diff --git a/CHANGELOG.md b/CHANGELOG.md
index 3db61498..b817ee1f 100644
--- a/CHANGELOG.md
+++ b/CHANGELOG.md
@@ -15,9 +15,24 @@
   is not drawn, and still names the saved file and window.
 - The panel plots (`plot!(ax, panel)`, `plot!(ax, body)`, `plot(panel)`, `plot(body)`)
   take `border_color`, default `:black`, for the panel outlines.
+- `relaxation_factor_limit` gives the largest stable relaxation factor of the `LOOP` solver,
+  and `VSMSolution.f_attached_trailed` the attached trailed part of each panel's force.
 
 ### Changed
 
+- Default results now agree with the Python VSM v2.3.0, most visibly on swept and curved
+  wings: the circulation is `0.5 |V_eff| c Cl`; the panel chord and chordwise axis are
+  square to the bound vortex; `correct_aoa` and `is_with_attached_trailed_force` default to
+  `true`; the flow turning the forces includes a curved bound vortex correction; drag acts
+  along that flow; each wake ring follows its own panel's inflow; trailing vortex cores grow
+  with the distance from their line; `q_ref` and the reference directions leave out body
+  rates; attached trailed forces are shared per section boundary; `cd` beyond a
+  `POLAR_VECTORS` table is held, not extrapolated. V3 kite at 8°: `cl` 1.051 → 1.099, `cd`
+  0.491 → 0.458.
+- `max_iterations` defaults to 5000 and `relaxation_factor` to `nothing`, meaning 0.8 of
+  `relaxation_factor_limit`, which is lower than the old 0.03 on narrow or high-slope panels.
+- `*_unrefined_dist` coefficients and angles are area-weighted means of their panels.
+- `velocity_3D_trailing_vortex_semiinfinite!` returns zero on its own axis, not `NaN`.
 - BREAKING: `solve` and `calculate_results` are removed; use `solve!`, which returns the
   solver's `VSMSolution`, and `solve!(solver, body_aero, nothing)` to start from a fresh
   circulation as `solve` did. Keys that became other fields: `Fx`/`Fy`/`Fz` → `force`,
diff --git a/data/TUDELFT_V3_KITE/vsm_settings.yaml b/data/TUDELFT_V3_KITE/vsm_settings.yaml
index 5477aed9..50a8cf56 100644
--- a/data/TUDELFT_V3_KITE/vsm_settings.yaml
+++ b/data/TUDELFT_V3_KITE/vsm_settings.yaml
@@ -53,7 +53,7 @@ solver_settings:
     # --- Core Aerodynamic Model ---
     aerodynamic_model_type: VSM       # VSM=3D vortex method, LLT=lifting line theory
     solver_type: LOOP                 # LOOP=fixed-point iteration, NEWTON=Newton-Raphson
-    #correct_aoa: false               # Defaults to false, meaning the computed AoA at 3/4c is used
+    #correct_aoa: true                # false turns the forces by the flow at 3/4c instead of 1/4c
     
     # --- Physical Properties ---
     density: 1.225                   # [kg/m³] Air density (ISA sea level)
diff --git a/data/TUDELFT_V3_KITE/vsm_settings_coarse.yaml b/data/TUDELFT_V3_KITE/vsm_settings_coarse.yaml
index d3fb98e8..a53486df 100644
--- a/data/TUDELFT_V3_KITE/vsm_settings_coarse.yaml
+++ b/data/TUDELFT_V3_KITE/vsm_settings_coarse.yaml
@@ -57,7 +57,7 @@ solver_settings:
   # --- Core Aerodynamic Model ---
   aerodynamic_model_type: VSM # VSM=3D vortex method, LLT=lifting line theory
   solver_type: LOOP # LOOP=fixed-point iteration, NONLIN=nonlinear solver
-  #correct_aoa: false               # Defaults to false, meaning the computed AoA at 3/4c is used
+  #correct_aoa: true                # false turns the forces by the flow at 3/4c instead of 1/4c
 
   # --- Physical Properties ---
   density: 1.225 # [kg/m³] Air density (ISA sea level)
@@ -79,4 +79,3 @@ solver_settings:
   # --- Output Control ---
   calc_only_f_and_gamma: false # true=compute only forces & circulation
 
-  correct_aoa: false
diff --git a/docs/src/private_functions.md b/docs/src/private_functions.md
index c70033dc..7a407118 100644
--- a/docs/src/private_functions.md
+++ b/docs/src/private_functions.md
@@ -33,9 +33,15 @@ nose_down
 bisect_sign_change
 calculate_relative_alpha_and_relative_velocity
 update_effective_angle_of_attack!
+quarter_chord_AIC!
+curved_bound_correction!
+unrolled_ring_difference
+unit_bound_velocity
+unit_trailing_velocity
+thin_airfoil_strips
+relaxation_factor_limit
 induced_velocity_at
-attached_trailed_loads
-panel_force_moment
+attached_trailed_loads!
 calculate_stall_angle_list
 wing_span_flip
 spanwise_extent
@@ -64,6 +70,8 @@ dynamic_pressure
 flow_curvature_cm
 spanwise_flow_drag
 panel_force_directions
+panel_chord
+perpendicular_part
 prescribed_va_directions
 analysis_fields!
 find_center_of_pressure!
diff --git a/src/body_aerodynamics.jl b/src/body_aerodynamics.jl
index 948d5668..19643748 100644
--- a/src/body_aerodynamics.jl
+++ b/src/body_aerodynamics.jl
@@ -24,6 +24,10 @@ Main structure for calculating aerodynamic properties of bodies. Use the constru
 - `AIC::Array{Float64, 3}` = zeros(P, P, 3): control-point influence coefficients, the
                         matrix the circulation is solved against; component last so that
                         each `AIC[:, :, k]` slice is a contiguous BLAS matrix
+- `AIC_curved_correction::Array{Float64, 3}` = zeros(P, P, 3): stored
+    [`curved_bound_correction!`](@ref)
+- `curved_correction_core_radius` = NaN: core radius fraction `AIC_curved_correction` was
+    built for, `NaN` until built for the current shape
 - `AIC_aero_center::Array{Float64, 3}` = zeros(P, P, 3): aerodynamic-centre (LLT)
                         influence coefficients, used only for the corrected angle of attack
 - `projected_area::Float64` = 1.0: The area projected onto the xy-plane of the KA body frame [m²]
@@ -44,9 +48,11 @@ Main structure for calculating aerodynamic properties of bodies. Use the constru
     alpha_dist::MVector{P, T} = zeros(MVector{P, T})
     v_rel_dist::MVector{P, T} = zeros(MVector{P, T})
     pitch_rate_dist::MVector{P, T} = zeros(MVector{P, T})
-    work_vectors::NTuple{10, MVector{3, T}} = ntuple(_ -> zeros(MVector{3, T}), 10)
+    work_vectors::NTuple{11, MVector{3, T}} = ntuple(_ -> zeros(MVector{3, T}), 11)
     AIC::Array{T, 3} = zeros(T, P, P, 3)
     AIC_aero_center::Array{T, 3} = zeros(T, P, P, 3)
+    AIC_curved_correction::Array{T, 3} = zeros(T, P, P, 3)
+    curved_correction_core_radius::T = T(NaN)
     projected_area::T = one(T)
     c_ref::T = one(T)
     cache::Vector{PreallocationTools.LazyBufferCache{typeof(identity), typeof(identity)}} = [LazyBufferCache()]
@@ -360,6 +366,7 @@ function reinit!(body_aero::BodyAerodynamics{P, W, T};
     body_aero.alpha_dist .= 0.0
     body_aero.v_rel_dist .= 0.0
     body_aero.AIC .= 0.0
+    body_aero.curved_correction_core_radius = T(NaN)
     set_va!(body_aero, va_vec, omega)
     return nothing
 end
@@ -444,32 +451,18 @@ Returns: nothing
     evaluation_point = model == VSM ? :control_point : :aero_center
     evaluation_point_on_bound = model == LLT
 
-    # Allocate work vectors for this function (separate from those used by child functions)
-    velocity_induced = zeros(MVector{3, T})
-    tempvel = zeros(MVector{3, T})
-    va_unit = zeros(MVector{3, T})
-    U_2D = zeros(MVector{3, T})
-
-    # Python parity: one shared area-weighted wake vector for all panels.
-    panel_areas = [panel.chord * panel.width for panel in body_aero.panels]
-    va_vec_dist = zeros(T, length(body_aero.panels), 3)
-    @inbounds for i in 1:length(body_aero.panels), k in 1:3
-        va_vec_dist[i, k] = va_unit_dist[i, k] * va_dist[i]
-    end
-    wake_velocity = _compute_reference_velocity_from_distribution(
-        va_vec_dist,
-        length(body_aero.panels),
-        panel_areas
-    )
-    wake_speed = norm(wake_velocity)
-    wake_speed > 0.0 || throw(ArgumentError("Wake reference speed must be positive."))
-    va_unit .= wake_velocity ./ wake_speed
-    va = wake_speed
-    
+    # Scratch separate from the vectors the filament kernels use
+    work = body_aero.work_vectors
+    velocity_induced, tempvel, va_unit, U_2D = work[8], work[9], work[10], work[11]
+
+    all(>(0), va_dist) || throw(ArgumentError(
+        "Wake reference speed must be positive on every panel."))
+
     # Calculate influence coefficients
     for jring in eachindex(body_aero.panels)
         panel_jring = body_aero.panels[jring]
         filaments = panel_jring.filaments
+        va_unit .= matrix_row(va_unit_dist, jring)
         for icp in eachindex(body_aero.panels)
             panel_icp = body_aero.panels[icp]
             ep = evaluation_point == :control_point ? panel_icp.control_point :
@@ -480,7 +473,7 @@ Returns: nothing
                 filaments,
                 ep,
                 evaluation_point_on_bound,
-                va,
+                va_dist[jring],
                 va_unit,
                 one(T),
                 core_radius_fraction,
@@ -521,13 +514,157 @@ function calculate_circulation_distribution_elliptical_wing(gamma_i,
     return nothing
 end
 
+"""
+    thin_airfoil_strips(n_strips=24)
+
+Chordwise stations `x/c` and weights of the flat-plate bound vorticity
+`gamma(x) ~ sqrt((c - x) / x)`. The weights sum to one with their centroid at the quarter
+chord.
+"""
+function thin_airfoil_strips(n_strips=24)
+    phi = ((1:n_strips) .- 0.5) .* (π / 2) ./ n_strips
+    weights = cos.(phi) .^ 2
+    return sin.(phi) .^ 2, weights ./ sum(weights)
+end
+
+"""Velocity at `point` of unit circulation on a trailing filament from `x1` to `x2`."""
+function unit_trailing_velocity(filament::BoundFilament{T}, x1, x2, point, va,
+        work_vectors) where {T}
+    reinit!(filament, x1, x2)
+    velocity = work_vectors[8]
+    velocity_3D_trailing_vortex!(velocity, filament, point, one(T), va, work_vectors)
+    return SVector{3, T}(velocity)
+end
+
+"""Velocity at `point` of unit circulation on a bound filament from `x1` to `x2`."""
+function unit_bound_velocity(filament::BoundFilament{T}, x1, x2, point,
+        core_radius_fraction, work_vectors) where {T}
+    reinit!(filament, x1, x2)
+    velocity = work_vectors[8]
+    velocity_3D_bound_vortex!(velocity, filament, point, one(T), core_radius_fraction,
+        work_vectors)
+    return SVector{3, T}(velocity)
+end
+
+"""
+    unrolled_ring_difference(filament, work_vectors, point, panel, unrolled_1, unrolled_2,
+                             chord_leg, stations, weights, core_radius_fraction, va)
+
+Velocity at `point` induced by unit circulation on the horseshoe of `panel` less the same
+horseshoe unrolled to the bound points `unrolled_1` and `unrolled_2`, whose legs are
+`chord_leg`. Both are spread along the chord over the [`thin_airfoil_strips`](@ref)
+`stations` and `weights`. `filament` is scratch.
+"""
+function unrolled_ring_difference(filament::BoundFilament{T}, work_vectors, point, panel,
+        unrolled_1, unrolled_2, chord_leg, stations, weights, core_radius_fraction,
+        va) where {T}
+    bound_1 = SVector{3, T}(panel.bound_point_1)
+    bound_2 = SVector{3, T}(panel.bound_point_2)
+    te_1 = SVector{3, T}(panel.TE_point_1)
+    te_2 = SVector{3, T}(panel.TE_point_2)
+    unrolled_te_1 = unrolled_1 + chord_leg
+    unrolled_te_2 = unrolled_2 + chord_leg
+    velocity = zero(SVector{3, T})
+    for (x1, x2, sign) in ((unrolled_1, unrolled_te_1, 1), (unrolled_te_2, unrolled_2, 1),
+                           (bound_1, te_1, -1), (te_2, bound_2, -1))
+        velocity += sign * unit_trailing_velocity(filament, x1, x2, point, va, work_vectors)
+    end
+    for (station, weight) in zip(stations, weights)
+        along_leg = (station - 0.25) / 0.75
+        node_1 = bound_1 + along_leg * (te_1 - bound_1)
+        node_2 = bound_2 + along_leg * (te_2 - bound_2)
+        unrolled_node_1 = unrolled_1 + along_leg * chord_leg
+        unrolled_node_2 = unrolled_2 + along_leg * chord_leg
+        strip = unit_bound_velocity(filament, node_2, node_1, point, core_radius_fraction,
+            work_vectors) - unit_bound_velocity(filament, unrolled_node_2, unrolled_node_1,
+            point, core_radius_fraction, work_vectors)
+        for (x1, x2, sign) in ((node_1, te_1, 1), (te_2, node_2, 1),
+                               (unrolled_node_1, unrolled_te_1, -1),
+                               (unrolled_te_2, unrolled_node_2, -1))
+            strip += sign * unit_trailing_velocity(filament, x1, x2, point, va,
+                work_vectors)
+        end
+        velocity += weight * strip
+    end
+    return velocity
+end
+
+"""
+    curved_bound_correction!(correction, body_aero::BodyAerodynamics, core_radius_fraction,
+                             va_dist)
+
+Write into `correction` the `(i, j, k)` influence, per unit circulation of panel `j`, of
+the curvature, sweep and rib skew of a wing at the aerodynamic centre of its panel `i`:
+the actual vortex system less the one unrolled onto the tangent of panel `i`, with the
+bound vortex spread over the chord. Zero on a straight wing with perpendicular legs, and
+between panels of different wings.
+"""
+function curved_bound_correction!(correction, body_aero::BodyAerodynamics{P, W, T},
+        core_radius_fraction, va_dist) where {P, W, T}
+    correction .= 0
+    stations, weights = thin_airfoil_strips()
+    filament = BoundFilament{T}()
+    panels = body_aero.panels
+    for wing_idx in eachindex(body_aero.wings)
+        wing_panels = panel_range(body_aero, wing_idx)
+        node_position = cumsum([zero(T); [panels[j].width for j in wing_panels]])
+        next_panel = panels[min(first(wing_panels) + 1, last(wing_panels))]
+        first_panel = panels[first(wing_panels)]
+        is_forward = norm(first_panel.bound_point_1 - next_panel.bound_point_2) <
+            norm(first_panel.bound_point_2 - next_panel.bound_point_1)
+        sequence = is_forward ? one(T) : -one(T)
+        for (local_i, i) in enumerate(wing_panels)
+            centre = SVector{3, T}(panels[i].aero_center)
+            tangent = SVector{3, T}(panels[i].bound_point_1 - panels[i].bound_point_2) /
+                panels[i].width
+            position_i = (node_position[local_i] + node_position[local_i + 1]) / 2
+            for (local_j, j) in enumerate(wing_panels)
+                node_low = centre +
+                    sequence * (node_position[local_j] - position_i) * tangent
+                node_high = centre +
+                    sequence * (node_position[local_j + 1] - position_i) * tangent
+                unrolled_1, unrolled_2 = is_forward ? (node_high, node_low) :
+                    (node_low, node_high)
+                chord_leg = 0.75 * panels[j].chord * SVector{3, T}(panels[i].x_airf)
+                velocity = unrolled_ring_difference(filament, body_aero.work_vectors,
+                    centre, panels[j], unrolled_1, unrolled_2, chord_leg, stations, weights,
+                    core_radius_fraction, va_dist[i])
+                correction[i, j, :] .= velocity
+            end
+        end
+    end
+    return nothing
+end
+
+"""
+    quarter_chord_AIC!(target, body_aero::BodyAerodynamics, core_radius_fraction, va_dist,
+                       va_unit_dist)
+
+Write into `target` the influence matrices of the flow at the aerodynamic centres that
+turns the panel forces: the [`LLT`](@ref Model) matrices plus
+[`curved_bound_correction!`](@ref), kept in `body_aero` for one shape and core radius.
+"""
+function quarter_chord_AIC!(target, body_aero::BodyAerodynamics, core_radius_fraction,
+        va_dist, va_unit_dist)
+    calculate_AIC_matrices!(body_aero, LLT, core_radius_fraction, va_dist, va_unit_dist,
+        target)
+    if body_aero.curved_correction_core_radius != core_radius_fraction
+        curved_bound_correction!(body_aero.AIC_curved_correction, body_aero,
+            core_radius_fraction, va_dist)
+        body_aero.curved_correction_core_radius = core_radius_fraction
+    end
+    target .+= body_aero.AIC_curved_correction
+    return nothing
+end
+
 """
     update_effective_angle_of_attack!(alpha_corrected, body_aero::BodyAerodynamics, gamma,
                                       core_radius_fraction, va_vec_dist, va_dist,
                                       va_unit_dist)
 
