XFOIL Version 6.94 Calculated polar for: E222 (10.17%) 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 0.3807 0.00632 0.00198 -0.0906 0.8085 0.7965 0.500 0.4311 0.00635 0.00202 -0.0888 0.7786 0.8451 1.000 0.4786 0.00641 0.00207 -0.0863 0.7468 0.8905 1.500 0.5227 0.00647 0.00208 -0.0830 0.7135 0.9395 2.000 0.5890 0.00657 0.00208 -0.0850 0.6765 0.9919 2.500 0.6431 0.00687 0.00219 -0.0847 0.6378 1.0000 3.000 0.6950 0.00720 0.00238 -0.0840 0.5948 1.0000 3.500 0.7463 0.00760 0.00261 -0.0831 0.5494 1.0000 4.000 0.7967 0.00806 0.00292 -0.0820 0.4996 1.0000 4.500 0.8465 0.00860 0.00329 -0.0809 0.4470 1.0000 5.000 0.8947 0.00928 0.00375 -0.0795 0.3858 1.0000 5.500 0.9398 0.01023 0.00433 -0.0778 0.3014 1.0000 6.000 0.9845 0.01129 0.00503 -0.0761 0.2241 1.0000 6.500 1.0284 0.01244 0.00588 -0.0743 0.1542 1.0000 7.000 1.0710 0.01371 0.00685 -0.0724 0.0952 1.0000 7.500 1.1067 0.01562 0.00830 -0.0694 0.0273 1.0000 8.000 1.1447 0.01725 0.00994 -0.0665 0.0139 1.0000 8.500 1.1791 0.01907 0.01195 -0.0630 0.0118 1.0000 9.000 1.2107 0.02086 0.01396 -0.0591 0.0109 1.0000 9.500 1.2328 0.02293 0.01624 -0.0539 0.0103 1.0000 10.000 1.2492 0.02537 0.01889 -0.0482 0.0099 1.0000 10.500 1.2631 0.02828 0.02207 -0.0428 0.0097 1.0000 11.000 1.2754 0.03181 0.02589 -0.0379 0.0096 1.0000 11.500 1.2852 0.03606 0.03049 -0.0336 0.0096 1.0000 12.000 1.2904 0.04067 0.03545 -0.0300 0.0094 1.0000 12.500 1.2890 0.04581 0.04092 -0.0273 0.0091 1.0000 13.000 1.2774 0.05274 0.04827 -0.0256 0.0091 1.0000 13.500 1.2575 0.06131 0.05726 -0.0260 0.0092 1.0000 14.000 1.2250 0.07278 0.06920 -0.0294 0.0094 1.0000 14.500 1.1845 0.08742 0.08429 -0.0367 0.0096 1.0000 15.000 1.1384 0.10582 0.10310 -0.0479 0.0098 1.0000 15.500 1.0888 0.12823 0.12586 -0.0625 0.0102 1.0000 16.000 1.0073 0.16528 0.16320 -0.0848 0.0115 1.0000