XFOIL Version 6.94 Calculated polar for: MH 116 9.84% 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 0.5061 0.00655 0.00230 -0.1150 0.7556 0.9306 0.500 0.5734 0.00660 0.00225 -0.1174 0.7406 1.0000 1.000 0.6282 0.00668 0.00225 -0.1171 0.7257 1.0000 1.500 0.6832 0.00689 0.00228 -0.1169 0.7055 1.0000 2.000 0.7380 0.00702 0.00240 -0.1167 0.6916 1.0000 2.500 0.7922 0.00716 0.00248 -0.1162 0.6719 1.0000 3.000 0.8461 0.00740 0.00266 -0.1158 0.6503 1.0000 3.500 0.8994 0.00757 0.00281 -0.1152 0.6273 1.0000 4.000 0.9517 0.00786 0.00307 -0.1145 0.6004 1.0000 4.500 1.0023 0.00820 0.00332 -0.1134 0.5659 1.0000 5.000 1.0526 0.00858 0.00369 -0.1123 0.5267 1.0000 5.500 1.0986 0.00921 0.00413 -0.1104 0.4707 1.0000 6.000 1.1413 0.01008 0.00473 -0.1081 0.3978 1.0000 6.500 1.1801 0.01125 0.00557 -0.1054 0.3205 1.0000 7.000 1.2157 0.01268 0.00658 -0.1022 0.2374 1.0000 8.000 1.2844 0.01557 0.00892 -0.0958 0.1188 1.0000 8.500 1.3131 0.01712 0.01026 -0.0916 0.0799 1.0000 9.000 1.3365 0.01887 0.01199 -0.0868 0.0435 1.0000 9.500 1.3483 0.02147 0.01445 -0.0805 0.0163 1.0000 10.000 1.3593 0.02428 0.01743 -0.0746 0.0090 1.0000 11.000 1.3622 0.03221 0.02607 -0.0632 0.0062 1.0000 11.500 1.3651 0.03668 0.03088 -0.0589 0.0056 1.0000 12.000 1.3649 0.04202 0.03660 -0.0551 0.0054 1.0000 12.500 1.3615 0.04821 0.04319 -0.0520 0.0051 1.0000 13.000 1.3524 0.05532 0.05074 -0.0499 0.0051 1.0000 13.500 1.3358 0.06335 0.05919 -0.0492 0.0048 1.0000 14.000 1.3136 0.07316 0.06941 -0.0503 0.0047 1.0000 14.500 1.2863 0.08482 0.08149 -0.0538 0.0047 1.0000 15.000 1.2572 0.09792 0.09494 -0.0598 0.0046 1.0000 15.500 1.2229 0.11340 0.11078 -0.0685 0.0047 1.0000 16.000 1.1873 0.13078 0.12848 -0.0797 0.0048 1.0000 16.500 1.1422 0.15273 0.15072 -0.0944 0.0052 1.0000 17.000 1.0900 0.18269 0.18084 -0.1108 0.0058 1.0000 17.500 1.0594 0.20937 0.20747 -0.1220 0.0064 1.0000