XFOIL Version 6.94 Calculated polar for: GOE 385 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 0.2770 0.01185 0.00362 -0.0482 0.6516 0.0723 1.000 0.3886 0.01139 0.00307 -0.0477 0.6125 0.0876 1.500 0.4615 0.00886 0.00312 -0.0510 0.5940 1.0000 2.000 0.5136 0.00906 0.00313 -0.0500 0.5719 1.0000 2.500 0.5664 0.00927 0.00319 -0.0492 0.5514 1.0000 3.000 0.6198 0.00945 0.00329 -0.0484 0.5260 1.0000 3.500 0.6728 0.00965 0.00335 -0.0477 0.4903 1.0000 4.000 0.7255 0.00999 0.00355 -0.0470 0.4496 1.0000 4.500 0.7774 0.01049 0.00385 -0.0463 0.4091 1.0000 5.000 0.8290 0.01112 0.00430 -0.0456 0.3763 1.0000 5.500 0.8801 0.01176 0.00478 -0.0449 0.3441 1.0000 6.000 0.9322 0.01238 0.00537 -0.0443 0.3266 1.0000 6.500 0.9838 0.01300 0.00595 -0.0436 0.3070 1.0000 7.000 1.0347 0.01366 0.00659 -0.0430 0.2905 1.0000 7.500 1.0854 0.01427 0.00726 -0.0422 0.2722 1.0000 8.000 1.1352 0.01489 0.00793 -0.0414 0.2520 1.0000 8.500 1.1843 0.01551 0.00865 -0.0405 0.2310 1.0000 9.000 1.2318 0.01624 0.00943 -0.0395 0.2003 1.0000 9.500 1.2732 0.01750 0.01051 -0.0377 0.1455 1.0000 10.000 1.2985 0.02009 0.01273 -0.0341 0.0902 1.0000 10.500 1.3273 0.02210 0.01462 -0.0309 0.0618 1.0000 11.000 1.3442 0.02443 0.01700 -0.0261 0.0533 1.0000 11.500 1.3494 0.02755 0.02023 -0.0209 0.0483 1.0000 12.000 1.3539 0.03129 0.02411 -0.0175 0.0456 1.0000 12.500 1.3545 0.03600 0.02892 -0.0154 0.0422 1.0000 13.000 1.3592 0.04081 0.03395 -0.0142 0.0393 1.0000 13.500 1.3620 0.04600 0.03926 -0.0137 0.0369 1.0000 14.000 1.3640 0.05149 0.04488 -0.0135 0.0342 1.0000 14.500 1.3686 0.05657 0.05011 -0.0127 0.0330 1.0000 15.000 1.3668 0.06288 0.05665 -0.0136 0.0305 1.0000 15.500 1.3647 0.06932 0.06329 -0.0145 0.0287 1.0000 16.000 1.3600 0.07659 0.07082 -0.0165 0.0270 1.0000 16.500 1.3583 0.08324 0.07758 -0.0176 0.0255 1.0000 17.000 1.3451 0.09261 0.08732 -0.0213 0.0240 1.0000 17.500 1.3410 0.10040 0.09521 -0.0243 0.0227 1.0000 18.000 1.3241 0.11081 0.10596 -0.0288 0.0216 1.0000 18.500 1.3077 0.12187 0.11729 -0.0347 0.0201 1.0000 19.000 1.2989 0.13119 0.12673 -0.0393 0.0189 1.0000 19.500 1.2763 0.14379 0.13970 -0.0462 0.0185 1.0000 20.000 1.2510 0.15803 0.15428 -0.0549 0.0175 1.0000