Unify RamAirWing and Wing into Wing

.github/workflows/CI.yml

@@ -23,7 +23,7 @@ jobs:
       matrix:
         version:
           - '1.10'
-          - '1'
+          - '1.11'
         os:
           - ubuntu-latest
         build_is_production_build:

Project.toml

@@ -38,16 +38,11 @@ VortexStepMethodControlPlotsExt = "ControlPlots"
 VortexStepMethodMakieExt = "Makie"
 
 [compat]
-Aqua = "0.8"
-BenchmarkTools = "1"
-CSV = "0.10"
 Colors = "0.13"
 ControlPlots = "0.2.5"
-DataFrames = "1.7"
 DefaultApplication = "1"
 DelimitedFiles = "1"
 DifferentiationInterface = "0.7.4"
-Documenter = "1.8"
 FiniteDiff = "2.27.0"
 Interpolations = "0.15, 0.16"
 LaTeXStrings = "1"
@@ -60,28 +55,13 @@ Parameters = "0.12"
 Pkg = "1"
 PreallocationTools = "0.4.31"
 PrecompileTools = "1.2.1"
-Random = "1.10.0"
 RecursiveArrayTools = "3 - 3.36.0"
 SciMLBase = "2.77.0"
 Serialization = "1"
 StaticArrays = "1"
 Statistics = "1"
 StructMapping = "0.2.3"
-Test = "1"
 Timers = "0.1"
 Xfoil = "1.1.0"
 YAML = "0.4.13"
 julia = "1.10, 1.11"
-
-[extras]
-Aqua = "4c88cf16-eb10-579e-8560-4a9242c79595"
-BenchmarkTools = "6e4b80f9-dd63-53aa-95a3-0cdb28fa8baf"
-CSV = "336ed68f-0bac-5ca0-87d4-7b16caf5d00b"
-ControlPlots = "23c2ee80-7a9e-4350-b264-8e670f12517c"
-DataFrames = "a93c6f00-e57d-5684-b7b6-d8193f3e46c0"
-Documenter = "e30172f5-a6a5-5a46-863b-614d45cd2de4"
-Random = "9a3f8284-a2c9-5f02-9a11-845980a1fd5c"
-Test = "8dfed614-e22c-5e08-85e1-65c5234f0b40"
-
-[targets]
-test = ["Test", "DataFrames", "CSV", "Documenter", "BenchmarkTools", "ControlPlots", "Aqua", "Random"]

README.md

@@ -19,9 +19,9 @@ if you haven't already. On Linux, make sure that Python3 and Matplotlib are inst
 ```
 sudo apt install python3-matplotlib
 ```
-Furthermore, the packages `TestEnv` and `ControlPlots` must be installed globally:
+Furthermore, the package `ControlPlots` must be installed globally:
 ```
-julia -e 'using Pkg; Pkg.add("TestEnv"); Pkg.add("ControlPlots")'
+julia -e 'using Pkg; Pkg.add("ControlPlots")'
 ```
 
 Before installing this software it is suggested to create a new project, for example like this:

bin/install

@@ -20,7 +20,7 @@ fi
 
 export JULIA_PKG_SERVER_REGISTRY_PREFERENCE=eager
 
-julia -e 'using Pkg; Pkg.add("TestEnv"); Pkg.add("ControlPlots")'
+julia -e 'using Pkg; Pkg.add("ControlPlots")'
 
 julia --project -e 'include("bin/install.jl")'
 

data/ram_air_kite/vsm_settings.yaml

@@ -10,13 +10,17 @@ PanelDistribution:
 InitialGammaDistribution:
     ELLIPTIC:           Elliptic distribution
     ZEROS:              Constant distribution
+PanelGroupingMethod:
+    EQUAL_SIZE:         Divide panels into equally-sized sequential groups
+    REFINE:             Group refined panels by their original unrefined section
 
 wings:
     - name: main_wing
       n_panels: 40
       n_groups: 40
       spanwise_panel_distribution: LINEAR
       spanwise_direction: [0.0, 1.0, 0.0]
+      grouping_method: EQUAL_SIZE
       remove_nan: true
 solver_settings:
     n_panels: 40