 Write each panel's [`inflow_angle`](@ref) at its aerodynamic centre into
-`alpha_corrected`, the induced velocity taken from `gamma` on the LLT influence matrix.
+`alpha_corrected`, the induced velocity taken from `gamma` on the
+[`quarter_chord_AIC!`](@ref) matrices.
 """
 function update_effective_angle_of_attack!(alpha_corrected,
     body_aero::BodyAerodynamics,
@@ -538,9 +675,9 @@ function update_effective_angle_of_attack!(alpha_corrected,
     va_unit_dist)
 
     # Its own buffer: `AIC` holds the control-point matrix the circulation was solved
-    # against, so overwriting it here would leave post-solve readers on the LLT one.
-    calculate_AIC_matrices!(body_aero, LLT, core_radius_fraction, va_dist,
-                            va_unit_dist, body_aero.AIC_aero_center)
+    # against, so overwriting it here would leave post-solve readers on the other one.
+    quarter_chord_AIC!(body_aero.AIC_aero_center, body_aero, core_radius_fraction, va_dist,
+        va_unit_dist)
 
     induced_velocity = body_aero.cache[1][va_vec_dist]
     for k in 1:3
@@ -577,57 +714,52 @@ function induced_velocity_at(body_aero::BodyAerodynamics{P, W, T}, point, gamma,
 end
 
 """
-    attached_trailed_loads(body_aero::BodyAerodynamics, i, gamma, density,
-                           core_radius_fraction, reference_point)
-
-Kutta–Joukowski force [N] on the two chordwise trailed vortex segments of panel `i`,
-from its quarter-chord bound points to its trailing edge, and its moment [N·m] about
-`reference_point`, as `(; force, moment)`. Each segment carries `gamma[i]` and sees, at
-its three-quarter-chord point, the inflow turned by `body_aero.omega` plus the
-[`induced_velocity_at`](@ref) that point.
+    attached_trailed_loads!(panel_forces, m_body_3D, body_aero::BodyAerodynamics, gamma,
+                            density, core_radius_fraction, reference_point)
+
+Write into the `3×n` matrix `panel_forces` [N] the Kutta–Joukowski force on the chordwise
+trailed vortex legs between bound vortex and trailing edge, and add its moment [N·m] about
+`reference_point` to `m_body_3D`. Each section boundary carries the net circulation of its
+two panels and sees, at its three-quarter-chord point, their mean inflow plus the
+[`induced_velocity_at`](@ref) that point; its load goes half to each panel, or whole to a
+tip panel.
 """
-function attached_trailed_loads(body_aero::BodyAerodynamics{P, W, T}, i, gamma, density,
-        core_radius_fraction, reference_point) where {P, W, T}
-    panel = body_aero.panels[i]
+function attached_trailed_loads!(panel_forces, m_body_3D,
+        body_aero::BodyAerodynamics{P, W, T}, gamma, density, core_radius_fraction,
+        reference_point) where {P, W, T}
+    panel_forces .= 0
     three_quarter_chord = (0.75 - 0.25) / (1 - 0.25)
-    force = zero(SVector{3, T})
-    moment = zero(SVector{3, T})
-    segments = ((panel.bound_point_1, panel.TE_point_1, 1),
-                (panel.bound_point_2, panel.TE_point_2, -1))
-    for (bound_point, te_point, orientation) in segments
-        chordwise = SVector{3, T}(te_point) - SVector{3, T}(bound_point)
-        point = SVector{3, T}(bound_point) + three_quarter_chord * chordwise
-        arm = point - SVector{3, T}(panel.control_point)
-        inflow = SVector{3, T}(panel.va_vec) - cross(SVector{3, T}(body_aero.omega), arm)
-        velocity = inflow +
-            induced_velocity_at(body_aero, point, gamma, core_radius_fraction)
-        segment_force = (density * orientation * gamma[i]) * cross(velocity, chordwise)
-        force += segment_force
-        moment += cross(point - SVector{3, T}(reference_point), segment_force)
-    end
-    return (; force, moment)
-end
-
-"""
-    panel_force_moment(body_aero::BodyAerodynamics, i, loads, y_airf, gamma, density,
-                       core_radius_fraction, reference_point,
-                       is_with_attached_trailed_force)
-
-Force [N] and moment [N·m] about `reference_point` of panel `i`: its section `loads`
-from [`panel_loads`](@ref) acting at the aerodynamic centre, plus
-[`attached_trailed_loads`](@ref) when `is_with_attached_trailed_force`.
-Returns `(; force, moment)`.
-"""
-function panel_force_moment(body_aero::BodyAerodynamics{P, W, T}, i, loads, y_airf, gamma,
-        density, core_radius_fraction, reference_point,
-        is_with_attached_trailed_force) where {P, W, T}
-    arm = SVector{3, T}(body_aero.panels[i].aero_center) - SVector{3, T}(reference_point)
-    force = loads.force
-    moment = loads.pitching_moment .* y_airf .+ cross(arm, force)
-    is_with_attached_trailed_force || return (; force, moment)
-    attached = attached_trailed_loads(body_aero, i, gamma, density, core_radius_fraction,
-        reference_point)
-    return (; force=force + attached.force, moment=moment + attached.moment)
+    panels = body_aero.panels
+    for wing_idx in eachindex(body_aero.wings)
+        wing_panels = panel_range(body_aero, wing_idx)
+        for boundary in first(wing_panels):(last(wing_panels) + 1)
+            is_first = boundary == first(wing_panels)
+            is_last = boundary == last(wing_panels) + 1
+            before = is_first ? boundary : boundary - 1
+            after = is_last ? before : boundary
+            panel = panels[after]
+            bound_point = is_last ? panel.bound_point_2 : panel.bound_point_1
+            te_point = is_last ? panel.TE_point_2 : panel.TE_point_1
+            leg = SVector{3, T}(te_point) - SVector{3, T}(bound_point)
+            point = SVector{3, T}(bound_point) + three_quarter_chord * leg
+            velocity = (SVector{3, T}(panels[before].va_vec) +
+                        SVector{3, T}(panels[after].va_vec)) / 2 +
+                induced_velocity_at(body_aero, point, gamma, core_radius_fraction)
+            net_gamma = (is_last ? zero(T) : gamma[after]) -
+                (is_first ? zero(T) : gamma[before])
+            force = (density * net_gamma) * cross(velocity, leg)
+            moment = cross(point - SVector{3, T}(reference_point), force)
+            share = before == after ? one(T) : T(0.5)
+            @inbounds for k in 1:3
+                panel_forces[k, before] += share * force[k]
+                m_body_3D[k, before] += share * moment[k]
+                before == after && continue
+                panel_forces[k, after] += share * force[k]
+                m_body_3D[k, after] += share * moment[k]
+            end
+        end
+    end
+    return nothing
 end
 
 @inline function intersect_line_with_plane(
diff --git a/src/filament.jl b/src/filament.jl
index 023addec..0b367261 100644
--- a/src/filament.jl
+++ b/src/filament.jl
@@ -62,7 +62,8 @@ end
                                  XVP, gamma, va, work_vectors)
 
 Calculate induced velocity by a trailing vortex filament, with a Lamb–Oseen core
-radius grown over the axial distance of `XVP` from the filament start.
+radius grown over the distance of `XVP` from the filament's line. A point on that line
+gets no velocity.
 
 # Arguments
 - `XVP`: Control point coordinates
@@ -82,9 +83,15 @@ as implemented in KiteAeroDyn".
     work_vectors
 )
     r1 = work_vectors[1]
+    r1Xr0 = work_vectors[4]
     r1 .= XVP .- filament.x1
-    axial_distance = abs(dot3(r1, filament.r0)) / filament.length
-    epsilon = sqrt(4 * ALPHA0 * NU * axial_distance / va)
+    cross3!(r1Xr0, r1, filament.r0)
+    distance = norm3(r1Xr0) / filament.length
+    if distance <= 1e-10 * max(norm3(r1), filament.length)
+        vel .= 0.0
+        return nothing
+    end
+    epsilon = sqrt(4 * ALPHA0 * NU * distance / va)
     velocity_3D_vortex_segment!(vel, filament, XVP, gamma, epsilon, work_vectors)
 end
 
@@ -189,10 +196,13 @@ function reinit!(filament::SemiInfiniteFilament{T}, x1::AbstractVector,
 end
 
 """
-    velocity_3D_trailing_vortex_semiinfinite(filament::SemiInfiniteFilament, 
-                                             Vf, XVP, GAMMA, va, work_vectors)
+    velocity_3D_trailing_vortex_semiinfinite!(vel, filament::SemiInfiniteFilament,
+                                              Vf, XVP, GAMMA, va, work_vectors)
 
-Calculate induced velocity by a semi-infinite trailing vortex filament.
+Calculate induced velocity by a semi-infinite trailing vortex filament at `XVP`, for an
+inflow speed `va` > 0 [m/s]. Inside the core radius, grown over the distance from the
+filament's line, the velocity is that on the core boundary, with no upper limit as
+`XVP` nears the line. A point on the line gets no velocity.
 """
 function velocity_3D_trailing_vortex_semiinfinite!(
     vel,
@@ -208,22 +218,24 @@ function velocity_3D_trailing_vortex_semiinfinite!(
     GAMMA = -GAMMA * filament.filament_direction
     r1 .= XVP .- filament.x1
 
-    # Core radius, grown with the axial distance of `XVP` along `Vf`.
     d_r1_Vf = dot3(r1, Vf)
     nVf = norm3(Vf)
-    epsilon = sqrt(4 * ALPHA0 * NU * abs(d_r1_Vf) * nVf / va)
-
     cross3!(r1XVf, r1, Vf)
-
     nr1XVf = norm3(r1XVf)
     nr1 = norm3(r1)
-    if nr1XVf / nVf > epsilon
+    distance = nr1XVf / nVf
+    if distance < 1e-16
+        vel .= 0.0
+        return nothing
+    end
+    # Core radius, grown with the distance of `XVP` from the filament's line.
+    epsilon = sqrt(4 * ALPHA0 * NU * distance / va)
+
+    if distance > epsilon
         K = GAMMA / (4π) / (nr1XVf^2) * (1 + d_r1_Vf / nr1)
         @inbounds for k in 1:3
             vel[k] = K * r1XVf[k]
         end
-    elseif nr1XVf / nVf < 1e-12 * epsilon
-        vel .= 0.0
     else
         r1_proj = work_vectors[4]
         cross_tmp = work_vectors[5]
diff --git a/src/panel.jl b/src/panel.jl
index 531ee1be..42482e5f 100644
--- a/src/panel.jl
+++ b/src/panel.jl
@@ -201,7 +201,7 @@ function build_interps(section_1::Section, section_2::Section, remove_nan)
             cm = Vector{Float64}((aero_1[4] .+ aero_2[4]) ./ 2)
             knots = polar_knots(alphas)
             cl_i = linear_interpolation(knots, cl; extrapolation_bc=extrap_flat)
-            cd_i = linear_interpolation(knots, cd; extrapolation_bc=extrap_line)
+            cd_i = linear_interpolation(knots, cd; extrapolation_bc=extrap_flat)
             cm_i = linear_interpolation(knots, cm; extrapolation_bc=extrap_flat)
         else
             (aero_1 isa Tuple{Vector{Float64}, Vector{Float64}, Matrix{Float64}, Matrix{Float64}, Matrix{Float64}} &&
diff --git a/src/panel_aerodynamics.jl b/src/panel_aerodynamics.jl
index 8aabc7e8..24e3be84 100644
--- a/src/panel_aerodynamics.jl
+++ b/src/panel_aerodynamics.jl
@@ -19,8 +19,21 @@ length the span width.
 @inline panel_span_vector(le_1, te_1, le_2, te_2) =
     (0.75 .* le_1 .+ 0.25 .* te_1) .- (0.75 .* le_2 .+ 0.25 .* te_2)
 
-@inline panel_chord(le_1, te_1, le_2, te_2) =
-    0.5 * (smooth_norm(te_1 .- le_1) + smooth_norm(te_2 .- le_2))
+"""
+    panel_chord(le_1, te_1, le_2, te_2)
+
+Mean length of the two section chords, each with its component along the quarter-chord
+line removed.
+"""
+@inline function panel_chord(le_1, te_1, le_2, te_2)
+    span_vec = panel_span_vector(le_1, te_1, le_2, te_2)
+    span_unit = span_vec ./ smooth_norm(span_vec)
+    return 0.5 * (smooth_norm(perpendicular_part(te_1 .- le_1, span_unit)) +
+                  smooth_norm(perpendicular_part(te_2 .- le_2, span_unit)))
+end
+
+"""Part of `vec` perpendicular to the unit vector `unit`."""
+@inline perpendicular_part(vec, unit) = vec .- dot(vec, unit) .* unit
 
 """
     panel_chord_weight(width_prev, width_own, width_next)
@@ -47,9 +60,9 @@ end
     panel_axes(le_1, te_1, le_2, te_2, chord_weight=0.5, orient=1)
 
 Airfoil frame and size of the panel between two sections, as
-`(; x_airf, y_airf, z_airf, chord, width)`. `x_airf` is chordwise, `y_airf` runs
-along the quarter-chord line the bound vortex sits on — not square to `x_airf` on
-a swept panel — and `z_airf` is `x_airf × y_airf` normalised. `chord_weight`
+`(; x_airf, y_airf, z_airf, chord, width)`, an orthonormal frame: `y_airf` runs along
+the quarter-chord line the bound vortex sits on, `x_airf` is chordwise and square to
+it, and `z_airf` is `x_airf × y_airf`. `chord_weight`
 ([`panel_chord_weight`](@ref)) is section 1's share of the chord-direction blend;
 `orient` is `±1` and flips `y_airf` and `z_airf` together.
 """
@@ -59,8 +72,10 @@ a swept panel — and `z_airf` is `x_airf × y_airf` normalised. `chord_weight`
         lean .* ((te_1 .- te_2) .- (le_1 .- le_2))
     span_vec = panel_span_vector(le_1, te_1, le_2, te_2)
     width = smooth_norm(span_vec)
-    x_airf = chord_vec ./ smooth_norm(chord_vec)
-    y_airf = orient .* (span_vec ./ width)
+    span_unit = span_vec ./ width
+    chord_square = perpendicular_part(chord_vec, span_unit)
+    x_airf = chord_square ./ smooth_norm(chord_square)
+    y_airf = orient .* span_unit
     z_cross = cross(x_airf, span_vec)
     z_airf = orient .* (z_cross ./ smooth_norm(z_cross))
     return (; x_airf, y_airf, z_airf,
@@ -149,18 +164,18 @@ Panel dynamic pressure from its two section densities and the
     0.25 * (rho_1 + rho_2) * v_rel^2
 
 """
-    panel_force_directions(axes, alpha_dir, spanwise)
+    panel_force_directions(axes, alpha_dir)
 
