XFOIL Version 6.94 Calculated polar for: GOA 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1530 0.02877 0.01587 -0.0328 0.9999 0.3131 -2.500 -0.1010 0.02303 0.01207 -0.0267 0.9999 1.0001 -2.250 -0.0761 0.02306 0.01151 -0.0255 0.9999 1.0001 -2.000 -0.0518 0.02310 0.01112 -0.0244 0.9999 1.0001 -1.750 -0.0275 0.02316 0.01077 -0.0234 0.9999 1.0001 -1.500 -0.0032 0.02323 0.01055 -0.0224 0.9999 1.0001 -1.250 0.0212 0.02332 0.01045 -0.0214 0.9999 1.0001 -1.000 0.0460 0.02341 0.01047 -0.0203 0.9999 1.0001 -0.750 0.0715 0.02351 0.01066 -0.0192 0.9999 1.0001 -0.500 0.0988 0.02357 0.01104 -0.0181 0.9999 1.0001 0.000 0.2328 0.02727 0.01220 -0.0268 0.3657 1.0001 0.250 0.2597 0.02866 0.01311 -0.0258 0.3458 1.0001 0.500 0.2901 0.02996 0.01414 -0.0254 0.3328 1.0001 0.750 0.3220 0.03137 0.01532 -0.0253 0.3221 1.0001 1.250 0.3862 0.03432 0.01814 -0.0256 0.2972 1.0001 1.500 0.4190 0.03599 0.01984 -0.0260 0.2896 1.0001 1.750 0.4530 0.03765 0.02177 -0.0267 0.2892 1.0001 2.000 0.4865 0.03950 0.02391 -0.0274 0.2909 1.0001 2.250 0.5194 0.04137 0.02621 -0.0283 0.2946 1.0001 2.500 0.5514 0.04342 0.02879 -0.0294 0.3009 1.0001 2.750 0.5820 0.04591 0.03159 -0.0304 0.3073 1.0001 3.000 0.6130 0.04838 0.03464 -0.0325 0.3177 1.0001 3.250 0.6424 0.05130 0.03794 -0.0342 0.3283 1.0001 3.500 0.6709 0.05460 0.04162 -0.0366 0.3424 1.0001 3.750 0.6981 0.05801 0.04564 -0.0409 0.3628 1.0001 4.000 0.7211 0.06210 0.05019 -0.0462 0.3882 1.0001 4.250 0.7382 0.06673 0.05519 -0.0531 0.4200 1.0001 4.500 0.7491 0.07200 0.06069 -0.0611 0.4589 1.0001 4.750 0.7362 0.07776 0.06654 -0.0707 0.5070 1.0001 5.000 0.7190 0.08350 0.07222 -0.0793 0.5642 1.0001