docs/make.jl

@@ -1,9 +1,3 @@
-using Pkg
-if ("TestEnv" ∈ keys(Pkg.project().dependencies))
-    if ! ("Documents" ∈ keys(Pkg.project().dependencies))
-        using TestEnv; TestEnv.activate()
-    end
-end
 using ControlPlots
 using VortexStepMethod
 using Documenter

docs/src/glossary.md

@@ -5,6 +5,7 @@
 | AIC | Aerodynamic Influence Coefficient (AIC). The AIC matrix represents the relationship between the induced velocities or pressures on aerodynamic surfaces and the circulation strength or modal deformations of the lifting surfaces.|
 | inviscid | A fluid flow in which viscosity is considered negligible or zero. This means that there is no internal friction between the fluid layers, and the effects of viscosity on the flow are assumed to be insignificant. |
 | Panel | Flat surface element in 3D that approximate the contour of the aerodynamic body being studied.|
+| Panel Group | A collection of panels whose aerodynamic forces and moments are summed together. Groups can be defined using EQUAL_SIZE (sequential grouping) or REFINE (based on original unrefined structure) methods.|
 | Section |A wing section, also known as an airfoil or aerofoil, is the cross-sectional shape of an aircraft wing.|
 | Span | Distance from one wing tip to the other wing tip. |
 | Polar | The polar typically plots the coefficient of lift (CL) against the coefficient of drag (CD), with the angle of attack as a parameter along the curve. |

docs/src/index.md

@@ -18,9 +18,9 @@ if you haven't already. On Linux, make sure that Python3 and Matplotlib are inst
 ```
 sudo apt install python3-matplotlib
 ```
-Furthermore, the packages `TestEnv` and `ControlPlots` must be installed globally:
+Furthermore, the package `ControlPlots` must be installed globally:
 ```
-julia -e 'using Pkg; Pkg.add("TestEnv"); Pkg.add("ControlPlots")'
+julia -e 'using Pkg; Pkg.add("ControlPlots")'
 ```
 
 Before installing this software it is suggested to create a new project, for example like this:

docs/src/tips_and_tricks.md

@@ -10,6 +10,38 @@ The following bodies can be simulated:
 To build the geometry of a RAM-air kite, a 3D .obj file can be used as input. In addition a `.dat` file is needed.
 It should have two columns, one for the `x` and one for the `y` coordinate of the 2D polar that is used.
 
+## Panel Grouping Methods
+When creating a wing, you can specify how panels should be grouped for moment and force calculations using the `grouping_method` parameter. Two methods are available:
+
+### EQUAL_SIZE (Default)
+Divides refined panels into equally-sized sequential groups. This is the original behavior.
+
+```julia
+wing = Wing(40; n_groups=4, grouping_method=EQUAL_SIZE)
+```
+
+In this example, with 40 panels and 4 groups, each group will contain 10 consecutive panels (panels 1-10, 11-20, 21-30, 31-40).
+
+### REFINE
+Groups refined panels back to their original unrefined section. This is useful when you want group moments and forces to represent the original wing structure, regardless of panel refinement.
+
+```julia
+# Create wing with 4 unrefined sections (3 panels)
+wing = Wing(40; n_groups=3, grouping_method=REFINE)
+add_section!(wing, [0, 5, 0], [1, 5, 0], INVISCID)    # Section 1
+add_section!(wing, [0, 2.5, 0], [1, 2.5, 0], INVISCID) # Section 2
+add_section!(wing, [0, 0, 0], [1, 0, 0], INVISCID)     # Section 3
+add_section!(wing, [0, -5, 0], [1, -5, 0], INVISCID)   # Section 4
+```
+
+**Important:** When using `REFINE`, `n_groups` must equal the number of unrefined panels (number of sections - 1). The solver will automatically map each refined panel to its closest original unrefined panel and sum their moments and forces accordingly.
+
+This is particularly useful for:
+- LEI kites where you want loads per rib
+- Wings with discrete control surfaces
+- Cases where physical structure doesn't align with uniform panel distribution
+- Dynamic simulations where you have fewer structural segments than panels needed for accurate VSM aerodynamics. For example, a 6-segment structural model can be combined with 40-panel aerodynamics by using `n_groups=6` and `grouping_method=REFINE` to map aerodynamic loads back to the structural segments.
+
 ## RAM-air kite model
 If running the example `ram_air_kite.jl` fails, try to run the `cleanup.jl` script and then try again. Background: this example caches the calculated polars. Reading cached polars can fail after an update.
 