-Lift and drag unit vectors, `(; dir_lift, dir_drag)`. `alpha_dir` is the angle that
-turns the force, `spanwise` the wing's spanwise direction.
+Lift and drag unit vectors, `(; dir_lift, dir_drag)`. `alpha_dir` is the angle of the
+flow that turns the force: drag runs along it in the section plane and lift square to
+it and to `y_airf`.
 """
-@inline function panel_force_directions(axes, alpha_dir, spanwise)
+@inline function panel_force_directions(axes, alpha_dir)
     (; y_airf, z_airf) = axes
     dir_inflow = cos(alpha_dir) .* cross(y_airf, z_airf) .+ sin(alpha_dir) .* z_airf
     lift_cross = cross(dir_inflow, y_airf)
-    dir_lift = lift_cross ./ smooth_norm(lift_cross)
-    drag_cross = cross(spanwise, dir_lift)
-    return (; dir_lift, dir_drag=drag_cross ./ smooth_norm(drag_cross))
+    return (; dir_lift=lift_cross ./ smooth_norm(lift_cross),
+            dir_drag=dir_inflow ./ smooth_norm(dir_inflow))
 end
 
 """Panel pitching moment per unit span about `y_airf`, positive nose-up."""
diff --git a/src/settings.jl b/src/settings.jl
index 4062d56d..f4a8c2cc 100644
--- a/src/settings.jl
+++ b/src/settings.jl
@@ -153,12 +153,12 @@ Solver configuration, used within [`VSMSettings`](@ref).
     [`LLT`](@ref Model) (default `VSM`)
 - `solver_type`: `"LOOP"` or `"NONLIN"` (default `"LOOP"`)
 - `density`: Air density (kg/m^3) (default `1.225`)
-- `max_iterations`: Maximum solver iterations (default `1500`)
+- `max_iterations`: Maximum solver iterations (default `5000`)
 - `rtol`: Relative tolerance on the fixed-point residual (default `1e-5`)
 - `tol_reference_error`: Reference error tolerance
     (default `0.001`)
-- `relaxation_factor`: Convergence relaxation factor
-    (default `0.03`)
+- `relaxation_factor`: Convergence relaxation factor, `nothing` for 0.8 of
+    [`relaxation_factor_limit`](@ref) (default `nothing`)
 - `is_with_artificial_viscosity`: Enable Li/Gaunaa post-stall
     artificial viscosity (default `false`)
 - `artificial_viscosity_factor`: Viscosity scaling coefficient k
@@ -174,24 +174,24 @@ Solver configuration, used within [`VSMSettings`](@ref).
 - `calc_only_f_and_gamma`: Leave the analysis fields of [`VSMSolution`](@ref) (`lift`,
     `cl`, their distributions, span and centers of pressure) at their last values
     (default `false`)
-- `correct_aoa`: Perform angle of attack correction (VSM model only)
-    (default `false`)
+- `correct_aoa`: Turn the panel forces by the flow at the quarter chord rather than at the
+    control point (VSM model only) (default `true`)
 - `flow_curvature`: Add the thin-airfoil pitch-rate moment increment to each
     section (default `false`)
 - `is_with_viscous_drag_correction`: Add the spanwise-flow viscous drag and side
     force to each section (default `false`)
 - `is_with_attached_trailed_force`: Add the force on the chordwise trailed vortex
-    segments bound to each panel (default `false`)
+    segments bound to each panel (default `true`)
 """
 @with_kw mutable struct SolverSettings
     n_panels::Int64 = 40
     aerodynamic_model_type::Model = VSM
     solver_type::String = "LOOP"    # type of solver
     density::Float64 = 1.225                # air density  [kg/m³] 
-    max_iterations::Int64 = 1500
+    max_iterations::Int64 = 5000
     rtol::Float64 = 1e-5                    # relative residual tolerance [-]
     tol_reference_error::Float64 = 0.001
-    relaxation_factor::Float64 = 0.03       # relaxation factor for convergence
+    relaxation_factor::Union{Nothing, Float64} = nothing # relaxation factor for convergence
     is_with_artificial_viscosity::Bool = false  # Li/Gaunaa post-stall artificial viscosity
     artificial_viscosity_factor::Float64 = 0.035 # viscosity scaling coefficient k
     type_initial_gamma_distribution::InitialGammaDistribution = ZEROS # see: [`InitialGammaDistribution`](@ref)
@@ -199,10 +199,10 @@ Solver configuration, used within [`VSMSettings`](@ref).
     core_radius_fraction::Float64 = 0.05
     mu::Float64 = 1.81e-5                   # dynamic viscosity [N·s/m²]
     calc_only_f_and_gamma::Bool=false       # skip the analysis fields of VSMSolution
-    correct_aoa::Bool=false                 # perform aoa correction
+    correct_aoa::Bool=true                  # turn the forces by the quarter-chord flow
     flow_curvature::Bool=false              # thin-airfoil pitch-rate moment increment
     is_with_viscous_drag_correction::Bool=false # spanwise-flow viscous force (Gaunaa 2024)
-    is_with_attached_trailed_force::Bool=false # chordwise bound-vortex force (Gaunaa 2026)
+    is_with_attached_trailed_force::Bool=true # chordwise bound-vortex force (Gaunaa 2026)
 end
 
 """
diff --git a/src/solver.jl b/src/solver.jl
index a8011d04..f4a9a3ca 100644
--- a/src/solver.jl
+++ b/src/solver.jl
@@ -19,6 +19,8 @@ Struct for storing the solution of the [`solve!`](@ref) function. Must contain a
 - panel_moment_dist::Vector{Float64}: Pitching moment per unit span about y_airf [Nm/m]
 - `f_body_3D`::Matrix{Float64}: Matrix of the aerodynamic forces (x, y, z vectors) [N]
 - `m_body_3D`::Matrix{Float64}: Matrix of the aerodynamic moments [Nm]
+- `f_attached_trailed`::Matrix{Float64}: Part of `f_body_3D` from
+    [`attached_trailed_loads!`](@ref) [N]
 - `gamma_distribution`::Union{Nothing, Vector{Float64}}: Vector containing the panel circulations.
 - force::MVec3: Aerodynamic force vector in the KA frame [N]
 - moment::MVec3: Aerodynamic moments [Mx, My, Mz] around the reference point [Nm]
@@ -45,12 +47,17 @@ Struct for storing the solution of the [`solve!`](@ref) function. Must contain a
 - `panel_cp_locations`¹::Vector{MVec3}: Center of pressure of each panel [m]
 - `moment_dist`::Vector{Float64}: Pitching moments around the spanwise vector of each panel. [Nm]
 - `moment_coeff_dist`::Vector{Float64}: Pitching moment coefficient around the spanwise vector of each panel. [-]
-- `moment_unrefined_dist`::MVector{U, Float64}: Averaged moments for unrefined sections [Nm]
-- `cl_unrefined_dist`::MVector{U, Float64}: Averaged lift coefficients for unrefined sections [-]
-- `cd_unrefined_dist`::MVector{U, Float64}: Averaged drag coefficients for unrefined sections [-]
-- `cm_unrefined_dist`::MVector{U, Float64}: Averaged airfoil moment coefficients for unrefined sections [-]
+- `moment_unrefined_dist`::MVector{U, Float64}: Pitching moment per unrefined section,
+  summed over its panels [Nm]
+- `cl_unrefined_dist`::MVector{U, Float64}: Area-weighted mean lift coefficient
+  per unrefined section [-]
+- `cd_unrefined_dist`::MVector{U, Float64}: Area-weighted mean drag coefficient
+  per unrefined section [-]
+- `cm_unrefined_dist`::MVector{U, Float64}: Area-weighted mean airfoil moment coefficient
+  per unrefined section [-]
 - `moment_coeff_unrefined_dist`::MVector{U, Float64}: Summed `moment_frac`-referenced pitching-moment coefficient per unrefined section [-]
-- `alpha_unrefined_dist`::MVector{U, Float64}: Averaged angles of attack for unrefined sections [rad]
+- `alpha_unrefined_dist`::MVector{U, Float64}: Area-weighted mean angle of attack
+  per unrefined section [rad]
 - `solver_status`::SolverStatus: enum, see [`SolverStatus`](@ref)
 
 ¹ Kept at its last value when the solver's `calc_only_f_and_gamma` is set.
@@ -74,6 +81,7 @@ Struct for storing the solution of the [`solve!`](@ref) function. Must contain a
     panel_moment_dist::Vector{T} = zeros(T, P)
     f_body_3D::Matrix{T} = zeros(T, 3, P)
     m_body_3D::Matrix{T} = zeros(T, 3, P)
+    f_attached_trailed::Matrix{T} = zeros(T, 3, P)
     gamma_distribution::Union{Nothing, Vector{T}} = nothing
     force::MVector{3, T} = zeros(MVector{3, T})
     moment::MVector{3, T} = zeros(MVector{3, T})
@@ -152,10 +160,11 @@ Main solver structure for the Vortex Step Method. See also: [`solve!`](@ref)
 ## General settings
 - `aerodynamic_model_type`::Model = VSM: The model type, see: [`Model`](@ref)
 - density::Float64 = 1.225: Air density [kg/m³] 
-- `max_iterations`::Int64 = 1500
+- `max_iterations`::Int64 = 5000
 - `rtol`::Float64 = 1e-5: Relative tolerance on the fixed-point residual
 - `tol_reference_error`::Float64 = 0.001
-- `relaxation_factor`::Float64 = 0.03: Relaxation factor for convergence 
+- `relaxation_factor`::Union{Nothing, Float64} = nothing: Relaxation factor for convergence,
+    `nothing` for 0.8 of [`relaxation_factor_limit`](@ref)
 
 ## Artificial viscosity settings
 - `is_with_artificial_viscosity`::Bool = false: Enable the Li/Gaunaa spanwise artificial
@@ -175,8 +184,8 @@ Main solver structure for the Vortex Step Method. See also: [`solve!`](@ref)
     increment `-(π/4) q̂` to each section, see: [`flow_curvature_cm`](@ref)
 - `is_with_viscous_drag_correction`::Bool = false: Add the spanwise-flow viscous drag
     and side force to each section, see: [`spanwise_flow_drag`](@ref)
-- `is_with_attached_trailed_force`::Bool = false: Add the force on the chordwise trailed
-    vortex segments bound to each panel, see: [`attached_trailed_loads`](@ref)
+- `is_with_attached_trailed_force`::Bool = true: Add the force on the chordwise trailed
+    vortex segments bound to each panel, see: [`attached_trailed_loads!`](@ref)
 - `reference_point`::MVec3 = [0.0, 0.0, 0.0]: Moment reference point in body frame
 
 ## Solution
@@ -187,10 +196,10 @@ sol::VSMSolution = VSMSolution(): The result of calling [`solve!`](@ref)
     solver_type::SolverType = LOOP
     aerodynamic_model_type::Model = VSM
     density::T = T(1.225)
-    max_iterations::Int64 = 1500
+    max_iterations::Int64 = 5000
     rtol::T = T(1e-5)
     tol_reference_error::T = T(0.001)
-    relaxation_factor::T = T(0.03)
+    relaxation_factor::Union{Nothing, T} = nothing
 
     # Nonlin solver fields
     atol::T = T(1e-5)
@@ -210,16 +219,16 @@ sol::VSMSolution = VSMSolution(): The result of calling [`solve!`](@ref)
     core_radius_fraction::T = T(0.05)
     mu::T = T(1.81e-5)
     is_only_f_and_gamma_output::Bool = false
-    correct_aoa::Bool = false
+    correct_aoa::Bool = true
     flow_curvature::Bool = false
     is_with_viscous_drag_correction::Bool = false
-    is_with_attached_trailed_force::Bool = false
+    is_with_attached_trailed_force::Bool = true
     reference_point::MVector{3, T} = zeros(MVector{3, T})
 
     # Intermediate results
     lr::LoopResult{P, T} = LoopResult{P, T}()
     br::BaseResult{P, T} = BaseResult{P, T}()
-    cache::Vector{PreallocationTools.LazyBufferCache{typeof(identity), typeof(identity)}} = [LazyBufferCache() for _ in 1:8]
+    cache::Vector{PreallocationTools.LazyBufferCache{typeof(identity), typeof(identity)}} = [LazyBufferCache() for _ in 1:6]
     cache_base::Vector{PreallocationTools.LazyBufferCache{typeof(identity), typeof(identity)}}  = [LazyBufferCache()]
     cache_lin::Vector{PreallocationTools.LazyBufferCache{typeof(identity), typeof(identity)}} = [LazyBufferCache() for _ in 1:4]
 
@@ -445,50 +454,56 @@ function calc_forces!(solver::Solver{P, U, T}, body_aero::BodyAerodynamics;
     projected_area = body_aero.projected_area
     c_ref = body_aero.c_ref
 
-    for (wing_idx, wing) in enumerate(body_aero.wings)
-        spanwise_unit = SVector{3, T}(wing.spanwise_direction)
-        for i in panel_range(body_aero, wing_idx)
-            panel = panels[i]
-            panel_area = panel.chord * panel.width
-            area_all_panels += panel_area
-            panel_areas[i] = panel_area
-
-            axes = panel_axes(panel)
-            dirs = panel_force_directions(axes, alpha_corrected[i], spanwise_unit)
-            c_span = zero(T)
-            if solver.is_with_viscous_drag_correction
-                viscous = spanwise_flow_drag(v_rel_dist[i], solver.lr.v_span_dist[i],
-                    panel.chord, density, solver.mu)
-                cd_dist[i] += viscous.delta_cd
-                c_span = viscous.c_span
-            end
-            loads = panel_loads(axes, dirs,
-                dynamic_pressure(density, density, v_rel_dist[i]),
-                cl_dist[i], cd_dist[i], cm_dist[i]; c_span)
-            lift[i] = loads.lift
-            drag[i] = loads.drag
-            panel_moment_dist[i] = loads.moment
-
-            (; force, moment) = panel_force_moment(body_aero, i, loads, axes.y_airf,
-                gamma_new, density, solver.core_radius_fraction, reference_point,
-                solver.is_with_attached_trailed_force)
-            @inbounds for k in 1:3
-                solver.sol.f_body_3D[k, i] = force[k]
-                solver.sol.m_body_3D[k, i] = moment[k]
-            end
-
-            arm = (moment_frac - 0.25) * panel.chord
-            moment_dist[i] = dot(loads.force, axes.z_airf) * arm + loads.pitching_moment
+    for (i, panel) in enumerate(panels)
+        panel_area = panel.chord * panel.width
+        area_all_panels += panel_area
+        panel_areas[i] = panel_area
+
+        axes = panel_axes(panel)
+        dirs = panel_force_directions(axes, alpha_corrected[i])
+        c_span = zero(T)
+        if solver.is_with_viscous_drag_correction
+            viscous = spanwise_flow_drag(v_rel_dist[i], solver.lr.v_span_dist[i],
+                panel.chord, density, solver.mu)
+            cd_dist[i] += viscous.delta_cd
+            c_span = viscous.c_span
         end
+        loads = panel_loads(axes, dirs,
+            dynamic_pressure(density, density, v_rel_dist[i]),
+            cl_dist[i], cd_dist[i], cm_dist[i]; c_span)
+        lift[i] = loads.lift
+        drag[i] = loads.drag
+        panel_moment_dist[i] = loads.moment
+
+        arm_to_panel = SVector{3, T}(panel.aero_center) - SVector{3, T}(reference_point)
+        moment = loads.pitching_moment .* axes.y_airf .+ cross(arm_to_panel, loads.force)
+        @inbounds for k in 1:3
+            solver.sol.f_body_3D[k, i] = loads.force[k]
+            solver.sol.m_body_3D[k, i] = moment[k]
+        end
+
+        arm = (moment_frac - 0.25) * panel.chord
+        moment_dist[i] = dot(loads.force, axes.z_airf) * arm + loads.pitching_moment
+    end
+    if solver.is_with_attached_trailed_force
+        attached_trailed_loads!(solver.sol.f_attached_trailed, solver.sol.m_body_3D,
+            body_aero, gamma_new, density, solver.core_radius_fraction, reference_point)
+        solver.sol.f_body_3D .+= solver.sol.f_attached_trailed
+    else
+        solver.sol.f_attached_trailed .= 0
     end
 
