NACA 2.5411 (naca2411-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 2.5411 (naca2411-il) Reynolds number: 100,000 Max Cl/Cd: 52.92 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca2411-il-100000.txt Download as CSV file: xf-naca2411-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 2.5411
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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4120 0.09791 0.09263 -0.0317 1.0000 0.1404
-9.000 -0.4572 0.09559 0.09052 -0.0371 1.0000 0.1453
-8.750 -0.4522 0.09116 0.08613 -0.0361 1.0000 0.1470
-8.500 -0.4232 0.08852 0.08344 -0.0323 1.0000 0.1514
-8.250 -0.4354 0.08603 0.08104 -0.0325 1.0000 0.1578
-8.000 -0.4738 0.08349 0.07868 -0.0332 1.0000 0.1596
-7.750 -0.5209 0.07922 0.07447 -0.0366 1.0000 0.1614
-7.500 -0.5938 0.05461 0.04901 -0.0433 1.0000 0.0894
-7.250 -0.5886 0.05118 0.04549 -0.0415 1.0000 0.0873
-7.000 -0.5889 0.04724 0.04131 -0.0397 1.0000 0.0867
-6.750 -0.5871 0.04329 0.03702 -0.0377 1.0000 0.0863
-6.500 -0.5816 0.03963 0.03298 -0.0358 1.0000 0.0859
-6.250 -0.5729 0.03620 0.02911 -0.0339 1.0000 0.0858
-6.000 -0.5612 0.03321 0.02560 -0.0321 1.0000 0.0865
-5.750 -0.5468 0.03107 0.02283 -0.0301 1.0000 0.0885
-5.500 -0.5303 0.02905 0.02079 -0.0290 1.0000 0.0921
-5.250 -0.5119 0.02783 0.01943 -0.0277 1.0000 0.0963
-5.000 -0.4926 0.02637 0.01754 -0.0262 1.0000 0.1006
-4.750 -0.4735 0.02510 0.01620 -0.0251 1.0000 0.1065
-4.500 -0.4539 0.02428 0.01519 -0.0238 1.0000 0.1138
-4.250 -0.4345 0.02332 0.01421 -0.0227 1.0000 0.1216
-4.000 -0.4148 0.02266 0.01333 -0.0214 1.0000 0.1314
-3.750 -0.3958 0.02200 0.01277 -0.0203 1.0000 0.1423
-3.500 -0.3768 0.02139 0.01220 -0.0191 1.0000 0.1545
-3.250 -0.3480 0.02096 0.01177 -0.0198 0.9970 0.1732
-3.000 -0.3063 0.02054 0.01149 -0.0230 0.9891 0.2008
-2.750 -0.2646 0.02022 0.01128 -0.0261 0.9813 0.2397
-2.500 -0.2245 0.01981 0.01116 -0.0288 0.9726 0.2889
-2.250 -0.1886 0.01935 0.01103 -0.0308 0.9631 0.3520
-2.000 -0.1496 0.01880 0.01105 -0.0332 0.9553 0.4453
-1.750 -0.1176 0.01807 0.01115 -0.0338 0.9455 0.6067
-1.500 -0.0171 0.01750 0.01149 -0.0457 0.9474 0.9733
-1.250 0.0496 0.01772 0.01140 -0.0537 0.9427 1.0000
-1.000 0.0875 0.01775 0.01124 -0.0561 0.9303 1.0000
-0.750 0.1274 0.01778 0.01110 -0.0587 0.9190 1.0000
-0.500 0.1808 0.01771 0.01089 -0.0636 0.9120 1.0000
-0.250 0.2174 0.01766 0.01073 -0.0653 0.8994 1.0000
0.000 0.2590 0.01756 0.01054 -0.0678 0.8887 1.0000
0.250 0.3138 0.01724 0.01015 -0.0727 0.8817 1.0000
0.500 0.3493 0.01708 0.00994 -0.0738 0.8691 1.0000
0.750 0.3877 0.01685 0.00967 -0.0754 0.8575 1.0000
1.000 0.4354 0.01643 0.00921 -0.0785 0.8492 1.0000
1.250 0.4658 0.01628 0.00903 -0.0785 0.8352 1.0000
1.500 0.4960 0.01615 0.00886 -0.0784 0.8211 1.0000
