nonzero Cd0

OpenFAST/IEA-22-280-RWT/Airfoils/IEA-22-280-RWT_AeroDyn15_Polar_02.dat

@@ -16,10 +16,10 @@ AF02_BL.txt              BL_file     ! The file name including the boundary laye
 True                     InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
 !........................................
 -3.578268                alpha0      ! 0-lift angle of attack, depends on airfoil.
-15.519696                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
--14.163607               alpha2      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
+15.703500                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
+-14.335951               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. (-)
-0.787596                 C_nalpha    ! Slope of the 2D normal force coefficient curve. (1/rad)
+0.783426                 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]
@@ -37,7 +37,7 @@ Default                  A5          ! Constant in the expression of K'''_q,Cm_q
 0.411993                 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.
 -0.411993                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.
+0.332085                 Cd0         ! 2D drag coefficient value at 0-lift.
 -0.005386                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]

OpenFAST/IEA-22-280-RWT/Airfoils/IEA-22-280-RWT_AeroDyn15_Polar_03.dat

@@ -16,10 +16,10 @@ AF03_BL.txt              BL_file     ! The file name including the boundary laye
 True                     InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
 !........................................
 -3.569649                alpha0      ! 0-lift angle of attack, depends on airfoil.
-13.354014                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
--12.079828               alpha2      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
+13.635096                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
+-12.229694               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. (-)
-1.857559                 C_nalpha    ! Slope of the 2D normal force coefficient curve. (1/rad)
+1.840246                 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]
@@ -37,7 +37,7 @@ Default                  A5          ! Constant in the expression of K'''_q,Cm_q
 0.473496                 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.
 -0.561627                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.
+0.288843                 Cd0         ! 2D drag coefficient value at 0-lift.
 -0.018158                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]

OpenFAST/IEA-22-280-RWT/Airfoils/IEA-22-280-RWT_AeroDyn15_Polar_04.dat

@@ -16,10 +16,10 @@ AF04_BL.txt              BL_file     ! The file name including the boundary laye
 True                     InclUAdata  ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
 !........................................
 -3.567983                alpha0      ! 0-lift angle of attack, depends on airfoil.
-12.805715                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
--11.696569               alpha2      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
+13.181649                alpha1      ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
+-11.859292               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. (-)
-3.195582                 C_nalpha    ! Slope of the 2D normal force coefficient curve. (1/rad)
+3.153877                 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]
@@ -37,7 +37,7 @@ Default                  A5          ! Constant in the expression of K'''_q,Cm_q
 0.801422                 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.
 -0.524438                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.
+0.235318                 Cd0         ! 2D drag coefficient value at 0-lift.
 -0.033969                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]