docs/src/types.md

@@ -7,6 +7,7 @@ Model
 WingType
 AeroModel
 PanelDistribution
+PanelGroupingMethod
 InitialGammaDistribution
 SolverStatus
 ```
@@ -24,17 +25,17 @@ AeroData
 ```
 
 ## Wing Geometry, Panel and Aerodynamics
-A body is constructed of one or more abstract wings. An abstract wing can be a Wing or a RamAirWing. 
-A Wing/ RamAirWing has one or more sections.
+A body is constructed of one or more abstract wings. All wings are of type Wing. 
+A Wing has one or more sections and can be created from YAML files or OBJ geometry.
 ```@docs
 Section
 Section(LE_point::PosVector, TE_point::PosVector, aero_model)
 Wing
 Wing(n_panels::Int; spanwise_distribution::PanelDistribution=LINEAR,
      spanwise_direction::PosVector=MVec3([0.0, 1.0, 0.0]))
-RamAirWing
-RamAirWing(obj_path, dat_path; alpha=0.0, crease_frac=0.75, wind_vel=10., mass=1.0, 
-         n_panels=54, n_sections=n_panels+1, spanwise_distribution=UNCHANGED, 
+ObjWing
+ObjWing(obj_path, dat_path; alpha=0.0, crease_frac=0.75, wind_vel=10., mass=1.0,
+         n_panels=54, n_sections=n_panels+1, spanwise_distribution=UNCHANGED,
          spanwise_direction=[0.0, 1.0, 0.0])
 BodyAerodynamics
 ```

examples/Project.toml

@@ -0,0 +1,8 @@
+[deps]
+CSV = "336ed68f-0bac-5ca0-87d4-7b16caf5d00b"
+ControlPlots = "23c2ee80-7a9e-4350-b264-8e670f12517c"
+DataFrames = "a93c6f00-e57d-5684-b7b6-d8193f3e46c0"
+GLMakie = "e9467ef8-e4e7-5192-8a1a-b1aee30e663a"
+LinearAlgebra = "37e2e46d-f89d-539d-b4ee-838fcccc9c8e"
+REPL = "3fa0cd96-eef1-5676-8a61-b3b8758bbffb"
+VortexStepMethod = "ed3cd733-9f0f-46a9-93e0-89b8d4998dd9"

examples/bench.jl

@@ -2,12 +2,6 @@ using LinearAlgebra
 using ControlPlots
 using VortexStepMethod
 
-using Pkg
-
-if !("CSV" ∈ keys(Pkg.project().dependencies))
-    using TestEnv
-    TestEnv.activate()
-end
 
 # Step 1: Define wing parameters
 n_panels = 20          # Number of panels
@@ -56,9 +50,9 @@ println("Rectangular wing, solve:")
 @time solve(vsm_solver, body_aero, nothing)
 
 # Create wing geometry
-wing = RamAirWing(
-    joinpath("data", "ram_air_kite", "ram_air_kite_body.obj"), 
-    joinpath("data", "ram_air_kite", "ram_air_kite_foil.dat"); 
+wing = ObjWing(
+    joinpath("data", "ram_air_kite", "ram_air_kite_body.obj"),
+    joinpath("data", "ram_air_kite", "ram_air_kite_foil.dat");
     prn=false
 )
 body_aero = BodyAerodynamics([wing])

examples/menu.jl

@@ -1,8 +1,3 @@
-using Pkg
-if ! ("ControlPlots" ∈ keys(Pkg.project().dependencies))
-    using TestEnv; TestEnv.activate()
-end
-
 using ControlPlots
 using VortexStepMethod
 using REPL.TerminalMenus

examples/ram_air_kite.jl

@@ -1,4 +1,4 @@
-using ControlPlots
+using GLMakie
 using VortexStepMethod
 using LinearAlgebra
 
