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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0527 0.02697 0.01756 -0.0172 0.9989 0.6782 -2.750 -0.0193 0.02433 0.01597 -0.0162 0.9989 1.0011 -2.500 0.0152 0.02449 0.01559 -0.0185 0.9989 1.0011 -2.250 0.0439 0.02474 0.01564 -0.0189 0.9989 1.0011 -2.000 0.0713 0.02504 0.01587 -0.0189 0.9989 1.0011 -1.750 0.0981 0.02541 0.01623 -0.0189 0.9989 1.0011 -1.500 0.1244 0.02585 0.01673 -0.0187 0.9989 1.0011 -1.250 0.1507 0.02634 0.01738 -0.0185 0.9989 1.0011 -1.000 0.1774 0.02690 0.01824 -0.0184 0.9989 1.0011 -0.750 0.2645 0.02746 0.01678 -0.0232 0.4905 1.0011 -0.500 0.2917 0.02900 0.01747 -0.0224 0.4233 1.0011 -0.250 0.3214 0.03024 0.01829 -0.0222 0.3857 1.0011 0.000 0.3521 0.03146 0.01923 -0.0222 0.3591 1.0011 0.250 0.3839 0.03272 0.02036 -0.0225 0.3390 1.0011 0.500 0.4150 0.03420 0.02165 -0.0226 0.3235 1.0011 0.750 0.4466 0.03557 0.02309 -0.0228 0.3099 1.0011 1.000 0.4776 0.03747 0.02485 -0.0231 0.2997 1.0011 1.250 0.5097 0.03907 0.02671 -0.0237 0.2901 1.0011 1.500 0.5407 0.04099 0.02866 -0.0241 0.2824 1.0011 1.750 0.5706 0.04350 0.03114 -0.0245 0.2756 1.0011 2.000 0.6026 0.04571 0.03378 -0.0255 0.2705 1.0011 2.250 0.6336 0.04831 0.03672 -0.0265 0.2659 1.0011 2.500 0.6631 0.05105 0.03970 -0.0274 0.2608 1.0011 2.750 0.6914 0.05405 0.04283 -0.0281 0.2565 1.0011 3.000 0.7173 0.05790 0.04663 -0.0286 0.2517 1.0011 3.250 0.7442 0.06184 0.05106 -0.0303 0.2507 1.0011 3.500 0.7692 0.06652 0.05604 -0.0319 0.2514 1.0011 3.750 0.7873 0.07499 0.06561 -0.0392 0.2699 1.0011 4.000 0.8101 0.08066 0.07122 -0.0397 0.2760 1.0011 4.250 0.3781 0.10413 0.09667 -0.0555 0.5993 1.0011 4.500 0.3807 0.10687 0.09936 -0.0542 0.5782 1.0011 4.750 0.4049 0.11151 0.10395 -0.0551 0.5613 1.0011 5.000 0.4377 0.11633 0.10875 -0.0561 0.5388 1.0011