XFOIL Version 6.94 Calculated polar for: Eif385mod2 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 0.1118 0.03557 0.02217 -0.0468 0.3665 0.3435 -2.750 0.1403 0.03540 0.02169 -0.0466 0.3512 0.3469 -2.500 0.1695 0.03535 0.02133 -0.0464 0.3417 0.3524 -2.250 0.2015 0.03520 0.02107 -0.0466 0.3350 0.3577 -2.000 0.2344 0.03519 0.02090 -0.0468 0.3293 0.3626 -1.750 0.2680 0.03531 0.02085 -0.0473 0.3246 0.3707 -1.500 0.3014 0.03524 0.02082 -0.0478 0.3208 0.3819 -1.250 0.3371 0.03551 0.02103 -0.0487 0.3177 0.3941 -1.000 0.3727 0.03572 0.02139 -0.0497 0.3155 0.4123 -0.750 0.4055 0.03579 0.02179 -0.0502 0.3138 0.4361 -0.500 0.4344 0.03548 0.02232 -0.0501 0.3119 0.5131 -0.250 0.4778 0.03509 0.02288 -0.0519 0.3091 1.0007 0.000 0.5095 0.03610 0.02373 -0.0522 0.3067 1.0007 0.250 0.5403 0.03719 0.02470 -0.0525 0.3050 1.0007 0.500 0.5706 0.03838 0.02582 -0.0527 0.3045 1.0007 0.750 0.6003 0.03967 0.02708 -0.0530 0.3051 1.0007 1.000 0.6296 0.04106 0.02848 -0.0533 0.3060 1.0007 1.250 0.6583 0.04258 0.03003 -0.0535 0.3073 1.0007 1.500 0.6864 0.04424 0.03172 -0.0538 0.3090 1.0007 1.750 0.7138 0.04607 0.03358 -0.0541 0.3109 1.0007 2.000 0.7405 0.04810 0.03564 -0.0544 0.3129 1.0007 2.250 0.7666 0.05045 0.03798 -0.0548 0.3150 1.0007 2.500 0.7916 0.05175 0.03965 -0.0550 0.3196 1.0007 2.750 0.8116 0.05434 0.04274 -0.0555 0.3276 1.0007 3.000 0.8318 0.05733 0.04592 -0.0559 0.3336 1.0007 3.250 0.8521 0.06046 0.04914 -0.0562 0.3382 1.0007 3.500 0.8749 0.06391 0.05251 -0.0567 0.3414 1.0007 3.750 0.8600 0.06946 0.05899 -0.0572 0.3614 1.0007 4.000 0.8760 0.07351 0.06311 -0.0582 0.3710 1.0007 4.750 0.3235 0.11408 0.10609 -0.0474 0.6577 1.0007 5.000 0.3327 0.11617 0.10807 -0.0468 0.6424 1.0007