XFOIL Version 6.94 Calculated polar for: EPPLER 266 AIRFOIL 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 0.3091 0.01389 0.00850 -0.0518 0.6261 0.8283 0.500 0.3610 0.01403 0.00858 -0.0521 0.6233 0.8351 1.000 0.4122 0.01396 0.00854 -0.0515 0.6204 0.8384 1.500 0.4641 0.01405 0.00865 -0.0512 0.6173 0.8418 2.000 0.5169 0.01408 0.00868 -0.0512 0.6137 0.8454 2.500 0.5713 0.01409 0.00869 -0.0518 0.6104 0.8491 3.000 0.6281 0.01413 0.00869 -0.0528 0.6069 0.8528 3.500 0.6837 0.01444 0.00900 -0.0539 0.6025 0.8553 4.000 0.7320 0.01421 0.00888 -0.0530 0.5980 0.8577 4.500 0.7822 0.01413 0.00888 -0.0524 0.5925 0.8605 5.000 0.8384 0.01395 0.00872 -0.0530 0.5870 0.8630 5.500 0.8945 0.01406 0.00885 -0.0538 0.5812 0.8654 6.000 0.9409 0.01384 0.00879 -0.0526 0.5739 0.8684 6.500 0.9965 0.01348 0.00846 -0.0530 0.5658 0.8710 7.000 1.0435 0.01321 0.00830 -0.0518 0.5548 0.8743 7.500 1.0869 0.01261 0.00773 -0.0495 0.5394 0.8770 8.000 1.1185 0.01227 0.00750 -0.0450 0.5205 0.8804 8.500 1.1414 0.01220 0.00755 -0.0390 0.4988 0.8844 9.000 1.1452 0.01252 0.00781 -0.0295 0.4593 0.8897 9.500 1.1232 0.01384 0.00888 -0.0165 0.4019 0.8965 10.000 1.0904 0.01605 0.01091 -0.0036 0.3495 0.9040 10.500 1.0572 0.01944 0.01413 0.0067 0.3032 0.9122 11.000 1.0342 0.02314 0.01770 0.0142 0.2584 0.9203 11.500 1.0213 0.02679 0.02123 0.0196 0.2198 0.9298 12.000 1.0196 0.03092 0.02520 0.0217 0.1757 0.9392 12.500 1.0353 0.03538 0.02950 0.0199 0.1316 0.9472 13.500 1.0866 0.04577 0.03951 0.0105 0.0468 0.9613 14.000 1.1049 0.05176 0.04536 0.0062 0.0179 0.9670