remove unsteady aero params from first 10% blade

OpenFAST/IEA-15-240-RWT-Monopile/IEA-15-240-RWT-Monopile_AeroDyn15.dat

@@ -37,6 +37,8 @@ default                IndToler    - Convergence tolerance for BEMT nonlinear so
 ======  Beddoes-Leishman Unsteady Airfoil Aerodynamics Options  ===================================== [used only when AFAeroMod=2]
 3                      UAMod       - Unsteady Aero Model Switch (switch) {1=Baseline model (Original), 2=Gonzalez's variant (changes in Cn,Cc,Cm), 3=Minnema/Pierce variant (changes in Cc and Cm)} [used only when AFAeroMod=2]
 True                   FLookup     - Flag to indicate whether a lookup for f' will be calculated (TRUE) or whether best-fit exponential equations will be used (FALSE); if FALSE S1-S4 must be provided in airfoil input files (flag) [used only when AFAeroMod=2]
+0.1                    UAStartRad  - Starting radius for dynamic stall (fraction of rotor radius) [used only when AFAeroMod=2]
+1.0                    UAEndRad    - Ending radius for dynamic stall (fraction of rotor radius) [used only when AFAeroMod=2]
 ======  Airfoil Information =========================================================================
 1                      AFTabMod    - Interpolation method for multiple airfoil tables {1=1D interpolation on AoA (first table only); 2=2D interpolation on AoA and Re; 3=2D interpolation on AoA and UserProp} (-)
 1                      InCol_Alfa  - The column in the airfoil tables that contains the angle of attack (-)

OpenFAST/IEA-15-240-RWT/Airfoils/IEA-15-240-RWT_AeroDyn15_Polar_00.dat

@@ -13,41 +13,7 @@ AF00_BL.txt              BL_file     ! The file name including the boundary laye
 ! ------------------------------------------------------------------------------
 3.000000                 Re          ! Reynolds number in millions
 0                        Ctrl        ! Control setting (must be 0 for current AirfoilInfo)
-True                     InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
-!........................................
-0.000000                 alpha0      ! 0-lift angle of attack, depends on airfoil.
-0.000000                 alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
-0.000000                 alpha2      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
-1.000000                 eta_e       ! Recovery factor in the range [0.85 - 0.95] used only for UAMOD=1, it is set to 1 in the code when flookup=True. (-)
-0.000000                 C_nalpha    ! Slope of the 2D normal force coefficient curve. (1/rad)
-Default                  T_f0        ! Initial value of the time constant associated with Df in the expression of Df and f. [default = 3]
-Default                  T_V0        ! Initial value of the time constant associated with the vortex lift decay process; it is used in the expression of Cvn. It depends on Re,M, and airfoil class. [default = 6]
-Default                  T_p         ! Boundary-layer,leading edge pressure gradient time constant in the expression of Dp. It should be tuned based on airfoil experimental data. [default = 1.7]
-Default                  T_VL        ! Initial value of the time constant associated with the vortex advection process; it represents the non-dimensional time in semi-chords, needed for a vortex to travel from LE to trailing edge (TE); it is used in the expression of Cvn. It depends on Re, M (weakly), and airfoil. [valid range = 6 - 13, default = 11]
-Default                  b1          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.14]
-Default                  b2          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.53]
-Default                  b5          ! Constant in the expression of K'''_q,Cm_q^nc, and k_m,q.  [from  experimental results, defaults to 5]
-Default                  A1          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.3]
-Default                  A2          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.7]
-Default                  A5          ! Constant in the expression of K'''_q,Cm_q^nc, and k_m,q. [from experimental results, defaults to 1]
-0.000000                 S1          ! Constant in the f curve best-fit for alpha0<=AOA<=alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S2          ! Constant in the f curve best-fit for         AOA> alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S3          ! Constant in the f curve best-fit for alpha2<=AOA< alpha0; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S4          ! Constant in the f curve best-fit for         AOA< alpha2; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 Cn1         ! Critical value of C0n at leading edge separation. It should be extracted from airfoil data at a given Mach and Reynolds number. It can be calculated from the static value of Cn at either the break in the pitching moment or the loss of chord force at the onset of stall. It is close to the condition of maximum lift of the airfoil at low Mach numbers.
-0.000000                 Cn2         ! As Cn1 for negative AOAs.
-Default                  St_sh       ! Strouhal's shedding frequency constant.  [default = 0.19]
-0.000000                 Cd0         ! 2D drag coefficient value at 0-lift.
-0.000000                 Cm0         ! 2D pitching moment coefficient about 1/4-chord location, at 0-lift, positive if nose up. [If the aerodynamics coefficients table does not include a column for Cm, this needs to be set to 0.0]
-0.000000                 k0          ! Constant in the \hat(x)_cp curve best-fit; = (\hat(x)_AC-0.25).  [ignored if UAMod<>1]
-0.000000                 k1          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k2          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k3          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k1_hat      ! Constant in the expression of Cc due to leading edge vortex effects.  [ignored if UAMod<>1]
-Default                  x_cp_bar    ! Constant in the expression of \hat(x)_cp^v. [ignored if UAMod<>1, default = 0.2]
-Default                  UACutout    ! Angle of attack above which unsteady aerodynamics are disabled (deg). [Specifying the string "Default" sets UACutout to 45 degrees]
-Default                  filtCutOff  ! Cut-off frequency (-3 dB corner frequency) for low-pass filtering the AoA input to UA, as well as the 1st and 2nd derivatives (Hz) [default = 20]
-!........................................
+False                    InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
 ! Table of aerodynamics coefficients
 200                      NumAlf      ! Number of data lines in the following table
 !    Alpha      Cl      Cd        Cm

