XFOIL Version 6.94 Calculated polar for: manu4/33 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 -0.2946 0.06896 0.06289 0.0461 0.9999 0.8533 -2.750 -0.3549 0.06641 0.06049 0.0575 0.9999 0.8310 -2.500 -0.4168 0.06355 0.05772 0.0683 0.9999 0.8088 -2.250 -0.4533 0.06088 0.05509 0.0715 0.9999 0.7696 -2.000 -0.1626 0.05754 0.05055 -0.0084 0.7952 0.4546 -1.500 0.0149 0.05119 0.04069 -0.0345 0.5016 0.3716 -1.250 0.0593 0.04983 0.03861 -0.0387 0.4877 0.3578 -1.000 0.1107 0.04867 0.03667 -0.0438 0.4763 0.3515 -0.750 0.1689 0.04761 0.03493 -0.0502 0.4671 0.3533 -0.500 0.2287 0.04653 0.03337 -0.0567 0.4576 0.3653 -0.250 0.3017 0.04593 0.03195 -0.0660 0.4484 0.3794 0.000 0.3710 0.04510 0.03077 -0.0744 0.4396 0.4043 0.250 0.4588 0.04389 0.02942 -0.0865 0.4313 0.4906 0.500 0.6292 0.04210 0.02878 -0.1168 0.4254 1.0001 0.750 0.6807 0.04333 0.02963 -0.1220 0.4247 1.0001 1.000 0.7256 0.04456 0.03057 -0.1258 0.4245 1.0001 1.250 0.7652 0.04582 0.03163 -0.1284 0.4248 1.0001 1.500 0.8009 0.04718 0.03282 -0.1303 0.4254 1.0001 1.750 0.8339 0.04864 0.03416 -0.1317 0.4264 1.0001 2.000 0.8578 0.04978 0.03539 -0.1313 0.4284 1.0001 2.250 0.8784 0.05109 0.03683 -0.1304 0.4312 1.0001 2.500 0.8976 0.05265 0.03852 -0.1295 0.4343 1.0001 2.750 0.9160 0.05443 0.04040 -0.1286 0.4379 1.0001 3.000 0.9328 0.05640 0.04245 -0.1275 0.4416 1.0001 3.250 0.9490 0.05855 0.04465 -0.1264 0.4449 1.0001 3.500 0.9653 0.06090 0.04703 -0.1254 0.4477 1.0001 3.750 0.9838 0.06342 0.04956 -0.1248 0.4501 1.0001 4.000 0.9641 0.06598 0.05252 -0.1191 0.4565 1.0001 4.500 0.9404 0.07344 0.06030 -0.1116 0.4694 1.0001 4.750 0.9316 0.07747 0.06441 -0.1089 0.4764 1.0001 5.000 0.8172 0.08743 0.07466 -0.0987 0.5015 1.0001