XFOIL Version 6.96 Calculated polar for: EPPLER E1212 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 -0.2299 0.00932 0.00305 -0.0628 0.5313 0.1853 -4.500 -0.1731 0.00920 0.00289 -0.0624 0.5125 0.1926 -4.000 -0.1156 0.00919 0.00278 -0.0620 0.4946 0.1977 -3.500 -0.0588 0.00907 0.00261 -0.0615 0.4774 0.2045 -3.000 -0.0015 0.00904 0.00253 -0.0612 0.4614 0.2105 -2.500 0.0560 0.00908 0.00247 -0.0608 0.4451 0.2147 -2.000 0.1124 0.00900 0.00235 -0.0603 0.4295 0.2231 -1.500 0.1699 0.00900 0.00231 -0.0600 0.4153 0.2292 -1.000 0.2272 0.00900 0.00227 -0.0597 0.4019 0.2359 -0.500 0.2840 0.00901 0.00226 -0.0593 0.3880 0.2444 0.500 0.3978 0.00910 0.00229 -0.0586 0.3624 0.2627 1.000 0.4545 0.00917 0.00235 -0.0582 0.3507 0.2737 1.500 0.5106 0.00928 0.00243 -0.0578 0.3382 0.2876 2.000 0.5672 0.00931 0.00252 -0.0575 0.3291 0.3080 2.500 0.6226 0.00940 0.00266 -0.0570 0.3181 0.3368 3.000 0.6785 0.00944 0.00280 -0.0566 0.3090 0.3773 3.500 0.7332 0.00949 0.00300 -0.0561 0.2999 0.4382 4.000 0.7878 0.00949 0.00321 -0.0555 0.2914 0.5196 4.500 0.8407 0.00947 0.00348 -0.0547 0.2826 0.6315 5.000 0.8903 0.00929 0.00377 -0.0531 0.2751 0.8010 5.500 0.9639 0.00932 0.00412 -0.0563 0.2664 1.0000 6.000 1.0155 0.00962 0.00437 -0.0552 0.2589 1.0000 6.500 1.0661 0.01000 0.00470 -0.0540 0.2512 1.0000 7.000 1.1180 0.01031 0.00500 -0.0530 0.2452 1.0000 7.500 1.1668 0.01081 0.00542 -0.0517 0.2370 1.0000 8.000 1.2186 0.01111 0.00575 -0.0508 0.2319 1.0000 8.500 1.2672 0.01157 0.00618 -0.0494 0.2253 1.0000 9.000 1.3150 0.01205 0.00666 -0.0480 0.2198 1.0000 9.500 1.3629 0.01248 0.00710 -0.0466 0.2144 1.0000 10.000 1.4051 0.01313 0.00772 -0.0444 0.2079 1.0000