XFOIL Version 6.94 Calculated polar for: GOE 100 (SOPWITH) 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 0.3022 0.00832 0.00203 -0.0554 0.7936 0.2136 0.500 0.3536 0.00802 0.00190 -0.0542 0.7297 0.3325 1.000 0.3933 0.00674 0.00207 -0.0497 0.6276 0.9648 1.500 0.4765 0.00768 0.00225 -0.0556 0.4886 0.9984 2.000 0.5279 0.00822 0.00242 -0.0548 0.4117 1.0000 2.500 0.5740 0.00879 0.00259 -0.0528 0.3422 1.0000 3.000 0.6212 0.00932 0.00284 -0.0510 0.2937 1.0000 3.500 0.6693 0.00987 0.00314 -0.0494 0.2523 1.0000 4.000 0.7187 0.01037 0.00348 -0.0482 0.2239 1.0000 4.500 0.7687 0.01091 0.00391 -0.0470 0.2015 1.0000 5.000 0.8190 0.01148 0.00437 -0.0459 0.1847 1.0000 5.500 0.8693 0.01207 0.00492 -0.0448 0.1730 1.0000 6.000 0.9193 0.01271 0.00553 -0.0438 0.1620 1.0000 6.500 0.9698 0.01330 0.00614 -0.0428 0.1550 1.0000 7.000 1.0194 0.01398 0.00679 -0.0418 0.1448 1.0000 7.500 1.0689 0.01467 0.00757 -0.0407 0.1398 1.0000 8.000 1.1186 0.01527 0.00823 -0.0397 0.1323 1.0000 8.500 1.1689 0.01574 0.00879 -0.0388 0.1232 1.0000 9.000 1.2176 0.01635 0.00948 -0.0378 0.1157 1.0000 9.500 1.2677 0.01677 0.00999 -0.0371 0.1010 1.0000 10.000 1.2932 0.01998 0.01263 -0.0332 0.0137 1.0000 10.500 1.3248 0.02231 0.01519 -0.0297 0.0090 1.0000 11.000 1.3519 0.02464 0.01779 -0.0259 0.0077 1.0000 11.500 1.3659 0.02728 0.02068 -0.0204 0.0069 1.0000 12.000 1.3598 0.03108 0.02475 -0.0135 0.0066 1.0000 12.500 1.3597 0.03503 0.02898 -0.0093 0.0063 1.0000 13.000 1.3520 0.04043 0.03465 -0.0070 0.0063 1.0000 13.500 1.3350 0.04819 0.04272 -0.0075 0.0062 1.0000 14.000 1.3135 0.05854 0.05338 -0.0116 0.0061 1.0000 14.500 1.2871 0.07097 0.06609 -0.0177 0.0062 1.0000 15.000 1.2576 0.08385 0.07921 -0.0235 0.0062 1.0000 15.500 1.2325 0.09578 0.09132 -0.0285 0.0063 1.0000 16.000 1.2112 0.10704 0.10277 -0.0331 0.0062 1.0000 16.500 1.1984 0.11695 0.11283 -0.0371 0.0062 1.0000 17.000 1.1897 0.12611 0.12213 -0.0407 0.0063 1.0000 17.500 1.1819 0.13556 0.13176 -0.0448 0.0063 1.0000 18.500 1.1661 0.15530 0.15186 -0.0544 0.0064 1.0000 19.500 1.1370 0.18193 0.17897 -0.0706 0.0062 1.0000