I'm currently trying to do an initial design for a propeller. In order to do this I'm trying to use Xrotor. Xrotor allows the user to enter certain information about both the propeller geometry, flight conditions and the airfoil lift and drag data after which it performs its calculations. Xrotor requires the following values for the lift and drag data (example values are given):
======================================================================== 1) Zero-lift alpha (deg): 0.00 7) Minimum Cd : 0.0070 2) d(Cl)/d(alpha) : 6.280 8) Cl at minimum Cd : 0.150 3) d(Cl)/d(alpha)@stall : 0.100 9) d(Cd)/d(Cl**2) : 0.0040 4) Maximum Cl : 2.00 10) Reference Re number : 2000000. 5) Minimum Cl : -1.50 11) Re scaling exponent : -0.2000 6) Cl increment to stall: 0.200 12) Cm : -0.100 13) Mcrit : 0.620 ========================================================================
This information can be obtained from the drag polars of the airfoil.
The 360 degree polar was made using JBlade. The only value that Xrotor requests that I am not sure on how to calculate is the
d(Cd)/d(Cl^2) figure. I'm fairly confident that this can be obtained using the $C_d,C_l$ drag polar, but I'm not completely confident on how I should go about it. What is the correct way of determining this value?