Rename wa to body_aero

Author: 1-Bart-1

README.md

@@ -117,7 +117,7 @@ add_section!(wing,
     INVISCID)
 
 # Step 3: Initialize aerodynamics
-wa = BodyAerodynamics([wing])
+body_aero = BodyAerodynamics([wing])
 
 # Set inflow conditions
 vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a

docs/src/examples.md

@@ -54,19 +54,19 @@ add_section!(wing,
 
 #### Step 4: Initialize aerodynamics
 ```
-wa = BodyAerodynamics([wing])
+body_aero = BodyAerodynamics([wing])
 ```
 We need to pass here an array of wing objects, because a body can have multiple wings.
 
 ###### Set inflow conditions
 ```julia
 vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
-set_va!(wa, vel_app, [0, 0, 0.1])
+set_va!(body_aero, vel_app, [0, 0, 0.1])
 ```
 #### Step 5: Plot the geometry
 ```julia
 plot_geometry(
-      wa,
+      body_aero,
       "Rectangular_wing_geometry";
       data_type=".pdf",
       save_path=".",
@@ -86,8 +86,8 @@ vsm_solver = Solver(aerodynamic_model_type=VSM)
 
 #### Step 7: Solve using both methods
 ```
-results_llt = solve(llt_solver, wa)
-results_vsm = solve(vsm_solver, wa)
+results_llt = solve(llt_solver, body_aero)
+results_vsm = solve(vsm_solver, body_aero)
 ```
 
 ##### Print results comparison
@@ -103,7 +103,7 @@ println("Projected area = $(round(results_vsm["projected_area"], digits=4)) m²"
 
 #### Step 8: Plot spanwise distributions
 ```julia
-y_coordinates = [panel.aero_center[2] for panel in wa.panels]
+y_coordinates = [panel.aero_center[2] for panel in body_aero.panels]
 
 plot_distribution(
     [y_coordinates, y_coordinates],
@@ -121,7 +121,7 @@ You should see a plot like this:
 angle_range = range(0, 20, 20)
 plot_polars(
     [llt_solver, vsm_solver],
-    [wa, wa],
+    [body_aero, body_aero],
     ["LLT", "VSM"];
     angle_range,
     angle_type="angle_of_attack",

docs/src/index.md

@@ -115,7 +115,7 @@ add_section!(wing,
     INVISCID)
 
 # Step 3: Initialize aerodynamics
-wa = BodyAerodynamics([wing])
+body_aero = BodyAerodynamics([wing])
 
 # Set inflow conditions
 vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a

examples/bench.jl

@@ -32,29 +32,29 @@ add_section!(wing,
     INVISCID)
 
 # Step 3: Initialize aerodynamics
-wa = BodyAerodynamics([wing])
+body_aero = BodyAerodynamics([wing])
 
 # Set inflow conditions
 vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
-set_va!(wa, vel_app)
+set_va!(body_aero, vel_app)
 
 # Step 4: Initialize solvers for both LLT and VSM methods
-P = length(wa.panels)
-llt_solver = Solver{P}(aerodynamic_model_type=LLT)
-vsm_solver = Solver{P}(aerodynamic_model_type=VSM)
+P = length(body_aero.panels)
+llt_solver = Solver(body_aero; aerodynamic_model_type=LLT)
+vsm_solver = Solver(body_aero; aerodynamic_model_type=VSM)
 
 # Step 5: Solve using both methods
-results_vsm = solve(vsm_solver, wa)
-sol = solve!(vsm_solver, wa)
-results_vsm_base = solve_base!(vsm_solver, wa)
+results_vsm = solve(vsm_solver, body_aero)
+sol = solve!(vsm_solver, body_aero)
+results_vsm_base = solve_base!(vsm_solver, body_aero)
 println("Rectangular wing, solve_base!:")
-@time results_vsm_base = solve_base!(vsm_solver, wa)
+@time results_vsm_base = solve_base!(vsm_solver, body_aero)
 # time Python: 32.0  ms Ryzen 7950x
 # time Julia:   0.42 ms Ryzen 7950x
 println("Rectangular wing, solve!:")
-@time sol = solve!(vsm_solver, wa)
+@time sol = solve!(vsm_solver, body_aero)
 println("Rectangular wing, solve:")
-@time solve(vsm_solver, wa)
+@time solve(vsm_solver, body_aero)
 
