USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 100,000 Max Cl/Cd: 59.84 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa41-il-100000.txt Download as CSV file: xf-usa41-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: USA 41 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3437 0.09995 0.09508 -0.0259 1.0000 0.0714
-7.750 -0.3564 0.09980 0.09507 -0.0273 1.0000 0.0721
-7.500 -0.3597 0.09889 0.09427 -0.0328 1.0000 0.0726
-7.250 -0.3388 0.09113 0.08644 -0.0239 1.0000 0.0768
-7.000 -0.3377 0.08889 0.08427 -0.0232 1.0000 0.0804
-6.750 -0.3398 0.08725 0.08273 -0.0253 1.0000 0.0844
-6.500 -0.3400 0.08676 0.08230 -0.0345 1.0000 0.0861
-6.250 -0.3389 0.08197 0.07760 -0.0295 1.0000 0.0877
-6.000 -0.3351 0.07903 0.07470 -0.0253 1.0000 0.0904
-5.750 -0.3320 0.07670 0.07241 -0.0248 1.0000 0.0939
-5.500 -0.3132 0.07494 0.07052 -0.0398 1.0000 0.1004
-5.250 -0.3171 0.07096 0.06668 -0.0324 1.0000 0.1018
-5.000 -0.3144 0.06844 0.06422 -0.0286 1.0000 0.1049
-4.750 -0.2864 0.06542 0.06099 -0.0400 1.0000 0.1147
-4.500 -0.2888 0.06244 0.05817 -0.0337 1.0000 0.1171
-4.250 -0.2619 0.05963 0.05516 -0.0408 1.0000 0.1289
-4.000 -0.2600 0.05683 0.05246 -0.0365 1.0000 0.1317
-3.750 -0.2317 0.05386 0.04929 -0.0423 1.0000 0.1433
-3.500 -0.2046 0.05233 0.04749 -0.0460 1.0000 0.1560
-3.250 -0.1972 0.04836 0.04369 -0.0439 1.0000 0.1583
-3.000 -0.1725 0.04612 0.04126 -0.0466 1.0000 0.1716
-2.750 -0.1359 0.04283 0.03788 -0.0505 0.9953 0.1865
-2.500 -0.0893 0.03944 0.03429 -0.0563 0.9877 0.2013
-2.250 -0.0156 0.03240 0.02616 -0.0666 0.9837 0.1305
-2.000 0.0367 0.02666 0.01929 -0.0707 0.9783 0.0984
-1.750 0.0847 0.02395 0.01607 -0.0746 0.9716 0.0938
-1.500 0.1290 0.02202 0.01363 -0.0774 0.9621 0.0923
-1.250 0.1751 0.02083 0.01213 -0.0806 0.9527 0.0967
-1.000 0.2233 0.01969 0.01082 -0.0841 0.9437 0.1018
-0.750 0.2659 0.01881 0.01001 -0.0866 0.9319 0.1083
-0.500 0.3109 0.01803 0.00930 -0.0895 0.9207 0.1233
-0.250 0.3619 0.01723 0.00878 -0.0933 0.9123 0.2228
0.000 0.4091 0.01624 0.00820 -0.0969 0.9015 0.3388
0.250 0.4644 0.01438 0.00778 -0.1017 0.8912 1.0000
0.500 0.5048 0.01420 0.00736 -0.1033 0.8765 1.0000
0.750 0.5413 0.01404 0.00703 -0.1041 0.8604 1.0000
1.000 0.5752 0.01389 0.00672 -0.1043 0.8434 1.0000
1.250 0.6054 0.01381 0.00651 -0.1039 0.8248 1.0000
1.500 0.6328 0.01379 0.00637 -0.1030 0.8039 1.0000
1.750 0.6614 0.01375 0.00621 -0.1023 0.7849 1.0000
2.000 0.6878 0.01382 0.00617 -0.1013 0.7640 1.0000