-    # Python parity: normalize with area-weighted reference velocity for distributed inflow.
+    # The inflow as handed in, without the turn rate: area-weighted when it is distributed.
     va_ref_vec = solver.sol.va_ref_vec
-    va_ref_vec .= _compute_reference_velocity_from_distribution(
-        solver.sol.va_vec_dist,
-        length(panels),
-        panel_areas
-    )
+    if body_aero.has_distributed_va
+        va_ref_vec .= _compute_reference_velocity_from_distribution(
+            solver.sol.va_vec_dist,
+            length(panels),
+            panel_areas
+        )
+    else
+        va_ref_vec .= body_aero.va_vec
+    end
     va_ref = norm(va_ref_vec)
     va_ref > 0.0 || throw(ArgumentError("Reference freestream magnitude must be positive."))
     q_ref = 0.5 * density * va_ref^2
@@ -538,42 +553,45 @@ function calc_forces!(solver::Solver{P, U, T}, body_aero::BodyAerodynamics;
                     original_section_idx = wing.refined_panel_mapping[local_panel_idx]
                     target_unrefined_idx = section_range[original_section_idx]
 
+                    area = panel.chord * panel.width
+
                     # Accumulate coefficients and moments
                     moment_unrefined_dist[target_unrefined_idx] += moment_dist[panel_idx]
-                    cl_unrefined_dist[target_unrefined_idx] += solver.sol.cl_dist[panel_idx]
-                    cd_unrefined_dist[target_unrefined_idx] += solver.sol.cd_dist[panel_idx]
-                    cm_unrefined_dist[target_unrefined_idx] += solver.sol.cm_dist[panel_idx]
+                    cl_unrefined_dist[target_unrefined_idx] +=
+                        solver.sol.cl_dist[panel_idx] * area
+                    cd_unrefined_dist[target_unrefined_idx] +=
+                        solver.sol.cd_dist[panel_idx] * area
+                    cm_unrefined_dist[target_unrefined_idx] +=
+                        solver.sol.cm_dist[panel_idx] * area
                     moment_coeff_unrefined_dist[target_unrefined_idx] += solver.sol.moment_coeff_dist[panel_idx]
-                    alpha_unrefined_dist[target_unrefined_idx] += solver.sol.alpha_dist[panel_idx]
+                    alpha_unrefined_dist[target_unrefined_idx] +=
+                        solver.sol.alpha_dist[panel_idx] * area
 
                     # Accumulate geometry
                     x_airf_unrefined_dist[target_unrefined_idx] .+= panel.x_airf
                     y_airf_unrefined_dist[target_unrefined_idx] .+= panel.y_airf
                     z_airf_unrefined_dist[target_unrefined_idx] .+= panel.z_airf
                     va_vec_unrefined_dist[target_unrefined_idx] .+= panel.va_vec
-                    chord_unrefined_dist[target_unrefined_idx] += panel.chord
+                    chord_unrefined_dist[target_unrefined_idx] += area
                     width_unrefined_dist[target_unrefined_idx] += panel.width
 
                     unrefined_count_dist[target_unrefined_idx] += 1
                 end
 
-                # Average coefficients and geometry. width and
-                # moment_coeff_unrefined_dist stay summed (extensive).
                 for target_unrefined_idx in section_range
                     if unrefined_count_dist[target_unrefined_idx] > 0
                         count = unrefined_count_dist[target_unrefined_idx]
-                        moment_unrefined_dist[target_unrefined_idx] /= count
-                        cl_unrefined_dist[target_unrefined_idx] /= count
-                        cd_unrefined_dist[target_unrefined_idx] /= count
-                        cm_unrefined_dist[target_unrefined_idx] /= count
-                        alpha_unrefined_dist[target_unrefined_idx] /= count
+                        area = chord_unrefined_dist[target_unrefined_idx]
+                        cl_unrefined_dist[target_unrefined_idx] /= area
+                        cd_unrefined_dist[target_unrefined_idx] /= area
+                        cm_unrefined_dist[target_unrefined_idx] /= area
+                        alpha_unrefined_dist[target_unrefined_idx] /= area
                         x_airf_unrefined_dist[target_unrefined_idx] ./= count
                         y_airf_unrefined_dist[target_unrefined_idx] ./= count
                         z_airf_unrefined_dist[target_unrefined_idx] ./= count
                         va_vec_unrefined_dist[target_unrefined_idx] ./= count
-                        chord_unrefined_dist[target_unrefined_idx] /= count
-                        # width_unrefined_dist is NOT averaged - it is the
-                        # sum of panel widths in the unrefined section
+                        chord_unrefined_dist[target_unrefined_idx] =
+                            area / width_unrefined_dist[target_unrefined_idx]
                     end
                 end
             end
@@ -632,8 +650,10 @@ function analysis_fields!(sol::VSMSolution, body_aero::BodyAerodynamics, density
     for (wing_idx, wing) in enumerate(body_aero.wings)
         spanwise_unit = SVector{3}(wing.spanwise_direction)
         for i in panel_range(body_aero, wing_idx)
+            attached = SVector{3}(sol.f_attached_trailed[1, i],
+                sol.f_attached_trailed[2, i], sol.f_attached_trailed[3, i])
             force = SVector{3}(sol.f_body_3D[1, i], sol.f_body_3D[2, i],
-                sol.f_body_3D[3, i])
+                sol.f_body_3D[3, i]) - attached
             panel_va = SVector{3}(sol.va_vec_dist[i, 1], sol.va_vec_dist[i, 2],
                 sol.va_vec_dist[i, 3])
             va_panel = norm(panel_va)
@@ -642,10 +662,13 @@ function analysis_fields!(sol::VSMSolution, body_aero::BodyAerodynamics, density
             body_dirs = prescribed_va_directions(panel_va, reference_spanwise)
 
             lift += dot(force, body_dirs.dir_lift) *
-                dot(body_dirs.dir_lift, reference_dirs.dir_lift)
-            drag += panel_drag * dot(panel_va, va_ref_unit) / va_panel
+                dot(body_dirs.dir_lift, reference_dirs.dir_lift) +
+                dot(attached, reference_dirs.dir_lift)
+            drag += panel_drag * dot(panel_va, va_ref_unit) / va_panel +
+                dot(attached, va_ref_unit)
             side += dot(force, body_dirs.dir_side) *
-                dot(body_dirs.dir_side, reference_dirs.dir_side)
+                dot(body_dirs.dir_side, reference_dirs.dir_side) +
+                dot(attached, reference_dirs.dir_side)
 
             q_area = dynamic_pressure(density, density, va_panel) * sol.panel_area_dist[i]
             sol.cl_distribution[i] = dot(force, wing_dirs.dir_lift) / q_area
@@ -704,7 +727,8 @@ function solve_base!(solver::Solver{P, U, T}, body_aero::BodyAerodynamics, gamma
     # Initialize variables
     panels = body_aero.panels
     n_panels = length(panels)
-    relaxation_factor = solver.relaxation_factor
+    relaxation_factor = something(solver.relaxation_factor,
+        0.8 * relaxation_factor_limit(panels, solver.aerodynamic_model_type))
     
     # Clear arrays
     solver.sol._y_airf_dist .= 0
@@ -786,9 +810,7 @@ end
     relative_velocity_dist,
     y_airf_dist,
     relative_velocity_crossz,
-    v_acrossz_dist,
     z_airf_dist,
-    va_magw_dist,
     cl_dist,
     chord_dist,
 )
@@ -807,12 +829,6 @@ end
         relative_velocity_crossz[i,1] = ay*bz - az*by
         relative_velocity_crossz[i,2] = az*bx - ax*bz
         relative_velocity_crossz[i,3] = ax*by - ay*bx
-        ax = va_vec_dist[i,1]
-        ay = va_vec_dist[i,2]
-        az = va_vec_dist[i,3]
-        v_acrossz_dist[i,1] = ay*bz - az*by
-        v_acrossz_dist[i,2] = az*bx - ax*bz
-        v_acrossz_dist[i,3] = ax*by - ay*bx
     end
 
     @inbounds for i in 1:n_panels
@@ -827,19 +843,40 @@ end
             relative_velocity_crossz[i,1]^2 +
             relative_velocity_crossz[i,2]^2 +
             relative_velocity_crossz[i,3]^2)
-        va_magw_dist[i] = smooth_sqrt(
-            v_acrossz_dist[i,1]^2 +
-            v_acrossz_dist[i,2]^2 +
-            v_acrossz_dist[i,3]^2)
     end
 
     for (i, (panel, alpha)) in enumerate(zip(panels, solver.lr.alpha_dist))
         cl_dist[i] = calculate_cl(panel, alpha)
     end
-    gamma_out .= 0.5 .* solver.lr.v_rel_dist.^2 ./ va_magw_dist .* cl_dist .* chord_dist
+    gamma_out .= 0.5 .* solver.lr.v_rel_dist .* cl_dist .* chord_dist
     return nothing
 end
 
+"""
+    relaxation_factor_limit(panels, model::Model)
+
+Largest relaxation factor the `LOOP` solver is stable at in attached flow (Li, Gaunaa,
+Pirrung & Lønbæk, TORQUE 2026, Eqs. 10-11), from the steepest positive lift slope and
+chord-to-width ratio over `panels`, halved for `VSM` and clipped to `[1e-3, 1]`.
+"""
+function relaxation_factor_limit(panels, model::Model)
+    step = deg2rad(1.0)
+    ratio = 0.0
+    for panel in panels
+        steepest = 0.0
+        cl_low = calculate_cl(panel, deg2rad(-40.0))
+        for alpha in deg2rad(-40.0):step:deg2rad(50.0)
+            cl_high = calculate_cl(panel, alpha + step)
+            slope = (cl_high - cl_low) / step
+            isfinite(slope) && (steepest = max(steepest, slope))
+            cl_low = cl_high
+        end
+        ratio = max(ratio, panel.chord / panel.width * steepest)
+    end
+    limit = 2 / (1 + 0.25 * ratio)
+    return clamp(model == VSM ? limit / 2 : limit, 1e-3, 1.0)
+end
+
 """
     build_spanwise_laplacian!(laplacian, n_panels)
 
@@ -953,14 +990,12 @@ function gamma_loop!(
     solver.lr.alpha_dist .= body_aero.alpha_dist
     solver.lr.v_rel_dist .= body_aero.v_rel_dist
     
-    va_magw_dist             = solver.cache[1][solver.lr.v_rel_dist]
-    gamma                    = solver.cache[2][solver.lr.gamma_new]
-    abs_gamma_new            = solver.cache[3][solver.lr.gamma_new]
-    induced_velocity_all     = solver.cache[4][va_vec_dist]
-    relative_velocity_dist   = solver.cache[5][va_vec_dist]
-    relative_velocity_crossz = solver.cache[6][va_vec_dist]
-    v_acrossz_dist           = solver.cache[7][va_vec_dist]
-    cl_dist                  = solver.cache[8][solver.lr.gamma_new]
+    gamma                    = solver.cache[1][solver.lr.gamma_new]
+    abs_gamma_new            = solver.cache[2][solver.lr.gamma_new]
+    induced_velocity_all     = solver.cache[3][va_vec_dist]
+    relative_velocity_dist   = solver.cache[4][va_vec_dist]
+    relative_velocity_crossz = solver.cache[5][va_vec_dist]
+    cl_dist                  = solver.cache[6][solver.lr.gamma_new]
 
     AIC_x = @view body_aero.AIC[:, :, 1]
     AIC_y = @view body_aero.AIC[:, :, 2]
@@ -988,8 +1023,8 @@ function gamma_loop!(
             AIC_x, AIC_y, AIC_z,
             velocity_view_x, velocity_view_y, velocity_view_z,
             va_vec_dist, induced_velocity_all, relative_velocity_dist,
-            y_airf_dist, relative_velocity_crossz, v_acrossz_dist,
-            z_airf_dist, va_magw_dist, cl_dist, chord_dist,
+            y_airf_dist, relative_velocity_crossz,
+            z_airf_dist, cl_dist, chord_dist,
         )
         @inbounds for i in 1:n_panels
             residual[i] -= gamma_iter[i]
@@ -1008,8 +1043,8 @@ function gamma_loop!(
                     AIC_x, AIC_y, AIC_z,
                     velocity_view_x, velocity_view_y, velocity_view_z,
                     va_vec_dist, induced_velocity_all, relative_velocity_dist,
-                    y_airf_dist, relative_velocity_crossz, v_acrossz_dist,
-                    z_airf_dist, va_magw_dist, cl_dist, chord_dist,
+                    y_airf_dist, relative_velocity_crossz,
+                    z_airf_dist, cl_dist, chord_dist,
                 )
                 inv_step = 1.0 / step
                 for i in 1:n_panels
@@ -1035,8 +1070,8 @@ function gamma_loop!(
                     AIC_x, AIC_y, AIC_z,
                     velocity_view_x, velocity_view_y, velocity_view_z,
                     va_vec_dist, induced_velocity_all, relative_velocity_dist,
-                    y_airf_dist, relative_velocity_crossz, v_acrossz_dist,
-                    z_airf_dist, va_magw_dist, cl_dist, chord_dist,
+                    y_airf_dist, relative_velocity_crossz,
+                    z_airf_dist, cl_dist, chord_dist,
                 )
                 @inbounds for i in 1:n_panels
                     residual_perturbed[i] -= gamma_perturbed[i]
@@ -1098,9 +1133,7 @@ function gamma_loop!(
                 relative_velocity_dist,
                 y_airf_dist,
                 relative_velocity_crossz,
-                v_acrossz_dist,
                 z_airf_dist,
-                va_magw_dist,
                 cl_dist,
                 chord_dist,
             )
@@ -1215,7 +1248,8 @@ function make_dual_shadow(solver::Solver{P, U, Float64},
         max_iterations = solver.max_iterations,
         rtol = TD(solver.rtol),
         tol_reference_error = TD(solver.tol_reference_error),
-        relaxation_factor = TD(solver.relaxation_factor),
+        relaxation_factor = isnothing(solver.relaxation_factor) ? nothing :
+            TD(solver.relaxation_factor),
         atol = TD(solver.atol),
         is_with_artificial_viscosity = solver.is_with_artificial_viscosity,
         artificial_viscosity_factor = TD(solver.artificial_viscosity_factor),
diff --git a/src/wake.jl b/src/wake.jl
index 971702bb..1b65ae9c 100644
--- a/src/wake.jl
+++ b/src/wake.jl
@@ -3,7 +3,7 @@
     frozen_wake(body_aero::BodyAerodynamics, va_vec_dist)
 
 Update the filaments of the panels with frozen wake model.
-Uses one shared wake vector computed from area-weighted distributed inflow.
+Each panel's wake follows its own row of `va_vec_dist`.
 
 Replaces older filaments if present by checking length of filaments.
 
