NASA SC(2)-1010 AIRFOIL (sc21010-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-1010 AIRFOIL (sc21010-il) Reynolds number: 50,000 Max Cl/Cd: 36.06 at α=1.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc21010-il-50000-n5.txt Download as CSV file: xf-sc21010-il-50000-n5.csv |
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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.5031 0.10168 0.09366 -0.0419 0.9999 0.0673
-12.250 -0.5090 0.09717 0.08923 -0.0431 0.9999 0.0669
-12.000 -0.5173 0.09243 0.08457 -0.0444 0.9999 0.0664
-11.750 -0.5296 0.08728 0.07951 -0.0459 0.9999 0.0659
-11.500 -0.5488 0.08139 0.07372 -0.0478 0.9999 0.0652
-11.250 -0.5759 0.07500 0.06742 -0.0506 0.9999 0.0643
-11.000 -0.5996 0.06801 0.06038 -0.0559 0.9999 0.0636
-10.500 -0.6135 0.05625 0.04807 -0.0638 0.9999 0.0636
-10.250 -0.6046 0.05150 0.04293 -0.0670 0.9999 0.0648
-10.000 -0.5884 0.04706 0.03796 -0.0703 0.9999 0.0672
-9.750 -0.5655 0.04271 0.03284 -0.0739 0.9999 0.0697
-9.500 -0.5426 0.03982 0.02969 -0.0752 0.9999 0.0716
-9.250 -0.5195 0.03778 0.02748 -0.0760 0.9999 0.0743
-9.000 -0.4943 0.03586 0.02525 -0.0769 0.9999 0.0788
-8.750 -0.4664 0.03377 0.02261 -0.0780 0.9999 0.0831
-8.500 -0.4428 0.03224 0.02113 -0.0780 0.9999 0.0867
-8.250 -0.4173 0.03100 0.01972 -0.0783 0.9999 0.0932
-8.000 -0.3920 0.02973 0.01828 -0.0782 0.9999 0.0993
-7.750 -0.3670 0.02871 0.01723 -0.0781 0.9999 0.1067
-7.500 -0.3412 0.02773 0.01620 -0.0782 0.9999 0.1165
-7.250 -0.3142 0.02681 0.01519 -0.0785 0.9999 0.1288
-7.000 -0.2860 0.02587 0.01433 -0.0794 0.9999 0.1465
-6.750 -0.2549 0.02488 0.01347 -0.0810 0.9999 0.1745
-6.500 -0.2188 0.02371 0.01270 -0.0842 0.9999 0.2294
-6.250 -0.1814 0.02272 0.01245 -0.0875 0.9999 0.3493
-6.000 -0.1541 0.02261 0.01271 -0.0875 0.9999 0.4489
-5.750 -0.1307 0.02281 0.01302 -0.0864 0.9999 0.5116
-5.500 -0.1088 0.02309 0.01331 -0.0849 0.9999 0.5569
-5.250 -0.0883 0.02341 0.01363 -0.0830 0.9999 0.5919
-5.000 -0.0682 0.02375 0.01395 -0.0811 0.9999 0.6232
-4.750 -0.0522 0.02415 0.01437 -0.0781 0.9999 0.6496
-4.500 -0.0332 0.02448 0.01466 -0.0760 0.9999 0.6754
-4.250 -0.0150 0.02476 0.01491 -0.0738 0.9999 0.6966
-4.000 0.0059 0.02496 0.01507 -0.0724 0.9999 0.7153
-3.750 0.0274 0.02513 0.01519 -0.0713 0.9999 0.7322
-3.500 0.0497 0.02529 0.01531 -0.0704 0.9999 0.7482
-3.250 0.0725 0.02545 0.01543 -0.0698 0.9999 0.7636
-3.000 0.0953 0.02562 0.01558 -0.0692 0.9999 0.7787
