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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0595 0.05766 0.02961 -0.0119 0.9989 1.0011 -2.750 -0.0354 0.05755 0.02916 -0.0123 0.9989 1.0011 -2.500 -0.0114 0.05753 0.02889 -0.0126 0.9989 1.0011 -2.250 0.0126 0.05760 0.02877 -0.0127 0.9989 1.0011 -2.000 0.0364 0.05774 0.02881 -0.0128 0.9989 1.0011 -1.750 0.0598 0.05795 0.02898 -0.0128 0.9989 1.0011 -1.500 0.0830 0.05825 0.02930 -0.0127 0.9989 1.0011 -1.250 0.1057 0.05862 0.02975 -0.0126 0.9989 1.0011 -1.000 0.1279 0.05908 0.03036 -0.0125 0.9989 1.0011 -0.750 0.1496 0.05964 0.03113 -0.0123 0.9989 1.0011 -0.500 0.1706 0.06031 0.03207 -0.0122 0.9989 1.0011 -0.250 0.1908 0.06113 0.03322 -0.0121 0.9989 1.0011 0.000 0.2100 0.06213 0.03460 -0.0121 0.9989 1.0011 0.250 0.2275 0.06339 0.03630 -0.0123 0.9989 1.0011 0.500 0.2424 0.06502 0.03841 -0.0126 0.9989 1.0011 0.750 0.2532 0.06726 0.04109 -0.0133 0.9989 1.0011 1.000 0.2586 0.07036 0.04448 -0.0145 0.9989 1.0011 1.250 0.2585 0.07425 0.04849 -0.0163 0.9989 1.0011 1.500 0.2570 0.07847 0.05267 -0.0186 0.9989 1.0011 1.750 0.2571 0.08257 0.05668 -0.0209 0.9989 1.0011 2.000 0.2594 0.08645 0.06046 -0.0232 0.9989 1.0011 2.250 0.2635 0.09013 0.06404 -0.0254 0.9989 1.0011 2.500 0.2690 0.09366 0.06748 -0.0275 0.9989 1.0011 2.750 0.2756 0.09709 0.07082 -0.0295 0.9989 1.0011 3.000 0.2832 0.10048 0.07413 -0.0314 0.9989 1.0011 3.250 0.2915 0.10382 0.07738 -0.0334 0.9989 1.0011 3.500 0.3003 0.10713 0.08060 -0.0352 0.9989 1.0011 3.750 0.3095 0.11040 0.08379 -0.0371 0.9989 1.0011 4.000 0.3191 0.11364 0.08695 -0.0389 0.9989 1.0011 4.250 0.3290 0.11687 0.09011 -0.0407 0.9989 1.0011 4.500 0.3392 0.12009 0.09325 -0.0425 0.9989 1.0011 4.750 0.3496 0.12331 0.09639 -0.0443 0.9989 1.0011 5.000 0.3603 0.12651 0.09954 -0.0461 0.9989 1.0011