@@ -20,34 +20,14 @@ function frozen_wake!(body_aero::BodyAerodynamics, va_vec_dist)
         throw(ArgumentError(
             "va_vec_dist must be shape ($(n_panels), 3), got $(size(va_vec_dist))"))
 
-    panel_areas = [panel.chord * panel.width for panel in body_aero.panels]
-    wake_velocity = _compute_reference_velocity_from_distribution(
-        va_vec_dist,
-        n_panels,
-        panel_areas
-    )
-    wake_speed = norm(wake_velocity)
-    if wake_speed <= 0.0
+    for (i, panel) in enumerate(body_aero.panels)
+        wake_velocity = matrix_row(va_vec_dist, i)
+        wake_speed = norm(wake_velocity)
         # Zero apparent flow: keep existing wake state and avoid NaN/Inf updates.
-        return nothing
-    end
-    wake_direction = wake_velocity / wake_speed
-
-    for panel in body_aero.panels
-        reinit!(
-            panel.filaments[4],
-            panel.TE_point_1, 
-            wake_direction, 
-            wake_speed, 
-            1
-        )
-        reinit!(
-            panel.filaments[5], 
-            panel.TE_point_2, 
-            wake_direction, 
-            wake_speed, 
-            -1
-        )
+        wake_speed <= 0.0 && continue
+        wake_direction = wake_velocity / wake_speed
+        reinit!(panel.filaments[4], panel.TE_point_1, wake_direction, wake_speed, 1)
+        reinit!(panel.filaments[5], panel.TE_point_2, wake_direction, wake_speed, -1)
     end
     return nothing
 end
diff --git a/test/body_aerodynamics/test_body_aerodynamics.jl b/test/body_aerodynamics/test_body_aerodynamics.jl
index a47650a6..c61a8164 100644
--- a/test/body_aerodynamics/test_body_aerodynamics.jl
+++ b/test/body_aerodynamics/test_body_aerodynamics.jl
@@ -225,15 +225,23 @@ end
                 end
 
                 @test isapprox(evaluation_point, expected_controlpoints[i]["coordinates"], atol=1e-4)
-                @test isapprox(panel.chord, expected_controlpoints[i]["chord"], atol=1e-4)
-                @test isapprox(panel.z_airf, expected_controlpoints[i]["normal"], atol=1e-4)
-                @test isapprox(panel.x_airf, expected_controlpoints[i]["tangential"], atol=1e-4)
+                # The thesis chord line is the raw leading-to-trailing-edge direction;
+                # panels take its part square to the bound vortex.
+                span_axis = expected_controlpoints[i]["airf_coord"][:, 3]
+                raw_tangential = expected_controlpoints[i]["tangential"]
+                along_span = dot(raw_tangential, span_axis)
+                tangential = normalize(raw_tangential .- along_span .* span_axis)
+                normal = cross(tangential, span_axis)
+                chord = expected_controlpoints[i]["chord"] * sqrt(max(0.0, 1 - along_span^2))
+                @test isapprox(panel.chord, chord, atol=1e-4)
+                @test isapprox(panel.z_airf, normal, atol=1e-4)
+                @test isapprox(panel.x_airf, tangential, atol=1e-4)
                 @test isapprox(
                     hcat(panel.z_airf, panel.x_airf, panel.y_airf),
-                    expected_controlpoints[i]["airf_coord"],
+                    hcat(normal, tangential, span_axis),
                     atol=1e-4
                 )
-                
+
                 if model === VSM
                     @test isapprox(
                         panel.aero_center,
@@ -332,6 +340,7 @@ end
         core_radius_fraction=core_radius_fraction,
         solver_type=LOOP,
         correct_aoa=true,
+        is_with_attached_trailed_force=false,
         atol=1e-8,
         rtol=1e-8
     )
@@ -340,6 +349,7 @@ end
         core_radius_fraction=core_radius_fraction,
         solver_type=NONLIN,
         correct_aoa=true,
+        is_with_attached_trailed_force=false,
         atol=1e-8,
         rtol=1e-8
     )
@@ -350,20 +360,20 @@ end
 
         @test all(isapprox.(nonlin_sol.gamma_distribution, loop_sol.gamma_distribution; atol=1e-4))
 
-        @test loop_sol.force.x ≈ -117.96518414816444 atol=1e-4
+        @test loop_sol.force.x ≈ -117.83764441623495 atol=1e-4
         @test loop_sol.force.y ≈ 0.0 atol=1e-10
         @test loop_sol.force.z ≈ 1481.996390329679 atol=1e-4 rtol= 1e-4
 
         @test loop_sol.moment.x ≈ -1481.996390329678 atol=1e-4 rtol= 1e-4
         @test loop_sol.moment.y ≈ 0.0 atol=1e-10
-        @test loop_sol.moment.z ≈ -117.9651841481644 atol=1e-4
+        @test loop_sol.moment.z ≈ -117.83764441623481 atol=1e-4
 
         @test loop_sol.force_coeffs[1] ≈ -0.039050322560956294 atol=1e-4 # CFx
         @test loop_sol.force_coeffs[2] ≈ 0.0                   atol=1e-4 # CFy
         @test loop_sol.force_coeffs[3] ≈ 0.49055973654418716   atol=3e-4 # CFz
         @test loop_sol.force_coeffs[3] / loop_sol.force_coeffs[1] ≈ loop_sol.force[3] / loop_sol.force[1]
-        @test loop_sol.moment_dist[1] ≈ -0.0006683569356186426 atol=1e-8
-        @test loop_sol.moment_coeff_dist[1] ≈ -2.212405554436003e-7 atol=1e-9
+        @test loop_sol.moment_dist[1] ≈ -0.0006766862950545562 atol=1e-8
+        @test loop_sol.moment_coeff_dist[1] ≈ -2.2428476412484825e-7 atol=1e-9
         @test loop_sol.moment_dist[1] / loop_sol.moment_dist[2] ≈ loop_sol.moment_coeff_dist[1] / loop_sol.moment_coeff_dist[2]
 
