replace Umag with v_a

Author: ufechner7

README.md

@@ -85,7 +85,7 @@ Apart from the wing geometry there is no input file yet, the input has to be def
 n_panels = 20          # Number of panels
 span = 20.0            # Wing span [m]
 chord = 1.0            # Chord length [m]
-Umag = 20.0            # Magnitude of inflow velocity [m/s]
+v_a = 20.0            # Magnitude of inflow velocity [m/s]
 density = 1.225        # Air density [kg/m³]
 alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
@@ -107,7 +107,7 @@ add_section!(wing,
 wa = WingAerodynamics([wing])
 
 # Set inflow conditions
-vel_app = [cos(alpha), 0.0, sin(alpha)] .* Umag
+vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
 set_va!(wa, (vel_app, 0.0))  # Second parameter is yaw rate
 ```
 It is possible to import the wing geometry using an `.obj` file as shown in the example `ram_air_kite.jl`.

examples/ram_air_kite.jl

@@ -33,7 +33,7 @@ VSM_with_stall_correction = Solver(
 )
 
 # Setting velocity conditions
-Umag = 15.0
+v_a = 15.0
 aoa = 15.0
 side_slip = 0.0
 yaw_rate = 0.0
@@ -42,7 +42,7 @@ vel_app = [
     cos(aoa_rad) * cos(side_slip),
     sin(side_slip),
     sin(aoa_rad)
-] * Umag
+] * v_a
 wing_aero.va = vel_app
 
 # Plotting geometry
@@ -67,7 +67,7 @@ plot_distribution(
     [CAD_y_coordinates, CAD_y_coordinates],
     [results, results_with_stall],
     ["VSM", "VSM with stall correction"];
-    title="CAD_spanwise_distributions_alpha_$(round(aoa, digits=1))_beta_$(round(side_slip, digits=1))_yaw_$(round(yaw_rate, digits=1))_Umag_$(round(Umag, digits=1))",
+    title="CAD_spanwise_distributions_alpha_$(round(aoa, digits=1))_beta_$(round(side_slip, digits=1))_yaw_$(round(yaw_rate, digits=1))_v_a_$(round(v_a, digits=1))",
     data_type=".pdf",
     is_save=false,
     is_show=true
@@ -85,7 +85,7 @@ plot_polars(
     angle_of_attack=0,
     side_slip=0,
     yaw_rate=0,
-    Umag=10,
+    v_a=10,
     title="tutorial_testing_stall_model_n_panels_$(wing.n_panels)_distribution_$(wing.spanwise_panel_distribution)",
     data_type=".pdf",
     is_save=false,

examples/rectangular_wing.jl

@@ -6,7 +6,7 @@ using VortexStepMethod
 n_panels = 20          # Number of panels
 span = 20.0            # Wing span [m]
 chord = 1.0            # Chord length [m]
-Umag = 20.0            # Magnitude of inflow velocity [m/s]
+v_a = 20.0            # Magnitude of inflow velocity [m/s]
 density = 1.225        # Air density [kg/m³]
 alpha_deg = 30.0       # Angle of attack [degrees]
 alpha = deg2rad(alpha_deg)
@@ -28,7 +28,7 @@ add_section!(wing,
 wa = WingAerodynamics([wing])
 
 # Set inflow conditions
-vel_app = [cos(alpha), 0.0, sin(alpha)] .* Umag
+vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
 set_va!(wa, (vel_app, 0.0))  # Second parameter is yaw rate
 
 # Step 4: Initialize solvers for both LLT and VSM methods
@@ -81,7 +81,7 @@ plot_polars(
     ["LLT", "VSM"],
     angle_range=angle_range,
     angle_type="angle_of_attack",
-    Umag=Umag,
+    v_a=v_a,
     title="Rectangular Wing Polars"
 )
 nothing

examples/stall_model.jl

@@ -46,7 +46,7 @@ VSM_with_stall_correction = Solver(
 )
 
 # Setting velocity conditions
-Umag = 15.0
+v_a = 15.0
 aoa = 17.0
 side_slip = 0.0
 yaw_rate = 0.0
@@ -55,7 +55,7 @@ vel_app = [
     cos(aoa_rad) * cos(side_slip),
     sin(side_slip),
     sin(aoa_rad)
-] * Umag
+] * v_a
 wing_aero_CAD_19ribs.va = vel_app
 
 # # Plotting geometry
@@ -81,7 +81,7 @@ plot_distribution(
     [CAD_y_coordinates, CAD_y_coordinates],
     [results, results_with_stall],
     ["VSM", "VSM with stall correction"];
-    title="CAD_spanwise_distributions_alpha_$(round(aoa, digits=1))_beta_$(round(side_slip, digits=1))_yaw_$(round(yaw_rate, digits=1))_Umag_$(round(Umag, digits=1))",
+    title="CAD_spanwise_distributions_alpha_$(round(aoa, digits=1))_beta_$(round(side_slip, digits=1))_yaw_$(round(yaw_rate, digits=1))_v_a_$(round(v_a, digits=1))",
     data_type=".pdf",
     save_path=joinpath(save_folder, "spanwise_distributions"),
     is_save=false,
@@ -112,7 +112,7 @@ plot_polars(
     angle_of_attack=0,
     side_slip=0,
     yaw_rate=0,
-    Umag=10,
+    v_a=10,
     title="tutorial_testing_stall_model_n_panels_$(n_panels)_distribution_$(spanwise_panel_distribution)",
     data_type=".pdf",
     save_path=joinpath(save_folder, "polars"),

src/plotting.jl

@@ -477,7 +477,7 @@ Generate polar data for aerodynamic analysis over a range of angles.
