XFOIL Version 6.94 Calculated polar for: GOE 496 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 0.6224 0.00951 0.00330 -0.1459 0.7606 0.4341 0.500 0.6758 0.00946 0.00334 -0.1450 0.7391 0.4973 1.000 0.7277 0.00927 0.00341 -0.1439 0.7165 0.6004 2.000 0.8310 0.00885 0.00350 -0.1410 0.6709 1.0000 2.500 0.8828 0.00911 0.00356 -0.1398 0.6431 1.0000 3.000 0.9276 0.00941 0.00356 -0.1370 0.5839 1.0000 3.500 0.9712 0.00998 0.00374 -0.1342 0.5176 1.0000 4.000 1.0135 0.01072 0.00411 -0.1315 0.4445 1.0000 4.500 1.0527 0.01174 0.00466 -0.1284 0.3558 1.0000 5.000 1.0818 0.01360 0.00574 -0.1239 0.2253 1.0000 5.500 1.1215 0.01473 0.00658 -0.1212 0.1730 1.0000 6.000 1.1359 0.01761 0.00849 -0.1147 0.0225 1.0000 6.500 1.1712 0.01875 0.00960 -0.1110 0.0057 1.0000 7.000 1.2058 0.01973 0.01071 -0.1072 0.0054 1.0000 7.500 1.2377 0.02086 0.01198 -0.1031 0.0054 1.0000 8.000 1.2672 0.02216 0.01345 -0.0988 0.0055 1.0000 8.500 1.2940 0.02368 0.01515 -0.0944 0.0057 1.0000 9.000 1.3173 0.02551 0.01717 -0.0899 0.0060 1.0000 9.500 1.3354 0.02781 0.01969 -0.0852 0.0062 1.0000 10.000 1.3474 0.03072 0.02280 -0.0805 0.0065 1.0000 10.500 1.3517 0.03450 0.02678 -0.0759 0.0067 1.0000 11.000 1.3517 0.03908 0.03153 -0.0720 0.0069 1.0000 11.500 1.3656 0.04275 0.03543 -0.0694 0.0074 1.0000 12.000 1.3694 0.04784 0.04076 -0.0669 0.0079 1.0000 12.500 1.3752 0.05330 0.04638 -0.0648 0.0085 1.0000 13.000 1.3921 0.05795 0.05122 -0.0630 0.0092 1.0000 13.500 1.4346 0.06281 0.05641 -0.0603 0.0111 1.0000 14.000 1.3090 0.08011 0.07483 -0.0618 0.0112 1.0000 14.500 1.2559 0.09432 0.09002 -0.0657 0.0125 1.0000 15.000 1.2332 0.10460 0.10077 -0.0691 0.0139 1.0000 15.500 1.1878 0.11868 0.11535 -0.0774 0.0140 1.0000 16.000 1.1382 0.13542 0.13256 -0.0886 0.0140 1.0000 16.500 1.1032 0.15075 0.14819 -0.0991 0.0139 1.0000 17.000 1.0700 0.16712 0.16479 -0.1101 0.0137 1.0000 17.500 0.9581 0.21758 0.21560 -0.1385 0.0128 1.0000 18.500 0.9473 0.25173 0.24975 -0.1506 0.0267 1.0000 19.000 0.9594 0.25989 0.25794 -0.1525 0.0264 1.0000 19.500 0.9514 0.27495 0.27298 -0.1603 0.0241 1.0000 20.000 0.9611 0.28316 0.28120 -0.1630 0.0231 1.0000 20.500 0.9639 0.29491 0.29296 -0.1677 0.0220 1.0000 21.000 0.9704 0.30490 0.30297 -0.1713 0.0202 1.0000 21.500 0.9832 0.31213 0.31023 -0.1727 0.0194 1.0000 22.000 0.9825 0.32561 0.32370 -0.1785 0.0177 1.0000 22.500 0.9904 0.33436 0.33247 -0.1813 0.0166 1.0000 23.000 0.9954 0.34561 0.34374 -0.1849 0.0159 1.0000 23.500 1.0011 0.35635 0.35449 -0.1885 0.0144 1.0000 24.000 1.0084 0.36503 0.36320 -0.1911 0.0137 1.0000 24.500 1.0126 0.37686 0.37504 -0.1948 0.0132 1.0000 25.000 1.0179 0.38792 0.38612 -0.1982 0.0118 1.0000 25.500 1.0236 0.39726 0.39549 -0.2010 0.0113 1.0000 26.000 1.0291 0.40718 0.40542 -0.2036 0.0110 1.0000 26.500 1.0324 0.41968 0.41795 -0.2075 0.0105 1.0000 27.000 1.0368 0.43043 0.42872 -0.2105 0.0098 1.0000 27.500 1.0410 0.44065 0.43896 -0.2134 0.0092 1.0000 28.000 1.0450 0.45008 0.44842 -0.2159 0.0089 1.0000 28.500 1.0488 0.46059 0.45896 -0.2186 0.0088 1.0000 29.000 1.0511 0.47197 0.47036 -0.2219 0.0088 1.0000 29.500 1.0537 0.48416 0.48257 -0.2251 0.0084 1.0000 30.000 1.0564 0.49509 0.49353 -0.2279 0.0081 1.0000 30.500 1.0584 0.50556 0.50403 -0.2307 0.0080 1.0000 31.000 1.0602 0.51617 0.51468 -0.2335 0.0079 1.0000 31.500 1.0617 0.52686 0.52540 -0.2362 0.0076 1.0000 32.000 1.0626 0.53701 0.53558 -0.2388 0.0075 1.0000 32.500 1.0634 0.54741 0.54601 -0.2415 0.0074 1.0000 33.000 1.0638 0.55766 0.55629 -0.2441 0.0072 1.0000 33.500 1.0638 0.56763 0.56629 -0.2467 0.0071 1.0000 34.000 1.0635 0.57772 0.57642 -0.2493 0.0072 1.0000 34.500 1.0628 0.58764 0.58637 -0.2518 0.0071 1.0000 35.000 1.0616 0.59721 0.59599 -0.2542 0.0070 1.0000 35.500 1.0602 0.60687 0.60568 -0.2566 0.0070 1.0000