XFOIL Version 6.94 Calculated polar for: GOE 92 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 ------- -------- --------- --------- -------- ------- ------- 1.500 0.5921 0.00010 -0.00771 -0.0658 0.3661 0.0509 2.000 0.6421 0.00010 -0.00778 -0.0651 0.3546 0.0526 2.500 0.6924 0.00010 -0.00781 -0.0644 0.3461 0.0582 3.000 0.7457 0.00004 -0.00599 -0.0649 0.3379 1.0000 3.500 0.7951 0.00005 -0.00613 -0.0641 0.3295 1.0000 4.000 0.8443 0.00007 -0.00624 -0.0633 0.3228 1.0000 4.500 0.8939 0.00009 -0.00624 -0.0626 0.3156 1.0000 5.000 0.9432 0.00012 -0.00624 -0.0619 0.3101 1.0000 5.500 0.9920 0.00016 -0.00626 -0.0612 0.3028 1.0000 6.000 1.0407 0.00019 -0.00616 -0.0604 0.2960 1.0000 6.500 1.0889 0.00024 -0.00608 -0.0596 0.2900 1.0000 7.000 1.1364 0.00030 -0.00596 -0.0588 0.2834 1.0000 7.500 1.1832 0.00036 -0.00578 -0.0578 0.2759 1.0000 8.000 1.2287 0.00043 -0.00561 -0.0567 0.2671 1.0000 8.500 1.2728 0.00050 -0.00541 -0.0553 0.2563 1.0000 9.000 1.3148 0.00059 -0.00521 -0.0537 0.2425 1.0000 9.500 1.3526 0.00072 -0.00503 -0.0516 0.2162 1.0000 10.000 1.3783 0.00106 -0.00480 -0.0478 0.1796 1.0000 10.500 1.3940 0.00162 -0.00427 -0.0429 0.1610 1.0000 11.000 1.4080 0.00235 -0.00345 -0.0384 0.1497 1.0000 11.500 1.4191 0.00341 -0.00226 -0.0343 0.1414 1.0000 12.000 1.4263 0.00502 -0.00052 -0.0304 0.1328 1.0000 12.500 1.4279 0.00753 0.00213 -0.0268 0.1244 1.0000 13.000 1.4263 0.01110 0.00588 -0.0239 0.1144 1.0000 13.500 1.4153 0.01658 0.01153 -0.0215 0.1046 1.0000 14.000 1.3987 0.02385 0.01900 -0.0199 0.0883 1.0000 14.500 1.3473 0.03636 0.03158 -0.0189 0.0619 1.0000 15.500 1.2287 0.06126 0.05698 -0.0177 0.0550 1.0000 16.000 1.1631 0.07205 0.06798 -0.0184 0.0501 1.0000 16.500 1.1124 0.08225 0.07831 -0.0212 0.0425 1.0000 17.000 1.0748 0.09136 0.08760 -0.0245 0.0365 1.0000 17.500 1.0358 0.10115 0.09750 -0.0288 0.0300 1.0000 18.000 1.0014 0.11095 0.10747 -0.0336 0.0281 1.0000 18.500 0.9870 0.11806 0.11471 -0.0371 0.0277 1.0000 19.000 0.9608 0.12737 0.12411 -0.0421 0.0207 1.0000