 - `angle_of_attack`: Initial angle of attack in radians
 - `side_slip`: Initial side slip angle in radians
 - `yaw_rate`: Yaw rate
-- `Umag`: Magnitude of velocity
+- `v_a`: Magnitude of velocity
 
 # Returns
 - Tuple of polar data array and Reynolds number
@@ -490,7 +490,7 @@ function generate_polar_data(
     angle_of_attack=0.0,
     side_slip=0.0,
     yaw_rate=0.0,
-    Umag=10.0
+    v_a=10.0
 )
     n_panels = length(wing_aero.panels)
     n_angles = length(angle_range)
@@ -529,7 +529,7 @@ function generate_polar_data(
                 cos(α) * cos(β),
                 sin(β),
                 sin(α)
-            ] * Umag
+            ] * v_a
         )
 
         # Solve and store results
@@ -566,7 +566,7 @@ end
 """
     plot_polars(solver_list, wing_aero_list, label_list;
     literature_path_list=String[], angle_range=range(0, 20, 2), angle_type="angle_of_attack", angle_of_attack=0.0,
-    side_slip=0.0, yaw_rate=0.0, Umag=10.0, title="polar", data_type=".pdf", save_path=nothing, 
+    side_slip=0.0, yaw_rate=0.0, v_a=10.0, title="polar", data_type=".pdf", save_path=nothing, 
     is_save=true,
     is_show=true)
 
@@ -593,7 +593,7 @@ function plot_polars(
     angle_of_attack=0.0,
     side_slip=0.0,
     yaw_rate=0.0,
-    Umag=10.0,
+    v_a=10.0,
     title="polar",
     data_type=".pdf",
     save_path=nothing,
@@ -617,7 +617,7 @@ function plot_polars(
             angle_of_attack=angle_of_attack,
             side_slip=side_slip,
             yaw_rate=yaw_rate,
-            Umag=Umag
+            v_a=v_a
         )
         push!(polar_data_list, polar_data)
         # Update label with Reynolds number

src/solver.jl

@@ -112,7 +112,7 @@ function solve(solver::Solver, wing_aero::WingAerodynamics, gamma_distribution=n
 
     @debug "Initial gamma_new: $gamma_initial"
     # Run main iteration loop
-    converged, gamma_new, alpha_array, Umag_array = gamma_loop(
+    converged, gamma_new, alpha_array, v_a_array = gamma_loop(
         solver,
         wing_aero,
         gamma_initial,
@@ -130,7 +130,7 @@ function solve(solver::Solver, wing_aero::WingAerodynamics, gamma_distribution=n
     # Try again with reduced relaxation factor if not converged
     if !converged && relaxation_factor > 1e-3
         @warn "Running again with half the relaxation_factor = $(relaxation_factor/2)"
-        converged, gamma_new, alpha_array, Umag_array = gamma_loop(
+        converged, gamma_new, alpha_array, v_a_array = gamma_loop(
             solver,
             wing_aero,
             gamma_initial,
@@ -156,7 +156,7 @@ function solve(solver::Solver, wing_aero::WingAerodynamics, gamma_distribution=n
         solver.core_radius_fraction,
         solver.mu,
         alpha_array,
-        Umag_array,
+        v_a_array,
         chord_array,
         x_airf_array,
         y_airf_array,
@@ -196,7 +196,7 @@ function gamma_loop(
 )
     converged = false
     alpha_array = wing_aero.alpha_array
-    Umag_array = wing_aero.Umag_array
+    v_a_array = wing_aero.v_a_array
 
     iters = 0
     for i in 1:solver.max_iterations
@@ -218,11 +218,11 @@ function gamma_loop(
         v_tangential_array = vec(sum(y_airf_array .* relative_velocity_array, dims=2))
         alpha_array .= atan.(v_normal_array, v_tangential_array)
 
