XFOIL Version 6.94 Calculated polar for: E214 (11.1%) 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 0.5886 0.00867 0.00338 -0.1404 0.8026 0.7436 0.500 0.6428 0.00869 0.00336 -0.1397 0.7780 0.7754 1.000 0.6952 0.00875 0.00339 -0.1386 0.7541 0.8110 1.500 0.7439 0.00881 0.00348 -0.1367 0.7298 0.8533 2.500 0.8461 0.00893 0.00358 -0.1342 0.6827 1.0000 3.000 0.9014 0.00928 0.00377 -0.1343 0.6586 1.0000 3.500 0.9536 0.00961 0.00405 -0.1336 0.6332 1.0000 4.000 1.0053 0.01000 0.00433 -0.1328 0.6075 1.0000 4.500 1.0543 0.01035 0.00469 -0.1314 0.5783 1.0000 5.000 1.1025 0.01077 0.00507 -0.1299 0.5487 1.0000 5.500 1.1486 0.01124 0.00550 -0.1280 0.5154 1.0000 6.000 1.1926 0.01179 0.00603 -0.1257 0.4783 1.0000 6.500 1.2338 0.01245 0.00662 -0.1229 0.4356 1.0000 7.000 1.2703 0.01332 0.00735 -0.1194 0.3820 1.0000 7.500 1.3008 0.01447 0.00829 -0.1149 0.3182 1.0000 8.000 1.3249 0.01588 0.00945 -0.1095 0.2527 1.0000 8.500 1.3442 0.01771 0.01094 -0.1037 0.1844 1.0000 9.000 1.3615 0.01979 0.01273 -0.0979 0.1252 1.0000 9.500 1.3789 0.02195 0.01469 -0.0926 0.0854 1.0000 10.000 1.3960 0.02421 0.01690 -0.0876 0.0608 1.0000 10.500 1.4122 0.02666 0.01940 -0.0828 0.0419 1.0000 11.000 1.4138 0.03043 0.02316 -0.0771 0.0233 1.0000 13.000 1.3993 0.05175 0.04539 -0.0629 0.0121 1.0000 13.500 1.4008 0.05777 0.05171 -0.0617 0.0116 1.0000 14.000 1.4004 0.06445 0.05870 -0.0610 0.0111 1.0000 14.500 1.3981 0.07184 0.06643 -0.0610 0.0107 1.0000 15.000 1.3916 0.08026 0.07523 -0.0619 0.0104 1.0000 15.500 1.3800 0.08987 0.08524 -0.0642 0.0102 1.0000 16.000 1.3639 0.10069 0.09644 -0.0680 0.0100 1.0000 16.500 1.3433 0.11301 0.10916 -0.0736 0.0098 1.0000 17.000 1.3183 0.12706 0.12360 -0.0811 0.0098 1.0000 17.500 1.2842 0.14433 0.14129 -0.0916 0.0099 1.0000 18.000 1.2355 0.16767 0.16507 -0.1071 0.0105 1.0000 18.500 1.1693 0.20105 0.19875 -0.1286 0.0119 1.0000