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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1334 0.02927 0.02044 -0.0214 0.7452 0.3854 -2.750 0.1621 0.03124 0.02296 -0.0509 0.7132 1.0000 -2.500 0.1803 0.03172 0.02320 -0.0500 0.7046 1.0000 -2.250 0.2027 0.03221 0.02356 -0.0508 0.6939 1.0000 -2.000 0.2198 0.03278 0.02391 -0.0495 0.6859 1.0000 -1.750 0.2431 0.03352 0.02457 -0.0509 0.6773 1.0000 -1.500 0.2628 0.03422 0.02514 -0.0507 0.6692 1.0000 -1.250 0.2810 0.03504 0.02582 -0.0501 0.6620 1.0000 -1.000 0.3022 0.03600 0.02673 -0.0512 0.6533 1.0000 -0.750 0.3199 0.03685 0.02746 -0.0504 0.6467 1.0000 -0.500 0.3385 0.03803 0.02857 -0.0507 0.6409 1.0000 -0.250 0.3567 0.03938 0.02990 -0.0519 0.6340 1.0000 0.000 0.3726 0.04045 0.03088 -0.0511 0.6277 1.0000 0.250 0.3874 0.04174 0.03210 -0.0504 0.6221 1.0000 0.500 0.3990 0.04365 0.03404 -0.0517 0.6175 1.0000 0.750 0.4082 0.04545 0.03583 -0.0518 0.6135 1.0000 1.000 0.4175 0.04705 0.03739 -0.0510 0.6094 1.0000 1.250 0.4333 0.04821 0.03843 -0.0489 0.6045 1.0000 1.500 0.4290 0.05069 0.04094 -0.0488 0.6026 1.0000 1.750 0.4268 0.05299 0.04322 -0.0479 0.6030 1.0000 2.000 0.4261 0.05521 0.04540 -0.0466 0.6042 1.0000