VortexStepMethod.jl/examples/rectangular_wing.jl at c8264df68b96e710879511e58fce172e2f1b6410 · OpenSourceAWE/VortexStepMethod.jl · GitHub

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using Pkg
if Base.active_project() != joinpath(@__DIR__, "Project.toml")
    Pkg.activate(@__DIR__)
end
using LinearAlgebra
using GLMakie
using MakieControlPlots
using VortexStepMethod

PLOT = true
SAVE_ALL = false
USE_TEX = false
OUTPUT_DIR = joinpath(dirname(@__DIR__), "output")

# Step 1: Define wing parameters
n_panels = 20          # Number of panels
span = 20.0            # Wing span [m]
chord = 1.0            # Chord length [m]
va = 20.0              # Magnitude of inflow velocity [m/s]
density = 1.225        # Air density [kg/m³]
alpha_deg = 30.0       # Angle of attack [degrees]
alpha = deg2rad(alpha_deg)

# Step 2: Create wing geometry with linear panel distribution
wing = Wing(n_panels, spanwise_distribution=LINEAR)

# Add wing sections - defining only tip sections with inviscid airfoil model
add_section!(wing,
    [0.0, span/2, 0.0],   # Left tip LE
    [chord, span/2, 0.0],  # Left tip TE
    INVISCID)
add_section!(wing,
    [0.0, -span/2, 0.0],  # Right tip LE
    [chord, -span/2, 0.0], # Right tip TE
    INVISCID)

# Refine mesh
refine!(wing)

# Step 3: Initialize aerodynamics
body_aero = BodyAerodynamics([wing])

# Set inflow conditions
va_vec = apparent_wind(alpha, 0.0, va)
set_va!(body_aero, va_vec, [0, 0, 0.1])

# Step 4: Initialize solvers for both LLT and VSM methods
llt_solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=LLT)
vsm_solver = Solver(wing.n_panels, wing.n_unrefined_sections; aerodynamic_model_type=VSM)

# Step 5: Solve using both methods
results_llt = solve(llt_solver, body_aero)
results_vsm = solve(vsm_solver, body_aero)
@time solve(llt_solver, body_aero)
@time solve(vsm_solver, body_aero)

# Print results comparison
println("\nLifting Line Theory Results:")
println("CL = $(round(results_llt["cl"], digits=4))")
println("CD = $(round(results_llt["cd"], digits=4))")
println("\nVortex Step Method Results:")
println("CL = $(round(results_vsm["cl"], digits=4))")
println("CD = $(round(results_vsm["cd"], digits=4))")
println("Projected area = $(round(results_vsm["projected_area"], digits=4)) m²")

# Step 6: Plot geometry
PLOT && plot_geometry(
      body_aero,
      "Rectangular wing geometry";
      save_path=OUTPUT_DIR,
      is_save=false || SAVE_ALL,
      is_show=true,
      use_tex=USE_TEX
)

# Step 7: Plot spanwise distributions
y_coordinates = [panel.aero_center[2] for panel in body_aero.panels]

PLOT && plot_distribution(
    [y_coordinates, y_coordinates],
    [results_vsm, results_llt],
    ["VSM", "LLT"],
    title="Spanwise Distributions",
    save_path=OUTPUT_DIR,
    is_save=false || SAVE_ALL,
    use_tex=USE_TEX
)

# Step 8: Plot polar curves
angle_range = range(0, 20, 20)
PLOT && plot_polars(
    [llt_solver, vsm_solver],
    [body_aero, body_aero],
    ["LLT", "VSM"];
    angle_range,
    angle_type="angle_of_attack",
    v_a=va,
    title="Rectangular Wing Polars",
    save_path=OUTPUT_DIR,
    is_save=false || SAVE_ALL,
    use_tex=USE_TEX
)
nothing