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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0621 0.05538 0.04718 -0.0031 0.4254 0.8266 -2.750 -0.1303 0.05351 0.04569 0.0108 0.4305 0.8093 -2.500 -0.1860 0.05139 0.04383 0.0214 0.4344 0.7905 -2.250 -0.2117 0.04960 0.04206 0.0249 0.4335 0.7627 -2.000 -0.1386 0.05185 0.04344 -0.0003 0.4213 0.6066 -1.750 0.0072 0.05193 0.04131 -0.0385 0.4072 0.3091 -1.500 0.0542 0.05017 0.03882 -0.0431 0.4000 0.2931 -1.250 0.0954 0.04834 0.03670 -0.0460 0.3925 0.2896 -1.000 0.1462 0.04672 0.03455 -0.0508 0.3852 0.2842 -0.750 0.2095 0.04544 0.03254 -0.0580 0.3783 0.2822 -0.500 0.2663 0.04429 0.03103 -0.0637 0.3751 0.2863 -0.250 0.3357 0.04351 0.02984 -0.0721 0.3724 0.3029 0.000 0.4158 0.04284 0.02876 -0.0829 0.3700 0.3304 0.250 0.5245 0.04145 0.02735 -0.0997 0.3682 0.4245 0.500 0.5943 0.04051 0.02718 -0.1084 0.3677 0.5671 0.750 0.7151 0.04078 0.02854 -0.1290 0.3682 1.0001 1.000 0.7545 0.04208 0.02961 -0.1316 0.3695 1.0001 1.250 0.7894 0.04342 0.03077 -0.1333 0.3711 1.0001 1.500 0.8206 0.04483 0.03203 -0.1341 0.3729 1.0001 1.750 0.8487 0.04631 0.03339 -0.1344 0.3744 1.0001 2.000 0.8747 0.04786 0.03484 -0.1343 0.3755 1.0001 2.250 0.8989 0.04947 0.03641 -0.1338 0.3765 1.0001 2.500 0.9210 0.05116 0.03806 -0.1331 0.3772 1.0001 2.750 0.9428 0.05306 0.03991 -0.1323 0.3782 1.0001 3.000 0.9565 0.05405 0.04118 -0.1300 0.3816 1.0001 3.250 0.9679 0.05569 0.04307 -0.1276 0.3864 1.0001 3.500 0.9781 0.05784 0.04541 -0.1254 0.3921 1.0001 3.750 0.9926 0.06023 0.04788 -0.1239 0.3974 1.0001 4.000 1.0136 0.06304 0.05062 -0.1236 0.4016 1.0001 4.250 0.9943 0.06522 0.05338 -0.1178 0.4146 1.0001 4.500 1.0043 0.06843 0.05664 -0.1164 0.4228 1.0001 4.750 0.9629 0.07287 0.06153 -0.1096 0.4407 1.0001 5.000 0.9978 0.07628 0.06484 -0.1123 0.4508 1.0001