XFOIL Version 6.94 Calculated polar for: GOE 118 (MVA MK.7) 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 0.3625 0.01228 0.00390 -0.0784 0.6168 0.1616 0.500 0.4184 0.01208 0.00370 -0.0782 0.6009 0.1683 1.000 0.4745 0.01189 0.00357 -0.0781 0.5862 0.1778 1.500 0.5303 0.01166 0.00349 -0.0780 0.5698 0.1987 2.000 0.5790 0.01016 0.00364 -0.0771 0.5555 0.6910 2.500 0.6443 0.00979 0.00373 -0.0781 0.5405 1.0000 3.000 0.6986 0.01000 0.00382 -0.0775 0.5202 1.0000 3.500 0.7535 0.01032 0.00408 -0.0772 0.5059 1.0000 4.000 0.8075 0.01059 0.00429 -0.0766 0.4861 1.0000 4.500 0.8616 0.01091 0.00460 -0.0761 0.4685 1.0000 5.000 0.9149 0.01127 0.00492 -0.0755 0.4486 1.0000 5.500 0.9677 0.01163 0.00528 -0.0748 0.4263 1.0000 6.000 1.0197 0.01203 0.00570 -0.0741 0.4020 1.0000 6.500 1.0701 0.01251 0.00613 -0.0731 0.3708 1.0000 7.000 1.1192 0.01314 0.00671 -0.0720 0.3363 1.0000 7.500 1.1647 0.01404 0.00749 -0.0704 0.2882 1.0000 8.000 1.1805 0.01754 0.00991 -0.0655 0.1058 1.0000 8.500 1.1982 0.02058 0.01256 -0.0607 0.0345 1.0000 9.000 1.2236 0.02250 0.01451 -0.0567 0.0269 1.0000 9.500 1.2402 0.02482 0.01701 -0.0521 0.0243 1.0000 10.000 1.2468 0.02830 0.02064 -0.0482 0.0227 1.0000 10.500 1.2533 0.03247 0.02499 -0.0458 0.0211 1.0000 11.000 1.2492 0.03815 0.03078 -0.0442 0.0201 1.0000 11.500 1.2505 0.04348 0.03630 -0.0428 0.0193 1.0000 12.000 1.2551 0.04846 0.04143 -0.0413 0.0183 1.0000 12.500 1.2607 0.05312 0.04611 -0.0395 0.0172 1.0000 13.000 1.2727 0.05709 0.05020 -0.0369 0.0164 1.0000 13.500 1.2852 0.06139 0.05475 -0.0351 0.0155 1.0000 14.000 1.3005 0.06544 0.05899 -0.0329 0.0149 1.0000 14.500 1.3185 0.06947 0.06318 -0.0306 0.0145 1.0000 15.000 1.3314 0.07449 0.06843 -0.0287 0.0142 1.0000 15.500 1.3357 0.08128 0.07555 -0.0272 0.0140 1.0000 16.000 1.2807 0.09563 0.09080 -0.0327 0.0150 1.0000 16.500 1.2603 0.10635 0.10192 -0.0357 0.0161 1.0000 17.000 1.2276 0.11989 0.11583 -0.0409 0.0165 1.0000