NASA SC(2)-0010 AIRFOIL (sc20010-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0010 AIRFOIL (sc20010-il) Reynolds number: 100,000 Max Cl/Cd: 33.05 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20010-il-100000.txt Download as CSV file: xf-sc20010-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0010 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6893 0.08899 0.08394 -0.0031 1.0000 0.1582
-8.750 -0.6920 0.08471 0.07969 -0.0050 1.0000 0.1643
-8.500 -0.7683 0.07836 0.07302 -0.0127 1.0000 0.1689
-8.250 -0.7914 0.05805 0.05170 -0.0167 1.0000 0.0952
-8.000 -0.7962 0.05003 0.04281 -0.0148 1.0000 0.0822
-7.750 -0.7839 0.04568 0.03821 -0.0139 1.0000 0.0813
-7.500 -0.7702 0.04165 0.03383 -0.0127 1.0000 0.0800
-7.250 -0.7544 0.03800 0.02975 -0.0114 1.0000 0.0795
-7.000 -0.7363 0.03485 0.02610 -0.0101 1.0000 0.0800
-6.750 -0.7161 0.03261 0.02330 -0.0087 1.0000 0.0825
-6.500 -0.6948 0.02993 0.02049 -0.0081 1.0000 0.0862
-6.250 -0.6710 0.02819 0.01863 -0.0075 1.0000 0.0904
-6.000 -0.6471 0.02675 0.01678 -0.0064 1.0000 0.0959
-5.750 -0.6232 0.02489 0.01498 -0.0060 1.0000 0.1026
-5.500 -0.5987 0.02385 0.01368 -0.0051 1.0000 0.1105
-5.250 -0.5746 0.02229 0.01229 -0.0045 1.0000 0.1191
-5.000 -0.5507 0.02110 0.01108 -0.0037 1.0000 0.1289
-4.750 -0.5273 0.02022 0.01014 -0.0028 1.0000 0.1398
-4.500 -0.5054 0.01914 0.00926 -0.0018 1.0000 0.1528
-4.250 -0.4846 0.01819 0.00844 -0.0005 1.0000 0.1685
-4.000 -0.4660 0.01719 0.00772 0.0011 1.0000 0.1910
-3.750 -0.4494 0.01603 0.00699 0.0029 1.0000 0.2381
-3.500 -0.4460 0.01349 0.00646 0.0069 1.0000 0.5323
-3.250 -0.4369 0.01356 0.00715 0.0126 1.0000 0.7305
-3.000 -0.4232 0.01383 0.00745 0.0168 1.0000 0.7866
-2.750 -0.4089 0.01413 0.00775 0.0208 1.0000 0.8249
-2.500 -0.3939 0.01443 0.00802 0.0246 1.0000 0.8572
-2.250 -0.3745 0.01477 0.00831 0.0277 1.0000 0.8861
-2.000 -0.3423 0.01526 0.00870 0.0284 1.0000 0.9152
-1.750 -0.2572 0.01618 0.00942 0.0199 1.0000 0.9465
-1.500 -0.1251 0.01660 0.00959 0.0017 1.0000 0.9719
-1.250 -0.0584 0.01637 0.00926 -0.0063 1.0000 0.9850
-1.000 -0.0074 0.01609 0.00894 -0.0117 1.0000 0.9946
-0.750 0.0251 0.01591 0.00876 -0.0141 1.0000 1.0000
-0.500 0.0194 0.01603 0.00889 -0.0098 1.0000 1.0000
-0.250 0.0098 0.01615 0.00901 -0.0049 1.0000 1.0000
0.000 0.0000 0.01619 0.00905 0.0000 1.0000 1.0000
0.250 -0.0098 0.01615 0.00901 0.0049 1.0000 1.0000
0.500 -0.0194 0.01603 0.00889 0.0098 1.0000 1.0000
0.750 -0.0251 0.01591 0.00876 0.0141 1.0000 1.0000
1.000 0.0074 0.01609 0.00894 0.0117 0.9946 1.0000
1.250 0.0584 0.01636 0.00926 0.0063 0.9850 1.0000
1.500 0.1252 0.01659 0.00958 -0.0017 0.9719 1.0000
1.750 0.2569 0.01618 0.00942 -0.0199 0.9465 1.0000
2.000 0.3423 0.01526 0.00870 -0.0284 0.9152 1.0000
2.250 0.3745 0.01477 0.00831 -0.0276 0.8861 1.0000
2.500 0.3938 0.01443 0.00802 -0.0246 0.8573 1.0000
2.750 0.4088 0.01413 0.00775 -0.0208 0.8249 1.0000
3.000 0.4232 0.01383 0.00745 -0.0168 0.7867 1.0000
3.250 0.4369 0.01356 0.00714 -0.0126 0.7306 1.0000
3.500 0.4459 0.01349 0.00646 -0.0069 0.5328 1.0000
3.750 0.4493 0.01603 0.00699 -0.0029 0.2381 1.0000
4.000 0.4659 0.01718 0.00772 -0.0011 0.1911 1.0000
4.250 0.4845 0.01819 0.00844 0.0005 0.1685 1.0000
4.500 0.5054 0.01914 0.00925 0.0018 0.1528 1.0000
4.750 0.5273 0.02022 0.01014 0.0028 0.1398 1.0000
5.000 0.5507 0.02110 0.01108 0.0037 0.1289 1.0000
5.250 0.5746 0.02229 0.01229 0.0045 0.1191 1.0000
5.500 0.5987 0.02384 0.01368 0.0051 0.1105 1.0000
5.750 0.6232 0.02489 0.01498 0.0060 0.1026 1.0000
6.000 0.6471 0.02675 0.01678 0.0064 0.0959 1.0000
6.250 0.6710 0.02819 0.01863 0.0075 0.0904 1.0000
6.500 0.6948 0.02994 0.02050 0.0081 0.0863 1.0000
6.750 0.7161 0.03261 0.02330 0.0087 0.0825 1.0000
7.000 0.7363 0.03485 0.02610 0.0101 0.0800 1.0000
7.250 0.7545 0.03800 0.02974 0.0114 0.0795 1.0000
7.500 0.7702 0.04165 0.03383 0.0127 0.0800 1.0000
7.750 0.7840 0.04568 0.03821 0.0139 0.0813 1.0000
8.000 0.7963 0.05002 0.04281 0.0148 0.0822 1.0000
8.250 0.7916 0.05806 0.05170 0.0167 0.0952 1.0000
8.500 0.7680 0.07836 0.07303 0.0126 0.1688 1.0000
8.750 0.6921 0.08476 0.07974 0.0048 0.1642 1.0000
9.000 0.6912 0.08884 0.08380 0.0033 0.1581 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0010 AIRFOIL (sc20010-il)