Merge pull request #340 from OpenSourceAWE/agent/153-remove-the-parameter-body-aero-from-the-

Build a Solver from VSMSettings or from panel and section counts, deprecate the body_aero constructors, and check the sizes in solve!

Author: 1-Bart-1

CHANGELOG.md

@@ -4,6 +4,8 @@
 
 ### Added
 
+- `Solver(settings)` and `Solver(n_panels, n_unrefined_sections)` build a solver without
+  a `BodyAerodynamics`; keyword arguments override the settings.
 - `set_va!(body_aero, va_vec, omega; reference_point)` turns the body about
   `reference_point` [m] instead of the origin. The point is stored on
   `BodyAerodynamics`, starts at the origin, and is kept by later `set_va!`, `reinit!`
@@ -24,9 +26,15 @@
 - The Makie `plot!` methods for a `Panel` or a `BodyAerodynamics` return a
   `Vector{Makie.AbstractPlot}` instead of a `Vector{Any}`; for a `BodyAerodynamics`
   drawn as flat panels it is one flat list rather than a list per panel.
+- `Solver(body_aero; kwargs...)` and `Solver(body_aero, settings)` are deprecated and warn
+  on use; build the solver with `Solver(settings)` or
+  `Solver(n_panels, n_unrefined_sections)` instead.
 
 ### Fixed
 
+- `solve!` and `solve` throw a `DimensionMismatch` naming both sizes for a `body_aero` whose
+  panel or unrefined-section count differs from the solver's, where they failed on a
+  broadcast partway through or silently left section results at zero.
 - `set_va!(body_aero, settings)` applies `condition.yaw_rate` as a turn rate about the
   body z axis; it was read from the settings file and ignored.
 - The `VSMSolution` docstring gives `lift_dist`, `drag_dist` and `panel_moment_dist` in

data/TUDELFT_V3_KITE/vsm_settings_coarse.yaml

@@ -44,7 +44,7 @@ condition:
 # Define wing geometry files and discretization parameters
 wings:
   - name: V3_Kite # Wing identifier for output labeling
-    geometry_file: data/TUDELFT_V3_KITE/aero_geometry.yaml
+    geometry_file: data/TUDELFT_V3_KITE/aero_geometry_coarse_discretisation.yaml
     n_panels: 54 # Total number of panels along wingspan
     spanwise_panel_distribution: SPLIT_PROVIDED # Panel spacing algorithm
     spanwise_direction: [0.0, 1.0, 0.0] # Unit vector defining wingspan direction

docs/src/examples.md

@@ -84,8 +84,8 @@ julia> set_va!(body_aero, va_vec, [0, 0, 0.1])
 #### Step 5: Initialize solvers for both LLT and VSM methods
 
 ```julia
-julia> llt_solver = Solver(body_aero; aerodynamic_model_type=LLT)
-julia> vsm_solver = Solver(body_aero; aerodynamic_model_type=VSM)
+julia> llt_solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=LLT)
+julia> vsm_solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=VSM)
 ```
 
 #### Step 6: Solve using both methods

docs/src/private_functions.md

@@ -12,6 +12,9 @@ settings_range
 
 ### Solver, forces and circulation
 ```@docs
+n_unrefined_sections
+solver_kwargs
+check_dimensions
 calculate_AIC_matrices!
 gamma_loop!
 build_spanwise_laplacian!

docs/src/settings.md

@@ -11,7 +11,7 @@ settings = VSMSettings("my/vsm_settings.yaml"; data_prefix=false)  # as written
 
 wing = Wing(settings)
 body_aero = BodyAerodynamics([wing])
-solver = Solver(body_aero, settings)
+solver = Solver(settings)
 set_va!(body_aero, settings)
 ```
 

examples/V3_kite.jl

@@ -42,7 +42,7 @@ wing = Wing(settings)
 refine!(wing)
 body_aero = BodyAerodynamics([wing])
 VortexStepMethod.reinit!(body_aero)
-solver = Solver(body_aero, settings)
+solver = Solver(settings)
 solver.reference_point .= REFERENCE_POINT
 
 if DEFORM
@@ -75,7 +75,7 @@ if NEURALFOIL
     refine!(wing_nf)
     body_nf = BodyAerodynamics([wing_nf])
     VortexStepMethod.reinit!(body_nf)
-    solver_nf = Solver(body_nf, settings_nf)
+    solver_nf = Solver(settings_nf)
     solver_nf.reference_point .= REFERENCE_POINT
 
     # Reading the generated directory instead of the OBJ shows the airfoils the polar

examples/V3_neuralfoil.jl

@@ -81,14 +81,18 @@ wing_cfd = Wing(settings_cfd)
 refine!(wing_cfd)
 body_cfd = BodyAerodynamics([wing_cfd])
 VortexStepMethod.reinit!(body_cfd)
-solver_cfd = Solver(body_cfd, settings_cfd)
+solver_cfd = Solver(settings_cfd)
 
 println("Creating wing with NeuralFoil polars...")
 wing_nf = Wing(nf_yaml; n_panels=50, spanwise_distribution=LINEAR)
 refine!(wing_nf)
 body_nf = BodyAerodynamics([wing_nf])
 VortexStepMethod.reinit!(body_nf)
-solver_nf = Solver(body_nf, settings_cfd)
+settings_nf = VSMSettings("TUDELFT_V3_KITE/vsm_settings.yaml")
+settings_nf.wings[1].geometry_file = nf_yaml
+settings_nf.solver_settings.relaxation_factor = RELAXATION
+settings_nf.solver_settings.artificial_damping = ARTIFICIAL_DAMPING
+solver_nf = Solver(settings_nf)
 
