reduce allocations more (#44)

* reduce allocations more

* add dat file

* add bench file

* improve test

* rename Umag to v_a

* cant remove calculate_velocity_induced_single_ring_semiinfinite! allocs

* two non-allocating functions in code_bench

* add simple script for alloc testing

* fix calculations

* three non-allocating functions

* make tests pass again

* non-allocating gamma loop

* more efficient polars

* use symbols instead of strings

* switch to Symbol

* tests pass again

* working test bench

* make tests and examples work

* remove unused dependency

* add bench.jl

* add bench to the menu

* less output

* fix comment

* add plot option to example

---------

Co-authored-by: Uwe Fechner <[email protected]>

Author: 1-Bart-1

.gitignore

@@ -5,4 +5,3 @@ results/TUDELFT_V3_LEI_KITE/polars/$C_L$ vs $C_D$.pdf
 *.bin
 results/TUDELFT_V3_LEI_KITE/polars/tutorial_testing_stall_model_n_panels_54_distribution_split_provided.pdf
 docs/build/
-

examples/bench.jl

@@ -0,0 +1,46 @@
+using LinearAlgebra
+using ControlPlots
+using VortexStepMethod
+
+plot = true
+
+# Step 1: Define wing parameters
+n_panels = 20          # Number of panels
+span = 20.0            # Wing span [m]
+chord = 1.0            # Chord length [m]
+v_a = 20.0            # Magnitude of inflow velocity [m/s]
+density = 1.225        # Air density [kg/m³]
+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)
+
+# Add wing sections - defining only tip sections with inviscid airfoil model
+add_section!(wing, 
+    [0.0, span/2, 0.0],   # Left tip LE 
+    [chord, span/2, 0.0],  # Left tip TE
+    :inviscid)
+add_section!(wing, 
+    [0.0, -span/2, 0.0],  # Right tip LE
+    [chord, -span/2, 0.0], # Right tip TE
+    :inviscid)
+
+# Step 3: Initialize aerodynamics
+wa = BodyAerodynamics([wing])
+
+# Set inflow conditions
+vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
+set_va!(wa, (vel_app, 0.0))  # Second parameter is yaw rate
+
+# Step 4: Initialize solvers for both LLT and VSM methods
+llt_solver = Solver(aerodynamic_model_type=:LLT)
+vsm_solver = Solver(aerodynamic_model_type=:VSM)
+
+# Step 5: Solve using both methods
+results_vsm = solve(vsm_solver, wa)
+@time results_vsm = solve(vsm_solver, wa)
+# time Python: 32.0 ms
+# time Julia:   0.7 ms
+
+nothing

examples/menu.jl

@@ -3,6 +3,7 @@ using REPL.TerminalMenus
 options = ["rectangular_wing = include(\"rectangular_wing.jl\")",
            "ram_air_kite = include(\"ram_air_kite.jl\")",
            "stall_model = include(\"stall_model.jl\")",
+           "bench = include(\"bench.jl\")",
            "quit"]
 
 function example_menu()

examples/ram_air_kite.jl

@@ -8,17 +8,19 @@ end
 using CSV
 using DataFrames
 
+plot = true
+
 # Create wing geometry
 wing = KiteWing("data/ram_air_kite_body.obj", "data/ram_air_kite_foil.dat")
 body_aero = BodyAerodynamics([wing])
 
 # Create solvers
 VSM = Solver(
-    aerodynamic_model_type="VSM",
+    aerodynamic_model_type=:VSM,
     is_with_artificial_damping=false
 )
 VSM_with_stall_correction = Solver(
-    aerodynamic_model_type="VSM",
+    aerodynamic_model_type=:VSM,
     is_with_artificial_damping=true
 )
 
@@ -36,7 +38,7 @@ vel_app = [
 body_aero.va = vel_app
 
 # Plotting geometry
-plot_geometry(
+plot && plot_geometry(
     body_aero,
     "";
     data_type=".svg",
@@ -53,7 +55,7 @@ plot_geometry(
 
 CAD_y_coordinates = [panel.aerodynamic_center[2] for panel in body_aero.panels]
 
-plot_distribution(
+plot && plot_distribution(
     [CAD_y_coordinates],
     [results],
     ["VSM"];
@@ -63,7 +65,7 @@ plot_distribution(
     is_show=true
 )
 
-plot_polars(
+plot && plot_polars(
     [VSM],
     [body_aero],
     [

examples/rectangular_wing.jl

@@ -2,6 +2,8 @@ using LinearAlgebra
 using ControlPlots
 using VortexStepMethod
 
+plot = true
+
 # Step 1: Define wing parameters
 n_panels = 20          # Number of panels
 span = 20.0            # Wing span [m]
@@ -12,17 +14,17 @@ 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_panel_distribution=:linear)
 
 # Add wing sections - defining only tip sections with inviscid airfoil model
 add_section!(wing, 
     [0.0, span/2, 0.0],   # Left tip LE 
     [chord, span/2, 0.0],  # Left tip TE
-    "inviscid")
+    :inviscid)
 add_section!(wing, 
     [0.0, -span/2, 0.0],  # Right tip LE
     [chord, -span/2, 0.0], # Right tip TE
-    "inviscid")
+    :inviscid)
 
