# Aerodyn dynamic stall model

Good afternoon,
I have never used the dynamic stall model of Aerodyn and I am also still not so used with unsteady airfoils aerodynamics.
In the airfoils data file if InclUAdata is set to true, there are 32 coefficients to be given (example is shown below)

True InclUAdata ! Is unsteady aerodynamics data included in this table? If TRUE, then include 30 UA coefficients below this line
!..
-4.2 alpha0 ! 0-lift angle of attack, depends on airfoil.
8 alpha1 ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg)
-8 alpha2 ! Angle of attack at f=0.7, (approximately the stall angle) for AOA<alpha0. (deg)
1 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. (-)
6.2047 C_nalpha ! Slope of the 2D normal force coefficient curve. (1/rad)
3 T_f0 ! Initial value of the time constant associated with Df in the expression of Df and f’’. [default = 3]
6 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]
1.7 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]
11 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]
0.14 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]
0.53 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]
5 b5 ! Constant in the expression of K’’’_q,Cm_q^nc, and k_m,q. [from experimental results, defaults to 5]
0.3 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]
0.7 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]
1 A5 ! Constant in the expression of K’’’_q,Cm_q^nc, and k_m,q. [from experimental results, defaults to 1]
0 S1 ! Constant in the f curve best-fit for alpha0<=AOA<=alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
0 S2 ! Constant in the f curve best-fit for AOA> alpha1; by definition it depends on the airfoil. [ignored if UAMod<>1]
0 S3 ! Constant in the f curve best-fit for alpha2<=AOA< alpha0; by definition it depends on the airfoil. [ignored if UAMod<>1]
0 S4 ! Constant in the f curve best-fit for AOA< alpha2; by definition it depends on the airfoil. [ignored if UAMod<>1]
1.4144 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.
-0.5324 Cn2 ! As Cn1 for negative AOAs.
0.19 St_sh ! Strouhal’s shedding frequency constant. [default = 0.19]
0.006 Cd0 ! 2D drag coefficient value at 0-lift.
-0.121 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 k0 ! Constant in the \hat(x)_cp curve best-fit; = (\hat(x)_AC-0.25). [ignored if UAMod<>1]
0 k1 ! Constant in the \hat(x)_cp curve best-fit. [ignored if UAMod<>1]
0 k2 ! Constant in the \hat(x)_cp curve best-fit. [ignored if UAMod<>1]
0 k3 ! Constant in the \hat(x)_cp curve best-fit. [ignored if UAMod<>1]
0 k1_hat ! Constant in the expression of Cc due to leading edge vortex effects. [ignored if UAMod<>1]
0.2 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]
10 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]

So my question will be very basic as a beginner. I am wondering how the user determines these coefficients for a given blade. It seems that some of them can be calculated directly from the static polars. Perhaps some other have to be tuned properly by an experimented user.
Many thanks for any tips and any help.
Best regards.
Florence Haudin.

Hi Florence,

Setting up the unsteady aerodynamics parameters can be a bit challenging. It’s important to note that from the 32 parameters requested, only 8 parameters must be calculated by the user. The other parameters can be left as “default” or 0. The 8 parameters of interest are: alpha0, alpha1, alpha2, C_nalpha, Cn1, Cn2, Cd0, and Cm0.

These 8 parameters are calculated based on the airfoil static polars.

I have modified one MATLAB function originally written by Srinivas Guntur. The new MATLAB file uses a neutral input format (the original one was using AeroDyn 14 files as input).

Attached you can find the MATLAB function (AeroDyn_UA_coefficients.m) and one example (Example_airfoil.xlsx) from OC6 phase III project. Note that the main MATLAB function has one dependency (intersections.m). Once downloaded and stored, you can simply run the MATLAB function with the example provided as:

blade_polars = xlsread('Example_airfoil.xlsx'); UA_coefficients = AeroDyn_UA_coefficients(blade_polars);

The 8 coefficients are stored in the function output (UA_coefficients) and also printed on the screen automatically:

I hope that helps!

Roger