-2.750 0.1187 0.02578 0.01573 -0.0688 0.9999 0.7934
-2.500 0.1421 0.02596 0.01590 -0.0684 0.9999 0.8080
-2.250 0.1656 0.02615 0.01610 -0.0682 0.9999 0.8228
-2.000 0.1892 0.02634 0.01632 -0.0680 0.9999 0.8378
-1.750 0.2202 0.02654 0.01654 -0.0693 0.9956 0.8528
-1.500 0.2528 0.02672 0.01676 -0.0708 0.9898 0.8677
-1.250 0.2872 0.02692 0.01703 -0.0727 0.9842 0.8833
-1.000 0.3175 0.02703 0.01722 -0.0738 0.9769 0.8994
-0.750 0.3529 0.02719 0.01748 -0.0758 0.9706 0.9175
-0.500 0.3812 0.02719 0.01763 -0.0765 0.9611 0.9377
-0.250 0.4111 0.02722 0.01781 -0.0776 0.9499 0.9709
0.000 0.4521 0.02753 0.01827 -0.0809 0.9391 1.0001
0.250 0.5004 0.02786 0.01878 -0.0854 0.9284 1.0001
0.500 0.5470 0.02794 0.01905 -0.0891 0.9132 1.0001
0.750 0.5915 0.02755 0.01890 -0.0915 0.8894 1.0001
1.000 0.6499 0.02578 0.01739 -0.0939 0.8504 1.0001
1.250 0.6916 0.02410 0.01594 -0.0932 0.8050 1.0001
1.500 0.7200 0.02311 0.01516 -0.0910 0.7473 1.0001
1.750 0.7796 0.02162 0.01226 -0.0904 0.4342 1.0001
2.000 0.7883 0.02420 0.01343 -0.0878 0.2440 1.0001
2.250 0.8071 0.02618 0.01477 -0.0870 0.1691 1.0001
2.500 0.8309 0.02777 0.01608 -0.0868 0.1378 1.0001
2.750 0.8575 0.02920 0.01744 -0.0868 0.1199 1.0001
3.000 0.8868 0.03060 0.01882 -0.0872 0.1067 1.0001
3.250 0.9196 0.03198 0.02024 -0.0880 0.0968 1.0001
3.500 0.9541 0.03350 0.02185 -0.0889 0.0886 1.0001
3.750 0.9871 0.03515 0.02346 -0.0898 0.0823 1.0001
4.000 1.0216 0.03717 0.02566 -0.0907 0.0774 1.0001
4.250 1.0515 0.03904 0.02774 -0.0911 0.0723 1.0001
4.500 1.0803 0.04120 0.02996 -0.0915 0.0690 1.0001
4.750 1.1077 0.04387 0.03294 -0.0915 0.0665 1.0001
5.000 1.1306 0.04649 0.03606 -0.0907 0.0634 1.0001
5.250 1.1511 0.04908 0.03900 -0.0898 0.0607 1.0001
5.500 1.1700 0.05196 0.04219 -0.0888 0.0592 1.0001
5.750 1.1863 0.05507 0.04559 -0.0877 0.0581 1.0001
6.000 1.2007 0.05833 0.04909 -0.0865 0.0569 1.0001
6.250 1.2056 0.06216 0.05341 -0.0841 0.0560 1.0001
6.500 1.2040 0.06622 0.05805 -0.0811 0.0551 1.0001
6.750 1.1981 0.07040 0.06271 -0.0782 0.0542 1.0001
7.000 1.1882 0.07466 0.06737 -0.0753 0.0537 1.0001
7.250 1.1737 0.07899 0.07203 -0.0724 0.0535 1.0001
7.500 1.1526 0.08321 0.07651 -0.0694 0.0535 1.0001
7.750 1.1294 0.08801 0.08154 -0.0675 0.0537 1.0001
8.000 1.1045 0.09353 0.08728 -0.0673 0.0539 1.0001
8.250 1.0784 0.10005 0.09397 -0.0689 0.0543 1.0001
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-1010 AIRFOIL (sc21010-il)