XFOIL Version 6.94 Calculated polar for: EPPLER 694 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 0.3925 0.02484 0.01995 -0.1243 0.7954 0.8326 0.500 0.4539 0.02448 0.01950 -0.1249 0.7932 0.8484 1.000 0.5195 0.02386 0.01881 -0.1263 0.7918 0.8600 1.500 0.5510 0.02416 0.01911 -0.1218 0.7822 0.8735 2.000 0.6241 0.02315 0.01805 -0.1248 0.7811 0.8832 4.000 0.9060 0.01874 0.01371 -0.1354 0.7675 0.9070 5.000 1.0707 0.01639 0.01150 -0.1456 0.7584 0.9163 5.500 1.1295 0.01590 0.01112 -0.1463 0.7474 0.9217 6.000 1.2033 0.01519 0.01048 -0.1499 0.7340 0.9257 6.500 1.2249 0.01552 0.01094 -0.1438 0.7137 0.9315 7.000 1.2401 0.01611 0.01168 -0.1368 0.6896 0.9376 7.500 1.2546 0.01674 0.01244 -0.1299 0.6583 0.9440 8.000 1.2938 0.01690 0.01248 -0.1272 0.6090 0.9489 8.500 1.3159 0.01801 0.01327 -0.1219 0.5404 0.9553 9.000 1.3166 0.02039 0.01529 -0.1139 0.4634 0.9635 9.500 1.3154 0.02339 0.01794 -0.1066 0.3881 0.9774 10.000 1.3185 0.02684 0.02105 -0.1010 0.3142 1.0000 10.500 1.3281 0.03049 0.02439 -0.0969 0.2487 1.0000 11.000 1.3425 0.03404 0.02771 -0.0939 0.1910 1.0000 11.500 1.3541 0.03802 0.03136 -0.0909 0.1295 1.0000 12.000 1.3674 0.04206 0.03516 -0.0885 0.0831 1.0000 12.500 1.3710 0.04714 0.03991 -0.0854 0.0337 1.0000 13.000 1.3723 0.05280 0.04553 -0.0824 0.0162 1.0000 13.500 1.3816 0.05790 0.05082 -0.0806 0.0136 1.0000 14.000 1.3901 0.06331 0.05646 -0.0792 0.0121 1.0000 14.500 1.3897 0.06995 0.06329 -0.0780 0.0113 1.0000 15.000 1.3934 0.07636 0.06996 -0.0774 0.0108 1.0000 15.500 1.3960 0.08309 0.07693 -0.0772 0.0104 1.0000 16.000 1.3993 0.08986 0.08391 -0.0774 0.0100 1.0000 16.500 1.4042 0.09642 0.09067 -0.0780 0.0098 1.0000 17.000 1.4112 0.10268 0.09713 -0.0788 0.0095 1.0000 17.500 1.4200 0.10860 0.10324 -0.0797 0.0094 1.0000 18.000 1.4290 0.11446 0.10933 -0.0808 0.0093 1.0000 18.500 1.4331 0.12127 0.11646 -0.0828 0.0093 1.0000 19.000 1.4295 0.12958 0.12515 -0.0861 0.0093 1.0000 19.500 1.4170 0.13981 0.13584 -0.0913 0.0094 1.0000 20.000 1.3912 0.15325 0.14984 -0.0994 0.0096 1.0000 20.500 1.3504 0.17114 0.16833 -0.1119 0.0100 1.0000 21.000 1.3108 0.19054 0.18818 -0.1265 0.0104 1.0000 21.500 1.2688 0.21322 0.21119 -0.1438 0.0109 1.0000