         @test loop_sol.solver_status == FEASIBLE
diff --git a/test/body_aerodynamics/test_curved_bound_correction.jl b/test/body_aerodynamics/test_curved_bound_correction.jl
new file mode 100644
index 00000000..a78c97c5
--- /dev/null
+++ b/test/body_aerodynamics/test_curved_bound_correction.jl
@@ -0,0 +1,71 @@
+using LinearAlgebra
+using Test
+using VortexStepMethod
+using VortexStepMethod: curved_bound_correction!, quarter_chord_AIC!, reinit!,
+    thin_airfoil_strips
+
+"""Constant-chord wing of `n_panels` panels from the section points of `le_points`."""
+function wing_through(le_points, n_panels)
+    wing = Wing(n_panels; spanwise_distribution=LINEAR)
+    for le in le_points
+        add_section!(wing, le, le .+ [1.0, 0.0, 0.0], INVISCID)
+    end
+    refine!(wing)
+    return wing
+end
+
+arc_points(radius=4.0, half_angle=55.0, n_sections=21) = [
+    [0.0, radius * sind(angle), radius * cosd(angle)]
+    for angle in range(half_angle, -half_angle; length=n_sections)]
+
+@testset "Curved bound vortex correction" begin
+    @testset "the strips hold a thin-airfoil load centred on the quarter chord" begin
+        stations, weights = thin_airfoil_strips()
+        @test length(stations) == 24
+        @test sum(weights) ≈ 1
+        @test sum(stations .* weights) ≈ 0.25 atol = 2e-3
+        @test all(0 .< stations .< 1)
+    end
+
+    n_panels = 20
+    va_dist = fill(15.0, n_panels)
+
+    @testset "it vanishes on a straight wing" begin
+        straight = wing_through([[0.0, 4.0, 0.0], [0.0, -4.0, 0.0]], n_panels)
+        body_aero = BodyAerodynamics([straight])
+        correction = zeros(n_panels, n_panels, 3)
+        curved_bound_correction!(correction, body_aero, 0.05, va_dist)
+        @test maximum(abs, correction) < 1e-10
+    end
+
+    @testset "on an arc it is kept per shape and rebuilt after a shape change" begin
+        wing = wing_through(arc_points(), n_panels)
+        body_aero = BodyAerodynamics([wing])
+        solver = Solver(n_panels, wing.n_unrefined_sections)
+        set_va!(body_aero, [15.0, 0.0, 2.0])
+        solve!(solver, body_aero)
+        @test maximum(abs, body_aero.AIC_curved_correction) > 1e-3
+        @test body_aero.curved_correction_core_radius == solver.core_radius_fraction
+
+        body_aero.AIC_curved_correction .= 0.0
+        solve!(solver, body_aero)
+        @test iszero(body_aero.AIC_curved_correction)
+
+        reinit!(body_aero)
+        @test isnan(body_aero.curved_correction_core_radius)
+        set_va!(body_aero, [15.0, 0.0, 2.0])
+        solve!(solver, body_aero)
+        @test maximum(abs, body_aero.AIC_curved_correction) > 1e-3
+    end
+
+    @testset "a second wing does not enter another wing's correction" begin
+        low = wing_through([point .+ [0.0, 0.0, -6.0] for point in arc_points()], 10)
+        high = wing_through(arc_points(), 10)
+        body_aero = BodyAerodynamics([low, high])
+        correction = zeros(20, 20, 3)
+        curved_bound_correction!(correction, body_aero, 0.05, fill(15.0, 20))
+        @test iszero(correction[1:10, 11:20, :])
+        @test iszero(correction[11:20, 1:10, :])
+        @test !iszero(correction[1:10, 1:10, :])
+    end
+end
diff --git a/test/data/README.md b/test/data/README.md
index e69de29b..399fd338 100644
--- a/test/data/README.md
+++ b/test/data/README.md
@@ -0,0 +1,2 @@
+
+`python_reference/reference.yaml` holds inputs and results of the Python VSM v2.3.0. `python_reference/generate_reference.py` writes it, run in an environment with that version installed.
diff --git a/test/data/python_reference/generate_reference.py b/test/data/python_reference/generate_reference.py
new file mode 100644
index 00000000..71006be6
--- /dev/null
+++ b/test/data/python_reference/generate_reference.py
@@ -0,0 +1,184 @@
+"""Write reference.yaml: inputs and results of the Python VSM v2.3.0 (awegroup/Vortex-Step-Method
+at a2a91cb), read by test/verification/test_python_reference.jl.
+
+Usage: python generate_reference.py <path to the Python repository> <output yaml>
+"""
+import sys
+from pathlib import Path
+
+import numpy as np
+import yaml
+from VSM.core.BodyAerodynamics import BodyAerodynamics
+from VSM.core.Solver import Solver
+from VSM.core.WingGeometry import Wing
+
+ALPHA_GRID = np.deg2rad(np.arange(-10.0, 31.0, 2.0))
+INVISCID = np.column_stack(
+    (ALPHA_GRID, 2 * np.pi * ALPHA_GRID, 0 * ALPHA_GRID, 0 * ALPHA_GRID)
+)
+
+
+def stalling_polar():
+    cl = np.where(
+        ALPHA_GRID < np.deg2rad(14.0),
+        2 * np.pi * ALPHA_GRID,
+        2 * np.pi * np.deg2rad(14.0) - 1.5 * (ALPHA_GRID - np.deg2rad(14.0)),
+    )
+    return np.column_stack(
+        (ALPHA_GRID, cl, 0.01 + 0.6 * ALPHA_GRID**2, -0.08 - 0.1 * ALPHA_GRID)
+    )
+
+
+def straight_sections():
+    return [
+        ([0.0, y, 0.0], [1.0, y, 0.0], INVISCID) for y in (5.0, -5.0)
+    ]
+
+
+def swept_sections():
+    """Tapered, swept, twisted wing with dihedral."""
+    sections = []
+    for y in np.linspace(4.0, -4.0, 9):
+        fraction = abs(y) / 4.0
+        le = np.array([1.2 * fraction, y, 0.4 * fraction])
+        chord = 1.5 - 0.7 * fraction
+        twist = np.deg2rad(4.0 * (1.0 - fraction))
+        te = le + chord * np.array([np.cos(twist), 0.0, -np.sin(twist)])
+        sections.append((le.tolist(), te.tolist(), INVISCID))
+    return sections
+
+
+def arc_sections():
+    """Circular-arc wing with anhedral and a tapered chord."""
+    sections = []
+    for angle in np.deg2rad(np.linspace(55.0, -55.0, 13)):
+        le = np.array([0.3 * abs(angle), 4.0 * np.sin(angle), 4.0 * np.cos(angle)])
+        chord = 1.2 - 0.4 * abs(angle)
+        sections.append((le.tolist(), (le + [chord, 0.0, 0.0]).tolist(), stalling_polar()))
+    return sections
+
+
+def v3_sections(python_repo):
+    """The V3 kite of the Python repository, with its Breukels polars on ALPHA_GRID."""
+    path = (
+        Path(python_repo)
+        / "data/TUDELFT_V3_KITE/CAD_derived_geometry/aero_geometry_CAD_breukels_regression.yaml"
+    )
+    body = BodyAerodynamics.instantiate(n_panels=18, file_path=path)
+    sections = []
+    for section in body.wings[0].sections:
+        polar = section.polar_data
+        resampled = np.column_stack(
+            [ALPHA_GRID] + [np.interp(ALPHA_GRID, polar[:, 0], polar[:, k]) for k in (1, 2, 3)]
+        )
+        sections.append(
+            (list(map(float, section.LE_point)), list(map(float, section.TE_point)), resampled)
+        )
+    return sections
+
+
+def plain(value):
+    return np.asarray(value, dtype=float).tolist()
+
+
+def run_case(name, sections, n_panels, va, keep_matrices=False, **options):
+    wing = Wing(n_panels=n_panels, spanwise_panel_distribution="uniform")
+    for le, te, polar in sections:
+        wing.add_section(np.array(le), np.array(te), polar)
+    body = BodyAerodynamics([wing])
+    rates = options.pop("body_rates", None)
+    pivot = options.pop("pivot", [0.0, 0.0, 0.0])
+    va_array = np.asarray(va, dtype=float)
+    if va_array.ndim == 2:
+        va_array = np.array([va_array[0] + (va_array[1] - va_array[0]) * i / (n_panels - 1)
+                             for i in range(n_panels)])
+    if rates is None:
+        body.va = va_array
+    else:
+        type(body).va.fset(body, va=va_array, body_rates=rates, body_axis=np.eye(3),
+                           reference_point=np.array(pivot))
+    reference_point = np.array([0.3, 0.1, 1.0])
+    solver = Solver(allowed_error=1e-10, max_iterations=400000,
+                    reference_point=reference_point, **options)
+    results = solver.solve(body)
+    assert results["gamma_converged"], name
+    panels = body.panels
+    va_panels = np.array([p.va for p in panels])
+    speeds = np.linalg.norm(va_panels, axis=1)
+    units = va_panels / speeds[:, None]
+    model = options.get("aerodynamic_model_type", "VSM")
+    AIC = np.array(body.compute_AIC_matrices(model, solver.core_radius_fraction, speeds, units))
+    AIC_qc = np.array(
+        body.compute_quarter_chord_AIC_matrices(solver.core_radius_fraction, speeds, units)
+    )
+    return {
+        "name": name,
+        "n_panels": n_panels,
+        "sections": [{"LE": list(le), "TE": list(te), "polar": plain(polar)}
+                     for le, te, polar in sections],
+        "va": plain(va),
+        "body_rates": None if rates is None else plain(rates),
+        "pivot": plain(pivot),
+        "reference_point": plain(reference_point),
+        "options": options,
+        "relaxation_factor": float(solver.relaxation_factor_used),
+        "chord": plain([p.chord for p in panels]),
+        "width": plain([p.width for p in panels]),
+        "normal": plain([p.x_airf for p in panels]),
+        "chordwise": plain([p.y_airf for p in panels]),
+        "spanwise": plain([p.z_airf for p in panels]),
+        "panel_va": plain(va_panels),
+        "AIC": plain(AIC) if keep_matrices else None,
+        "AIC_quarter_chord": plain(AIC_qc) if keep_matrices else None,
+        "gamma": plain(results["gamma_distribution"]),
+        "alpha": np.ravel(results["alpha_uncorrected"]).tolist(),
+        "alpha_at_ac": np.ravel(results["alpha_at_ac"]).tolist(),
+        "force": [float(results[key]) for key in ("Fx", "Fy", "Fz")],
+        "moment": [float(results[key]) for key in ("Mx", "My", "Mz")],
+        "force_coefficients": [float(results[key]) for key in ("cl", "cd", "cs")],
+        "moment_coefficients": [float(results[key]) for key in ("cmx", "cmy", "cmz")],
+        "lift_drag_side": [float(results[key]) for key in ("lift", "drag", "side")],
+        "F_distribution": plain(results["F_distribution"]),
+        "M_distribution": plain(results["M_distribution"]),
+        "cl_distribution": np.ravel(results["cl_distribution"]).tolist(),
+        "cd_distribution": np.ravel(results["cd_distribution"]).tolist(),
+        "q_ref": float(results["q_ref"]),
+        "va_ref": plain(results["va_ref"]),
+        "projected_area": float(results["projected_area"]),
+        "area_all_panels": float(results["area_all_panels"]),
+    }
+
+
+def main(python_repo, output):
+    alpha = np.deg2rad(8.0)
+    va = (15.0 * np.array([np.cos(alpha), 0.0, np.sin(alpha)])).tolist()
+    sideslip = (15.0 * np.array([0.97, 0.15, 0.19])).tolist()
+    arc = arc_sections()
+    cases = [
+        run_case("straight", straight_sections(), 20, va),
+        run_case("swept", swept_sections(), 16, va, keep_matrices=True),
+        run_case("arc", arc, 12, va, keep_matrices=True),
+        run_case("arc_sideslip", arc, 12, sideslip),
+        run_case("arc_llt", arc, 12, va, aerodynamic_model_type="LLT"),
+        run_case("arc_viscous", arc, 12, va, is_with_viscous_drag_correction=True),
+        run_case("arc_rates", arc, 12, va, body_rates=[0.2, -0.3, 0.5], pivot=[0.1, 0.0, 2.0]),
+        run_case("arc_distributed", arc, 12, [[14.0, 0.0, 2.0], [17.0, 1.0, 2.5]]),
+        run_case("v3", v3_sections(python_repo), 18, va),
+    ]
+    defaults = Solver()
+    reference = {
+        "python_commit": "a2a91cb02c05cef4c0201dfe2a666eba8e941f0f",
+        "defaults": {
+            "max_iterations": defaults.max_iterations,
+            "core_radius_fraction": defaults.core_radius_fraction,
+            "is_aoa_corrected": defaults.is_aoa_corrected,
+            "is_with_attached_trailed_vortex_force": defaults.is_with_attached_trailed_vortex_force,
+        },
+        "cases": cases,
+    }
+    with open(output, "w") as file:
+        yaml.safe_dump(reference, file, default_flow_style=None, width=100000)
+
+
+if __name__ == "__main__":
+    main(sys.argv[1], sys.argv[2])
diff --git a/test/data/python_reference/reference.yaml b/test/data/python_reference/reference.yaml
new file mode 100644
index 00000000..687ad0bf
--- /dev/null
+++ b/test/data/python_reference/reference.yaml
@@ -0,0 +1,4250 @@
+cases:
+- AIC: null
+  AIC_quarter_chord: null
+  F_distribution:
+  - [-2.887539187258424, 5.708457924577283, 29.979824628664126]
+  - [-4.218357836845189, 0.16133479746965784, 39.73583292456661]
+  - [-5.042778210591447, 0.0332178650300636, 44.83317169170296]
+  - [-5.5853754368271105, 0.009638625329559582, 47.83723211325796]
+  - [-5.954457422021959, 0.0032715176569471376, 49.73965904400908]
+  - [-6.2098108677240855, 0.001191310713323829, 50.993506978188485]
+  - [-6.3863023586325, 0.0004365941402538206, 51.83144043272808]
+  - [-6.504857133236032, 0.00015085053837829267, 52.38142150799156]
+  - [-6.578042644967964, 4.5207003300686746e-05, 52.715823388524754]
+  - [-6.613009786274694, 8.972447764424828e-06, 52.874227822582]
+  - [-6.613009786274697, -8.972447764435725e-06, 52.874227822582]
+  - [-6.578042644967968, -4.520700330071204e-05, 52.715823388524775]
+  - [-6.504857133236027, -0.00015085053837828128, 52.38142150799152]
+  - [-6.386302358632502, -0.0004365941402537566, 51.83144043272811]
+  - [-6.209810867724093, -0.001191310713323746, 50.993506978188556]
+  - [-5.954457422021958, -0.0032715176569469862, 49.739659044009024]
+  - [-5.585375436827114, -0.009638625329559565, 47.83723211325801]
+  - [-5.0427782105914485, -0.03321786503006368, 44.83317169170296]
+  - [-4.218357836845191, -0.16133479746965806, 39.735832924566616]
+  - [-2.887539187258423, -5.708457924577283, 29.979824628664115]
+  M_distribution:
+  - [145.16934535934578, 4.252780819790024, 15.995863286811453]
+  - [165.06929745454292, 6.189317747376229, 17.578785681768878]
+  - [163.67544675717406, 7.279421534579333, 18.421088507922306]
+  - [150.69731424422187, 7.97531514100856, 17.5982700074037]
+  - [131.81352462114248, 8.44062067072332, 15.780784351303819]
+  - [109.63730021559348, 8.759117590407186, 13.351629455427778]
+  - [85.522345796253, 8.977708533815314, 10.537595359106739]
+  - [60.23880139037925, 9.123858949706078, 7.480653585963706]
+  - [34.265337697983874, 9.213810743644801, 4.2757480623806625]
+  - [7.931145269302673, 9.256716930468576, 0.9919555055426981]
+  - [-18.505990833819073, 9.256716930468578, -2.3145574627976377]
+  - [-44.808502375688846, 9.213810743644807, -5.591356591374257]
+  - [-70.71508569197752, 9.12385894970607, -8.781625012610908]
+  - [-95.8886338827987, 8.977708533815317, -11.814855830833244]
+  - [-119.83600161123137, 8.759117590407197, -14.593591628972614]
+  - [-141.76145642994413, 8.440620670723316, -16.971675835708208]
+  - [-160.2647606668736, 7.975315141008566, -18.715345094769134]
+  - [-172.64208109551467, 7.279421534579335, -19.429644150040602]
+  - [-173.01646403945625, 6.189317747376231, -18.42245724913792]
+  - [-151.16531028507856, 4.252780819790021, -16.573371124263126]
+  alpha: [0.06973832415549032, 0.0927265788761257, 0.10450711920813573, 0.11145057205176885, 0.11585440100720554, 0.11876177617970121, 0.12070767210809984, 0.12198640610154685, 0.12276457955697803, 0.12313338719204592, 0.12313338719204592, 0.12276457955697809, 0.12198640610154673, 0.12070767210809989, 0.11876177617970139, 0.11585440100720541, 0.11145057205176898, 0.10450711920813573, 0.09272657887612573, 0.06973832415549029]
+  alpha_at_ac: [0.09687898156817969, 0.10584763130087901, 0.11203281409291167, 0.1162409043007095, 0.11914935773514786, 0.12118157398744796, 0.12259578497133697, 0.12355033768000329, 0.12414151059551481, 0.12442449755255906, 0.12442449755255912, 0.12414151059551484, 0.12355033768000331, 0.12259578497133694, 0.12118157398744793, 0.11914935773514797, 0.11624090430070945, 0.11203281409291173, 0.10584763130087904, 0.09687898156817967]
+  area_all_panels: 10.0
+  body_rates: null
+  cd_distribution: [0.018513830205032514, 0.01956930344472821, 0.018060383662961955, 0.016342523842863627, 0.014885229957889741, 0.013749761074527455, 0.012906591585472759, 0.012314024856003546, 0.011937851832838778, 0.0117553434586278, 0.01175534345862776, 0.011937851832838778, 0.012314024856003546, 0.012906591585472797, 0.013749761074527493, 0.014885229957889639, 0.01634252384286368, 0.01806038366296193, 0.0195693034447282, 0.01851383020503251]
+  chord: [1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0, 1.0]
+  chordwise:
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  - [1.0, 0.0, 0.0]
+  cl_distribution: [0.4328349736790805, 0.5791180501985475, 0.6543491268625418, 0.6987060111814373, 0.7268233716525071, 0.745371431922717, 0.7577758687410246, 0.7659220039311215, 0.7708769163866988, 0.7732245498702444, 0.7732245498702444, 0.770876916386699, 0.7659220039311209, 0.7577758687410251, 0.745371431922718, 0.7268233716525062, 0.6987060111814379, 0.6543491268625418, 0.5791180501985476, 0.4328349736790804]
+  force: [-111.96106176875882, 0.0, 945.8442810644318]
+  force_coefficients: [0.6909542731559715, 0.015067295934023494, 0.0]
+  gamma: [3.267733154548073, 4.3577345602982405, 4.9171287274668805, 5.246853887118394, 5.455912998890128, 5.593876206102452, 5.6861788078552875, 5.746815436361271, 5.783707126398217, 5.801189065252356, 5.8011890652523554, 5.783707126398217, 5.746815436361271, 5.686178807855288, 5.593876206102453, 5.455912998890127, 5.2468538871183945, 4.9171287274668805, 4.3577345602982405, 3.2677331545480737]
+  lift_drag_side: [952.2213576930732, 20.76461720907613, 0.0]
+  moment: [-94.58442810644368, 158.93733732303883, -11.196106176875904]
+  moment_coefficients: [-0.06863269159651242, 0.11532868014370165, -0.008124158677098162]
+  n_panels: 20
+  name: straight
+  normal:
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  - [0.0, 0.0, 1.0]
+  options: {}
+  panel_va:
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  - [14.854021031123555, 0.0, 2.0875965144009814]
+  pivot: [0.0, 0.0, 0.0]
+  projected_area: 10.0
+  q_ref: 137.8125
+  reference_point: [0.3, 0.1, 1.0]
+  relaxation_factor: 0.19316240560417844
+  sections:
+  - LE: [0.0, 5.0, 0.0]
+    TE: [1.0, 5.0, 0.0]
+    polar:
+    - [-0.17453292519943295, -1.096622711232151, -0.0, -0.0]
+    - [-0.13962634015954636, -0.8772981689857208, -0.0, -0.0]
+    - [-0.10471975511965978, -0.6579736267392906, -0.0, -0.0]
+    - [-0.06981317007977318, -0.4386490844928604, -0.0, -0.0]
+    - [-0.03490658503988659, -0.2193245422464302, -0.0, -0.0]
+    - [0.0, 0.0, 0.0, 0.0]
+    - [0.03490658503988659, 0.2193245422464302, 0.0, 0.0]
+    - [0.06981317007977318, 0.4386490844928604, 0.0, 0.0]
+    - [0.10471975511965978, 0.6579736267392906, 0.0, 0.0]
+    - [0.13962634015954636, 0.8772981689857208, 0.0, 0.0]
+    - [0.17453292519943295, 1.096622711232151, 0.0, 0.0]
+    - [0.20943951023931956, 1.3159472534785812, 0.0, 0.0]
+    - [0.24434609527920614, 1.5352717957250113, 0.0, 0.0]
+    - [0.2792526803190927, 1.7545963379714415, 0.0, 0.0]
+    - [0.3141592653589793, 1.9739208802178716, 0.0, 0.0]
+    - [0.3490658503988659, 2.193245422464302, 0.0, 0.0]
+    - [0.3839724354387525, 2.4125699647107317, 0.0, 0.0]
+    - [0.4188790204786391, 2.6318945069571624, 0.0, 0.0]
+    - [0.4537856055185257, 2.8512190492035923, 0.0, 0.0]
+    - [0.4886921905584123, 3.0705435914500225, 0.0, 0.0]
+    - [0.5235987755982988, 3.2898681336964524, 0.0, 0.0]
+  - LE: [0.0, -5.0, 0.0]
+    TE: [1.0, -5.0, 0.0]
+    polar:
+    - [-0.17453292519943295, -1.096622711232151, -0.0, -0.0]
+    - [-0.13962634015954636, -0.8772981689857208, -0.0, -0.0]
+    - [-0.10471975511965978, -0.6579736267392906, -0.0, -0.0]
+    - [-0.06981317007977318, -0.4386490844928604, -0.0, -0.0]
+    - [-0.03490658503988659, -0.2193245422464302, -0.0, -0.0]
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+    - [-0.03490658503988659, 0.1954715829000001, 0.195002256408, 0.10162101249999997]
+    - [0.0, 0.5760560950800001, 0.1245390976, 0.22238077769999998]
+    - [0.03490658503988659, 0.9396355243880001, 0.195002256408, 0.10162101249999997]
+    - [0.06981317007977318, 1.2969637352880001, 0.406391732832, -0.26065828310000005]
+    - [0.10471975511965978, 1.6587945922440004, 0.758707526872, -0.8644571091000002]
+    - [0.13962634015954636, 2.03588195972, 1.251949638528, -1.7097754655000001]
+    - [0.17453292519943295, 2.43897970218, 1.8861180678, -2.7966133523000005]
+    - [0.20943951023931956, 2.8788416840880005, 2.661212814688, -4.124970769500001]
+    - [0.24434609527920614, 3.3662217699080004, 3.5772338791919998, -5.694847717100001]
+    - [0.2792526803190927, 3.9118738241040005, 4.6341812613119995, -7.506244195100001]
+    - [0.3141592653589793, 4.526551711140001, 5.832054961048, -9.559160203500001]
+    - [0.3490658503988659, 5.221009295480001, 7.1708549783999995, -11.853595742300001]
+    - [0.3839724354387525, 0.2602236057454936, 0.10513716130256759, -14.389550811500001]
+    - [0.4188790204786391, 0.3022642316338267, 0.13457670646612352, -17.167025411100003]
+    - [0.4537856055185257, 0.3454411777407038, 0.1684829198167133, -20.186019541100002]
+    - [0.4886921905584123, 0.3892095647956368, 0.20694639649469954, -23.446533201500003]
+    - [0.5235987755982988, 0.43301270189221924, 0.24999999999999992, -26.948566392300002]
+  - LE: [-0.12594856150986303, -4.095312460948618, 8.381752674216157]
+    TE: [0.9480319715306028, -4.119176793915081, 8.387506882641508]
+    polar:
+    - [-0.17453292519943295, -1.70376580514, 1.7561540103999997, -2.7088904081000003]
+    - [-0.13962634015954636, -1.1586845734640001, 1.167675268552, -1.6552693685000002]
+    - [-0.10471975511965978, -0.6696127632040001, 0.709969580448, -0.8357863377000001]
+    - [-0.06981317007977318, -0.22668856695200001, 0.383036946088, -0.25044131570000006]
+    - [-0.03490658503988659, 0.17994982269999998, 0.186877365472, 0.10076569749999997]
+    - [0.0, 0.56016421316, 0.1214908386, 0.21783470189999998]
+    - [0.03490658503988659, 0.9238164118359999, 0.186877365472, 0.10076569749999997]
+    - [0.06981317007977318, 1.280768226136, 0.383036946088, -0.25044131570000006]
+    - [0.10471975511965978, 1.640881463468, 0.709969580448, -0.8357863377000001]
+    - [0.13962634015954636, 2.0140179312399997, 1.167675268552, -1.6552693685000002]
+    - [0.17453292519943295, 2.41003943686, 1.7561540103999997, -2.7088904081000003]
+    - [0.20943951023931956, 2.838807787736, 2.475405805992, -3.9966494565]
+    - [0.24434609527920614, 3.310184791276, 3.325430655328, -5.5185465137000005]
+    - [0.2792526803190927, 3.834032254888, 4.306228558408, -7.2745815797]
+    - [0.3141592653589793, 4.42021198598, 5.417799515232, -9.2647546545]
+    - [0.3490658503988659, 5.07858579196, 6.660143525799999, -11.489065738100003]
+    - [0.3839724354387525, 0.2602236057454936, 0.10513716130256759, -13.947514830500001]
+    - [0.4188790204786391, 0.3022642316338267, 0.13457670646612352, -16.640101931700002]
+    - [0.4537856055185257, 0.3454411777407038, 0.1684829198167133, -19.5668270417]
+    - [0.4886921905584123, 0.3892095647956368, 0.20694639649469954, -22.727690160500003]
+    - [0.5235987755982988, 0.43301270189221924, 0.24999999999999992, -26.122691288100004]
+  - LE: [0.10494809354160103, -4.110425010287237, 8.27069488190273]
+    TE: [0.9322352740642089, -4.136759541813347, 8.276769856483405]
+    polar:
+    - [-0.17453292519943295, -1.69553702646, 1.6310278393999993, -2.6211674639]
+    - [-0.13962634015954636, -1.1615180514160002, 1.0864971458719996, -1.6007632715]
+    - [-0.10471975511965978, -0.6794973784760001, 0.6629732731279998, -0.8071155663]
+    - [-0.06981317007977318, -0.2405052572880001, 0.3604562211679999, -0.24022434830000003]
+    - [-0.03490658503988659, 0.16442806249999992, 0.17894598999199995, 0.09991038249999996]
+    - [0.0, 0.54427233124, 0.11844257959999999, 0.21328862609999996]
+    - [0.03490658503988659, 0.907997299284, 0.17894598999199995, 0.09991038249999996]
+    - [0.06981317007977318, 1.264572716984, 0.3604562211679999, -0.24022434830000003]
+    - [0.10471975511965978, 1.622968334692, 0.6629732731279998, -0.8071155663]
+    - [0.13962634015954636, 1.9921539027600002, 1.0864971458719996, -1.6007632715]
+    - [0.17453292519943295, 2.38109917154, 1.6310278393999993, -2.6211674639]
+    - [0.20943951023931956, 2.798773891384, 2.2965653537119994, -3.8683281435000003]
+    - [0.24434609527920614, 3.2541478126440007, 3.083109688807999, -5.3422453103]
+    - [0.2792526803190927, 3.756190685672, 3.9906608446879988, -7.0429189643]
+    - [0.3141592653589793, 4.31387226082, 5.019218821351998, -8.9703491055]
+    - [0.3490658503988659, 4.93616228844, 6.168783618799997, -11.124535733899998]
+    - [0.3839724354387525, 0.2602236057454936, 0.10513716130256759, -13.505478849499998]
+    - [0.4188790204786391, 0.3022642316338267, 0.13457670646612352, -16.1131784523]
+    - [0.4537856055185257, 0.3454411777407038, 0.1684829198167133, -18.9476345423]
+    - [0.4886921905584123, 0.3892095647956368, 0.20694639649469954, -22.0088471195]
+    - [0.5235987755982988, 0.43301270189221924, 0.24999999999999992, -25.296816183900003]
+  spanwise:
+  - [0.6985791782360972, 0.1257972036168741, -0.704387815977217]
+  - [0.2368167927993375, 0.26734829279805294, -0.9340464105097788]
+  - [0.0903134721763781, 0.5947932656943928, -0.7987893638676257]
+  - [0.06184778630998915, 0.8016207198114119, -0.5946251532668189]
+  - [0.08379619477088915, 0.8384348873836808, -0.5385212506111842]
+  - [0.04369005614237356, 0.9571059846379837, -0.28642505680270375]
+  - [0.05440353769941419, 0.9398197044157173, -0.3373116930640595]
+  - [0.010735494483407174, 0.9999177943378356, -0.0070109717411059765]
+  - [0.023068000905146515, 0.9893600631793182, -0.1436472510006905]
+  - [-0.023068000905146515, 0.9893600631793182, 0.1436472510006905]
+  - [-0.01073549448340719, 0.9999177943378356, 0.007010971741105987]
+  - [-0.05440353769941417, 0.9398197044157173, 0.3373116930640594]
+  - [-0.0436900561423736, 0.9571059846379835, 0.28642505680270397]
+  - [-0.08379619477088893, 0.838434887383681, 0.5385212506111838]
+  - [-0.06184778630998925, 0.8016207198114125, 0.5946251532668182]
+  - [-0.09031347217637813, 0.5947932656943943, 0.7987893638676247]
+  - [-0.23681679279933743, 0.26734829279805244, 0.934046410509779]
+  - [-0.6985791782360949, 0.12579720361687494, 0.7043878159772193]
+  va: [14.854021031123555, 0.0, 2.0875965144009814]
+  va_ref: [14.854021031123555, 0.0, 2.0875965144009814]
+  width: [0.6258880488816809, 0.6279356959355253, 0.6245319211523185, 0.6346776890263534, 0.6292130498130764, 0.6339059064338606, 0.6315153516729275, 0.6343894445054834, 0.6347949776147365, 0.6347949776147365, 0.6343894445054825, 0.6315153516729277, 0.63390590643386, 0.6292130498130767, 0.634677689026354, 0.624531921152317, 0.6279356959355232, 0.6258880488816839]
+defaults: {core_radius_fraction: 0.05, is_aoa_corrected: true, is_with_attached_trailed_vortex_force: true, max_iterations: 5000}
+python_commit: a2a91cb02c05cef4c0201dfe2a666eba8e941f0f
diff --git a/test/filament/test_semi_infinite_filament.jl b/test/filament/test_semi_infinite_filament.jl
index 11014cae..33964ba5 100644
--- a/test/filament/test_semi_infinite_filament.jl
+++ b/test/filament/test_semi_infinite_filament.jl
@@ -83,7 +83,7 @@ end
             filament.va,
             work_vectors
         )
-        @test all(isnan.(induced_velocity))
+        @test iszero(induced_velocity)
 
