XFOIL Version 6.96 Calculated polar for: NASA SC(2)-1010 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.4713 0.09711 0.09177 -0.0323 0.9999 0.1508 -11.000 -0.5097 0.09510 0.08994 -0.0333 0.9999 0.1569 -10.750 -0.5475 0.09193 0.08697 -0.0333 0.9999 0.1586 -10.500 -0.5041 0.08896 0.08389 -0.0292 0.9999 0.1624 -10.250 -0.5098 0.08690 0.08189 -0.0268 0.9999 0.1667 -10.000 -0.5999 0.05598 0.05009 -0.0640 0.9999 0.0964 -9.750 -0.5781 0.04536 0.03850 -0.0718 0.9999 0.0826 -9.500 -0.5525 0.04074 0.03342 -0.0751 0.9999 0.0825 -9.250 -0.5224 0.03634 0.02847 -0.0787 0.9999 0.0838 -9.000 -0.4926 0.03288 0.02464 -0.0811 0.9999 0.0849 -8.750 -0.4633 0.03048 0.02201 -0.0826 0.9999 0.0869 -8.500 -0.4348 0.02894 0.02029 -0.0837 0.9999 0.0914 -8.250 -0.4025 0.02734 0.01822 -0.0851 0.9999 0.0960 -8.000 -0.3713 0.02541 0.01617 -0.0865 0.9999 0.1006 -7.750 -0.3418 0.02442 0.01509 -0.0873 0.9999 0.1083 -7.500 -0.3109 0.02310 0.01371 -0.0882 0.9999 0.1155 -7.250 -0.2807 0.02236 0.01285 -0.0890 0.9999 0.1269 -7.000 -0.2495 0.02130 0.01195 -0.0900 0.9999 0.1412 -6.750 -0.2157 0.02028 0.01119 -0.0918 0.9999 0.1656 -6.500 -0.1747 0.01905 0.01040 -0.0954 0.9999 0.2249 -6.250 -0.1267 0.01775 0.01035 -0.1007 0.9999 0.4469 -6.000 -0.1014 0.01802 0.01078 -0.1001 0.9999 0.5244 -5.750 -0.0787 0.01843 0.01122 -0.0988 0.9999 0.5714 -5.500 -0.0590 0.01887 0.01171 -0.0967 0.9999 0.6031 -5.250 -0.0407 0.01932 0.01221 -0.0943 0.9999 0.6278 -5.000 -0.0228 0.01980 0.01271 -0.0919 0.9999 0.6511 -4.750 -0.0027 0.02022 0.01311 -0.0901 0.9999 0.6732 -4.500 0.0149 0.02063 0.01354 -0.0878 0.9999 0.6912 -4.250 0.0321 0.02104 0.01396 -0.0853 0.9999 0.7087 -4.000 0.0471 0.02146 0.01442 -0.0823 0.9999 0.7266 -3.750 0.0625 0.02184 0.01482 -0.0795 0.9999 0.7445 -3.500 0.0800 0.02215 0.01515 -0.0773 0.9999 0.7612 -3.250 0.0997 0.02240 0.01540 -0.0759 0.9999 0.7766 -3.000 0.1204 0.02264 0.01564 -0.0747 0.9999 0.7916 -2.750 0.1418 0.02285 0.01586 -0.0738 0.9999 0.8061 -2.500 0.1641 0.02306 0.01609 -0.0732 0.9999 0.8205 -2.250 0.1864 0.02327 0.01631 -0.0726 0.9999 0.8352 -2.000 0.2088 0.02346 0.01654 -0.0721 0.9999 0.8501 -1.750 0.2313 0.02364 0.01676 -0.0717 0.9999 0.8649 -1.500 0.2538 0.02383 0.01699 -0.0714 0.9999 0.8801 -1.250 0.2760 0.02400 0.01722 -0.0710 0.9999 0.8960 -1.000 0.2979 0.02418 0.01746 -0.0707 0.9999 0.9130 -0.750 0.3195 0.02437 0.01772 -0.0704 0.9999 0.9320 -0.500 0.3537 0.02438 0.01784 -0.0725 0.9907 0.9533 -0.250 0.3923 0.02432 0.01789 -0.0754 0.9737 0.9971 0.000 0.4571 0.02464 0.01832 -0.0832 0.9587 1.0001 0.250 0.5286 0.02399 0.01781 -0.0905 0.9320 1.0001 0.500 0.5976 0.02274 0.01675 -0.0965 0.9065 1.0001 0.750 0.6684 0.02050 0.01471 -0.1011 0.8784 1.0001 1.000 0.7290 0.01798 0.01242 -0.1034 0.8474 1.0001 1.250 0.7741 0.01588 0.01049 -0.1028 0.7953 1.0001 1.500 0.8002 0.01669 0.00871 -0.0979 0.3093 1.0001 1.750 0.8139 0.01949 0.01024 -0.0961 0.1633 1.0001 2.000 0.8387 0.02110 0.01158 -0.0958 0.1358 1.0001 2.250 0.8669 0.02257 0.01290 -0.0962 0.1198 1.0001 2.500 0.8983 0.02436 0.01444 -0.0971 0.1088 1.0001 2.750 0.9308 0.02559 0.01579 -0.0979 0.0998 1.0001 3.000 0.9659 0.02770 0.01778 -0.0994 0.0934 1.0001 3.250 0.9991 0.02929 0.01957 -0.1001 0.0879 1.0001 3.500 1.0309 0.03115 0.02143 -0.1010 0.0827 1.0001 3.750 1.0640 0.03434 0.02477 -0.1020 0.0802 1.0001 4.000 1.0918 0.03648 0.02737 -0.1015 0.0784 1.0001 4.250 1.1174 0.03886 0.03019 -0.1009 0.0758 1.0001 4.500 1.1411 0.04196 0.03375 -0.1000 0.0749 1.0001 4.750 1.1611 0.04581 0.03810 -0.0986 0.0758 1.0001 5.000 1.1782 0.05021 0.04293 -0.0971 0.0776 1.0001 5.250 1.1934 0.05513 0.04815 -0.0958 0.0794 1.0001 5.500 1.1126 0.05257 0.04704 -0.0776 0.0917 1.0001 6.000 1.1807 0.08939 0.08499 -0.0885 0.1699 1.0001 6.250 1.1011 0.09196 0.08814 -0.0842 0.1664 1.0001 6.500 1.0620 0.09754 0.09383 -0.0842 0.1642 1.0001