XFOIL Version 6.96 Calculated polar for: GOE 116 (MVA MK.3) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.3772 0.09338 0.09026 -0.0348 1.0000 0.0687 -8.250 -0.3094 0.07950 0.07683 -0.0471 0.9948 0.0779 -8.000 -0.3668 0.08732 0.08429 -0.0368 1.0000 0.0731 -7.000 -0.3515 0.03644 0.03260 -0.0949 0.9620 0.0436 -6.750 -0.3238 0.02848 0.02373 -0.0996 0.9516 0.0352 -6.500 -0.2890 0.02596 0.02042 -0.1011 0.9416 0.0307 -6.250 -0.2629 0.02237 0.01635 -0.1018 0.9294 0.0292 -6.000 -0.2358 0.02021 0.01384 -0.1017 0.9175 0.0281 -5.750 -0.2093 0.01863 0.01201 -0.1013 0.9063 0.0274 -5.500 -0.1831 0.01741 0.01057 -0.1007 0.8960 0.0274 -5.250 -0.1571 0.01649 0.00948 -0.1002 0.8864 0.0279 -5.000 -0.1305 0.01581 0.00859 -0.0998 0.8759 0.0291 -4.750 -0.1035 0.01517 0.00770 -0.0995 0.8669 0.0325 -4.500 -0.0762 0.01451 0.00681 -0.0991 0.8591 0.0364 -4.250 -0.0489 0.01372 0.00596 -0.0989 0.8505 0.0529 -4.000 -0.0204 0.01422 0.00652 -0.0987 0.8433 0.1412 -3.750 0.0077 0.01516 0.00741 -0.0986 0.8350 0.1600 -3.500 0.0350 0.01574 0.00790 -0.0985 0.8284 0.1743 -3.250 0.0627 0.01593 0.00788 -0.0984 0.8210 0.1818 -3.000 0.0895 0.01601 0.00796 -0.0984 0.8151 0.1927 -2.750 0.1167 0.01602 0.00796 -0.0985 0.8081 0.2031 -2.500 0.1441 0.01600 0.00776 -0.0984 0.8029 0.2112 -2.250 0.1720 0.01566 0.00738 -0.0986 0.7961 0.2132 -2.000 0.1997 0.01543 0.00705 -0.0986 0.7907 0.2156 -1.750 0.2276 0.01522 0.00677 -0.0987 0.7850 0.2172 -1.500 0.2556 0.01502 0.00650 -0.0988 0.7792 0.2187 -1.250 0.2834 0.01487 0.00625 -0.0987 0.7747 0.2203 -1.000 0.3115 0.01474 0.00609 -0.0989 0.7685 0.2220 -0.750 0.3394 0.01464 0.00591 -0.0989 0.7635 0.2235 -0.500 0.3670 0.01443 0.00573 -0.0990 0.7586 0.2259 -0.250 0.3947 0.01430 0.00567 -0.0991 0.7529 0.2290 0.000 0.4225 0.01421 0.00558 -0.0991 0.7485 0.2321 0.250 0.4503 0.01417 0.00557 -0.0992 0.7433 0.2350 0.500 0.4782 0.01414 0.00555 -0.0993 0.7381 0.2378 0.750 0.5058 0.01401 0.00549 -0.0993 0.7341 0.2415 1.000 0.5334 0.01400 0.00560 -0.0995 0.7286 0.2461 1.250 0.5612 0.01399 0.00564 -0.0995 0.7236 0.2514 1.500 0.5890 0.01392 0.00565 -0.0996 0.7199 0.2585 1.750 0.6166 0.01395 0.00583 -0.0998 0.7142 0.2682 2.000 0.6443 0.01390 0.00595 -0.0999 0.7093 0.2860 2.250 0.6687 0.01223 0.00607 -0.0990 0.7058 1.0000 2.500 0.6961 0.01244 0.00627 -0.0990 0.6993 1.0000 2.750 0.7240 0.01253 0.00631 -0.0988 0.6937 1.0000 3.000 0.7511 0.01268 0.00651 -0.0987 0.6867 1.0000 3.250 0.7787 0.01277 0.00661 -0.0985 0.6804 1.0000 3.500 0.8053 0.01270 0.00657 -0.0979 0.6692 1.0000 3.750 0.8310 0.01224 0.00598 -0.0966 0.6494 1.0000 4.000 0.8564 0.01203 0.00581 -0.0957 0.6305 1.0000 4.250 0.8814 0.01176 0.00556 -0.0945 0.6046 1.0000 4.500 0.9029 0.01151 0.00514 -0.0925 0.5390 1.0000 4.750 0.9037 0.01380 0.00580 -0.0882 0.2468 1.0000 5.000 0.9054 0.01673 0.00758 -0.0849 0.0442 1.0000 5.250 0.9260 0.01756 0.00849 -0.0836 0.0369 1.0000 5.500 0.9443 0.01859 0.00965 -0.0821 0.0320 1.0000 5.750 0.9623 0.01957 0.01081 -0.0805 0.0301 1.0000 6.000 0.9775 0.02076 0.01210 -0.0785 0.0289 1.0000 6.250 0.9908 0.02207 0.01349 -0.0762 0.0282 1.0000 6.500 1.0041 0.02349 0.01496 -0.0739 0.0279 1.0000 6.750 1.0196 0.02503 0.01654 -0.0717 0.0280 1.0000 7.000 1.0396 0.02662 0.01815 -0.0702 0.0274 1.0000 7.250 1.0614 0.02823 0.01977 -0.0693 0.0253 1.0000 7.500 1.0934 0.03046 0.02215 -0.0690 0.0266 1.0000 7.750 1.1329 0.03446 0.02640 -0.0694 0.0312 1.0000 16.000 0.7270 0.17689 0.17391 -0.0641 0.0427 1.0000 16.250 0.7215 0.18107 0.17808 -0.0670 0.0405 1.0000