NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il) Reynolds number: 50,000 Max Cl/Cd: 34.87 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ls413mod-il-50000.txt Download as CSV file: xf-ls413mod-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY LS(1)-0413MOD 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3632 0.10545 0.09819 -0.0158 1.0000 0.3440
-9.000 -0.3621 0.10275 0.09556 -0.0154 1.0000 0.3568
-8.750 -0.3664 0.10058 0.09348 -0.0148 1.0000 0.3707
-8.500 -0.3743 0.09899 0.09200 -0.0138 1.0000 0.3858
-8.250 -0.3453 0.09406 0.08703 -0.0136 1.0000 0.3918
-8.000 -0.3507 0.09182 0.08490 -0.0125 1.0000 0.4047
-7.250 -0.5517 0.05799 0.05062 -0.0455 1.0000 0.1546
-7.000 -0.5429 0.05408 0.04656 -0.0454 1.0000 0.1500
-6.750 -0.5306 0.04880 0.04031 -0.0481 1.0000 0.1401
-6.500 -0.5135 0.04554 0.03686 -0.0481 1.0000 0.1379
-6.250 -0.4934 0.04239 0.03335 -0.0486 1.0000 0.1354
-6.000 -0.4707 0.03956 0.03010 -0.0490 1.0000 0.1337
-5.750 -0.4465 0.03714 0.02728 -0.0493 1.0000 0.1334
-5.500 -0.4220 0.03521 0.02501 -0.0494 1.0000 0.1357
-5.250 -0.3964 0.03361 0.02300 -0.0494 1.0000 0.1396
-5.000 -0.3723 0.03197 0.02125 -0.0491 1.0000 0.1441
-4.750 -0.3484 0.03070 0.01995 -0.0484 1.0000 0.1496
-4.500 -0.3244 0.02963 0.01873 -0.0474 1.0000 0.1580
-4.250 -0.3020 0.02870 0.01785 -0.0462 1.0000 0.1724
-4.000 -0.2803 0.02760 0.01700 -0.0449 1.0000 0.1928
-3.750 -0.2530 0.02581 0.01584 -0.0454 1.0000 0.2578
-3.500 -0.2556 0.02499 0.01779 -0.0351 1.0000 0.6902
-3.250 -0.2655 0.02642 0.01922 -0.0242 1.0000 0.7564
-3.000 -0.2745 0.02732 0.02007 -0.0141 1.0000 0.8056
-2.750 -0.2808 0.02764 0.02031 -0.0052 1.0000 0.8503
-2.500 -0.2830 0.02745 0.01999 0.0024 1.0000 0.8941
-2.250 -0.0971 0.02805 0.01967 -0.0211 1.0000 1.0000
-2.000 -0.1023 0.02784 0.01939 -0.0182 1.0000 1.0000
-1.750 -0.1067 0.02764 0.01912 -0.0153 1.0000 1.0000
-1.500 -0.1101 0.02745 0.01887 -0.0125 1.0000 1.0000
-1.250 -0.1122 0.02728 0.01862 -0.0100 1.0000 1.0000
-1.000 -0.1124 0.02714 0.01842 -0.0077 1.0000 1.0000
-0.750 -0.1065 0.02715 0.01835 -0.0065 1.0000 1.0000
-0.500 -0.0937 0.02735 0.01846 -0.0065 1.0000 1.0000
-0.250 -0.0758 0.02774 0.01875 -0.0074 1.0000 1.0000
0.000 -0.0344 0.02872 0.01957 -0.0125 0.9926 1.0000
0.250 0.0159 0.02988 0.02058 -0.0189 0.9795 1.0000
0.500 0.0635 0.03098 0.02157 -0.0247 0.9662 1.0000
0.750 0.1105 0.03206 0.02256 -0.0303 0.9522 1.0000
1.000 0.1570 0.03312 0.02354 -0.0355 0.9377 1.0000
1.250 0.2028 0.03412 0.02450 -0.0405 0.9227 1.0000
1.500 0.2471 0.03506 0.02540 -0.0449 0.9068 1.0000
1.750 0.2893 0.03591 0.02625 -0.0488 0.8900 1.0000
2.000 0.3294 0.03670 0.02705 -0.0521 0.8722 1.0000
2.250 0.3688 0.03743 0.02780 -0.0551 0.8539 1.0000
2.500 0.4087 0.03809 0.02850 -0.0579 0.8353 1.0000
2.750 0.4505 0.03861 0.02908 -0.0606 0.8163 1.0000
3.000 0.4950 0.03890 0.02943 -0.0632 0.7972 1.0000
3.250 0.5429 0.03888 0.02950 -0.0658 0.7785 1.0000
3.500 0.5939 0.03846 0.02919 -0.0681 0.7605 1.0000
3.750 0.6201 0.03883 0.02966 -0.0678 0.7368 1.0000
4.000 0.6718 0.03762 0.02858 -0.0689 0.7164 1.0000
4.250 0.7273 0.03559 0.02668 -0.0693 0.6969 1.0000
4.500 0.7660 0.03445 0.02564 -0.0685 0.6727 1.0000
4.750 0.8141 0.03249 0.02379 -0.0680 0.6494 1.0000
5.000 0.8532 0.03113 0.02250 -0.0670 0.6214 1.0000
5.250 0.8930 0.02972 0.02111 -0.0659 0.5898 1.0000
5.500 0.9294 0.02865 0.01994 -0.0646 0.5529 1.0000
5.750 0.9605 0.02825 0.01935 -0.0633 0.5118 1.0000
6.000 0.9894 0.02837 0.01920 -0.0621 0.4696 1.0000
6.250 1.0141 0.02916 0.01977 -0.0611 0.4300 1.0000
6.500 1.0393 0.03022 0.02062 -0.0604 0.3960 1.0000
6.750 1.0659 0.03143 0.02160 -0.0601 0.3680 1.0000
7.000 1.0902 0.03285 0.02302 -0.0598 0.3444 1.0000
7.250 1.1174 0.03425 0.02428 -0.0598 0.3247 1.0000
7.500 1.1435 0.03590 0.02592 -0.0598 0.3085 1.0000
7.750 1.1667 0.03773 0.02789 -0.0596 0.2944 1.0000
8.000 1.1884 0.03970 0.03003 -0.0592 0.2819 1.0000
8.250 1.2127 0.04179 0.03218 -0.0592 0.2716 1.0000
8.500 1.2328 0.04409 0.03471 -0.0587 0.2627 1.0000
8.750 1.2493 0.04672 0.03761 -0.0580 0.2548 1.0000
9.000 1.2676 0.04905 0.04011 -0.0574 0.2463 1.0000
9.250 1.2792 0.05219 0.04354 -0.0563 0.2405 1.0000
9.500 1.2760 0.05600 0.04784 -0.0541 0.2359 1.0000
10.000 1.2870 0.06268 0.05496 -0.0514 0.2245 1.0000
10.250 1.2597 0.06788 0.06058 -0.0484 0.2226 1.0000
10.500 1.2184 0.07397 0.06696 -0.0454 0.2230 1.0000
10.750 1.1660 0.08206 0.07524 -0.0446 0.2251 1.0000
11.000 1.1177 0.09213 0.08540 -0.0471 0.2270 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il)