XFOIL Version 6.94 Calculated polar for: GOE 281 (DAIMLER XII) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3390 0.01236 0.00470 -0.0661 0.7406 0.0592 0.500 0.3930 0.01184 0.00405 -0.0652 0.7024 0.0615 1.000 0.4478 0.01150 0.00361 -0.0645 0.6702 0.0634 1.500 0.5028 0.01119 0.00327 -0.0640 0.6404 0.0670 2.000 0.5574 0.01109 0.00304 -0.0633 0.5996 0.0721 2.500 0.6022 0.00917 0.00321 -0.0614 0.5680 0.8036 3.000 0.6675 0.00919 0.00320 -0.0626 0.5249 1.0000 3.500 0.7197 0.00965 0.00330 -0.0618 0.4670 1.0000 4.000 0.7726 0.01019 0.00362 -0.0610 0.4331 1.0000 4.500 0.8250 0.01077 0.00403 -0.0603 0.4002 1.0000 5.000 0.8774 0.01135 0.00449 -0.0596 0.3666 1.0000 5.500 0.9296 0.01188 0.00492 -0.0589 0.3297 1.0000 6.000 0.9806 0.01251 0.00543 -0.0581 0.2854 1.0000 6.500 1.0295 0.01338 0.00609 -0.0571 0.2430 1.0000 7.000 1.0773 0.01440 0.00698 -0.0558 0.2133 1.0000 7.500 1.1235 0.01553 0.00799 -0.0543 0.1824 1.0000 8.000 1.1698 0.01652 0.00889 -0.0530 0.1571 1.0000 8.500 1.2152 0.01759 0.00994 -0.0514 0.1407 1.0000 9.000 1.2602 0.01860 0.01104 -0.0498 0.1285 1.0000 9.500 1.3021 0.01977 0.01221 -0.0479 0.1183 1.0000 10.000 1.3432 0.02092 0.01351 -0.0459 0.1094 1.0000 10.500 1.3830 0.02200 0.01471 -0.0438 0.0992 1.0000 11.000 1.4203 0.02315 0.01600 -0.0414 0.0850 1.0000 11.500 1.4499 0.02478 0.01750 -0.0379 0.0437 1.0000 12.000 1.4539 0.02798 0.02068 -0.0316 0.0319 1.0000 12.500 1.4559 0.03165 0.02455 -0.0267 0.0283 1.0000 13.000 1.4554 0.03606 0.02920 -0.0233 0.0255 1.0000 13.500 1.4471 0.04199 0.03536 -0.0216 0.0237 1.0000 14.000 1.4401 0.04855 0.04217 -0.0215 0.0224 1.0000 14.500 1.4193 0.05739 0.05117 -0.0228 0.0213 1.0000 15.000 1.4054 0.06591 0.06000 -0.0247 0.0202 1.0000 15.500 1.3887 0.07510 0.06941 -0.0272 0.0194 1.0000 16.000 1.3723 0.08428 0.07869 -0.0296 0.0185 1.0000 16.500 1.3586 0.09353 0.08817 -0.0324 0.0178 1.0000 17.000 1.3436 0.10361 0.09853 -0.0362 0.0169 1.0000 17.500 1.3327 0.11287 0.10798 -0.0397 0.0164 1.0000 18.000 1.3338 0.11888 0.11397 -0.0403 0.0154 1.0000 18.500 1.3113 0.13162 0.12708 -0.0472 0.0151 1.0000 19.000 1.2815 0.14703 0.14288 -0.0568 0.0142 1.0000 19.500 1.2603 0.16059 0.15675 -0.0646 0.0140 1.0000