Airfoil polars: the same data, parameterised two ways-505101500.511.5Re = 2e+05Re = 5e+05Re = 1e+06angle of attack (degrees)lift coefficient ClCl and Cd against alpha (black dot = stall)Re = 2e+05Re = 5e+05Re = 1e+0600.020.040.060.0800.511.5Re = 2e+05Re = 5e+05Re = 1e+06max L/D = 62at 4.9 degdrag coefficient Cdlift coefficient ClDrag polar: best glide is a tangent from the originRe = 2e+05Re = 5e+05Re = 1e+06drag coefficient Cd (dashed)00.020.040.060.08Airfoil polars: the same data, parameterised two ways