XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4721 0.06178 0.05458 0.0621 0.9999 0.7386 -2.750 -0.4542 0.06079 0.05340 0.0494 0.9999 0.6500 -2.500 -0.3783 0.06045 0.05242 0.0245 0.9999 0.5122 -2.250 -0.3221 0.05865 0.05007 0.0117 0.9999 0.4311 -2.000 -0.2803 0.05661 0.04765 0.0053 0.9999 0.3901 -1.750 -0.2436 0.05475 0.04552 0.0015 0.9999 0.3715 -1.500 -0.2141 0.05246 0.04330 0.0006 0.9999 0.3659 -1.250 -0.1787 0.05050 0.04131 -0.0015 0.9999 0.3572 -1.000 -0.0036 0.04370 0.03393 -0.0286 0.8450 0.3476 -0.750 -0.0037 0.04296 0.03327 -0.0236 0.6534 0.3477 -0.500 0.2047 0.03878 0.02529 -0.0539 0.5106 0.3936 -0.250 0.2798 0.03792 0.02381 -0.0632 0.4902 0.4387 0.000 0.3622 0.03526 0.02178 -0.0736 0.4731 0.6365 0.250 0.4975 0.03599 0.02197 -0.0967 0.4538 1.0001 0.500 0.5433 0.03683 0.02235 -0.1003 0.4486 1.0001 0.750 0.5869 0.03771 0.02287 -0.1035 0.4449 1.0001 1.000 0.6284 0.03863 0.02352 -0.1062 0.4423 1.0001 1.250 0.6685 0.03961 0.02433 -0.1088 0.4404 1.0001 1.500 0.7069 0.04066 0.02528 -0.1111 0.4393 1.0001 1.750 0.7429 0.04175 0.02633 -0.1130 0.4388 1.0001 2.000 0.7758 0.04288 0.02749 -0.1143 0.4390 1.0001 2.250 0.8057 0.04408 0.02877 -0.1151 0.4399 1.0001 2.500 0.8329 0.04535 0.03016 -0.1155 0.4413 1.0001 2.750 0.8580 0.04676 0.03173 -0.1155 0.4432 1.0001 3.000 0.8813 0.04831 0.03342 -0.1154 0.4452 1.0001 3.250 0.9027 0.04997 0.03522 -0.1150 0.4471 1.0001 3.500 0.9216 0.05180 0.03719 -0.1142 0.4488 1.0001 3.750 0.9390 0.05378 0.03934 -0.1133 0.4506 1.0001 4.000 0.9547 0.05590 0.04159 -0.1123 0.4521 1.0001 4.250 0.9702 0.05822 0.04402 -0.1113 0.4541 1.0001 4.500 0.9871 0.06069 0.04656 -0.1106 0.4562 1.0001 4.750 0.9696 0.06369 0.04999 -0.1061 0.4631 1.0001 5.000 0.9607 0.06738 0.05387 -0.1032 0.4702 1.0001