XFOIL Version 6.94 Calculated polar for: S2046 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 0.3826 0.00800 0.00314 -0.0922 0.8584 0.7437 0.500 0.4344 0.00789 0.00309 -0.0903 0.8361 0.7865 1.000 0.4862 0.00777 0.00302 -0.0886 0.8107 0.8214 1.500 0.5364 0.00762 0.00291 -0.0865 0.7825 0.8634 2.000 0.5908 0.00734 0.00275 -0.0852 0.7483 1.0000 2.500 0.6479 0.00755 0.00279 -0.0853 0.7047 1.0000 3.000 0.7015 0.00789 0.00289 -0.0844 0.6421 1.0000 3.500 0.7514 0.00853 0.00308 -0.0829 0.5486 1.0000 4.000 0.8007 0.00936 0.00349 -0.0816 0.4600 1.0000 4.500 0.8502 0.01022 0.00402 -0.0805 0.3858 1.0000 5.000 0.9001 0.01106 0.00462 -0.0795 0.3226 1.0000 5.500 0.9494 0.01197 0.00528 -0.0785 0.2587 1.0000 6.000 0.9960 0.01322 0.00613 -0.0772 0.1717 1.0000 6.500 1.0389 0.01502 0.00748 -0.0754 0.1006 1.0000 7.000 1.0841 0.01648 0.00891 -0.0737 0.0780 1.0000 7.500 1.1278 0.01805 0.01055 -0.0718 0.0656 1.0000 8.000 1.1668 0.02028 0.01281 -0.0693 0.0571 1.0000 8.500 1.2112 0.02164 0.01437 -0.0675 0.0513 1.0000 9.000 1.2477 0.02462 0.01745 -0.0650 0.0452 1.0000 9.500 1.2890 0.02605 0.01914 -0.0629 0.0409 1.0000 10.000 1.3251 0.02778 0.02099 -0.0604 0.0361 1.0000 11.000 1.3809 0.03308 0.02688 -0.0537 0.0283 1.0000 11.500 1.3901 0.03730 0.03131 -0.0488 0.0250 1.0000 12.000 1.3949 0.04082 0.03527 -0.0438 0.0233 1.0000 12.500 1.3936 0.04531 0.04016 -0.0398 0.0216 1.0000 13.000 1.3863 0.05070 0.04590 -0.0372 0.0203 1.0000 13.500 1.3768 0.05683 0.05229 -0.0365 0.0193 1.0000 14.000 1.3566 0.06546 0.06122 -0.0380 0.0186 1.0000 14.500 1.3192 0.07828 0.07447 -0.0432 0.0183 1.0000 15.000 1.2781 0.09391 0.09053 -0.0521 0.0183 1.0000 15.500 1.2297 0.11419 0.11126 -0.0656 0.0187 1.0000