         for point in test_points
             velocity_3D_trailing_vortex_semiinfinite!(
diff --git a/test/panel/test_panel.jl b/test/panel/test_panel.jl
index 05c8b42a..68e76de3 100644
--- a/test/panel/test_panel.jl
+++ b/test/panel/test_panel.jl
@@ -207,9 +207,23 @@ end
         @test abs(dot(axes.z_airf, axes.x_airf)) < 1e-12
     end
 
-    @testset "the frame closes as z = x cross y" begin
-        closure = cross(axes.x_airf, axes.y_airf)
-        @test isapprox(closure ./ norm(closure), axes.z_airf; atol=1e-12)
+    @testset "the frame is orthonormal with z = x cross y" begin
+        @test abs(dot(axes.x_airf, axes.y_airf)) < 1e-12
+        @test isapprox(cross(axes.x_airf, axes.y_airf), axes.z_airf; atol=1e-12)
+    end
+
+    @testset "the chord is the section chords' mean square to the bound vortex" begin
+        y_airf = span_vec ./ norm(span_vec)
+        square_chord(le, te) = norm((te .- le) .- dot(te .- le, y_airf) .* y_airf)
+        @test axes.chord ≈ (square_chord(le_1, te_1) + square_chord(le_2, te_2)) / 2
+        @test axes.chord < (norm(te_1 .- le_1) + norm(te_2 .- le_2)) / 2
+    end
+
+    @testset "drag runs along the flow and square to the lift" begin
+        alpha = deg2rad(9)
+        dirs = panel_force_directions(axes, alpha)
+        @test dirs.dir_drag ≈ cos(alpha) .* axes.x_airf .+ sin(alpha) .* axes.z_airf
+        @test abs(dot(dirs.dir_lift, dirs.dir_drag)) < 1e-12
     end
 
     @testset "orient flips y and z together" begin
@@ -227,7 +241,7 @@ end
             @test alpha ≈ crossflow_alpha(v_eff, x_airf, y_airf, z_airf)
             along_span = v_eff .+ 2.0 .* y_airf
             @test panel_inflow(oriented, along_span, along_span, zeros(3)).alpha ≈ alpha
-            dir_lift = panel_force_directions(oriented, alpha, y_airf).dir_lift
+            dir_lift = panel_force_directions(oriented, alpha).dir_lift
             v_crossflow = v_eff .- dot(v_eff, y_airf) .* y_airf
             @test abs(dot(dir_lift, v_crossflow)) < 1e-12
             @test abs(dot(dir_lift, y_airf)) < 1e-12
diff --git a/test/runtests.jl b/test/runtests.jl
index 39350918..6d695476 100644
--- a/test/runtests.jl
+++ b/test/runtests.jl
@@ -36,6 +36,7 @@ function include_selected_tests()
     end
     should_run_test("body_aerodynamics/test_body_aerodynamics.jl") && include("body_aerodynamics/test_body_aerodynamics.jl")
     should_run_test("body_aerodynamics/test_results.jl") && include("body_aerodynamics/test_results.jl")
+    should_run_test("body_aerodynamics/test_curved_bound_correction.jl") && include("body_aerodynamics/test_curved_bound_correction.jl")
     should_run_test("test_refinement_validation.jl") && include("test_refinement_validation.jl")
     should_run_test("filament/test_bound_filament.jl") && include("filament/test_bound_filament.jl")
     should_run_test("filament/test_semi_infinite_filament.jl") && include("filament/test_semi_infinite_filament.jl")
@@ -63,6 +64,7 @@ function include_selected_tests()
     should_run_test("solver/test_unrefined_dist.jl") && include("solver/test_unrefined_dist.jl")
     should_run_test("solver/test_stability.jl") && include("solver/test_stability.jl")
     should_run_test("verification/test_verification.jl") && include("verification/test_verification.jl")
+    should_run_test("verification/test_python_reference.jl") && include("verification/test_python_reference.jl")
     should_run_test("VortexStepMethod/test_VortexStepMethod.jl") && include("VortexStepMethod/test_VortexStepMethod.jl")
     should_run_test("wake/test_wake.jl") && include("wake/test_wake.jl")
     should_run_test("wing_geometry/test_wing_geometry.jl") && include("wing_geometry/test_wing_geometry.jl")
diff --git a/test/solver/solver_test_polar.csv b/test/solver/solver_test_polar.csv
index 7e043d9c..dd714098 100644
--- a/test/solver/solver_test_polar.csv
+++ b/test/solver/solver_test_polar.csv
@@ -1,4 +1,5 @@
 alpha,cl,cd,cm
+-20.0,-0.8,0.04,-0.02
 -10.0,0.1,0.02,-0.01
 0.0,1.0,0.01,0.0
 10.0,1.8,0.025,0.01
diff --git a/test/solver/test_attached_trailed_force.jl b/test/solver/test_attached_trailed_force.jl
index 77c820a8..3ed2bb19 100644
--- a/test/solver/test_attached_trailed_force.jl
+++ b/test/solver/test_attached_trailed_force.jl
@@ -1,5 +1,5 @@
 using VortexStepMethod
-using VortexStepMethod: attached_trailed_loads, calc_forces!, solver_kwargs
+using VortexStepMethod: calc_forces!, solver_kwargs
 using DifferentiationInterface
 using LinearAlgebra
 using Test
@@ -42,11 +42,11 @@ function lift_and_oswald(wing, aspect_ratio, model, is_with_attached_trailed_for
 end
 
 @testset "Attached trailed vortex force" begin
-    @testset "defaults to off and follows the solver settings" begin
-        @test VortexStepMethod.SolverSettings().is_with_attached_trailed_force == false
-        @test Solver(4, 2).is_with_attached_trailed_force == false
-        settings = VortexStepMethod.SolverSettings(is_with_attached_trailed_force=true)
-        @test solver_kwargs(settings).is_with_attached_trailed_force
+    @testset "defaults to on and follows the solver settings" begin
+        @test VortexStepMethod.SolverSettings().is_with_attached_trailed_force
+        @test Solver(4, 2).is_with_attached_trailed_force
+        settings = VortexStepMethod.SolverSettings(is_with_attached_trailed_force=false)
+        @test !solver_kwargs(settings).is_with_attached_trailed_force
     end
 
     # Gaunaa, Li & Pirrung 2026 §5: without the force CL is ~2 % high on the curved wing.
@@ -67,23 +67,19 @@ end
     wing, _ = curved_elliptic_wing(; sweep=0.1, anhedral=0.25, n_sections=21)
     va_vec = 20.0 .* [cosd(10), 0.0, sind(10)]
 
-    @testset "adds attached_trailed_loads to every panel in solve! and solve" begin
+    @testset "adds the leg forces to the panels in solve!" begin
         body_aero = BodyAerodynamics([wing])
         set_va!(body_aero, va_vec)
         reference_point = [0.1, 0.0, -0.2]
         solver_off = Solver(wing.n_panels, wing.n_unrefined_sections; use_gamma_prev=false,
-            reference_point)
+            reference_point, is_with_attached_trailed_force=false)
         solver_on = Solver(wing.n_panels, wing.n_unrefined_sections; use_gamma_prev=false,
             reference_point, is_with_attached_trailed_force=true)
         force_off = copy(solve!(solver_off, body_aero).f_body_3D)
-        moment_off = copy(solver_off.sol.m_body_3D)
+        @test iszero(solver_off.sol.f_attached_trailed)
         solve!(solver_on, body_aero)
-        for i in eachindex(body_aero.panels)
-            attached = attached_trailed_loads(body_aero, i, solver_on.lr.gamma_new,
-                solver_on.density, solver_on.core_radius_fraction, reference_point)
-            @test solver_on.sol.f_body_3D[:, i] ≈ force_off[:, i] .+ attached.force
-            @test solver_on.sol.m_body_3D[:, i] ≈ moment_off[:, i] .+ attached.moment
-        end
+        @test solver_on.sol.f_body_3D ≈ force_off .+ solver_on.sol.f_attached_trailed
+        @test !iszero(solver_on.sol.f_attached_trailed)
         @test (@allocated calc_forces!(solver_on, body_aero)) == 0
     end
 
diff --git a/test/solver/test_polar_vectors.jl b/test/solver/test_polar_vectors.jl
index 48439163..2544223a 100644
--- a/test/solver/test_polar_vectors.jl
+++ b/test/solver/test_polar_vectors.jl
@@ -38,7 +38,7 @@ end
         sol = solve_three_panel_wing(POLAR_VECTORS,
             linear_polar(4π, 0.1, -0.02), linear_polar(2π, 0.05, -0.01))
         @test sol.solver_status == FEASIBLE
-        @test sol.cl_dist ≈ [4π, 3π, 2π] .* sol.alpha_dist
+        @test sol.cl_dist ≈ [4π, 3π, 2π] .* sol.alpha_uncorrected
         @test sol.cd_dist ≈ [0.1, 0.075, 0.05]
         @test sol.cm_dist ≈ [-0.02, -0.015, -0.01]
         @testset "the stronger +y side rolls the wing positive about x" begin
diff --git a/test/solver/test_solver.jl b/test/solver/test_solver.jl
index 872848cc..7023a8a2 100644
--- a/test/solver/test_solver.jl
+++ b/test/solver/test_solver.jl
@@ -218,7 +218,7 @@ end
 
 @testset "solve! fills the reference inflow and the centers of pressure" begin
     body_aero = inviscid_plates_aero([(4.0, -4.0)])
-    solver = Solver(length(body_aero.panels), 2)
+    solver = Solver(length(body_aero.panels), 2; is_with_attached_trailed_force=false)
     sol = solve!(solver, body_aero)
     @test sol.va_ref_vec ≈ 12.0 .* [cosd(5), 0.0, sind(5)]
     @test sol.q_ref ≈ 0.5 * solver.density * 12.0^2
diff --git a/test/solver/test_stability.jl b/test/solver/test_stability.jl
index 758b68a1..9bff8d3f 100644
--- a/test/solver/test_stability.jl
+++ b/test/solver/test_stability.jl
@@ -93,7 +93,7 @@ end
 
     @testset "moments about the trailing edge: unstable trim" begin
         body_aero, solver = trimmable_wing(-0.05; reference_point=[1.0, 0.0, 0.0])
-        trim = only(trim_angle(solver, body_aero, beta, va))
+        trim = only(trim_angle(solver, body_aero, beta, va; alpha_tol=1e-6))
         trim_coeffs = coeffs_at_angles(solver, body_aero, trim.alpha, beta, va)
         @test abs(trim_coeffs[5]) < 1e-5
         @test trim.dCMy_dalpha > 0
diff --git a/test/solver/test_unrefined_dist.jl b/test/solver/test_unrefined_dist.jl
index 9bca8a5c..d601cead 100644
--- a/test/solver/test_unrefined_dist.jl
+++ b/test/solver/test_unrefined_dist.jl
@@ -2,6 +2,15 @@ using VortexStepMethod
 using LinearAlgebra
 using Test
 
+"""
+    area_mean(values, panels, indices)
+
+Mean of `values[indices]` weighted by the area of `panels[indices]`.
+"""
+area_mean(values, panels, indices) =
+    sum(values[i] * panels[i].chord * panels[i].width for i in indices) /
+    sum(panels[i].chord * panels[i].width for i in indices)
+
 @testset "Unrefined Arrays Tests" begin
     @testset "Unrefined section array aggregation" begin
         # Create a simple wing with unrefined sections
@@ -28,6 +37,7 @@ using Test
             va_vec = [10.0, 0.0, 0.0]
             set_va!(body_aero, va_vec)
             sol = solve!(solver, body_aero)
+            @test sol.solver_status == FEASIBLE
 
             # Test 1: Unrefined arrays exist and have correct size
             @test length(sol.cl_unrefined_dist) == wing.n_unrefined_sections
@@ -38,50 +48,33 @@ using Test
             @test !all(sol.cl_unrefined_dist .== 0.0)
             @test !all(sol.cd_unrefined_dist .== 0.0)
 
