XFOIL Version 6.94 Calculated polar for: S7012 8.75% 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 0.2636 0.00621 0.00165 -0.0556 0.7209 0.8445 0.500 0.3135 0.00630 0.00165 -0.0537 0.6792 0.8818 1.000 0.3604 0.00637 0.00164 -0.0512 0.6420 0.9276 1.500 0.4262 0.00646 0.00161 -0.0531 0.6053 0.9965 2.000 0.4827 0.00672 0.00170 -0.0534 0.5708 1.0000 2.500 0.5383 0.00701 0.00186 -0.0535 0.5394 1.0000 3.000 0.5935 0.00732 0.00205 -0.0534 0.5083 1.0000 3.500 0.6484 0.00765 0.00228 -0.0533 0.4765 1.0000 4.000 0.7032 0.00799 0.00259 -0.0531 0.4441 1.0000 4.500 0.7574 0.00837 0.00292 -0.0528 0.4025 1.0000 5.000 0.8092 0.00902 0.00332 -0.0523 0.3306 1.0000 5.500 0.8598 0.00989 0.00390 -0.0516 0.2523 1.0000 6.000 0.9084 0.01109 0.00472 -0.0509 0.1617 1.0000 6.500 0.9527 0.01292 0.00595 -0.0496 0.0561 1.0000 7.000 0.9958 0.01503 0.00790 -0.0477 0.0129 1.0000 7.500 1.0419 0.01655 0.00963 -0.0462 0.0107 1.0000 8.000 1.0827 0.01862 0.01195 -0.0441 0.0096 1.0000 8.500 1.1179 0.02132 0.01490 -0.0412 0.0093 1.0000 9.000 1.1506 0.02449 0.01835 -0.0381 0.0093 1.0000 9.500 1.1825 0.02826 0.02250 -0.0350 0.0096 1.0000 10.000 1.2067 0.03392 0.02879 -0.0314 0.0102 1.0000 10.500 1.2226 0.03811 0.03336 -0.0278 0.0097 1.0000 11.000 1.2004 0.04566 0.04164 -0.0208 0.0105 1.0000 11.500 1.1583 0.05480 0.05137 -0.0177 0.0113 1.0000 12.000 1.1189 0.06547 0.06245 -0.0207 0.0117 1.0000 12.500 1.0793 0.07883 0.07615 -0.0285 0.0119 1.0000 13.000 1.0388 0.09537 0.09298 -0.0399 0.0120 1.0000 13.500 0.9965 0.11608 0.11389 -0.0541 0.0122 1.0000