XFOIL Version 6.96 Calculated polar for: S1221 w/ 4 deg flap 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.500 -0.2854 0.10201 0.09672 -0.0057 1.0000 0.2589 -5.250 -0.2604 0.09631 0.09103 -0.0034 1.0000 0.2633 -5.000 -0.2612 0.09443 0.08927 -0.0017 1.0000 0.2694 -4.750 -0.2857 0.09550 0.09058 -0.0038 1.0000 0.2758 -4.500 -0.2748 0.09128 0.08644 -0.0004 1.0000 0.2784 -4.250 -0.2732 0.08908 0.08437 0.0016 1.0000 0.2822 -4.000 -0.2810 0.08803 0.08349 0.0027 1.0000 0.2870 -3.750 -0.1084 0.04699 0.04103 -0.0922 1.0000 0.1176 -3.500 -0.0247 0.03944 0.03282 -0.1099 1.0000 0.1150 -3.250 0.0601 0.03447 0.02675 -0.1252 1.0000 0.1182 -3.000 0.1008 0.03331 0.02558 -0.1305 0.9957 0.1228 -2.750 0.1949 0.03092 0.02287 -0.1429 0.9770 0.1473 -2.500 0.2807 0.03078 0.02395 -0.1515 0.9444 0.3128 -2.250 0.2969 0.03316 0.02677 -0.1432 0.9017 0.4174 -2.000 0.3105 0.03389 0.02763 -0.1352 0.8683 0.4712 -1.750 0.3442 0.03317 0.02676 -0.1323 0.8440 0.5194 -1.500 0.3905 0.03181 0.02512 -0.1339 0.8160 0.5514 -1.250 0.4417 0.03027 0.02329 -0.1360 0.7930 0.5738 -1.000 0.5116 0.02862 0.02118 -0.1424 0.7695 0.5947 -0.750 0.5773 0.02770 0.01973 -0.1490 0.7434 0.6106 -0.500 0.6343 0.02745 0.01903 -0.1547 0.7184 0.6238 -0.250 0.6894 0.02752 0.01868 -0.1606 0.6966 0.6363 0.000 0.7259 0.02779 0.01876 -0.1624 0.6785 0.6470 0.250 0.7657 0.02811 0.01888 -0.1653 0.6624 0.6574 0.500 0.8029 0.02856 0.01923 -0.1680 0.6480 0.6704 0.750 0.8315 0.02904 0.01965 -0.1685 0.6359 0.6827 1.000 0.8653 0.02950 0.01999 -0.1699 0.6251 0.6963 1.250 0.8951 0.03007 0.02053 -0.1708 0.6144 0.7114 1.500 0.9240 0.03071 0.02115 -0.1714 0.6057 0.7294 1.750 0.9523 0.03135 0.02178 -0.1720 0.5970 0.7498 2.000 0.9804 0.03206 0.02250 -0.1726 0.5892 0.7731 2.250 1.0040 0.03275 0.02326 -0.1724 0.5812 0.8006 2.500 1.0278 0.03312 0.02358 -0.1715 0.5754 0.8346 2.750 1.0363 0.03395 0.02468 -0.1690 0.5690 0.8727 3.000 1.0456 0.03421 0.02515 -0.1661 0.5634 1.0000 3.250 1.1102 0.03606 0.02681 -0.1758 0.5550 1.0000 3.500 1.1480 0.03812 0.02887 -0.1801 0.5475 1.0000 3.750 1.1873 0.03952 0.03012 -0.1830 0.5418 1.0000 4.000 1.2052 0.04196 0.03269 -0.1834 0.5359 1.0000 4.250 1.2241 0.04422 0.03502 -0.1836 0.5298 1.0000 4.500 1.2528 0.04579 0.03654 -0.1843 0.5249 1.0000 4.750 1.2618 0.04894 0.03988 -0.1836 0.5202 1.0000 5.000 1.2589 0.05302 0.04418 -0.1823 0.5149 1.0000 5.250 1.2734 0.05572 0.04694 -0.1820 0.5098 1.0000 5.500 1.3032 0.05734 0.04852 -0.1823 0.5053 1.0000 5.750 1.2447 0.06675 0.05827 -0.1793 0.5027 1.0000 6.000 1.1709 0.07805 0.06973 -0.1775 0.5036 1.0000 6.250 1.1371 0.08701 0.07875 -0.1791 0.5065 1.0000