XFOIL Version 6.94 Calculated polar for: manu4/26 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.2935 0.06683 0.05709 -0.0603 0.3866 0.9047 -2.750 0.2762 0.06536 0.05570 -0.0560 0.3862 0.8817 -2.500 0.2192 0.06412 0.05473 -0.0437 0.3883 0.8569 -2.250 0.1691 0.06224 0.05309 -0.0334 0.3896 0.8326 -2.000 0.0906 0.06022 0.05139 -0.0171 0.3931 0.8151 -1.750 0.0247 0.05793 0.04932 -0.0041 0.3957 0.7969 -1.500 0.0081 0.05648 0.04783 -0.0028 0.3959 0.7596 -1.250 0.1797 0.06262 0.05253 -0.0494 0.3863 0.5329 -1.000 0.2831 0.06252 0.05160 -0.0708 0.3829 0.4003 -0.750 0.3485 0.06162 0.05016 -0.0805 0.3823 0.3515 -0.500 0.4023 0.06080 0.04896 -0.0868 0.3822 0.3311 -0.250 0.4479 0.05993 0.04792 -0.0908 0.3826 0.3251 0.000 0.4947 0.05965 0.04738 -0.0951 0.3833 0.3262 0.250 0.5451 0.05990 0.04724 -0.1002 0.3842 0.3330 0.500 0.5955 0.05984 0.04706 -0.1053 0.3855 0.3424 0.750 0.6376 0.05971 0.04680 -0.1085 0.3882 0.3533 1.000 0.6799 0.05977 0.04699 -0.1121 0.3923 0.3719 1.250 0.7322 0.06016 0.04751 -0.1181 0.3972 0.4193 1.500 0.7938 0.06055 0.04832 -0.1260 0.4026 0.5316 1.750 0.8275 0.06163 0.04984 -0.1281 0.4072 0.6142 2.000 0.9399 0.06476 0.05391 -0.1482 0.4138 1.0001 2.250 0.9276 0.06560 0.05525 -0.1420 0.4224 1.0001 2.500 0.9228 0.06847 0.05833 -0.1380 0.4305 1.0001 2.750 0.9301 0.07179 0.06164 -0.1362 0.4374 1.0001 3.000 0.9581 0.07552 0.06515 -0.1379 0.4424 1.0001 3.250 0.8256 0.08092 0.07153 -0.1166 0.4616 1.0001 3.500 0.7802 0.08699 0.07780 -0.1104 0.4778 1.0001 4.500 0.1207 0.12603 0.11659 -0.0484 0.8809 0.3150 4.750 0.1483 0.12889 0.11927 -0.0525 0.8814 0.3235 5.000 0.1716 0.13018 0.12024 -0.0554 0.8726 0.3314