XFOIL Version 6.94 Calculated polar for: WORTMANN FX 049-915 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.4356 0.01535 0.00703 -0.0559 0.4725 0.0155 0.500 0.4857 0.01595 0.00775 -0.0549 0.4698 0.0524 1.000 0.5370 0.01643 0.00833 -0.0542 0.4684 0.0658 1.500 0.5889 0.01662 0.00861 -0.0536 0.4673 0.0823 2.000 0.6210 0.01535 0.00888 -0.0497 0.4661 0.6182 2.500 0.8215 0.01514 0.00953 -0.0810 0.4636 1.0000 3.000 0.8661 0.01561 0.01000 -0.0793 0.4616 1.0000 3.500 0.9110 0.01615 0.01054 -0.0777 0.4598 1.0000 4.000 0.9564 0.01670 0.01110 -0.0762 0.4581 1.0000 4.500 1.0290 0.01198 0.00562 -0.0770 0.4320 1.0000 5.000 1.0766 0.01116 0.00485 -0.0752 0.4094 1.0000 5.500 1.1232 0.01114 0.00475 -0.0736 0.3809 1.0000 6.000 1.1210 0.01417 0.00691 -0.0647 0.2185 1.0000 6.500 1.0666 0.01933 0.01134 -0.0490 0.0488 1.0000 7.000 1.0342 0.02355 0.01561 -0.0402 0.0063 1.0000 7.500 1.0336 0.02925 0.02150 -0.0403 0.0059 1.0000 8.000 1.0324 0.03436 0.02673 -0.0391 0.0061 1.0000 8.500 1.0332 0.03901 0.03149 -0.0376 0.0067 1.0000 9.000 1.0361 0.04344 0.03603 -0.0363 0.0073 1.0000 9.500 1.0426 0.04761 0.04030 -0.0351 0.0082 1.0000 10.000 1.0502 0.05174 0.04454 -0.0340 0.0085 1.0000 10.500 1.0577 0.05626 0.04919 -0.0333 0.0083 1.0000 11.000 1.0666 0.06074 0.05381 -0.0326 0.0081 1.0000 11.500 1.0735 0.06557 0.05879 -0.0321 0.0081 1.0000 12.000 1.0810 0.07043 0.06379 -0.0316 0.0082 1.0000 12.500 1.0806 0.07635 0.06982 -0.0314 0.0077 1.0000 13.000 1.0837 0.08197 0.07553 -0.0313 0.0078 1.0000 13.500 1.0845 0.08798 0.08162 -0.0313 0.0078 1.0000 14.000 1.0835 0.09372 0.08732 -0.0309 0.0072 1.0000 14.500 1.1041 0.09600 0.08952 -0.0293 0.0070 1.0000 15.000 0.9181 0.10259 0.09712 -0.0166 0.0083 1.0000 15.500 0.9278 0.10582 0.10038 -0.0160 0.0082 1.0000