XFOIL Version 6.94 Calculated polar for: GOE 137 (MVA H.15) 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 0.2869 0.01061 0.00294 -0.0555 0.6584 0.1807 0.500 0.3416 0.01016 0.00289 -0.0552 0.6350 0.3127 1.000 0.4097 0.00841 0.00295 -0.0571 0.6121 1.0000 1.500 0.4633 0.00859 0.00291 -0.0562 0.5848 1.0000 2.000 0.5177 0.00882 0.00297 -0.0555 0.5628 1.0000 2.500 0.5720 0.00904 0.00304 -0.0549 0.5352 1.0000 3.000 0.6262 0.00931 0.00316 -0.0542 0.5072 1.0000 3.500 0.6803 0.00963 0.00335 -0.0536 0.4813 1.0000 4.000 0.7343 0.00998 0.00360 -0.0530 0.4537 1.0000 4.500 0.7882 0.01039 0.00390 -0.0524 0.4274 1.0000 5.000 0.8415 0.01083 0.00426 -0.0518 0.4029 1.0000 5.500 0.8947 0.01131 0.00469 -0.0511 0.3790 1.0000 6.000 0.9472 0.01180 0.00516 -0.0504 0.3480 1.0000 6.500 0.9992 0.01235 0.00569 -0.0497 0.3168 1.0000 7.000 1.0489 0.01309 0.00631 -0.0487 0.2724 1.0000 7.500 1.0968 0.01406 0.00710 -0.0476 0.2233 1.0000 8.000 1.1429 0.01518 0.00805 -0.0463 0.1898 1.0000 8.500 1.1882 0.01631 0.00915 -0.0448 0.1697 1.0000 9.000 1.2318 0.01752 0.01037 -0.0433 0.1498 1.0000 9.500 1.2749 0.01870 0.01160 -0.0416 0.1272 1.0000 10.000 1.3173 0.01985 0.01270 -0.0399 0.0906 1.0000 10.500 1.3425 0.02232 0.01500 -0.0363 0.0664 1.0000 11.000 1.3583 0.02496 0.01771 -0.0315 0.0547 1.0000 11.500 1.3705 0.02765 0.02050 -0.0269 0.0470 1.0000 12.000 1.3802 0.03091 0.02387 -0.0235 0.0415 1.0000 12.500 1.3888 0.03481 0.02793 -0.0214 0.0366 1.0000 13.000 1.3945 0.03953 0.03282 -0.0205 0.0333 1.0000 13.500 1.3977 0.04495 0.03846 -0.0204 0.0302 1.0000 14.000 1.3937 0.05136 0.04505 -0.0207 0.0283 1.0000 14.500 1.3902 0.05791 0.05180 -0.0215 0.0265 1.0000 15.000 1.3816 0.06520 0.05927 -0.0227 0.0249 1.0000 15.500 1.3741 0.07279 0.06708 -0.0244 0.0234 1.0000 16.000 1.3660 0.08035 0.07472 -0.0259 0.0221 1.0000 16.500 1.3535 0.08948 0.08418 -0.0289 0.0208 1.0000 17.000 1.3444 0.09812 0.09297 -0.0319 0.0196 1.0000 17.500 1.3322 0.10732 0.10238 -0.0351 0.0186 1.0000 18.000 1.3109 0.11929 0.11468 -0.0411 0.0173 1.0000 19.000 1.2919 0.13792 0.13360 -0.0494 0.0160 1.0000 19.500 1.2672 0.15186 0.14790 -0.0576 0.0157 1.0000