XFOIL Version 6.94 Calculated polar for: EPPLER 403 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.3364 0.01950 0.01416 -0.0975 0.8433 0.6737 1.000 0.4165 0.01874 0.01337 -0.1023 0.8392 0.6797 1.500 0.4991 0.01772 0.01241 -0.1073 0.8367 0.6845 2.000 0.5795 0.01664 0.01142 -0.1115 0.8335 0.6885 3.500 0.7800 0.01383 0.00889 -0.1168 0.7989 0.7045 4.000 0.8438 0.01294 0.00813 -0.1177 0.7777 0.7087 4.500 0.9031 0.01242 0.00768 -0.1178 0.7398 0.7141 5.000 0.9726 0.01218 0.00726 -0.1201 0.6748 0.7197 5.500 1.0081 0.01296 0.00760 -0.1162 0.5822 0.7248 6.000 1.0209 0.01417 0.00837 -0.1083 0.4884 0.7295 6.500 1.0319 0.01568 0.00944 -0.1006 0.3950 0.7350 7.000 1.0518 0.01722 0.01059 -0.0951 0.3108 0.7413 7.500 1.0788 0.01878 0.01180 -0.0912 0.2373 0.7466 8.000 1.1071 0.02024 0.01304 -0.0876 0.1749 0.7517 8.500 1.1341 0.02201 0.01453 -0.0840 0.1147 0.7573 9.000 1.1606 0.02402 0.01627 -0.0806 0.0645 0.7641 9.500 1.1787 0.02686 0.01883 -0.0762 0.0225 0.7700 10.000 1.1996 0.02935 0.02148 -0.0721 0.0159 0.7759 10.500 1.2189 0.03202 0.02435 -0.0683 0.0141 0.7824 11.000 1.2363 0.03501 0.02753 -0.0647 0.0130 0.7890 11.500 1.2473 0.03848 0.03119 -0.0608 0.0125 0.7962 12.000 1.2583 0.04241 0.03534 -0.0575 0.0121 0.8041 12.500 1.2733 0.04648 0.03965 -0.0550 0.0119 0.8117 13.000 1.2885 0.05075 0.04424 -0.0528 0.0117 0.8198 13.500 1.3048 0.05555 0.04936 -0.0511 0.0116 0.8289 14.000 1.3159 0.06106 0.05529 -0.0497 0.0116 0.8386 14.500 1.3186 0.06781 0.06250 -0.0487 0.0117 0.8504 15.000 1.3103 0.07577 0.07097 -0.0484 0.0118 0.8634 15.500 1.2930 0.08510 0.08078 -0.0493 0.0119 0.8808 16.000 1.2645 0.09425 0.09042 -0.0498 0.0120 0.9706