-            # Test 3: Verify unrefined coefficients are averaged from refined panels
+            # Test 3: Unrefined coefficients are area-weighted means of their refined panels
             # refined_panel_mapping maps each refined panel to its unrefined section index
+            panels = body_aero.panels
             for unrefined_idx in 1:wing.n_unrefined_sections
                 # Find all refined panels that map to this unrefined section
                 refined_panel_indices = findall(x -> x == unrefined_idx, wing.refined_panel_mapping)
 
                 if !isempty(refined_panel_indices)
-                    # Calculate expected average from refined panel coefficients
-                    expected_cl = sum(sol.cl_dist[refined_panel_indices]) / length(refined_panel_indices)
-                    expected_cd = sum(sol.cd_dist[refined_panel_indices]) / length(refined_panel_indices)
-                    expected_cm = sum(sol.cm_dist[refined_panel_indices]) / length(refined_panel_indices)
-
-                    # Check if unrefined coefficients match expected averages
-                    # Handle NaN values that can occur in INVISCID models
-                    if isnan(expected_cl)
-                        @test isnan(sol.cl_unrefined_dist[unrefined_idx])
-                    else
-                        @test isapprox(sol.cl_unrefined_dist[unrefined_idx], expected_cl, rtol=1e-10)
-                    end
-                    if isnan(expected_cd)
-                        @test isnan(sol.cd_unrefined_dist[unrefined_idx])
-                    else
-                        @test isapprox(sol.cd_unrefined_dist[unrefined_idx], expected_cd, rtol=1e-10)
-                    end
-                    if isnan(expected_cm)
-                        @test isnan(sol.cm_unrefined_dist[unrefined_idx])
-                    else
-                        @test isapprox(sol.cm_unrefined_dist[unrefined_idx], expected_cm, rtol=1e-10)
-                    end
+                    expected_cl = area_mean(sol.cl_dist, panels, refined_panel_indices)
+                    expected_cd = area_mean(sol.cd_dist, panels, refined_panel_indices)
+                    expected_cm = area_mean(sol.cm_dist, panels, refined_panel_indices)
+
+                    @test isapprox(sol.cl_unrefined_dist[unrefined_idx], expected_cl, rtol=1e-10)
+                    @test isapprox(sol.cd_unrefined_dist[unrefined_idx], expected_cd, rtol=1e-10)
+                    @test isapprox(sol.cm_unrefined_dist[unrefined_idx], expected_cm, rtol=1e-10)
                 end
             end
 
             # Test 4: Verify physical consistency (lift coefficients should be positive at positive AoA)
-            # Skip test if values are NaN
-            if !any(isnan.(sol.cl_unrefined_dist))
-                @test all(sol.cl_unrefined_dist .> 0.0)
-            end
+            @test all(sol.cl_unrefined_dist .> 0.0)
 
         finally
             rm(settings_file; force=true)
         end
     end
 
-    @testset "Unrefined force consistency: coeff*width*chord" begin
+    @testset "Unrefined section loads equal the sums of their panel loads" begin
         # Verify that for each unrefined section:
         # coeff_unrefined * width_unrefined * chord_unrefined ≈
         #   sum(coeff_panel * width_panel * chord_panel) for all panels in section
@@ -102,6 +95,7 @@ using Test
             va_vec = [10.0, 0.0, 0.0]
             set_va!(body_aero, va_vec)
             sol = solve!(solver, body_aero)
+            @test sol.solver_status == FEASIBLE
 
             panels = body_aero.panels
             for unrefined_idx in 1:wing.n_unrefined_sections
@@ -133,17 +127,11 @@ using Test
                     panels[i].width for i in pidxs)
                 @test w_u ≈ expected_width rtol=1e-12
 
-                # Force-like quantities should be approximately
-                # equal (not exact due to averaging)
-                if !isnan(cl_u)
-                    @test cl_u * c_u * w_u ≈ sum_cl_cw rtol=0.05
-                end
-                if !isnan(cd_u)
-                    @test cd_u * c_u * w_u ≈ sum_cd_cw rtol=0.05
-                end
-                if !isnan(cm_u)
-                    @test cm_u * c_u * w_u ≈ sum_cm_cw rtol=0.05
-                end
+                @test cl_u * c_u * w_u ≈ sum_cl_cw rtol=1e-12
+                @test cd_u * c_u * w_u ≈ sum_cd_cw rtol=1e-12
+                @test cm_u * c_u * w_u ≈ sum_cm_cw rtol=1e-12
+                @test sol.moment_unrefined_dist[unrefined_idx] ≈
+                    sum(sol.moment_dist[i] for i in pidxs) rtol=1e-12
             end
         finally
             rm(settings_file; force=true)
@@ -176,6 +164,7 @@ using Test
                 va_vec = [10.0, 0.0, 0.0]
                 set_va!(body_aero, va_vec)
                 sol = solve!(solver, body_aero)
+                @test sol.solver_status == FEASIBLE
 
                 # Verify arrays have correct size
                 @test length(sol.cl_unrefined_dist) == wing.n_unrefined_sections
@@ -183,30 +172,18 @@ using Test
                 @test length(sol.cm_unrefined_dist) == wing.n_unrefined_sections
 
                 # Verify unrefined coefficients are computed correctly using mapping
+                panels = body_aero.panels
                 for unrefined_idx in 1:wing.n_unrefined_sections
                     refined_panel_indices = findall(x -> x == unrefined_idx, wing.refined_panel_mapping)
 
                     if !isempty(refined_panel_indices)
-                        expected_cl = sum(sol.cl_dist[refined_panel_indices]) / length(refined_panel_indices)
-                        expected_cd = sum(sol.cd_dist[refined_panel_indices]) / length(refined_panel_indices)
-                        expected_cm = sum(sol.cm_dist[refined_panel_indices]) / length(refined_panel_indices)
-
-                        # Handle NaN for all coefficients
-                        if isnan(expected_cl)
-                            @test isnan(sol.cl_unrefined_dist[unrefined_idx])
-                        else
-                            @test isapprox(sol.cl_unrefined_dist[unrefined_idx], expected_cl, rtol=1e-10)
-                        end
-                        if isnan(expected_cd)
-                            @test isnan(sol.cd_unrefined_dist[unrefined_idx])
-                        else
-                            @test isapprox(sol.cd_unrefined_dist[unrefined_idx], expected_cd, rtol=1e-10)
-                        end
-                        if isnan(expected_cm)
-                            @test isnan(sol.cm_unrefined_dist[unrefined_idx])
-                        else
-                            @test isapprox(sol.cm_unrefined_dist[unrefined_idx], expected_cm, rtol=1e-10)
-                        end
+                        expected_cl = area_mean(sol.cl_dist, panels, refined_panel_indices)
+                        expected_cd = area_mean(sol.cd_dist, panels, refined_panel_indices)
+                        expected_cm = area_mean(sol.cm_dist, panels, refined_panel_indices)
+
+                        @test isapprox(sol.cl_unrefined_dist[unrefined_idx], expected_cl, rtol=1e-10)
+                        @test isapprox(sol.cd_unrefined_dist[unrefined_idx], expected_cd, rtol=1e-10)
+                        @test isapprox(sol.cm_unrefined_dist[unrefined_idx], expected_cm, rtol=1e-10)
                     end
                 end
 
diff --git a/test/verification/test_python_reference.jl b/test/verification/test_python_reference.jl
new file mode 100644
index 00000000..e561d25f
--- /dev/null
+++ b/test/verification/test_python_reference.jl
@@ -0,0 +1,126 @@
+using LinearAlgebra
+using Test
+using VortexStepMethod
+using VortexStepMethod: calculate_AIC_matrices!, quarter_chord_AIC!, relaxation_factor_limit
+using YAML
+
+const PYTHON_REFERENCE = YAML.load_file(
+    joinpath(@__DIR__, "..", "data", "python_reference", "reference.yaml"))
+
+"""Rows of a nested list as a matrix."""
+as_matrix(rows) = reduce(vcat, permutedims(Float64.(row)) for row in rows)
+
+"""`POLAR_VECTORS` aero data of a reference section's `[alpha, cl, cd, cm]` table."""
+function polar_vectors(table)
+    polar = as_matrix(table)
+    return (polar[:, 1], polar[:, 2], polar[:, 3], polar[:, 4])
+end
+
+"""Body and solver of a reference `case`, with the Python call's inflow and options."""
+function reference_setup(case)
+    n_panels = case["n_panels"]
+    wing = Wing(n_panels; spanwise_distribution=LINEAR)
+    for section in case["sections"]
+        add_section!(wing, Float64.(section["LE"]), Float64.(section["TE"]), POLAR_VECTORS,
+            polar_vectors(section["polar"]))
+    end
+    refine!(wing)
+    body_aero = BodyAerodynamics([wing])
+    va = case["va"]
+    if first(va) isa AbstractVector
+        va_low, va_high = Float64.(va[1]), Float64.(va[2])
+        va_dist = [va_low[k] + (va_high[k] - va_low[k]) * (i - 1) / (n_panels - 1)
+                   for i in 1:n_panels, k in 1:3]
+        set_va!(body_aero, va_dist)
+    else
+        rates = case["body_rates"]
+        omega = isnothing(rates) ? zeros(3) : Float64.(rates)
+        set_va!(body_aero, Float64.(va), omega; reference_point=Float64.(case["pivot"]))
+    end
+    model = get(case["options"], "aerodynamic_model_type", "VSM") == "LLT" ? LLT : VSM
+    solver = Solver(n_panels, length(case["sections"]);
+        aerodynamic_model_type=model, rtol=1e-10, max_iterations=400_000,
+        reference_point=Float64.(case["reference_point"]),
+        is_with_viscous_drag_correction=get(case["options"],
+            "is_with_viscous_drag_correction", false))
+    return body_aero, solver, model
+end
+
+"""Julia's `(i, j, k)` influence array in the `(k, i, j)` nesting of the reference."""
+reference_order(matrices) = [matrices[k][i][j] for i in eachindex(matrices[1]),
+    j in eachindex(matrices[1]), k in eachindex(matrices)]
+
+@testset "The default solve agrees with the Python VSM v2.3.0" begin
+    @testset "Solver defaults are the Python ones" begin
+        defaults = PYTHON_REFERENCE["defaults"]
+        solver = Solver(4, 2)
+        @test solver.max_iterations == defaults["max_iterations"]
+        @test solver.core_radius_fraction == defaults["core_radius_fraction"]
+        @test solver.correct_aoa == defaults["is_aoa_corrected"]
+        @test solver.is_with_attached_trailed_force ==
+            defaults["is_with_attached_trailed_vortex_force"]
+        @test isnothing(solver.relaxation_factor)
+    end
+
+    @testset "$(case["name"])" for case in PYTHON_REFERENCE["cases"]
+        body_aero, solver, model = reference_setup(case)
+        panels = body_aero.panels
+        sol = solve!(solver, body_aero; log=false)
+        @test solver.lr.converged
+
+        force_scale = norm(case["force"])
+        moment_scale = norm(case["moment"])
+        @testset "panels have the perpendicular chord and an orthonormal frame" begin
+            @test [panel.chord for panel in panels] ≈ Float64.(case["chord"]) rtol = 1e-12
+            @test [panel.width for panel in panels] ≈ Float64.(case["width"]) rtol = 1e-9
+            for (i, panel) in enumerate(panels)
+                @test panel.x_airf ≈ Float64.(case["chordwise"][i]) atol = 1e-12
+                @test panel.y_airf ≈ Float64.(case["spanwise"][i]) atol = 1e-12
+                @test panel.z_airf ≈ Float64.(case["normal"][i]) atol = 1e-12
+            end
+        end
+        @testset "each panel feels the Python inflow" begin
+            @test sol.va_vec_dist ≈ as_matrix(case["panel_va"]) rtol = 1e-12
+        end
+        if !isnothing(case["AIC"])
+            @testset "influence matrices match" begin
+                @test body_aero.AIC ≈ reference_order(case["AIC"]) rtol = 1e-10
+                correction = zeros(size(body_aero.AIC))
+                quarter_chord_AIC!(correction, body_aero, solver.core_radius_fraction,
+                    solver.br.va_dist, solver.br.va_unit_dist)
+                @test correction ≈ reference_order(case["AIC_quarter_chord"]) rtol = 1e-10
+            end
+        end
+        @testset "circulation and angles match" begin
+            @test sol.gamma_distribution ≈ Float64.(case["gamma"]) rtol = 1e-6
+            @test solver.lr.alpha_dist ≈ Float64.(case["alpha"]) rtol = 1e-6
+            @test sol.alpha_dist ≈ Float64.(case["alpha_at_ac"]) rtol = 1e-6
+        end
+        @testset "panel forces and moments match" begin
+            forces = as_matrix(case["F_distribution"])'
+            moments = as_matrix(case["M_distribution"])'
+            @test maximum(abs, sol.f_body_3D .- forces) < 1e-6 * force_scale
+            @test maximum(abs, sol.m_body_3D .- moments) < 1e-6 * moment_scale
+            @test sol.cl_distribution ≈ Float64.(case["cl_distribution"]) rtol = 1e-6
+            @test sol.cd_distribution ≈ Float64.(case["cd_distribution"]) rtol = 1e-5
+        end
+        @testset "totals and coefficients match" begin
+            @test sol.force ≈ Float64.(case["force"]) atol = 1e-6 * force_scale
+            @test sol.moment ≈ Float64.(case["moment"]) atol = 1e-6 * moment_scale
+            @test [sol.lift, sol.drag, sol.side] ≈ Float64.(case["lift_drag_side"]) atol =
+                1e-6 * force_scale
+            @test [sol.cl, sol.cd, sol.cs] ≈ Float64.(case["force_coefficients"]) atol =
+                1e-6
+            @test sol.moment_coeffs ≈ Float64.(case["moment_coefficients"]) atol = 1e-5
+            @test sol.q_ref ≈ case["q_ref"] rtol = 1e-12
+            @test sol.projected_area ≈ case["projected_area"] rtol = 1e-12
+            @test sol.area_all_panels ≈ case["area_all_panels"] rtol = 1e-12
+        end
+        if case["name"] != "v3"
+            @testset "relaxation is 0.8 of the stability limit" begin
+                @test 0.8 * relaxation_factor_limit(panels, model) ≈
+                    case["relaxation_factor"] rtol = 1e-3
+            end
+        end
+    end
+end
diff --git a/test/verification/test_verification.jl b/test/verification/test_verification.jl
index cdf39467..ed5cb651 100644
--- a/test/verification/test_verification.jl
+++ b/test/verification/test_verification.jl
@@ -54,10 +54,10 @@ end
 Wing `CL` and `CD` at each angle of attack in `alphas` [deg], solved with the
 settings of the Python verification cases.
 """
-function lift_drag_polar(wing, model, alphas; va, relaxation_factor)
+function lift_drag_polar(wing, model, alphas; va, relaxation_factor, correct_aoa=true)
     body_aero = BodyAerodynamics([wing])
     solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=model,
-                    relaxation_factor, core_radius_fraction=1e-20)
+                    relaxation_factor, core_radius_fraction=1e-20, correct_aoa)
     CL = zeros(length(alphas))
     CD = zeros(length(alphas))
     for (i, alpha) in enumerate(alphas)
@@ -105,6 +105,19 @@ max_error(actual, expected) = maximum(abs.(actual .- expected))
         @test max_error(CD_llt, CD_vsm) < 1e-2
     end
 
+    @testset "elliptic AR 4 wing reaches Oswald efficiency 1 (Gaunaa et al. 2026)" begin
+        aspect_ratio = 4.0
+        span = aspect_ratio * π / 4
+        coordinates = generate_coordinates_el_wing(1.0, span, 41, "cos")
+        wing = wing_from_coordinates(coordinates, INVISCID)
+        oswald_efficiency(CL, CD) = only(CL)^2 / (π * aspect_ratio * only(CD))
+        polar(correct_aoa) = lift_drag_polar(wing, VSM, [10.0]; va=10.0,
+                                             relaxation_factor=0.05, correct_aoa)
+
+        @test oswald_efficiency(polar(true)...) ≈ 1 atol = 0.01
+        @test oswald_efficiency(polar(false)...) < 0.8
+    end
+
     @testset "curved Clark Y wing matches RANS" begin
         alphas = [9.0]
         coordinates = generate_coordinates_curved_wing(2.18, 6.969, π / 4, 4.673, 60,
diff --git a/test/wake/test_wake.jl b/test/wake/test_wake.jl
index 4c21405d..12c80f47 100644
--- a/test/wake/test_wake.jl
+++ b/test/wake/test_wake.jl
@@ -22,5 +22,15 @@ using VortexStepMethod
         # Test that the function accepts the right number of velocity vectors
         @test size(va_vec_dist, 1) == length(body_aero.panels)
         @test size(va_vec_dist, 2) == 3  # 3D velocity vectors
+
+        va_vec_dist[:, 2] .= range(-3.0, 3.0; length=n_panels)
+        VortexStepMethod.frozen_wake!(body_aero, va_vec_dist)
+        @testset "each panel's wake follows its own inflow" begin
+            for (i, panel) in enumerate(body_aero.panels), filament_idx in (4, 5)
+                wake = panel.filaments[filament_idx]
+                @test wake.va ≈ norm(va_vec_dist[i, :])
+                @test wake.direction ≈ normalize(va_vec_dist[i, :])
+            end
+        end
     end
 end