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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1895 0.05993 0.05204 -0.0069 0.3749 0.5385 -2.750 -0.1823 0.05747 0.04956 -0.0026 0.3696 0.5726 -2.500 -0.1758 0.05461 0.04669 0.0017 0.3646 0.6076 -2.250 -0.1660 0.05237 0.04432 0.0044 0.3599 0.6394 -2.000 -0.1456 0.04946 0.04119 0.0049 0.3547 0.6625 -1.750 -0.1163 0.04719 0.03882 0.0026 0.3494 0.6664 -1.250 0.1534 0.04443 0.03237 -0.0576 0.3355 0.2332 -1.000 0.2216 0.04288 0.03018 -0.0658 0.3325 0.2355 -0.750 0.2873 0.04159 0.02831 -0.0734 0.3305 0.2375 -0.500 0.3404 0.04070 0.02733 -0.0785 0.3293 0.2441 -0.250 0.4017 0.04023 0.02650 -0.0852 0.3285 0.2575 0.000 0.4647 0.04016 0.02635 -0.0926 0.3280 0.2839 0.250 0.5559 0.03949 0.02568 -0.1062 0.3279 0.3680 0.500 0.6121 0.03958 0.02609 -0.1123 0.3285 0.4592 0.750 0.6528 0.04010 0.02696 -0.1152 0.3293 0.5244 1.000 0.6893 0.04067 0.02793 -0.1172 0.3299 0.5872 1.250 0.7937 0.04206 0.03022 -0.1347 0.3300 1.0001 1.500 0.8236 0.04347 0.03148 -0.1353 0.3303 1.0001 1.750 0.8508 0.04492 0.03280 -0.1353 0.3305 1.0001 2.000 0.8765 0.04651 0.03427 -0.1350 0.3310 1.0001 2.250 0.8986 0.04719 0.03513 -0.1337 0.3341 1.0001 2.500 0.9173 0.04844 0.03664 -0.1321 0.3397 1.0001 2.750 0.9367 0.05021 0.03851 -0.1308 0.3447 1.0001 3.000 0.9562 0.05218 0.04050 -0.1296 0.3492 1.0001 3.250 0.9776 0.05449 0.04275 -0.1289 0.3528 1.0001 3.750 0.9995 0.05823 0.04711 -0.1243 0.3722 1.0001 4.000 1.0238 0.06145 0.05018 -0.1246 0.3778 1.0001 4.250 1.0176 0.06358 0.05289 -0.1206 0.3984 1.0001 4.500 1.0133 0.06706 0.05675 -0.1181 0.4227 1.0001 5.000 0.4541 0.11294 0.10481 -0.0987 0.8178 1.0001