XFOIL Version 6.94 Calculated polar for: manu04 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 -0.4995 0.06083 0.05546 0.0638 0.9999 0.6984 -2.750 -0.4924 0.05719 0.05191 0.0675 0.9999 0.7256 -2.500 -0.4864 0.05400 0.04882 0.0697 0.9999 0.7436 -2.250 -0.4761 0.05118 0.04609 0.0695 0.9999 0.7493 -1.750 0.0089 0.04175 0.03035 -0.0420 0.4173 0.2534 -1.500 0.0365 0.04052 0.02878 -0.0419 0.4065 0.2554 -1.250 0.0707 0.03946 0.02729 -0.0431 0.3978 0.2579 -1.000 0.1128 0.03833 0.02570 -0.0457 0.3905 0.2589 -0.750 0.1590 0.03737 0.02431 -0.0491 0.3824 0.2634 -0.500 0.2291 0.03672 0.02280 -0.0575 0.3732 0.2753 -0.250 0.2968 0.03604 0.02188 -0.0655 0.3651 0.3025 0.000 0.4051 0.03449 0.02033 -0.0822 0.3582 0.4164 0.250 0.4795 0.03352 0.02010 -0.0919 0.3550 0.5608 0.500 0.6127 0.03404 0.02115 -0.1150 0.3520 1.0001 0.750 0.6516 0.03492 0.02176 -0.1172 0.3515 1.0001 1.000 0.6869 0.03585 0.02251 -0.1187 0.3515 1.0001 1.250 0.7197 0.03686 0.02338 -0.1196 0.3518 1.0001 1.500 0.7495 0.03797 0.02438 -0.1200 0.3523 1.0001 1.750 0.7770 0.03889 0.02532 -0.1199 0.3534 1.0001 2.000 0.8030 0.03984 0.02636 -0.1195 0.3551 1.0001 2.250 0.8281 0.04091 0.02758 -0.1190 0.3574 1.0001 2.500 0.8524 0.04215 0.02893 -0.1185 0.3593 1.0001 2.750 0.8755 0.04350 0.03038 -0.1178 0.3609 1.0001 3.000 0.8974 0.04495 0.03194 -0.1169 0.3624 1.0001 3.250 0.9179 0.04646 0.03360 -0.1159 0.3640 1.0001 3.500 0.9372 0.04811 0.03536 -0.1146 0.3656 1.0001 3.750 0.9560 0.04989 0.03726 -0.1135 0.3680 1.0001 4.000 0.9755 0.05191 0.03936 -0.1125 0.3709 1.0001 4.250 0.9969 0.05408 0.04158 -0.1119 0.3742 1.0001 4.500 0.9992 0.05575 0.04384 -0.1088 0.3843 1.0001 4.750 1.0132 0.05827 0.04654 -0.1075 0.3909 1.0001 5.000 1.0368 0.06122 0.04945 -0.1077 0.3956 1.0001