XFOIL Version 6.96 Calculated polar for: EPPLER 1098 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.0558 0.00982 0.00356 -0.1094 0.6449 0.1287 -4.000 0.0513 0.00762 0.00255 -0.1117 0.6406 0.4874 -3.500 0.1111 0.00780 0.00264 -0.1128 0.6380 0.5119 -3.000 0.1705 0.00798 0.00279 -0.1138 0.6361 0.5306 -2.500 0.2298 0.00809 0.00289 -0.1147 0.6345 0.5406 -2.000 0.2891 0.00812 0.00287 -0.1158 0.6328 0.5463 -1.500 0.3482 0.00814 0.00288 -0.1169 0.6309 0.5503 -1.000 0.4075 0.00817 0.00288 -0.1179 0.6290 0.5535 -0.500 0.4667 0.00822 0.00288 -0.1191 0.6270 0.5563 0.000 0.5261 0.00830 0.00292 -0.1202 0.6252 0.5587 0.500 0.5853 0.00840 0.00301 -0.1214 0.6229 0.5619 1.000 0.6440 0.00847 0.00312 -0.1225 0.6210 0.5647 1.500 0.7019 0.00848 0.00318 -0.1234 0.6186 0.5677 2.000 0.7598 0.00849 0.00323 -0.1243 0.6153 0.5709 2.500 0.8180 0.00851 0.00324 -0.1252 0.6114 0.5740 3.000 0.8756 0.00857 0.00331 -0.1260 0.6061 0.5775 3.500 0.9315 0.00850 0.00335 -0.1265 0.6016 0.5812 4.000 0.9884 0.00851 0.00342 -0.1272 0.5973 0.5850 4.500 1.0457 0.00863 0.00354 -0.1280 0.5927 0.5889 5.000 1.1007 0.00863 0.00367 -0.1284 0.5881 0.5935 5.500 1.1556 0.00861 0.00377 -0.1287 0.5821 0.5987 6.000 1.2087 0.00863 0.00386 -0.1286 0.5730 0.6039 6.500 1.2605 0.00865 0.00396 -0.1283 0.5620 0.6096 7.000 1.3109 0.00870 0.00413 -0.1278 0.5464 0.6162 7.500 1.3544 0.00895 0.00437 -0.1260 0.5123 0.6230 8.000 1.3620 0.01010 0.00521 -0.1174 0.4382 0.6309 8.500 1.3602 0.01160 0.00648 -0.1076 0.3819 0.6396 9.000 1.3485 0.01373 0.00839 -0.0974 0.3258 0.6496 9.500 1.3387 0.01645 0.01096 -0.0891 0.2795 0.6610