-        Umag_array .= norm.(relative_velocity_crossz)
-        Umagw_array = norm.(Uinfcrossz_array)
+        v_a_array .= norm.(relative_velocity_crossz)
+        v_aw_array = norm.(Uinfcrossz_array)
 
         cl_array = [calculate_cl(panel, alpha) for (panel, alpha) in zip(panels, alpha_array)]
-        gamma_new .= 0.5 .* Umag_array.^2 ./ Umagw_array .* cl_array .* chord_array
+        gamma_new .= 0.5 .* v_a_array.^2 ./ v_aw_array .* cl_array .* chord_array
 
         # Apply damping if needed
         if solver.is_with_artificial_damping
@@ -257,7 +257,7 @@ function gamma_loop(
         @warn "NO convergence after $(solver.max_iterations) iterations"
     end
 
-    return converged, gamma_new, alpha_array, Umag_array
+    return converged, gamma_new, alpha_array, v_a_array
 end
 
 """

src/wing_aerodynamics.jl

@@ -15,7 +15,7 @@ mutable struct WingAerodynamics
     stall_angle_list::Vector{Float64}
 
     alpha_array::Vector{Float64}
-    Umag_array::Vector{Float64}
+    v_a_array::Vector{Float64}
     work_vectors::NTuple{10,Vector{Float64}}
 
     function WingAerodynamics(
@@ -58,7 +58,7 @@ mutable struct WingAerodynamics
         stall_angle_list = calculate_stall_angle_list(panels)
 
         alpha_array = zeros(n_panels)
-        Umag_array = zeros(n_panels)    
+        v_a_array = zeros(n_panels)    
         work_vectors = ntuple(_ -> Vector{Float64}(undef, 3), 10)
 
         new(
@@ -71,7 +71,7 @@ mutable struct WingAerodynamics
             nothing,  # alpha_corrected
             stall_angle_list,
             alpha_array,
-            Umag_array,
+            v_a_array,
             work_vectors
         )
     end
@@ -377,7 +377,7 @@ end
     calculate_results(wa::WingAerodynamics, gamma_new::Vector{Float64}, 
                      density::Float64, aerodynamic_model_type::String,
                      core_radius_fraction::Float64, mu::Float64,
-                     alpha_array::Vector{Float64}, Umag_array::Vector{Float64},
+                     alpha_array::Vector{Float64}, v_a_array::Vector{Float64},
                      chord_array::Vector{Float64}, x_airf_array::Matrix{Float64},
                      y_airf_array::Matrix{Float64}, z_airf_array::Matrix{Float64},
                      va_array::Matrix{Float64}, va_norm_array::Vector{Float64},
@@ -396,7 +396,7 @@ function calculate_results(wa::WingAerodynamics,
     core_radius_fraction::Float64,
     mu::Float64,
     alpha_array::Vector{Float64},
-    Umag_array::Vector{Float64},
+    v_a_array::Vector{Float64},
     chord_array::Vector{Float64},
     x_airf_array::Matrix{Float64},
     y_airf_array::Matrix{Float64},
@@ -422,9 +422,9 @@ function calculate_results(wa::WingAerodynamics,
     end
 
     # Calculate forces
-    lift = reshape((cl_array .* 0.5 .* density .* Umag_array.^2 .* chord_array), :, 1)
-    drag = reshape((cd_array .* 0.5 .* density .* Umag_array.^2 .* chord_array), :, 1)
-    moment = reshape((cm_array .* 0.5 .* density .* Umag_array.^2 .* chord_array), :, 1)
+    lift = reshape((cl_array .* 0.5 .* density .* v_a_array.^2 .* chord_array), :, 1)
+    drag = reshape((cd_array .* 0.5 .* density .* v_a_array.^2 .* chord_array), :, 1)
