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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0901 0.02894 0.01929 -0.0049 0.9989 0.8133 -2.750 -0.0195 0.02706 0.01697 -0.0166 0.9989 1.0011 -2.500 0.0110 0.02721 0.01674 -0.0177 0.9989 1.0011 -2.250 0.0389 0.02743 0.01678 -0.0180 0.9989 1.0011 -2.000 0.0659 0.02772 0.01698 -0.0180 0.9989 1.0011 -1.750 0.0921 0.02808 0.01732 -0.0179 0.9989 1.0011 -1.500 0.1180 0.02850 0.01780 -0.0177 0.9989 1.0011 -1.250 0.1436 0.02899 0.01841 -0.0175 0.9989 1.0011 -1.000 0.1691 0.02956 0.01920 -0.0174 0.9989 1.0011 -0.750 0.1948 0.03028 0.02032 -0.0174 0.9989 1.0011 -0.500 0.3045 0.03096 0.01948 -0.0258 0.5517 1.0011 -0.250 0.3311 0.03270 0.02039 -0.0244 0.4870 1.0011 0.000 0.3611 0.03424 0.02153 -0.0241 0.4470 1.0011 0.250 0.3921 0.03577 0.02282 -0.0241 0.4176 1.0011 0.500 0.4236 0.03736 0.02427 -0.0243 0.3949 1.0011 0.750 0.4548 0.03910 0.02582 -0.0245 0.3768 1.0011 1.000 0.4878 0.04082 0.02776 -0.0253 0.3617 1.0011 1.250 0.5185 0.04293 0.02981 -0.0256 0.3502 1.0011 1.500 0.5514 0.04510 0.03233 -0.0268 0.3403 1.0011 1.750 0.5822 0.04744 0.03472 -0.0274 0.3313 1.0011 2.000 0.6125 0.05038 0.03773 -0.0283 0.3252 1.0011 2.250 0.6440 0.05349 0.04133 -0.0304 0.3201 1.0011 2.500 0.6740 0.05709 0.04529 -0.0326 0.3171 1.0011 2.750 0.7019 0.06110 0.04965 -0.0349 0.3146 1.0011 3.000 0.7272 0.06572 0.05463 -0.0377 0.3142 1.0011 3.250 0.7490 0.07113 0.06038 -0.0412 0.3166 1.0011 3.500 0.7670 0.07692 0.06642 -0.0444 0.3204 1.0011 3.750 0.7827 0.08324 0.07298 -0.0483 0.3296 1.0011 4.000 0.7911 0.09215 0.08210 -0.0542 0.3514 1.0011 4.250 0.2944 0.10407 0.09550 -0.0528 0.7456 1.0011 4.750 0.3202 0.11036 0.10171 -0.0537 0.7108 1.0011 5.000 0.3275 0.11299 0.10430 -0.0533 0.6926 1.0011