XFOIL Version 6.94 Calculated polar for: FAUVEL 14% AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1623 0.03184 0.02339 -0.0581 0.7542 1.0000 -2.750 0.1815 0.03234 0.02366 -0.0579 0.7443 1.0000 -2.500 0.2021 0.03288 0.02402 -0.0583 0.7339 1.0000 -2.250 0.2206 0.03354 0.02446 -0.0578 0.7246 1.0000 -2.000 0.2416 0.03425 0.02504 -0.0585 0.7146 1.0000 -1.750 0.2588 0.03501 0.02562 -0.0574 0.7068 1.0000 -1.500 0.2801 0.03600 0.02653 -0.0588 0.6988 1.0000 -1.250 0.2983 0.03692 0.02734 -0.0587 0.6907 1.0000 -1.000 0.3148 0.03791 0.02818 -0.0577 0.6838 1.0000 -0.750 0.3323 0.03922 0.02947 -0.0590 0.6761 1.0000 -0.500 0.3481 0.04046 0.03063 -0.0589 0.6703 1.0000 -0.250 0.3637 0.04161 0.03167 -0.0577 0.6650 1.0000 0.000 0.3741 0.04337 0.03341 -0.0585 0.6596 1.0000 0.250 0.3816 0.04509 0.03510 -0.0583 0.6545 1.0000 0.500 0.3886 0.04677 0.03674 -0.0576 0.6506 1.0000 0.750 0.3924 0.04865 0.03858 -0.0568 0.6489 1.0000 1.000 0.3930 0.05058 0.04047 -0.0556 0.6476 1.0000 1.250 0.3831 0.05273 0.04259 -0.0537 0.6483 1.0000 1.500 0.3531 0.05512 0.04495 -0.0501 0.6548 1.0000 1.750 0.3432 0.05717 0.04694 -0.0475 0.6609 1.0000 2.750 -0.0651 0.05143 0.04102 0.0115 1.0000 1.0000 3.000 -0.0235 0.05461 0.04408 0.0053 0.9832 1.0000 3.750 0.1193 0.06518 0.05435 -0.0125 0.8645 1.0000 4.000 0.1503 0.06781 0.05691 -0.0144 0.8356 1.0000 4.250 0.1821 0.07092 0.05993 -0.0164 0.8113 1.0000 4.500 0.2064 0.07356 0.06251 -0.0171 0.7899 1.0000 5.000 0.2345 0.07694 0.06575 -0.0155 0.7509 1.0000