NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il) Reynolds number: 100,000 Max Cl/Cd: 39.16 at α=0.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc21006-il-100000.txt Download as CSV file: xf-sc21006-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-1006 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.4124 0.10007 0.09526 -0.0238 0.9999 0.0929
-11.500 -0.4216 0.09723 0.09249 -0.0249 0.9999 0.0963
-11.250 -0.5066 0.10306 0.09799 -0.0220 0.9999 0.0886
-11.000 -0.5054 0.10023 0.09519 -0.0216 0.9999 0.0920
-10.750 -0.5129 0.09782 0.09287 -0.0220 0.9999 0.0954
-10.500 -0.5329 0.09627 0.09146 -0.0233 0.9999 0.0972
-10.250 -0.5454 0.09215 0.08745 -0.0361 0.9999 0.0980
-10.000 -0.5267 0.08893 0.08420 -0.0209 0.9999 0.1020
-9.750 -0.5247 0.08598 0.08125 -0.0209 0.9999 0.1064
-9.500 -0.5288 0.07889 0.07420 -0.0423 0.9999 0.1122
-9.250 -0.5224 0.07735 0.07274 -0.0296 0.9999 0.1148
-9.000 -0.5080 0.07023 0.06546 -0.0470 0.9999 0.1261
-8.750 -0.5026 0.06822 0.06357 -0.0383 0.9999 0.1287
-8.500 -0.4855 0.06302 0.05827 -0.0453 0.9999 0.1416
-8.250 -0.4705 0.05960 0.05482 -0.0465 0.9999 0.1562
-8.000 -0.4538 0.05634 0.05154 -0.0475 0.9999 0.1715
-7.750 -0.3477 0.03600 0.02911 -0.0793 0.9999 0.0828
-7.500 -0.2995 0.02964 0.02187 -0.0849 0.9999 0.0677
-7.250 -0.2581 0.02587 0.01745 -0.0882 0.9999 0.0647
-7.000 -0.2212 0.02351 0.01457 -0.0901 0.9999 0.0672
-6.750 -0.1860 0.02160 0.01221 -0.0915 0.9999 0.0723
-6.500 -0.1526 0.01983 0.01041 -0.0927 0.9999 0.0781
-6.250 -0.1181 0.01847 0.00905 -0.0941 0.9999 0.0923
-6.000 -0.0786 0.01699 0.00770 -0.0966 0.9999 0.1238
-5.750 -0.0225 0.01437 0.00703 -0.1038 0.9999 0.5332
-5.500 -0.0003 0.01448 0.00736 -0.1018 0.9999 0.6307
-5.250 0.0191 0.01471 0.00763 -0.0991 0.9999 0.6870
-5.000 0.0386 0.01488 0.00779 -0.0966 0.9999 0.7277
-4.750 0.0567 0.01499 0.00788 -0.0938 0.9999 0.7597
-4.500 0.0741 0.01504 0.00789 -0.0909 0.9999 0.7916
-4.250 0.0912 0.01504 0.00787 -0.0880 0.9999 0.8256
-4.000 0.1051 0.01488 0.00771 -0.0844 0.9999 0.8576
-3.750 0.1207 0.01460 0.00741 -0.0815 0.9999 0.8918
-3.500 0.1337 0.01410 0.00692 -0.0782 0.9999 0.9287
-3.250 0.1391 0.01355 0.00642 -0.0740 0.9999 1.0001
-3.000 0.1798 0.01373 0.00648 -0.0778 0.9999 1.0001
-2.750 0.2178 0.01393 0.00661 -0.0808 0.9999 1.0001
-2.500 0.2538 0.01415 0.00678 -0.0834 0.9999 1.0001
-2.250 0.2882 0.01439 0.00701 -0.0855 0.9999 1.0001
-2.000 0.3212 0.01464 0.00727 -0.0873 0.9999 1.0001
-1.750 0.3531 0.01491 0.00756 -0.0887 0.9999 1.0001
-1.500 0.3841 0.01520 0.00791 -0.0900 0.9999 1.0001
-1.250 0.4142 0.01552 0.00833 -0.0910 0.9999 1.0001
-1.000 0.4435 0.01585 0.00877 -0.0919 0.9999 1.0001
-0.750 0.4722 0.01621 0.00926 -0.0926 0.9999 1.0001
-0.500 0.5003 0.01660 0.00981 -0.0932 0.9999 1.0001
-0.250 0.5278 0.01702 0.01042 -0.0937 0.9999 1.0001
0.000 0.5548 0.01748 0.01116 -0.0941 0.9999 1.0001
0.250 0.7022 0.01793 0.00809 -0.1089 0.0966 1.0001
0.500 0.7269 0.01942 0.00957 -0.1079 0.0799 1.0001
0.750 0.7534 0.02142 0.01139 -0.1074 0.0719 1.0001
1.000 0.7832 0.02288 0.01297 -0.1072 0.0641 1.0001
1.250 0.8143 0.02496 0.01509 -0.1072 0.0607 1.0001
1.500 0.8463 0.02751 0.01783 -0.1071 0.0600 1.0001
1.750 0.8779 0.03022 0.02111 -0.1062 0.0627 1.0001
2.000 0.9070 0.03363 0.02514 -0.1050 0.0661 1.0001
2.250 0.9330 0.03704 0.02902 -0.1037 0.0674 1.0001
3.000 1.0179 0.05594 0.05094 -0.0974 0.1846 1.0001
3.250 1.0331 0.06014 0.05506 -0.0961 0.1682 1.0001
3.500 1.0375 0.06336 0.05866 -0.0943 0.1476 1.0001
5.000 0.9745 0.08317 0.07940 -0.0733 0.1114 1.0001
5.250 0.9456 0.08769 0.08416 -0.0722 0.1107 1.0001
5.500 0.9173 0.09214 0.08871 -0.0710 0.1104 1.0001
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Polar data table (+)
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