 # Create wing geometry
 wing = RamAirWing("data/ram_air_kite_body.obj", "data/ram_air_kite_foil.dat")

examples/ram_air_kite.jl

@@ -22,7 +22,7 @@ if DEFORM
 end
 
 # Create solvers
-vsm_solver = Solver(body_aero;
+solver = Solver(body_aero;
     aerodynamic_model_type=VSM,
     is_with_artificial_damping=false,
     rtol=1e-5,
@@ -45,7 +45,7 @@ set_va!(body_aero, vel_app)
 if LINEARIZE
     println("Linearize")
     jac, res = VortexStepMethod.linearize(
-        vsm_solver, 
+        solver, 
         body_aero, 
         wing, 
         [zeros(4); vel_app; zeros(3)]; 
@@ -54,7 +54,7 @@ if LINEARIZE
         omega_idxs=8:10,
         moment_frac=0.1)
     @time jac, res = VortexStepMethod.linearize(
-        vsm_solver, 
+        solver, 
         body_aero, 
         wing, 
         [zeros(4); vel_app; zeros(3)]; 
@@ -82,8 +82,8 @@ PLOT && plot_geometry(
 
 # Solving and plotting distributions
 println("Solve")
-results = VortexStepMethod.solve(vsm_solver, body_aero; log=true)
-@time results = solve(vsm_solver, body_aero; log=true)
+results = VortexStepMethod.solve(solver, body_aero; log=true)
+@time results = solve(solver, body_aero; log=true)
 
 body_y_coordinates = [panel.aero_center[2] for panel in body_aero.panels]
 
@@ -99,7 +99,7 @@ PLOT && plot_distribution(
 )
 
 PLOT && plot_polars(
-    [vsm_solver],
+    [solver],
     [body_aero],
     [
         "VSM from Ram Air Kite OBJ and DAT file",

examples/rectangular_wing.jl

@@ -28,22 +28,22 @@ add_section!(wing,
     INVISCID)
 
 # Step 3: Initialize aerodynamics
-wa = BodyAerodynamics([wing])
+body_aero = BodyAerodynamics([wing])
 
 # Set inflow conditions
 vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
-set_va!(wa, vel_app, [0, 0, 0.1])
+set_va!(body_aero, vel_app, [0, 0, 0.1])
 
 # Step 4: Initialize solvers for both LLT and VSM methods
-P = length(wa.panels)
-llt_solver = Solver{P}(aerodynamic_model_type=LLT)
-vsm_solver = Solver{P}(aerodynamic_model_type=VSM)
+P = length(body_aero.panels)
+llt_solver = Solver(body_aero; aerodynamic_model_type=LLT)
+vsm_solver = Solver(body_aero; aerodynamic_model_type=VSM)
 
 # Step 5: Solve using both methods
-results_llt = solve(llt_solver, wa)
-@time results_llt = solve(llt_solver, wa)
-results_vsm = solve(vsm_solver, wa)
-@time results_vsm = solve(vsm_solver, wa)
+results_llt = solve(llt_solver, body_aero)
+@time results_llt = solve(llt_solver, body_aero)
+results_vsm = solve(vsm_solver, body_aero)
+@time results_vsm = solve(vsm_solver, body_aero)
 
 # Print results comparison
 println("\nLifting Line Theory Results:")
@@ -56,7 +56,7 @@ println("Projected area = $(round(results_vsm["projected_area"], digits=4)) m²"
 
 # Step 6: Plot geometry
 PLOT && plot_geometry(
-      wa,
+      body_aero,
       "Rectangular_wing_geometry";
       data_type=".pdf",
       save_path=".",
@@ -66,7 +66,7 @@ PLOT && plot_geometry(
 )
 
 # Step 7: Plot spanwise distributions
-y_coordinates = [panel.aero_center[2] for panel in wa.panels]
+y_coordinates = [panel.aero_center[2] for panel in body_aero.panels]
 
 PLOT && plot_distribution(
     [y_coordinates, y_coordinates],
@@ -80,7 +80,7 @@ PLOT && plot_distribution(
 angle_range = range(0, 20, 20)
 PLOT && plot_polars(
     [llt_solver, vsm_solver],
-    [wa, wa],
+    [body_aero, body_aero],
     ["LLT", "VSM"];
     angle_range,
     angle_type="angle_of_attack",

examples/stall_model.jl

@@ -42,15 +42,15 @@ 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
-body_aero_CAD_19ribs = BodyAerodynamics([CAD_wing])
+body_aero = BodyAerodynamics([CAD_wing])
 