2.250 0.7144 0.01392 0.00615 -0.1004 0.7438 1.0000
2.500 0.7414 0.01406 0.00618 -0.0996 0.7247 1.0000
2.750 0.7665 0.01429 0.00633 -0.0986 0.7038 1.0000
3.000 0.7925 0.01453 0.00647 -0.0978 0.6849 1.0000
3.250 0.8184 0.01482 0.00665 -0.0970 0.6668 1.0000
3.500 0.8432 0.01516 0.00699 -0.0961 0.6477 1.0000
3.750 0.8683 0.01550 0.00729 -0.0952 0.6299 1.0000
4.000 0.8936 0.01587 0.00761 -0.0945 0.6130 1.0000
4.250 0.9189 0.01626 0.00796 -0.0937 0.5966 1.0000
4.500 0.9442 0.01667 0.00838 -0.0929 0.5808 1.0000
4.750 0.9691 0.01712 0.00884 -0.0921 0.5653 1.0000
5.000 0.9937 0.01760 0.00936 -0.0913 0.5499 1.0000
5.250 1.0185 0.01810 0.00991 -0.0906 0.5353 1.0000
5.500 1.0430 0.01861 0.01052 -0.0898 0.5205 1.0000
5.750 1.0668 0.01909 0.01111 -0.0888 0.5048 1.0000
6.000 1.0885 0.01927 0.01132 -0.0872 0.4816 1.0000
6.250 1.1077 0.01915 0.01114 -0.0850 0.4507 1.0000
6.500 1.1257 0.01910 0.01111 -0.0828 0.4185 1.0000
6.750 1.1431 0.01915 0.01129 -0.0806 0.3831 1.0000
7.000 1.1579 0.01935 0.01149 -0.0779 0.3298 1.0000
7.250 1.1521 0.02207 0.01274 -0.0731 0.1186 1.0000
7.500 1.1562 0.02478 0.01504 -0.0696 0.0801 1.0000
7.750 1.1649 0.02668 0.01697 -0.0667 0.0698 1.0000
8.000 1.1746 0.02839 0.01869 -0.0641 0.0624 1.0000
8.250 1.1860 0.03020 0.02054 -0.0616 0.0580 1.0000
8.500 1.2023 0.03198 0.02241 -0.0597 0.0548 1.0000
8.750 1.2228 0.03408 0.02448 -0.0584 0.0521 1.0000
9.000 1.2501 0.03736 0.02768 -0.0584 0.0488 1.0000
9.250 1.2715 0.03947 0.03010 -0.0571 0.0469 1.0000
9.500 1.2946 0.04251 0.03344 -0.0561 0.0463 1.0000
9.750 1.3129 0.04588 0.03719 -0.0546 0.0463 1.0000
10.000 1.3261 0.04957 0.04132 -0.0528 0.0467 1.0000
10.250 1.3346 0.05360 0.04576 -0.0507 0.0474 1.0000
10.500 1.3381 0.05757 0.05012 -0.0484 0.0477 1.0000
10.750 1.3366 0.06143 0.05435 -0.0458 0.0478 1.0000
11.000 1.3306 0.06532 0.05858 -0.0431 0.0480 1.0000
11.250 1.3210 0.06932 0.06286 -0.0405 0.0483 1.0000
11.500 1.3102 0.07378 0.06753 -0.0381 0.0487 1.0000
11.750 1.3081 0.07709 0.07107 -0.0361 0.0501 1.0000
12.000 1.2786 0.07974 0.07404 -0.0335 0.0508 1.0000
12.250 1.2388 0.08423 0.07888 -0.0335 0.0516 1.0000
12.500 1.1878 0.09236 0.08738 -0.0377 0.0542 1.0000
12.750 1.1545 0.10111 0.09632 -0.0429 0.0564 1.0000
13.000 1.1269 0.11014 0.10542 -0.0486 0.0577 1.0000
13.250 1.1093 0.11855 0.11388 -0.0532 0.0592 1.0000
13.500 0.9049 0.12931 0.12512 -0.0582 0.0576 1.0000
13.750 0.8917 0.13620 0.13200 -0.0614 0.0596 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 41 AIRFOIL (usa41-il)