Clean up duplicate/ old files (#6)

* del old airtable

* del old blax

* air --> airfoil_data + add additional airfoils

docs/src/aero/drag.md

@@ -43,7 +43,7 @@ aerodynamics.blsys(simi,lami,wake,direct, Mach, uinv,
                       cfm, cfm_thm, cfm_dsm, cfm_uem,
                       dim, dim_thm, dim_dsm, dim_uem)
 
-aerodynamics.blax2(ndim, n,ite, xi, bi, rni, uinv, Reyn, Mach, fexcr)
+aerodynamics.blax(ndim, n,ite, xi, bi, rni, uinv, Reyn, Mach, fexcr)
 
 aerodynamics.blvar(simi,lami,wake, Reyn,Mach, fexcr,
                       x, θ ,δs ,ue )

src/aero/aero.jl

@@ -35,11 +35,11 @@ jdim::Int = 360
 
 # Aerofoil calculations
 include("airfoil.jl")
-include("airtable2.jl")
+include("airtable.jl")
 include("airfun.jl")
 
 __abs_path_prefix__ = dirname(@__DIR__)
-airfoil_data = joinpath(__abs_path_prefix__,"air/C.air")
+airfoil_data = joinpath(__abs_path_prefix__,"airfoil_data/C.air")
 
 airsection = airtable(airfoil_data);
 
@@ -53,7 +53,6 @@ include("wingsc.jl")
 include("fusebl.jl")
 include("axisol.jl")
 include("blax.jl")
-include("blax2.jl")
 include("blsys.jl")
 
 # Trefftz plane CDi calcs

src/aero/airfun.jl

@@ -3,7 +3,7 @@ export airfun
 """
     airfun(cl, τ, Mach, air::airfoil)
 
-Looks up airfoil performance data at specified conditions, as precomputed and found in `./src/air/C.air`.
+Looks up airfoil performance data at specified conditions, as precomputed and found in `./src/airfoil_data/`.
 
 !!! details "🔃 Inputs and Outputs"
       **Inputs:**      

src/aero/airtable.jl

@@ -1,22 +1,31 @@
-using Dierckx
-
 """
-Reads airfoil file and outputs a matrix and spline objects
+    airtable(fname)
+
+Reads airfoil file and outputs the matrix and splines as an `airfoil`.
+
+The airfoil data is stored as a function of three variables, typically
+Mach number ``\\mathrm{Ma}``, lift coefficient ``c_l``, and thickness to chord ratio ``\\tau``.
+
+    cdf(Ma, cl, τ), cdp(Ma, cl, τ), cm(Ma, cl, τ)
+
+!!! details "🔃 Inputs and Outputs"
+    **Inputs:**
+    - `fname::String`: Path to file.
 
-!!! compat "Future Changes"
-      Deprecated by faster [`airtable`](@ref) in `airtable2.jl`. Will be removed in a future revision.
+    **Outputs:**
+    - `airf::TASOPT.aerodynamics.airfoil`: struct with airfoil performance characteristics.
 
 """
 function airtable(fname)
 
-    # Read file
-    f = open(fname)
+# Read airfoil data
+f = open(fname)
 
     nAMa::Int, nAcl::Int, nAτ::Int, nAfun::Int = [parse(Int,ss) for ss in split(readline(f))]  
 
-    AMa = zeros(nAMa)
-    Acl = zeros(nAcl)
-    Aτ  = zeros(nAτ)
+    AMa = zeros(nAMa)     # array of Mach numbers in database
+    Acl = zeros(nAcl)     # array of cl           in database     
+    Aτ  = zeros(nAτ)      # array of t/c          in database    
     A   = zeros(nAMa, nAcl, nAτ, nAfun)
 
     @inbounds for  i = 1:nAMa
@@ -29,7 +38,7 @@ function airtable(fname)
         Aτ[i] = parse(Float64, readline(f))
     end
 
