_section!(wing, 

docs/src/examples.md

@@ -37,7 +37,7 @@ alpha = deg2rad(alpha_deg)
 
 #### Step 3: Create wing geometry with linear panel distribution
 ```julia
-wing = Wing(n_panels, spanwise_panel_distribution=LINEAR)
+wing = Wing(n_panels, spanwise_distribution=LINEAR)
 ```
 
 ##### Add wing sections - defining only tip sections with inviscid airfoil model

docs/src/index.md

@@ -102,7 +102,7 @@ alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
 
 # Step 2: Create wing geometry with linear panel distribution
-wing = Wing(n_panels, spanwise_panel_distribution=LINEAR)
+wing = Wing(n_panels, spanwise_distribution=LINEAR)
 
 # Add wing sections - defining only tip sections with inviscid airfoil model
 add_section!(wing, 

docs/src/private_functions.md

@@ -5,7 +5,7 @@ CurrentModule = VortexStepMethod
 ## Private Functions
 ```@docs
 calculate_AIC_matrices!
-calculate_panel_properties
+update_panel_properties!
 calculate_inertia_tensor
 init!
 ```

docs/src/types.md

@@ -30,11 +30,11 @@ A Wing/ RamAirWing has one or more sections.
 Section
 Section(LE_point::Vector{Float64}, TE_point::Vector{Float64}, aero_model=nothing, aero_data=nothing)
 Wing
-Wing(n_panels::Int; spanwise_panel_distribution::PanelDistribution=LINEAR,
+Wing(n_panels::Int; spanwise_distribution::PanelDistribution=LINEAR,
      spanwise_direction::PosVector=MVec3([0.0, 1.0, 0.0]))
 RamAirWing
 RamAirWing(obj_path, dat_path; alpha=0.0, crease_frac=0.75, wind_vel=10., mass=1.0, 
-         n_panels=54, n_sections=n_panels+1, spanwise_panel_distribution=UNCHANGED, 
+         n_panels=54, n_sections=n_panels+1, spanwise_distribution=UNCHANGED, 
          spanwise_direction=[0.0, 1.0, 0.0])
 BodyAerodynamics
 ```

examples/bench.jl

@@ -19,7 +19,7 @@ alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
 
 # Step 2: Create wing geometry with linear panel distribution
-wing = Wing(n_panels, spanwise_panel_distribution=LINEAR)
+wing = Wing(n_panels, spanwise_distribution=LINEAR)
 
 # Add wing sections - defining only tip sections with inviscid airfoil model
 add_section!(wing, 

examples/ram_air_kite.jl

@@ -2,13 +2,15 @@ using ControlPlots
 using VortexStepMethod
 using LinearAlgebra
 
-PLOT = true
+PLOT = false
 USE_TEX = false
-DEFORM = false
+DEFORM = true
 
 # Create wing geometry
 wing = RamAirWing("data/ram_air_kite_body.obj", "data/ram_air_kite_foil.dat")
 body_aero = BodyAerodynamics([wing];)
+println("First init")
+@time VortexStepMethod.init!(body_aero)
 
 if DEFORM
     # Linear interpolation of alpha from 10° at one tip to 0° at the other
@@ -19,8 +21,10 @@ if DEFORM
     delta_end = deg2rad(0)
     theta_dist = [theta_start - i * (theta_start - theta_end)/(n_panels-1) for i in 0:(n_panels-1)]
     delta_dist = [delta_start - i * (delta_start - delta_end)/(n_panels-1) for i in 0:(n_panels-1)]
+    println("Deform")
     @time VortexStepMethod.deform!(wing, theta_dist, delta_dist)
-    @time VortexStepMethod.init!(body_aero)
+    println("Deform init")
+    @time VortexStepMethod.init!(body_aero; init_aero=false)
 end
 
 # Create solvers
@@ -86,7 +90,7 @@ PLOT && plot_polars(
     angle_of_attack=0,
     side_slip=0,
     v_a=10,
-    title="ram_kite_panels_$(wing.n_panels)_distribution_$(wing.spanwise_panel_distribution)",
+    title="ram_kite_panels_$(wing.n_panels)_distribution_$(wing.spanwise_distribution)",
     data_type=".pdf",
     is_save=false,
     is_show=true,

examples/rectangular_wing.jl

@@ -15,7 +15,7 @@ alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
 
 # Step 2: Create wing geometry with linear panel distribution
-wing = Wing(n_panels, spanwise_panel_distribution=LINEAR)
+wing = Wing(n_panels, spanwise_distribution=LINEAR)
 
 # Add wing sections - defining only tip sections with inviscid airfoil model
 add_section!(wing, 

examples/stall_model.jl

@@ -19,7 +19,7 @@ mkpath(save_folder)
 
 # Defining discretisation
 n_panels = 54
-spanwise_panel_distribution = SPLIT_PROVIDED
+spanwise_distribution = SPLIT_PROVIDED
 
 # Load rib data from CSV
 csv_file_path = joinpath(
@@ -38,7 +38,7 @@ for row in eachrow(df)
 end
 
 # Create wing geometry
-CAD_wing = Wing(n_panels; spanwise_panel_distribution)
+CAD_wing = Wing(n_panels; spanwise_distribution)
 for rib in rib_list
     add_section!(CAD_wing, rib[1], rib[2], rib[3], rib[4])
 end
@@ -124,7 +124,7 @@ PLOT && plot_polars(
     angle_of_attack=0,
     side_slip=0,
     v_a=10,
-    title="tutorial_testing_stall_model_n_panels_$(n_panels)_distribution_$(spanwise_panel_distribution)",
+    title="tutorial_testing_stall_model_n_panels_$(n_panels)_distribution_$(spanwise_distribution)",
     data_type=".pdf",
     save_path=joinpath(save_folder, "polars"),
     is_save=true,

mwes/mwe_03.jl

@@ -15,7 +15,7 @@ alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
 
 # Step 2: Create wing geometry with linear panel distribution
-wing = Wing(n_panels, spanwise_panel_distribution=LINEAR)
+wing = Wing(n_panels, spanwise_distribution=LINEAR)
 
 # Add wing sections - defining only tip sections with inviscid airfoil model
 add_section!(wing,