XFOIL Version 6.94 Calculated polar for: E211 (10.96%) 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 0.4092 0.00825 0.00309 -0.1014 0.8407 0.7449 0.500 0.4695 0.00830 0.00303 -0.1019 0.8055 0.7669 1.000 0.5242 0.00840 0.00306 -0.1013 0.7708 0.7876 1.500 0.5768 0.00855 0.00316 -0.1002 0.7382 0.8094 2.000 0.6277 0.00873 0.00328 -0.0989 0.7062 0.8350 2.500 0.6746 0.00886 0.00344 -0.0966 0.6743 0.8657 3.000 0.7171 0.00893 0.00358 -0.0933 0.6442 0.9104 3.500 0.7763 0.00904 0.00371 -0.0939 0.6111 1.0000 4.000 0.8303 0.00941 0.00398 -0.0938 0.5777 1.0000 4.500 0.8813 0.00980 0.00432 -0.0929 0.5423 1.0000 5.000 0.9297 0.01026 0.00469 -0.0915 0.5005 1.0000 5.500 0.9726 0.01089 0.00510 -0.0890 0.4306 1.0000 6.000 1.0122 0.01186 0.00575 -0.0861 0.3417 1.0000 6.500 1.0463 0.01340 0.00673 -0.0826 0.2288 1.0000 7.000 1.0792 0.01520 0.00798 -0.0792 0.1311 1.0000 7.500 1.1105 0.01712 0.00948 -0.0754 0.0557 1.0000 8.000 1.1340 0.01956 0.01174 -0.0700 0.0222 1.0000 8.500 1.1582 0.02150 0.01388 -0.0647 0.0194 1.0000 9.000 1.1820 0.02342 0.01598 -0.0598 0.0172 1.0000 9.500 1.1939 0.02638 0.01907 -0.0537 0.0156 1.0000 10.000 1.2115 0.02945 0.02236 -0.0488 0.0150 1.0000 10.500 1.2335 0.03274 0.02592 -0.0449 0.0146 1.0000 11.000 1.2559 0.03672 0.03023 -0.0415 0.0143 1.0000 11.500 1.2723 0.04162 0.03560 -0.0379 0.0142 1.0000 12.000 1.2749 0.04760 0.04209 -0.0341 0.0143 1.0000 12.500 1.2620 0.05484 0.04985 -0.0309 0.0145 1.0000 13.000 1.2364 0.06366 0.05916 -0.0295 0.0148 1.0000 13.500 1.2034 0.07438 0.07031 -0.0309 0.0151 1.0000 14.000 1.1775 0.08544 0.08171 -0.0342 0.0155 1.0000 14.500 1.1486 0.09705 0.09368 -0.0405 0.0158 1.0000 15.500 0.9666 0.16883 0.16629 -0.0853 0.0207 1.0000