1.750 0.5259 0.01603 0.00872 -0.0781 0.8068 1.0000
2.000 0.5548 0.01594 0.00860 -0.0778 0.7924 1.0000
2.250 0.5830 0.01588 0.00852 -0.0772 0.7777 1.0000
2.500 0.6102 0.01587 0.00848 -0.0765 0.7628 1.0000
2.750 0.6366 0.01589 0.00847 -0.0756 0.7476 1.0000
3.000 0.6623 0.01595 0.00852 -0.0747 0.7321 1.0000
3.250 0.6874 0.01605 0.00858 -0.0736 0.7164 1.0000
3.500 0.7120 0.01617 0.00868 -0.0724 0.7004 1.0000
3.750 0.7362 0.01631 0.00881 -0.0712 0.6840 1.0000
4.000 0.7600 0.01647 0.00894 -0.0699 0.6673 1.0000
4.250 0.7836 0.01664 0.00909 -0.0686 0.6503 1.0000
4.500 0.8071 0.01681 0.00923 -0.0673 0.6328 1.0000
4.750 0.8308 0.01696 0.00936 -0.0659 0.6146 1.0000
5.000 0.8542 0.01713 0.00948 -0.0645 0.5956 1.0000
5.250 0.8739 0.01734 0.00972 -0.0625 0.5745 1.0000
5.500 0.8952 0.01752 0.00988 -0.0608 0.5532 1.0000
5.750 0.9163 0.01770 0.01001 -0.0590 0.5312 1.0000
6.000 0.9348 0.01791 0.01023 -0.0568 0.5070 1.0000
6.250 0.9538 0.01814 0.01042 -0.0547 0.4826 1.0000
6.500 0.9712 0.01837 0.01065 -0.0524 0.4561 1.0000
6.750 0.9869 0.01865 0.01093 -0.0499 0.4268 1.0000
7.000 1.0011 0.01896 0.01120 -0.0471 0.3939 1.0000
7.250 1.0128 0.01939 0.01154 -0.0440 0.3544 1.0000
7.500 1.0215 0.02004 0.01202 -0.0406 0.3076 1.0000
7.750 1.0264 0.02108 0.01271 -0.0367 0.2561 1.0000
8.000 1.0297 0.02247 0.01371 -0.0329 0.2101 1.0000
8.250 1.0350 0.02393 0.01489 -0.0294 0.1751 1.0000
8.500 1.0421 0.02543 0.01612 -0.0264 0.1518 1.0000
8.750 1.0530 0.02685 0.01745 -0.0239 0.1336 1.0000
9.000 1.0661 0.02831 0.01883 -0.0218 0.1198 1.0000
9.250 1.0821 0.02990 0.02029 -0.0203 0.1083 1.0000
9.500 1.0990 0.03131 0.02171 -0.0188 0.0992 1.0000
9.750 1.1193 0.03309 0.02357 -0.0179 0.0912 1.0000
10.250 1.1653 0.03715 0.02777 -0.0170 0.0800 1.0000
10.500 1.1828 0.03892 0.02967 -0.0158 0.0758 1.0000
10.750 1.2093 0.04193 0.03264 -0.0164 0.0717 1.0000
11.000 1.2182 0.04413 0.03526 -0.0139 0.0701 1.0000
11.250 1.2246 0.04667 0.03817 -0.0114 0.0686 1.0000
11.500 1.2276 0.04919 0.04102 -0.0087 0.0671 1.0000
11.750 1.2282 0.05155 0.04361 -0.0058 0.0657 1.0000
12.000 1.2305 0.05397 0.04621 -0.0034 0.0643 1.0000
12.250 1.2366 0.05655 0.04887 -0.0018 0.0627 1.0000
12.500 1.2333 0.05987 0.05238 0.0004 0.0621 1.0000
12.750 1.2215 0.06325 0.05602 0.0032 0.0619 1.0000
13.000 1.2037 0.06653 0.05959 0.0058 0.0620 1.0000
13.250 1.1805 0.07025 0.06360 0.0078 0.0623 1.0000
13.500 1.1408 0.07512 0.06884 0.0087 0.0631 1.0000
13.750 1.0685 0.08462 0.07884 0.0054 0.0654 1.0000
14.000 1.0112 0.09602 0.09051 -0.0010 0.0678 1.0000
14.250 0.9683 0.10766 0.10226 -0.0080 0.0699 1.0000
14.500 0.9420 0.11779 0.11243 -0.0135 0.0713 1.0000
14.750 0.9283 0.12622 0.12087 -0.0172 0.0721 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 2.5411 (naca2411-il)