XFOIL Version 6.94 Calculated polar for: NACA4412c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0429 0.03309 0.02392 -0.0510 0.9999 0.1562 -2.750 -0.0169 0.03296 0.02351 -0.0508 0.9999 0.1658 -2.500 0.0317 0.03149 0.02211 -0.0551 0.9940 0.1828 -2.250 0.0954 0.02961 0.02019 -0.0618 0.9823 0.2073 -2.000 0.1585 0.02737 0.01825 -0.0682 0.9682 0.2400 -1.750 0.2346 0.02449 0.01572 -0.0765 0.9473 0.2986 -1.500 0.2817 0.02107 0.01426 -0.0785 0.9122 0.6638 -1.250 0.3218 0.01873 0.01183 -0.0766 0.8157 1.0001 -1.000 0.3453 0.01914 0.01063 -0.0726 0.5420 1.0001 -0.750 0.3697 0.02029 0.01070 -0.0718 0.4396 1.0001 -0.500 0.3980 0.02098 0.01087 -0.0718 0.4125 1.0001 -0.250 0.4272 0.02151 0.01106 -0.0720 0.3972 1.0001 0.000 0.4565 0.02207 0.01128 -0.0722 0.3868 1.0001 0.250 0.4861 0.02259 0.01155 -0.0724 0.3777 1.0001 0.500 0.5156 0.02326 0.01189 -0.0727 0.3703 1.0001 0.750 0.5456 0.02377 0.01227 -0.0731 0.3635 1.0001 1.000 0.5754 0.02439 0.01269 -0.0735 0.3575 1.0001 1.250 0.6054 0.02514 0.01321 -0.0739 0.3528 1.0001 1.500 0.6356 0.02591 0.01383 -0.0745 0.3491 1.0001 1.750 0.6657 0.02653 0.01443 -0.0749 0.3455 1.0001 2.000 0.6957 0.02722 0.01508 -0.0754 0.3420 1.0001 2.250 0.7256 0.02797 0.01576 -0.0759 0.3383 1.0001 2.500 0.7553 0.02881 0.01649 -0.0764 0.3346 1.0001 2.750 0.7849 0.02993 0.01746 -0.0770 0.3310 1.0001 3.000 0.8143 0.03076 0.01833 -0.0775 0.3281 1.0001 3.250 0.8435 0.03157 0.01922 -0.0779 0.3256 1.0001 3.500 0.8726 0.03247 0.02020 -0.0784 0.3231 1.0001 3.750 0.9016 0.03344 0.02122 -0.0788 0.3208 1.0001 4.000 0.9303 0.03447 0.02230 -0.0793 0.3185 1.0001 4.250 0.9588 0.03555 0.02339 -0.0797 0.3159 1.0001 4.500 0.9869 0.03682 0.02465 -0.0802 0.3132 1.0001 4.750 1.0143 0.03839 0.02624 -0.0806 0.3107 1.0001 5.000 1.0411 0.03950 0.02760 -0.0809 0.3087 1.0001