-    #Read in Reynolds number
+    # Read in Reynolds number
     ARe = parse(Float64, readline(f))
 
     #Read in airfoil data
@@ -41,21 +50,69 @@ function airtable(fname)
         end
     end
 
-    close(f)
+close(f)
 
-    # Create spline objects to make it faster for airfun.jl since this does not need to be computed each time
+# ---------------------
+     A_M = zeros(nAMa, nAcl, nAτ, nAfun)
+     A_τ = zeros(nAMa, nAcl, nAτ, nAfun)
+    A_cl = zeros(nAMa, nAcl, nAτ, nAfun)
+   A_M_τ = zeros(nAMa, nAcl, nAτ, nAfun)
+  A_M_cl = zeros(nAMa, nAcl, nAτ, nAfun)
+  A_cl_τ = zeros(nAMa, nAcl, nAτ, nAfun)
+A_M_cl_τ = zeros(nAMa, nAcl, nAτ, nAfun)
+# Iterate over the 3 functions - cdf, cdp and cm    
+for l = 1:nAfun
+    for k = 1:nAτ
+        for j = 1:nAcl
+            A_M[:, j, k, l] = spline(AMa, A[:, j, k, l]) # This gets the derivatives ∂cdf/∂M, ∂cdp/∂M, ∂cm/∂M
+        end
+    end
 
-    ∂cdf_∂M  = Array{Spline1D, 2}(undef, nAcl, nAτ)
-    ∂cdp_∂M  = Array{Spline1D, 2}(undef, nAcl, nAτ)
-     ∂cm_∂M  = Array{Spline1D, 2}(undef, nAcl, nAτ)
+    for k = 1:nAτ
+        for i = 1:nAMa
+            # for j = 1:nAcl
+                f   =   A[i, :, k, l]
+                f_M = A_M[i, :, k, l]
+                  A_cl[i, :, k, l] = spline(Acl, f)
+                A_M_cl[i, :, k, l] = spline(Acl, f_M)
+        end
+    end
 
-    @inbounds for  k = 1:nAτ
-        @inbounds for  j = 1:nAcl
-            ∂cdf_∂M[j, k] = Spline1D(AMa, A[:,j,k,1], k = 3)
-            ∂cdp_∂M[j, k] = Spline1D(AMa, A[:,j,k,2], k = 3)
-             ∂cm_∂M[j, k] = Spline1D(AMa, A[:,j,k,3], k = 3)
+    for j = 1:nAcl
+        for i = 1:nAMa
+            f   = A[i, j, :, l]
+            f_M = A_M[i, j, :, l]
+              A_τ[i, j, :, l] = spline(Aτ, f)
+            A_M_τ[i, j, :, l] = spline(Aτ, f_M)
+        end
+    end
+
+    for i = 1:nAMa
+        for j = 1:nAcl
+            f   = A_cl[i, j, :, l]
+            f_M = A_M_cl[i, j, :, l]
+              A_cl_τ[i, j, :, l] = spline(Aτ, f)
+            A_M_cl_τ[i, j, :, l] = spline(Aτ, f_M)
+        end
+        for k = 1:nAτ
+            f   = A_τ[i, :, k, l]
+            f_M = A_M_τ[i, :, k, l]
+              A_cl_τ[i, :, k, l] = 0.5*(  A_cl_τ[i, :, k, l] + spline(Acl, f))
+            A_M_cl_τ[i, :, k, l] = 0.5*(A_M_cl_τ[i, :, k, l] + spline(Acl, f_M))
         end
     end
 
-return nAMa, nAcl, nAτ, nAfun, AMa, Acl, Aτ, A, ∂cdf_∂M, ∂cdp_∂M, ∂cm_∂M
+
+end
+
+return  airfoil(AMa, Acl, Aτ, ARe,
+        A,
+        A_M,
+        A_τ,
+        A_cl,
+        A_M_τ,
+        A_M_cl,
+        A_cl_τ,
+        A_M_cl_τ)
 end
+