XFOIL Version 6.94 Calculated polar for: GOE 365 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5628 0.01043 0.00325 -0.0994 0.6784 0.0314 0.500 0.6132 0.01025 0.00300 -0.0981 0.6618 0.0348 1.000 0.6628 0.01007 0.00285 -0.0966 0.6449 0.0413 1.500 0.7116 0.00990 0.00274 -0.0951 0.6246 0.0600 2.000 0.8375 0.00803 0.00294 -0.1109 0.5854 1.0000 2.500 0.8745 0.00834 0.00301 -0.1070 0.5487 1.0000 3.000 0.9114 0.00875 0.00318 -0.1031 0.5115 1.0000 3.500 0.9477 0.00922 0.00343 -0.0992 0.4757 1.0000 4.000 0.9830 0.00975 0.00375 -0.0952 0.4407 1.0000 5.000 1.0309 0.01156 0.00478 -0.0831 0.3073 1.0000 5.500 1.0357 0.01303 0.00564 -0.0738 0.2165 1.0000 6.000 1.0623 0.01383 0.00627 -0.0686 0.1912 1.0000 6.500 1.0851 0.01491 0.00708 -0.0631 0.1556 1.0000 7.000 1.1160 0.01572 0.00783 -0.0592 0.1371 1.0000 7.500 1.1217 0.01790 0.00955 -0.0519 0.0694 1.0000 8.000 1.1305 0.02024 0.01164 -0.0458 0.0034 1.0000 8.500 1.1609 0.02144 0.01292 -0.0429 0.0031 1.0000 9.000 1.1902 0.02281 0.01438 -0.0400 0.0031 1.0000 9.500 1.2176 0.02438 0.01606 -0.0373 0.0032 1.0000 10.000 1.2428 0.02621 0.01805 -0.0346 0.0034 1.0000 10.500 1.2657 0.02832 0.02032 -0.0320 0.0036 1.0000 11.000 1.2885 0.03056 0.02270 -0.0297 0.0041 1.0000 11.500 1.3064 0.03331 0.02564 -0.0274 0.0044 1.0000 12.000 1.3184 0.03671 0.02926 -0.0250 0.0048 1.0000 12.500 1.3339 0.03994 0.03266 -0.0232 0.0053 1.0000 13.000 1.3390 0.04435 0.03731 -0.0215 0.0059 1.0000 13.500 1.3312 0.05044 0.04363 -0.0201 0.0062 1.0000 14.000 1.3390 0.05519 0.04859 -0.0195 0.0069 1.0000 14.500 1.3266 0.06274 0.05640 -0.0197 0.0075 1.0000 15.000 1.3007 0.07268 0.06661 -0.0211 0.0079 1.0000 15.500 1.2961 0.08009 0.07426 -0.0223 0.0086 1.0000 16.000 1.2737 0.09031 0.08473 -0.0248 0.0090 1.0000 16.500 1.2474 0.10123 0.09586 -0.0276 0.0096 1.0000