OpenFAST/IEA-15-240-RWT/Airfoils/IEA-15-240-RWT_AeroDyn15_Polar_01.dat

@@ -13,41 +13,7 @@ AF01_BL.txt              BL_file     ! The file name including the boundary laye
 ! ------------------------------------------------------------------------------
 3.000000                 Re          ! Reynolds number in millions
 0                        Ctrl        ! Control setting (must be 0 for current AirfoilInfo)
-True                     InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
-!........................................
-0.000000                 alpha0      ! 0-lift angle of attack, depends on airfoil.
-0.000000                 alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
-0.000000                 alpha2      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
-1.000000                 eta_e       ! Recovery factor in the range [0.85 - 0.95] used only for UAMOD=1, it is set to 1 in the code when flookup=True. (-)
-0.000000                 C_nalpha    ! Slope of the 2D normal force coefficient curve. (1/rad)
-Default                  T_f0        ! Initial value of the time constant associated with Df in the expression of Df and f. [default = 3]
-Default                  T_V0        ! Initial value of the time constant associated with the vortex lift decay process; it is used in the expression of Cvn. It depends on Re,M, and airfoil class. [default = 6]
-Default                  T_p         ! Boundary-layer,leading edge pressure gradient time constant in the expression of Dp. It should be tuned based on airfoil experimental data. [default = 1.7]
-Default                  T_VL        ! Initial value of the time constant associated with the vortex advection process; it represents the non-dimensional time in semi-chords, needed for a vortex to travel from LE to trailing edge (TE); it is used in the expression of Cvn. It depends on Re, M (weakly), and airfoil. [valid range = 6 - 13, default = 11]
-Default                  b1          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.14]
-Default                  b2          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.53]
-Default                  b5          ! Constant in the expression of K'''_q,Cm_q^nc, and k_m,q.  [from  experimental results, defaults to 5]
-Default                  A1          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.3]
-Default                  A2          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.7]
-Default                  A5          ! Constant in the expression of K'''_q,Cm_q^nc, and k_m,q. [from experimental results, defaults to 1]
-0.000000                 S1          ! Constant in the f curve best-fit for alpha0<=AOA<=alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S2          ! Constant in the f curve best-fit for         AOA> alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S3          ! Constant in the f curve best-fit for alpha2<=AOA< alpha0; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S4          ! Constant in the f curve best-fit for         AOA< alpha2; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 Cn1         ! Critical value of C0n at leading edge separation. It should be extracted from airfoil data at a given Mach and Reynolds number. It can be calculated from the static value of Cn at either the break in the pitching moment or the loss of chord force at the onset of stall. It is close to the condition of maximum lift of the airfoil at low Mach numbers.
-0.000000                 Cn2         ! As Cn1 for negative AOAs.
-Default                  St_sh       ! Strouhal's shedding frequency constant.  [default = 0.19]
-0.000000                 Cd0         ! 2D drag coefficient value at 0-lift.
-0.000000                 Cm0         ! 2D pitching moment coefficient about 1/4-chord location, at 0-lift, positive if nose up. [If the aerodynamics coefficients table does not include a column for Cm, this needs to be set to 0.0]
-0.000000                 k0          ! Constant in the \hat(x)_cp curve best-fit; = (\hat(x)_AC-0.25).  [ignored if UAMod<>1]
-0.000000                 k1          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k2          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k3          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k1_hat      ! Constant in the expression of Cc due to leading edge vortex effects.  [ignored if UAMod<>1]
-Default                  x_cp_bar    ! Constant in the expression of \hat(x)_cp^v. [ignored if UAMod<>1, default = 0.2]
-Default                  UACutout    ! Angle of attack above which unsteady aerodynamics are disabled (deg). [Specifying the string "Default" sets UACutout to 45 degrees]
-Default                  filtCutOff  ! Cut-off frequency (-3 dB corner frequency) for low-pass filtering the AoA input to UA, as well as the 1st and 2nd derivatives (Hz) [default = 20]
-!........................................
+False                    InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
 ! Table of aerodynamics coefficients
 200                      NumAlf      ! Number of data lines in the following table
 !    Alpha      Cl      Cd        Cm