 # Step 3: Initialize aerodynamics
 wa = BodyAerodynamics([wing])
@@ -32,16 +34,16 @@ vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
 set_va!(wa, (vel_app, 0.0))  # Second parameter is yaw rate
 
 # Step 4: Initialize solvers for both LLT and VSM methods
-llt_solver = Solver(aerodynamic_model_type="LLT")
-vsm_solver = Solver(aerodynamic_model_type="VSM")
+llt_solver = Solver(aerodynamic_model_type=:LLT)
+vsm_solver = Solver(aerodynamic_model_type=:VSM)
 
 # Step 5: Solve using both methods
+results_llt = solve(llt_solver, wa)
 @time results_llt = solve(llt_solver, wa)
-@time results_llt = solve(llt_solver, wa)
-@time results_vsm = solve(vsm_solver, wa)
+results_vsm = solve(vsm_solver, wa)
 @time results_vsm = solve(vsm_solver, wa)
 # time Python: 32.0ms
-# time Julia:   1.5ms
+# time Julia:   0.7ms
 
 # Print results comparison
 println("\nLifting Line Theory Results:")
@@ -53,20 +55,19 @@ println("CD = $(round(results_vsm["cd"], digits=4))")
 println("Projected area = $(round(results_vsm["projected_area"], digits=4)) m²")
 
 # Step 6: Plot geometry
-plot_geometry(
+plot && plot_geometry(
       wa,
       "Rectangular_wing_geometry";
       data_type=".pdf",
       save_path=".",
       is_save=false,
       is_show=true,
 )
-nothing
 
 # Step 7: Plot spanwise distributions
 y_coordinates = [panel.aerodynamic_center[2] for panel in wa.panels]
 
-plot_distribution(
+plot && plot_distribution(
     [y_coordinates, y_coordinates],
     [results_vsm, results_llt],
     ["VSM", "LLT"],
@@ -75,7 +76,7 @@ plot_distribution(
 
 # Step 8: Plot polar curves
 angle_range = range(0, 20, 20)
-plot_polars(
+plot && plot_polars(
     [llt_solver, vsm_solver],
     [wa, wa],
     ["LLT", "VSM"];

examples/stall_model.jl

@@ -8,14 +8,16 @@ end
 using CSV
 using DataFrames
 
+plot = true
+
 # Find root directory
 root_dir = dirname(@__DIR__)
 save_folder = joinpath(root_dir, "results", "TUDELFT_V3_LEI_KITE")
 mkpath(save_folder)
 
 # Defining discretisation
 n_panels = 54
-spanwise_panel_distribution = "split_provided"
+spanwise_panel_distribution = :split_provided
 
 # Load rib data from CSV
 csv_file_path = joinpath(
@@ -30,7 +32,7 @@ rib_list = []
 for row in eachrow(df)
     LE = [row.LE_x, row.LE_y, row.LE_z]
     TE = [row.TE_x, row.TE_y, row.TE_z]
-    push!(rib_list, (LE, TE, ("lei_airfoil_breukels", [row.d_tube, row.camber])))
+    push!(rib_list, (LE, TE, (:lei_airfoil_breukels, [row.d_tube, row.camber])))
 end
 
 # Create wing geometry
@@ -42,11 +44,11 @@ body_aero_CAD_19ribs = BodyAerodynamics([CAD_wing])
 
 # Create solvers
 VSM = Solver(
-    aerodynamic_model_type="VSM",
+    aerodynamic_model_type=:VSM,
     is_with_artificial_damping=false
 )
 VSM_with_stall_correction = Solver(
-    aerodynamic_model_type="VSM",
+    aerodynamic_model_type=:VSM,
     is_with_artificial_damping=true
 )
 
@@ -63,17 +65,17 @@ vel_app = [
 ] * v_a
 body_aero_CAD_19ribs.va = vel_app
 
-# # Plotting geometry
-# plot_geometry(
-#     body_aero_CAD_19ribs,
-#     "";
-#     data_type=".svg",
-#     save_path="",
-#     is_save=false,
-#     is_show=true,
-#     view_elevation=15,
-#     view_azimuth=-120
-# )
+# Plotting geometry
+plot && plot_geometry(
+    body_aero_CAD_19ribs,
+    "";
+    data_type=".svg",
+    save_path="",
+    is_save=false,
+    is_show=true,
+    view_elevation=15,
+    view_azimuth=-120
+)
 
 # Solving and plotting distributions
 results = solve(VSM, body_aero_CAD_19ribs)
@@ -82,7 +84,7 @@ results = solve(VSM, body_aero_CAD_19ribs)
 
 CAD_y_coordinates = [panel.aerodynamic_center[2] for panel in body_aero_CAD_19ribs.panels]
 
-plot_distribution(
+plot && plot_distribution(
     [CAD_y_coordinates, CAD_y_coordinates],
     [results, results_with_stall],
     ["VSM", "VSM with stall correction"];
@@ -103,7 +105,7 @@ path_cfd_lebesque = joinpath(
     "V3_CL_CD_RANS_Lebesque_2024_Rey_300e4.csv"
 )
 
-plot_polars(
+plot && plot_polars(
     [VSM, VSM_with_stall_correction],
     [body_aero_CAD_19ribs, body_aero_CAD_19ribs],
     [