XFOIL Version 6.94 Calculated polar for: NACA 2408 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2672 0.00541 0.00168 -0.0594 0.8749 1.0000 0.500 0.3182 0.00547 0.00156 -0.0579 0.8382 1.0000 1.000 0.3685 0.00564 0.00148 -0.0562 0.7842 1.0000 1.500 0.4194 0.00590 0.00149 -0.0548 0.7262 1.0000 2.000 0.4701 0.00627 0.00157 -0.0534 0.6589 1.0000 2.500 0.5203 0.00676 0.00172 -0.0521 0.5801 1.0000 3.000 0.5720 0.00722 0.00195 -0.0512 0.5173 1.0000 3.500 0.6233 0.00777 0.00225 -0.0503 0.4456 1.0000 4.000 0.6717 0.00874 0.00268 -0.0492 0.3146 1.0000 4.500 0.7147 0.01068 0.00361 -0.0476 0.1034 1.0000 5.000 0.7638 0.01181 0.00448 -0.0465 0.0526 1.0000 5.500 0.8144 0.01269 0.00543 -0.0455 0.0456 1.0000 6.000 0.8619 0.01397 0.00679 -0.0441 0.0401 1.0000 6.500 0.9097 0.01513 0.00806 -0.0428 0.0352 1.0000 7.000 0.9530 0.01700 0.01008 -0.0408 0.0298 1.0000 7.500 0.9925 0.01986 0.01304 -0.0385 0.0239 1.0000 8.000 1.0408 0.02076 0.01417 -0.0372 0.0202 1.0000 8.500 1.0805 0.02354 0.01712 -0.0351 0.0169 1.0000 9.000 1.1228 0.02566 0.01958 -0.0330 0.0149 1.0000 9.500 1.1614 0.02768 0.02184 -0.0309 0.0132 1.0000 10.000 1.1847 0.03268 0.02737 -0.0273 0.0120 1.0000 10.500 1.1946 0.03874 0.03418 -0.0223 0.0114 1.0000 11.000 1.1742 0.04626 0.04241 -0.0153 0.0111 1.0000 11.500 1.1343 0.05467 0.05135 -0.0108 0.0112 1.0000 12.000 1.0905 0.06581 0.06293 -0.0134 0.0113 1.0000 12.500 1.0466 0.08087 0.07833 -0.0232 0.0115 1.0000 13.000 0.9998 0.10104 0.09873 -0.0376 0.0119 1.0000