XFOIL Version 6.94 Calculated polar for: FX 61-184 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4978 0.01578 0.00960 -0.1110 0.6870 0.6497 0.500 0.5532 0.01578 0.00959 -0.1110 0.6795 0.6535 1.000 0.6077 0.01563 0.00949 -0.1112 0.6700 0.6575 1.500 0.6689 0.01545 0.00922 -0.1126 0.6625 0.6628 2.000 0.7263 0.01539 0.00917 -0.1138 0.6532 0.6676 2.500 0.7842 0.01509 0.00887 -0.1146 0.6438 0.6713 3.000 0.8391 0.01503 0.00887 -0.1147 0.6340 0.6751 3.500 0.8946 0.01491 0.00879 -0.1150 0.6232 0.6789 4.000 0.9510 0.01483 0.00874 -0.1155 0.6117 0.6831 4.500 1.0082 0.01467 0.00856 -0.1162 0.5978 0.6880 5.000 1.0603 0.01462 0.00860 -0.1160 0.5831 0.6916 5.500 1.1132 0.01466 0.00870 -0.1157 0.5697 0.6951 6.000 1.1645 0.01474 0.00883 -0.1152 0.5540 0.6989 6.500 1.2116 0.01491 0.00911 -0.1140 0.5357 0.7035 7.000 1.2576 0.01518 0.00944 -0.1127 0.5156 0.7085 7.500 1.3006 0.01549 0.00981 -0.1108 0.4945 0.7124 8.000 1.3345 0.01589 0.01025 -0.1072 0.4687 0.7165 8.500 1.3542 0.01646 0.01086 -0.1011 0.4392 0.7210 9.000 1.3646 0.01749 0.01183 -0.0940 0.4042 0.7261 9.500 1.3703 0.01903 0.01332 -0.0870 0.3668 0.7314 10.000 1.3673 0.02124 0.01546 -0.0800 0.3273 0.7368 10.500 1.3611 0.02426 0.01837 -0.0738 0.2861 0.7422 11.000 1.3542 0.02794 0.02190 -0.0688 0.2447 0.7479 11.500 1.3520 0.03180 0.02564 -0.0649 0.2103 0.7534 12.000 1.3490 0.03594 0.02972 -0.0616 0.1793 0.7594 12.500 1.3521 0.04007 0.03385 -0.0592 0.1506 0.7667 13.000 1.3516 0.04496 0.03867 -0.0574 0.1212 0.7744 13.500 1.3477 0.05048 0.04414 -0.0560 0.0937 0.7822 14.000 1.3461 0.05633 0.04996 -0.0554 0.0702 0.7912 14.500 1.3451 0.06245 0.05614 -0.0553 0.0560 0.8011 15.000 1.3452 0.06892 0.06273 -0.0558 0.0474 0.8139 15.500 1.3456 0.07555 0.06956 -0.0567 0.0416 0.8311 16.000 1.3451 0.08207 0.07635 -0.0573 0.0373 0.8644