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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.500 -0.0100 0.01032 0.00464 -0.1105 0.6738 0.5048 -4.000 0.0490 0.01065 0.00484 -0.1113 0.6714 0.5308 -3.500 0.1076 0.01123 0.00537 -0.1118 0.6689 0.5545 -3.000 0.1658 0.01160 0.00569 -0.1125 0.6667 0.5670 -2.500 0.2239 0.01171 0.00571 -0.1134 0.6647 0.5738 -2.000 0.2816 0.01169 0.00567 -0.1143 0.6625 0.5782 -1.500 0.3395 0.01174 0.00569 -0.1152 0.6601 0.5819 -1.000 0.3980 0.01178 0.00568 -0.1162 0.6577 0.5855 -0.500 0.4569 0.01188 0.00569 -0.1175 0.6558 0.5892 0.000 0.5160 0.01196 0.00571 -0.1187 0.6537 0.5920 0.500 0.5737 0.01216 0.00594 -0.1197 0.6515 0.5954 1.000 0.6299 0.01226 0.00611 -0.1204 0.6490 0.5986 1.500 0.6864 0.01239 0.00627 -0.1211 0.6460 0.6021 2.000 0.7438 0.01249 0.00638 -0.1221 0.6427 0.6057 2.500 0.8023 0.01253 0.00645 -0.1231 0.6397 0.6095 3.000 0.8602 0.01273 0.00669 -0.1241 0.6361 0.6133 3.500 0.9138 0.01270 0.00678 -0.1242 0.6304 0.6176 4.000 0.9727 0.01256 0.00662 -0.1252 0.6247 0.6219 4.500 1.0285 0.01260 0.00678 -0.1258 0.6196 0.6266 5.000 1.0816 0.01260 0.00693 -0.1258 0.6137 0.6323 6.000 1.1913 0.01255 0.00711 -0.1264 0.6016 0.6439 6.500 1.2473 0.01233 0.00695 -0.1268 0.5933 0.6511 7.000 1.2937 0.01212 0.00694 -0.1253 0.5813 0.6584 7.500 1.3398 0.01205 0.00704 -0.1239 0.5690 0.6678 8.000 1.3819 0.01198 0.00718 -0.1216 0.5527 0.6776 8.500 1.4144 0.01206 0.00737 -0.1176 0.5233 0.6889 9.000 1.4147 0.01284 0.00802 -0.1077 0.4705 0.7021 9.500 1.3882 0.01520 0.01012 -0.0951 0.4094 0.7178 10.000 1.3596 0.01878 0.01355 -0.0847 0.3567 0.7380