@@ -9,9 +9,9 @@ DEFORM = false
 LINEARIZE = false
 
 # Create wing geometry
-wing = RamAirWing(
-    joinpath("data", "ram_air_kite", "ram_air_kite_body.obj"), 
-    joinpath("data", "ram_air_kite", "ram_air_kite_foil.dat"); 
+wing = ObjWing(
+    joinpath("data", "ram_air_kite", "ram_air_kite_body.obj"),
+    joinpath("data", "ram_air_kite", "ram_air_kite_foil.dat");
     prn=PRN
 )
 body_aero = BodyAerodynamics([wing];)
@@ -21,7 +21,7 @@ println("First init")
 if DEFORM
     # Linear interpolation of alpha from 10° at one tip to 0° at the other
     println("Deform")
-    @time VortexStepMethod.smooth_deform!(wing, deg2rad.([10,20,10,0]), deg2rad.([-10,0,-10,0]))
+    @time group_deform!(wing, deg2rad.([10,20,10,0]), deg2rad.([-10,0,-10,0]); smooth=true)
     println("Deform init")
     @time VortexStepMethod.reinit!(body_aero; init_aero=false)
 end
@@ -118,4 +118,4 @@ PLOT && plot_polars(
     is_show=true,
     use_tex=USE_TEX
 )
-nothing
+nothing

examples/stall_model.jl

@@ -2,10 +2,6 @@ using ControlPlots
 using LinearAlgebra
 using VortexStepMethod
 
-using Pkg
-if ! ("CSV" ∈ keys(Pkg.project().dependencies))
-    using TestEnv; TestEnv.activate()
-end
 using CSV
 using DataFrames
 
@@ -38,7 +34,9 @@ for row in eachrow(df)
 end
 
 # Create wing geometry
-CAD_wing = Wing(n_panels; spanwise_distribution)
+# Using REFINE grouping method: n_groups should equal number of unrefined panels (18 sections = 18 panels)
+n_groups = length(rib_list) - 1
+CAD_wing = Wing(n_panels; spanwise_distribution, n_groups, grouping_method=REFINE)
 for rib in rib_list
     add_section!(CAD_wing, rib[1], rib[2], rib[3], rib[4])
 end
@@ -109,21 +107,23 @@ path_cfd_lebesque = joinpath(
     "V3_CL_CD_RANS_Lebesque_2024_Rey_300e4.csv"
 )
 
+# Only include literature data if file exists
+literature_paths = isfile(path_cfd_lebesque) ? [path_cfd_lebesque] : String[]
+labels = isfile(path_cfd_lebesque) ?
+    ["VSM CAD 19ribs", "VSM CAD 19ribs , with stall correction", "CFD_Lebesque Rey 30e5"] :
+    ["VSM CAD 19ribs", "VSM CAD 19ribs , with stall correction"]
+
 PLOT && plot_polars(
     [vsm_solver, VSM_with_stall_correction],
     [body_aero, body_aero],
-    [
-        "VSM CAD 19ribs",
-        "VSM CAD 19ribs , with stall correction",
-        "CFD_Lebesque Rey 30e5"
-    ];
-    literature_path_list=[path_cfd_lebesque],
+    labels;
+    literature_path_list=literature_paths,
     angle_range=range(0, 25, length=25),
     angle_type="angle_of_attack",
     angle_of_attack=0,
     side_slip=0,
     v_a=10,
-    title="tutorial_testing_stall_model_n_panels_$(n_panels)_distribution_$(spanwise_distribution)",
+    title="tutorial_testing_stall_model_n_panels_$(n_panels)_distribution_$(spanwise_distribution)_grouping_$(CAD_wing.grouping_method)",
     data_type=".pdf",
     save_path=joinpath(save_folder, "polars"),
     is_save=true,