XFOIL Version 6.94 Calculated polar for: WBL FX 71-089A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0286 0.03872 0.02257 0.0020 0.9989 1.0011 -2.750 -0.0031 0.03860 0.02207 0.0013 0.9989 1.0011 -2.500 0.0230 0.03862 0.02173 0.0007 0.9989 1.0011 -2.250 0.0490 0.03874 0.02153 0.0002 0.9989 1.0011 -2.000 0.0747 0.03894 0.02148 -0.0001 0.9989 1.0011 -1.750 0.1001 0.03921 0.02155 -0.0003 0.9989 1.0011 -1.500 0.1249 0.03956 0.02175 -0.0004 0.9989 1.0011 -1.250 0.1493 0.03998 0.02208 -0.0005 0.9989 1.0011 -1.000 0.1732 0.04047 0.02256 -0.0005 0.9989 1.0011 -0.750 0.1966 0.04105 0.02319 -0.0005 0.9989 1.0011 -0.500 0.2195 0.04173 0.02398 -0.0005 0.9989 1.0011 -0.250 0.2417 0.04255 0.02499 -0.0006 0.9989 1.0011 0.000 0.2628 0.04358 0.02632 -0.0008 0.9989 1.0011 0.250 0.2812 0.04507 0.02823 -0.0013 0.9989 1.0011 0.500 0.2901 0.04807 0.03176 -0.0028 0.9989 1.0011 0.750 0.2718 0.05508 0.03902 -0.0071 0.9989 1.0011 1.000 0.2603 0.06124 0.04517 -0.0115 0.9989 1.0011 1.250 0.5122 0.06657 0.05070 -0.0593 0.7916 1.0011 1.500 0.5719 0.06925 0.05332 -0.0643 0.7372 1.0011 1.750 0.5896 0.07322 0.05724 -0.0658 0.7173 1.0011 2.000 0.6087 0.07722 0.06120 -0.0674 0.6996 1.0011 2.250 0.6044 0.08166 0.06557 -0.0671 0.6910 1.0011 2.500 0.6179 0.08576 0.06960 -0.0680 0.6766 1.0011 2.750 0.6158 0.09019 0.07397 -0.0680 0.6715 1.0011 3.000 0.6101 0.09468 0.07838 -0.0679 0.6723 1.0011 3.250 0.6072 0.09928 0.08291 -0.0683 0.6759 1.0011 3.500 0.6090 0.10388 0.08744 -0.0693 0.6787 1.0011 3.750 0.6077 0.10834 0.09185 -0.0699 0.6830 1.0011 4.000 0.5959 0.11245 0.09587 -0.0697 0.6931 1.0011 4.250 0.6022 0.11705 0.10043 -0.0713 0.6984 1.0011 4.500 0.5919 0.12102 0.10434 -0.0715 0.7155 1.0011 4.750 0.5900 0.12558 0.10885 -0.0732 0.7378 1.0011 5.000 0.5646 0.12838 0.11159 -0.0721 0.7801 1.0011