+    moment = reshape((cm_array .* 0.5 .* density .* v_a_array.^2 .* chord_array), :, 1)
 
     # Calculate alpha corrections based on model type
     alpha_corrected = if aerodynamic_model_type == "VSM"

test/test_plotting.jl

@@ -7,7 +7,7 @@ function create_wa()
     n_panels = 20          # Number of panels
     span = 20.0            # Wing span [m]
     chord = 1.0            # Chord length [m]
-    Umag = 20.0            # Magnitude of inflow velocity [m/s]
+    v_a = 20.0            # Magnitude of inflow velocity [m/s]
     density = 1.225        # Air density [kg/m³]
     alpha_deg = 30.0       # Angle of attack [degrees]
     alpha = deg2rad(alpha_deg)
@@ -28,7 +28,7 @@ function create_wa()
     # Step 3: Initialize aerodynamics
     wa = WingAerodynamics([wing])
     # Set inflow conditions
-    vel_app = [cos(alpha), 0.0, sin(alpha)] .* Umag
+    vel_app = [cos(alpha), 0.0, sin(alpha)] .* v_a
     set_va!(wa, (vel_app, 0.0))  # Second parameter is yaw rate
     wa
 end
@@ -82,15 +82,15 @@ plt.ioff()
         @test fig isa plt.PyPlot.Figure
 
         # Step 8: Plot polar curves
-        Umag = 20.0            # Magnitude of inflow velocity [m/s]
+        v_a = 20.0            # Magnitude of inflow velocity [m/s]
         angle_range = range(0, 20, 20)
         fig = plot_polars(
             [llt_solver, vsm_solver],
             [wa, wa],
             ["LLT", "VSM"],
             angle_range=angle_range,
             angle_type="angle_of_attack",
-            Umag=Umag,
+            v_a=v_a,
             title="Rectangular Wing Polars",
             data_type=".pdf",
             save_path="/tmp",

test/test_wing_aerodynamics.jl

@@ -16,10 +16,10 @@ include("utils.jl")
     N = 40
     max_chord = 1.0
     span = 15.709  # AR = 20
-    Umag = 20.0
+    v_a = 20.0
     AR = span^2 / (π * span * max_chord / 4)
     aoa = deg2rad(5)
-    Uinf = [cos(aoa), 0.0, sin(aoa)] .* Umag
+    Uinf = [cos(aoa), 0.0, sin(aoa)] .* v_a
     model = "VSM"
 
     # Setup wing geometry
@@ -118,9 +118,9 @@ end
     @debug "N: $N"
     @debug "size(coord): $(size(coord))"
 
-    Umag = 20.0
+    v_a = 20.0
     aoa = 5.7106 * π / 180
-    Uinf = [cos(aoa), 0.0, sin(aoa)] .* Umag
+    Uinf = [cos(aoa), 0.0, sin(aoa)] .* v_a
 
     # Create wing geometry
     core_radius_fraction = 1e-20
@@ -215,9 +215,9 @@ end
         span = 17.0
         AR = span^2 / (π * span * max_chord / 4)
         @debug "AR: $AR"
-        Umag = 20.0
+        v_a = 20.0
         aoa = 5.7106 * π / 180
-        Uinf = [cos(aoa), 0.0, sin(aoa)] .* Umag
+        Uinf = [cos(aoa), 0.0, sin(aoa)] .* v_a
 
         coord = if wing_type == "rectangular"
             twist = range(-0.5, 0.5, length=N)

test/thesis_oriol_cayon.jl

@@ -74,10 +74,10 @@ function output_results(Fmag, aero_coeffs, ringvec, Uinf, controlpoints, Atot)
         SFtot += SideF[i] * norm(ringvec[i]["r0"])
     end
 
-    Umag = norm(Uinf)
-    CL = Ltot / (0.5 * Umag^2 * Atot * rho)
-    CD = Dtot / (0.5 * Umag^2 * Atot * rho)
-    CS = SFtot / (0.5 * Umag^2 * Atot * rho)
+    v_a = norm(Uinf)
+    CL = Ltot / (0.5 * v_a^2 * Atot * rho)
+    CD = Dtot / (0.5 * v_a^2 * Atot * rho)
+    CS = SFtot / (0.5 * v_a^2 * Atot * rho)
 
     return F_rel, F_gl, Ltot, Dtot, CL, CD, CS
 end