 # Create solvers
-P = length(body_aero_CAD_19ribs.panels)
-vsm_solver = Solver{P}(
+P = length(body_aero.panels)
+vsm_solver = Solver(body_aero; 
     aerodynamic_model_type=VSM,
     is_with_artificial_damping=false
 )
-VSM_with_stall_correction = Solver{P}(
+VSM_with_stall_correction = Solver(body_aero; 
     aerodynamic_model_type=VSM,
     is_with_artificial_damping=true
 )
@@ -66,11 +66,11 @@ vel_app = [
     sin(side_slip),
     sin(aoa_rad)
 ] * v_a
-set_va!(body_aero_CAD_19ribs, vel_app)
+set_va!(body_aero, vel_app)
 
 # Plotting geometry
 PLOT && plot_geometry(
-    body_aero_CAD_19ribs,
+    body_aero,
     "";
     data_type=".svg",
     save_path="",
@@ -82,11 +82,11 @@ PLOT && plot_geometry(
 )
 
 # Solving and plotting distributions
-results = solve(vsm_solver, body_aero_CAD_19ribs)
-@time results_with_stall = solve(VSM_with_stall_correction, body_aero_CAD_19ribs)
-@time results_with_stall = solve(VSM_with_stall_correction, body_aero_CAD_19ribs)
+results = solve(vsm_solver, body_aero)
+@time results_with_stall = solve(VSM_with_stall_correction, body_aero)
+@time results_with_stall = solve(VSM_with_stall_correction, body_aero)
 
-CAD_y_coordinates = [panel.aero_center[2] for panel in body_aero_CAD_19ribs.panels]
+CAD_y_coordinates = [panel.aero_center[2] for panel in body_aero.panels]
 
 PLOT && plot_distribution(
     [CAD_y_coordinates, CAD_y_coordinates],
@@ -112,7 +112,7 @@ path_cfd_lebesque = joinpath(
 
 PLOT && plot_polars(
     [vsm_solver, VSM_with_stall_correction],
-    [body_aero_CAD_19ribs, body_aero_CAD_19ribs],
+    [body_aero, body_aero],
     [
         "VSM CAD 19ribs",
         "VSM CAD 19ribs , with stall correction",

src/body_aerodynamics.jl

@@ -67,7 +67,8 @@ function BodyAerodynamics(
             section.LE_point .-= kite_body_origin
             section.TE_point .-= kite_body_origin
         end
-        if wing.spanwise_distribution == UNCHANGED
+        if wing.spanwise_distribution == UNCHANGED && !(wing.n_panels == length(wing.sections) - 1)
+            @warn "Changing wing.n_panels from $(wing.n_panels) to $(length(wing.sections) - 1)"
             wing.n_panels = length(wing.sections) - 1
             wing.refined_sections = wing.sections
         else

src/kite_geometry.jl

@@ -435,12 +435,15 @@ Constructor for a [RamAirWing](@ref) that allows to use an `.obj` and a `.dat` f
 - `spanwise_direction`=[0.0, 1.0, 0.0]
 - `remove_nan::Bool`: Wether to remove the NaNs from interpolations or not
 """
-function RamAirWing(obj_path, dat_path; crease_frac=0.9, wind_vel=10., mass=1.0, 
-                  n_panels=56, n_sections=n_panels+1, n_groups=n_panels÷4, spanwise_distribution=UNCHANGED, 
-                  spanwise_direction=[0.0, 1.0, 0.0], remove_nan=true, align_to_principal=false,
-                  alpha_range=deg2rad.(-5:1:20), delta_range=deg2rad.(-5:1:20), interp_steps=n_sections
-                  )
-
+function RamAirWing(
+    obj_path, dat_path; 
+    crease_frac=0.9, wind_vel=10., mass=1.0, 
+    n_panels=56, n_sections=n_panels+1, n_groups=4, spanwise_distribution=UNCHANGED, 
+    spanwise_direction=[0.0, 1.0, 0.0], remove_nan=true, align_to_principal=false,
+    alpha_range=deg2rad.(-5:1:20), delta_range=deg2rad.(-5:1:20), interp_steps=n_sections # TODO: check if interpolations are still needed
+)
+
+    !(n_panels % n_groups == 0) && throw(ArgumentError("Number of panels should be divisible by number of groups"))
     !isapprox(spanwise_direction, [0.0, 1.0, 0.0]) && throw(ArgumentError("Spanwise direction has to be [0.0, 1.0, 0.0], not $spanwise_direction"))
 
     # Load or create polars