NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il) Reynolds number: 50,000 Max Cl/Cd: 29.81 at α=0.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc21006-il-50000-n5.txt Download as CSV file: xf-sc21006-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.5167 0.11436 0.10703 -0.0210 0.9999 0.1164
-11.500 -0.5305 0.11260 0.10542 -0.0233 0.9999 0.1177
-11.250 -0.5272 0.10851 0.10140 -0.0230 0.9999 0.1186
-11.000 -0.5138 0.10010 0.09289 -0.0246 0.9999 0.0679
-10.750 -0.5163 0.09608 0.08893 -0.0251 0.9999 0.0590
-10.500 -0.5205 0.09256 0.08549 -0.0253 0.9999 0.0565
-10.000 -0.5249 0.08133 0.07449 -0.0361 0.9999 0.0497
-9.750 -0.5198 0.07691 0.07009 -0.0383 0.9999 0.0488
-9.500 -0.5130 0.07178 0.06487 -0.0421 0.9999 0.0478
-9.250 -0.5032 0.06618 0.05920 -0.0470 0.9999 0.0469
-9.000 -0.4885 0.06009 0.05292 -0.0529 0.9999 0.0459
-8.750 -0.4668 0.05360 0.04609 -0.0598 0.9999 0.0450
-8.500 -0.4372 0.04719 0.03914 -0.0669 0.9999 0.0444
-8.250 -0.4026 0.04156 0.03285 -0.0730 0.9999 0.0444
-8.000 -0.3662 0.03704 0.02762 -0.0777 0.9999 0.0460
-7.750 -0.3287 0.03353 0.02321 -0.0813 0.9999 0.0500
-7.500 -0.2968 0.03062 0.01993 -0.0833 0.9999 0.0528
-7.250 -0.2663 0.02840 0.01737 -0.0842 0.9999 0.0555
-7.000 -0.2369 0.02654 0.01512 -0.0845 0.9999 0.0594
-6.750 -0.2084 0.02512 0.01353 -0.0849 0.9999 0.0688
-6.500 -0.1801 0.02380 0.01201 -0.0849 0.9999 0.0788
-6.250 -0.1482 0.02245 0.01057 -0.0860 0.9999 0.0941
-6.000 -0.1090 0.02079 0.00919 -0.0894 0.9999 0.1546
-5.750 -0.0668 0.01890 0.00868 -0.0939 0.9999 0.4385
-5.500 -0.0447 0.01871 0.00888 -0.0919 0.9999 0.5734
-5.250 -0.0254 0.01873 0.00895 -0.0893 0.9999 0.6429
-5.000 -0.0073 0.01877 0.00896 -0.0864 0.9999 0.6952
-4.750 0.0112 0.01879 0.00890 -0.0837 0.9999 0.7418
-4.500 0.0289 0.01871 0.00876 -0.0810 0.9999 0.7804
-4.250 0.0476 0.01856 0.00848 -0.0786 0.9999 0.8146
-4.000 0.0672 0.01835 0.00817 -0.0767 0.9999 0.8478
-3.750 0.0860 0.01804 0.00780 -0.0747 0.9999 0.8814
-3.500 0.1041 0.01763 0.00736 -0.0727 0.9999 0.9203
-3.250 0.1175 0.01722 0.00697 -0.0702 0.9999 1.0001
-3.000 0.1530 0.01741 0.00699 -0.0727 0.9999 1.0001
-2.750 0.1870 0.01762 0.00707 -0.0748 0.9999 1.0001
-2.500 0.2197 0.01784 0.00722 -0.0765 0.9999 1.0001
-2.250 0.2515 0.01808 0.00743 -0.0780 0.9999 1.0001
-2.000 0.2824 0.01835 0.00769 -0.0793 0.9999 1.0001
-1.750 0.3126 0.01863 0.00799 -0.0804 0.9999 1.0001
-1.500 0.3421 0.01894 0.00835 -0.0813 0.9999 1.0001
-1.250 0.3710 0.01926 0.00877 -0.0821 0.9999 1.0001
-1.000 0.3994 0.01961 0.00928 -0.0828 0.9999 1.0001
-0.750 0.4273 0.01999 0.00982 -0.0834 0.9999 1.0001
-0.500 0.4548 0.02040 0.01041 -0.0839 0.9999 1.0001
-0.250 0.4818 0.02084 0.01108 -0.0843 0.9999 1.0001
0.000 0.5084 0.02132 0.01190 -0.0846 0.9999 1.0001
0.250 0.5346 0.02185 0.01274 -0.0849 0.9999 1.0001
0.500 0.5603 0.02242 0.01370 -0.0852 0.9999 1.0001
0.750 0.6982 0.02342 0.01115 -0.0972 0.0969 1.0001
1.000 0.7261 0.02496 0.01270 -0.0970 0.0756 1.0001
1.250 0.7560 0.02655 0.01440 -0.0971 0.0642 1.0001
1.500 0.7891 0.02831 0.01638 -0.0973 0.0584 1.0001
1.750 0.8213 0.03043 0.01857 -0.0976 0.0524 1.0001
2.000 0.8525 0.03240 0.02087 -0.0976 0.0469 1.0001
2.250 0.8831 0.03496 0.02381 -0.0974 0.0449 1.0001
2.500 0.9113 0.03781 0.02703 -0.0970 0.0435 1.0001
2.750 0.9372 0.04102 0.03065 -0.0962 0.0427 1.0001
3.000 0.9603 0.04455 0.03465 -0.0952 0.0422 1.0001
3.250 0.9791 0.04871 0.03921 -0.0939 0.0411 1.0001
3.500 0.9966 0.05197 0.04323 -0.0919 0.0396 1.0001
3.750 1.0111 0.05587 0.04775 -0.0899 0.0384 1.0001
4.000 1.0224 0.06020 0.05258 -0.0881 0.0382 1.0001
4.250 1.0310 0.06475 0.05755 -0.0863 0.0386 1.0001
4.500 1.0368 0.06944 0.06261 -0.0846 0.0390 1.0001
4.750 1.0400 0.07421 0.06769 -0.0832 0.0395 1.0001
5.000 1.0412 0.07905 0.07286 -0.0819 0.0400 1.0001
5.250 1.0415 0.08396 0.07793 -0.0808 0.0405 1.0001
5.750 0.9271 0.08520 0.07965 -0.0629 0.0410 1.0001
6.250 0.9884 0.10616 0.10084 -0.0834 0.0467 1.0001
6.500 0.9782 0.11279 0.10747 -0.0876 0.0482 1.0001
6.750 0.9725 0.11950 0.11416 -0.0923 0.0498 1.0001
7.000 0.9711 0.12525 0.11990 -0.0947 0.0518 1.0001
7.500 0.7987 0.12743 0.12223 -0.0799 0.0497 1.0001
7.750 0.7948 0.13203 0.12682 -0.0811 0.0514 1.0001
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-1006 AIRFOIL (sc21006-il)