OpenFAST/IEA-15-240-RWT/Airfoils/IEA-15-240-RWT_AeroDyn15_Polar_02.dat

@@ -13,41 +13,7 @@ AF02_BL.txt              BL_file     ! The file name including the boundary laye
 ! ------------------------------------------------------------------------------
 3.000000                 Re          ! Reynolds number in millions
 0                        Ctrl        ! Control setting (must be 0 for current AirfoilInfo)
-True                     InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
-!........................................
--3.932905                alpha0      ! 0-lift angle of attack, depends on airfoil.
-19.926944                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
--20.460148               alpha2      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
-1.000000                 eta_e       ! Recovery factor in the range [0.85 - 0.95] used only for UAMOD=1, it is set to 1 in the code when flookup=True. (-)
-0.836557                 C_nalpha    ! Slope of the 2D normal force coefficient curve. (1/rad)
-Default                  T_f0        ! Initial value of the time constant associated with Df in the expression of Df and f. [default = 3]
-Default                  T_V0        ! Initial value of the time constant associated with the vortex lift decay process; it is used in the expression of Cvn. It depends on Re,M, and airfoil class. [default = 6]
-Default                  T_p         ! Boundary-layer,leading edge pressure gradient time constant in the expression of Dp. It should be tuned based on airfoil experimental data. [default = 1.7]
-Default                  T_VL        ! Initial value of the time constant associated with the vortex advection process; it represents the non-dimensional time in semi-chords, needed for a vortex to travel from LE to trailing edge (TE); it is used in the expression of Cvn. It depends on Re, M (weakly), and airfoil. [valid range = 6 - 13, default = 11]
-Default                  b1          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.14]
-Default                  b2          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.53]
-Default                  b5          ! Constant in the expression of K'''_q,Cm_q^nc, and k_m,q.  [from  experimental results, defaults to 5]
-Default                  A1          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.3]
-Default                  A2          ! Constant in the expression of phi_alpha^c and phi_q^c.  This value is relatively insensitive for thin airfoils, but may be different for turbine airfoils. [from experimental results, defaults to 0.7]
-Default                  A5          ! Constant in the expression of K'''_q,Cm_q^nc, and k_m,q. [from experimental results, defaults to 1]
-0.000000                 S1          ! Constant in the f curve best-fit for alpha0<=AOA<=alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S2          ! Constant in the f curve best-fit for         AOA> alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S3          ! Constant in the f curve best-fit for alpha2<=AOA< alpha0; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.000000                 S4          ! Constant in the f curve best-fit for         AOA< alpha2; by definition it depends on the airfoil. [ignored if UAMod<>1]
-0.295202                 Cn1         ! Critical value of C0n at leading edge separation. It should be extracted from airfoil data at a given Mach and Reynolds number. It can be calculated from the static value of Cn at either the break in the pitching moment or the loss of chord force at the onset of stall. It is close to the condition of maximum lift of the airfoil at low Mach numbers.
--0.437208                Cn2         ! As Cn1 for negative AOAs.
-Default                  St_sh       ! Strouhal's shedding frequency constant.  [default = 0.19]
-0.333782                 Cd0         ! 2D drag coefficient value at 0-lift.
--0.008236                Cm0         ! 2D pitching moment coefficient about 1/4-chord location, at 0-lift, positive if nose up. [If the aerodynamics coefficients table does not include a column for Cm, this needs to be set to 0.0]
-0.000000                 k0          ! Constant in the \hat(x)_cp curve best-fit; = (\hat(x)_AC-0.25).  [ignored if UAMod<>1]
-0.000000                 k1          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k2          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k3          ! Constant in the \hat(x)_cp curve best-fit.  [ignored if UAMod<>1]
-0.000000                 k1_hat      ! Constant in the expression of Cc due to leading edge vortex effects.  [ignored if UAMod<>1]
-Default                  x_cp_bar    ! Constant in the expression of \hat(x)_cp^v. [ignored if UAMod<>1, default = 0.2]
-Default                  UACutout    ! Angle of attack above which unsteady aerodynamics are disabled (deg). [Specifying the string "Default" sets UACutout to 45 degrees]
-Default                  filtCutOff  ! Cut-off frequency (-3 dB corner frequency) for low-pass filtering the AoA input to UA, as well as the 1st and 2nd derivatives (Hz) [default = 20]
-!........................................
+False                    InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
 ! Table of aerodynamics coefficients
 200                      NumAlf      ! Number of data lines in the following table
 !    Alpha      Cl      Cd        Cm