 # Compare CFD-polar and NeuralFoil-polar wings against published references
 # (Poland 2025 RANS CFD and wind tunnel). `plot_polars` sweeps each solver over the

examples/bench.jl

@@ -41,8 +41,8 @@ va_vec = [cos(alpha), 0.0, sin(alpha)] .* va
 set_va!(body_aero, va_vec)
 
 # Step 4: Initialize solvers for both LLT and VSM methods
-llt_solver = Solver(body_aero; aerodynamic_model_type=LLT)
-vsm_solver = Solver(body_aero; aerodynamic_model_type=VSM)
+llt_solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=LLT)
+vsm_solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=VSM)
 
 # Step 5: Solve using both methods
 results_vsm = solve(vsm_solver, body_aero, nothing)
@@ -66,7 +66,7 @@ body_aero = BodyAerodynamics([wing])
 
 # Create solvers
 vsm_solver = Solver(
-    body_aero;
+    wing.n_panels, wing.n_unrefined_sections;
     aerodynamic_model_type=VSM,
     is_with_artificial_damping=false,
     solver_type=LOOP,

examples/billowing.jl

@@ -30,15 +30,12 @@ literature_paths = [
         "windtunnel_alpha_sweep_beta_00_0_Poland_2025_Rey_5e5.csv"),
 ]
 
-# Load solver settings (coarse: 54 panels, matches 10-section geometry)
-settings_data = VortexStepMethod.YAML.load_file(
-    joinpath(v3_dir, "vsm_settings_coarse.yaml"))
-condition_cfg = settings_data["condition"]
-solver_cfg = settings_data["solver_settings"]
-wing_cfg = settings_data["wings"][1]
-n_panels = wing_cfg["n_panels"]
+# Coarse settings: 54 panels on the 10-section geometry
+settings = VSMSettings(joinpath(v3_dir, "vsm_settings_coarse.yaml"); data_prefix=false)
+settings.wings[1].geometry_file = joinpath(project_dir, settings.wings[1].geometry_file)
+n_panels = settings.wings[1].n_panels
 
-BILLOWING_PCT = get(wing_cfg, "billowing_percentage", 0.0)
+BILLOWING_PCT = settings.wings[1].billowing_percentage
 
 labels = [
     "VSM flat",
@@ -49,9 +46,7 @@ labels = [
     "WindTunnel Re=5e5",
 ]
 
-# Load coarse geometry (10 structural rib sections)
-geom_data = VortexStepMethod.YAML.load_file(
-    joinpath(v3_dir, "aero_geometry_coarse_discretisation.yaml"))
+geom_data = VortexStepMethod.YAML.load_file(settings.wings[1].geometry_file)
 section_headers = geom_data["wing_sections"]["headers"]
 section_rows = geom_data["wing_sections"]["data"]
 
@@ -84,43 +79,13 @@ body_aero_bill = BodyAerodynamics([wing_bill])
 VortexStepMethod.reinit!(body_aero_bill)
 
 # --- Build solvers ---
-function make_solver(body_aero)
-    Solver(body_aero;
-        solver_type=(solver_cfg["solver_type"] == "NONLIN" ?
-            NONLIN : LOOP),
-        aerodynamic_model_type=getproperty(
-            VortexStepMethod,
-            Symbol(solver_cfg["aerodynamic_model_type"])),
-        density=solver_cfg["density"],
-        max_iterations=solver_cfg["max_iterations"],
-        rtol=solver_cfg["rtol"],
-        tol_reference_error=solver_cfg["tol_reference_error"],
-        relaxation_factor=solver_cfg["relaxation_factor"],
-        is_with_artificial_damping=solver_cfg["artificial_damping"],
-        artificial_damping=(
-            k2=solver_cfg["k2"], k4=solver_cfg["k4"]),
-        type_initial_gamma_distribution=getproperty(
-            VortexStepMethod,
-            Symbol(solver_cfg["type_initial_gamma_distribution"])),
-        use_gamma_prev=get(solver_cfg, "use_gamma_prev",
-                           get(solver_cfg, "use_gamme_prev", true)),
-        core_radius_fraction=solver_cfg["core_radius_fraction"],
-        mu=solver_cfg["mu"],
-        is_only_f_and_gamma_output=get(
-            solver_cfg, "calc_only_f_and_gamma", false),
-        correct_aoa=get(solver_cfg, "correct_aoa", false),
-        reference_point=get(solver_cfg, "reference_point",
-                            [0.422646, 0.0, 9.3667]),
-    )
-end
-
-solver_flat = make_solver(body_aero_flat)
-solver_bill = make_solver(body_aero_bill)
+solver_flat = Solver(settings; reference_point=[0.422646, 0.0, 9.3667])
+solver_bill = Solver(settings; reference_point=[0.422646, 0.0, 9.3667])
 
 # --- Set flight conditions ---
-wind_speed = condition_cfg["wind_speed"]
+wind_speed = settings.condition.wind_speed
 angle_of_attack_deg = 10.0
-sideslip_deg = condition_cfg["beta"]
+sideslip_deg = settings.condition.beta
 
 α0 = deg2rad(angle_of_attack_deg)
 β0 = deg2rad(sideslip_deg)

examples/linearize_check.jl

@@ -23,7 +23,7 @@ yaml = obj_to_yaml(
 wing = Wing(yaml; n_panels=16)
 body_aero = BodyAerodynamics([wing])
 
-solver = Solver(body_aero;
+solver = Solver(wing.n_panels, wing.n_unrefined_sections;
     aerodynamic_model_type=VSM,
     is_with_artificial_damping=false,
     rtol=1e-7,