XFOIL Version 6.94 Calculated polar for: EPPLER 715 AIRFOIL 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 0.2457 0.01354 0.00628 -0.0435 0.5731 0.5671 0.500 0.3026 0.01351 0.00632 -0.0437 0.5677 0.5770 1.000 0.3601 0.01362 0.00640 -0.0440 0.5631 0.5867 1.500 0.4172 0.01374 0.00651 -0.0443 0.5590 0.5953 2.000 0.4737 0.01401 0.00685 -0.0446 0.5540 0.6043 2.500 0.5303 0.01415 0.00711 -0.0449 0.5485 0.6138 3.000 0.5869 0.01421 0.00728 -0.0451 0.5428 0.6237 3.500 0.6440 0.01431 0.00739 -0.0453 0.5376 0.6351 4.000 0.6993 0.01454 0.00781 -0.0454 0.5312 0.6486 4.500 0.7545 0.01451 0.00800 -0.0454 0.5235 0.6654 5.000 0.8110 0.01433 0.00795 -0.0452 0.5168 0.6938 5.500 0.8602 0.01403 0.00829 -0.0439 0.5084 0.7990 6.000 0.9398 0.01362 0.00817 -0.0485 0.4982 1.0000 6.500 0.9946 0.01365 0.00825 -0.0482 0.4882 1.0000 7.000 1.0494 0.01343 0.00809 -0.0478 0.4759 1.0000 7.500 1.1023 0.01335 0.00817 -0.0471 0.4611 1.0000 8.000 1.1546 0.01323 0.00816 -0.0464 0.4421 1.0000 8.500 1.2045 0.01326 0.00828 -0.0453 0.4144 1.0000 9.000 1.2478 0.01376 0.00871 -0.0435 0.3691 1.0000 9.500 1.2750 0.01526 0.00993 -0.0398 0.3006 1.0000 10.000 1.2823 0.01755 0.01194 -0.0339 0.2407 1.0000 10.500 1.2651 0.02026 0.01453 -0.0253 0.2038 1.0000 11.000 1.2496 0.02445 0.01865 -0.0207 0.1716 1.0000 11.500 1.2382 0.02934 0.02346 -0.0181 0.1428 1.0000 12.000 1.2282 0.03444 0.02850 -0.0162 0.1184 1.0000 13.000 1.2134 0.04485 0.03882 -0.0136 0.0804 1.0000 13.500 1.2096 0.05024 0.04420 -0.0130 0.0663 1.0000 14.000 1.2057 0.05601 0.04996 -0.0129 0.0554 1.0000 14.500 1.2061 0.06162 0.05563 -0.0132 0.0462 1.0000 15.000 1.2050 0.06764 0.06172 -0.0138 0.0395 1.0000 15.500 1.2032 0.07405 0.06820 -0.0147 0.0342 1.0000 16.000 1.2015 0.08072 0.07497 -0.0161 0.0297 1.0000 16.500 1.2000 0.08767 0.08204 -0.0179 0.0259 1.0000 17.000 1.1979 0.09495 0.08946 -0.0200 0.0227 1.0000 17.500 1.1958 0.10242 0.09709 -0.0224 0.0200 1.0000 18.000 1.1935 0.11008 0.10488 -0.0251 0.0179 1.0000 18.500 1.1915 0.11792 0.11281 -0.0284 0.0160 1.0000 19.000 1.1875 0.12647 0.12163 -0.0323 0.0144 1.0000 19.500 1.1868 0.13422 0.12942 -0.0361 0.0130 1.0000 20.000 1.1796 0.14384 0.13937 -0.0412 0.0120 1.0000 20.500 1.1771 0.15238 0.14806 -0.0461 0.0111 1.0000 21.000 1.1771 0.16013 0.15593 -0.0504 0.0105 1.0000 21.500 1.1614 0.17212 0.16829 -0.0579 0.0101 1.0000 22.000 1.1446 0.18462 0.18110 -0.0659 0.0098 1.0000 22.500 1.1180 0.20031 0.19713 -0.0761 0.0098 1.0000