USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 100,000 Max Cl/Cd: 60.9 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa41-il-100000-n5.txt Download as CSV file: xf-usa41-il-100000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3436 0.10532 0.10043 -0.0295 1.0000 0.0388
-8.250 -0.3447 0.10306 0.09825 -0.0301 1.0000 0.0389
-8.000 -0.3462 0.10074 0.09601 -0.0302 1.0000 0.0389
-7.750 -0.3463 0.09810 0.09345 -0.0312 1.0000 0.0390
-7.500 -0.3441 0.09492 0.09037 -0.0320 1.0000 0.0391
-7.250 -0.3291 0.08995 0.08532 -0.0255 1.0000 0.0435
-7.000 -0.3307 0.08790 0.08336 -0.0249 1.0000 0.0458
-6.750 -0.3319 0.08571 0.08124 -0.0255 1.0000 0.0479
-6.500 -0.3335 0.08416 0.07979 -0.0319 1.0000 0.0505
-6.250 -0.3269 0.08142 0.07706 -0.0370 1.0000 0.0510
-6.000 -0.3301 0.07748 0.07321 -0.0328 1.0000 0.0520
-5.750 -0.3290 0.07503 0.07081 -0.0293 1.0000 0.0534
-5.500 -0.3241 0.07255 0.06835 -0.0288 0.9999 0.0550
-5.250 -0.2666 0.06234 0.05785 -0.0478 0.9919 0.0385
-5.000 -0.2302 0.06019 0.05563 -0.0558 0.9852 0.0649
-4.500 -0.1484 0.04361 0.03840 -0.0718 0.9729 0.0385
-4.250 -0.1144 0.03981 0.03441 -0.0765 0.9670 0.0416
-4.000 -0.0781 0.03575 0.03003 -0.0808 0.9602 0.0437
-3.750 -0.0366 0.02992 0.02352 -0.0858 0.9549 0.0437
-3.500 -0.0003 0.02674 0.01975 -0.0886 0.9485 0.0474
-3.250 0.0363 0.02383 0.01616 -0.0908 0.9424 0.0484
-3.000 0.0756 0.02170 0.01344 -0.0931 0.9378 0.0493
-2.750 0.1082 0.02044 0.01177 -0.0939 0.9295 0.0513
-2.500 0.1459 0.01906 0.01022 -0.0959 0.9242 0.0539
-2.250 0.1774 0.01812 0.00910 -0.0964 0.9149 0.0545
-2.000 0.2132 0.01728 0.00810 -0.0977 0.9076 0.0554
-1.750 0.2457 0.01662 0.00735 -0.0983 0.8978 0.0567
-1.500 0.2773 0.01608 0.00673 -0.0988 0.8872 0.0586
-1.250 0.3096 0.01562 0.00616 -0.0993 0.8764 0.0612
-1.000 0.3419 0.01521 0.00568 -0.0997 0.8649 0.0652
-0.750 0.3724 0.01482 0.00530 -0.0999 0.8520 0.0748
-0.500 0.4011 0.01443 0.00517 -0.0998 0.8373 0.1282
-0.250 0.4298 0.01438 0.00510 -0.0995 0.8212 0.1918
0.000 0.4586 0.01422 0.00492 -0.0994 0.8044 0.2239
0.250 0.4876 0.01407 0.00470 -0.0993 0.7870 0.2404
0.500 0.5168 0.01385 0.00451 -0.0992 0.7692 0.2716
0.750 0.5442 0.01290 0.00451 -0.0991 0.7517 0.5893
1.250 0.6103 0.01234 0.00424 -0.1003 0.7116 1.0000
1.500 0.6372 0.01249 0.00420 -0.0997 0.6920 1.0000
1.750 0.6634 0.01266 0.00423 -0.0991 0.6717 1.0000
2.000 0.6896 0.01285 0.00427 -0.0985 0.6524 1.0000
2.250 0.7156 0.01307 0.00435 -0.0979 0.6339 1.0000
2.500 0.7413 0.01330 0.00449 -0.0972 0.6158 1.0000
2.750 0.7670 0.01355 0.00464 -0.0966 0.5984 1.0000
3.000 0.7925 0.01381 0.00482 -0.0960 0.5821 1.0000
3.250 0.8180 0.01409 0.00506 -0.0954 0.5666 1.0000
3.500 0.8433 0.01438 0.00530 -0.0947 0.5516 1.0000
3.750 0.8686 0.01469 0.00557 -0.0941 0.5374 1.0000
4.000 0.8937 0.01500 0.00588 -0.0935 0.5235 1.0000
4.250 0.9187 0.01533 0.00623 -0.0929 0.5098 1.0000
4.500 0.9436 0.01567 0.00659 -0.0922 0.4965 1.0000
4.750 0.9684 0.01602 0.00697 -0.0916 0.4834 1.0000
5.000 0.9930 0.01639 0.00741 -0.0909 0.4705 1.0000
5.500 1.0419 0.01712 0.00831 -0.0895 0.4452 1.0000
5.750 1.0658 0.01750 0.00880 -0.0887 0.4315 1.0000
6.000 1.0895 0.01789 0.00934 -0.0878 0.4175 1.0000
6.250 1.1119 0.01827 0.00979 -0.0867 0.3995 1.0000
6.500 1.1300 0.01860 0.01010 -0.0849 0.3596 1.0000
6.750 1.1465 0.01917 0.01050 -0.0830 0.3068 1.0000
7.000 1.1616 0.02008 0.01112 -0.0811 0.2372 1.0000
7.250 1.1676 0.02224 0.01241 -0.0785 0.1084 1.0000
7.500 1.1735 0.02467 0.01434 -0.0757 0.0473 1.0000
7.750 1.1857 0.02626 0.01595 -0.0735 0.0370 1.0000
8.000 1.1974 0.02776 0.01758 -0.0712 0.0320 1.0000
8.250 1.2076 0.02927 0.01926 -0.0688 0.0285 1.0000
8.500 1.2170 0.03074 0.02093 -0.0663 0.0262 1.0000
8.750 1.2238 0.03233 0.02269 -0.0636 0.0248 1.0000
9.000 1.2278 0.03398 0.02446 -0.0605 0.0235 1.0000
9.250 1.2295 0.03585 0.02641 -0.0575 0.0223 1.0000
9.500 1.2313 0.03815 0.02876 -0.0546 0.0209 1.0000
9.750 1.2411 0.03984 0.03068 -0.0526 0.0199 1.0000
10.000 1.2521 0.04189 0.03290 -0.0507 0.0192 1.0000
10.250 1.2650 0.04420 0.03539 -0.0491 0.0186 1.0000
10.500 1.2781 0.04675 0.03817 -0.0476 0.0180 1.0000
10.750 1.2882 0.04945 0.04114 -0.0459 0.0175 1.0000
11.000 1.2938 0.05220 0.04415 -0.0442 0.0169 1.0000
11.250 1.2945 0.05497 0.04718 -0.0425 0.0161 1.0000
11.500 1.2924 0.05792 0.05037 -0.0410 0.0156 1.0000
11.750 1.2876 0.06111 0.05382 -0.0398 0.0152 1.0000
12.000 1.2808 0.06469 0.05765 -0.0389 0.0149 1.0000
12.250 1.2717 0.06862 0.06183 -0.0386 0.0147 1.0000
12.500 1.2607 0.07304 0.06651 -0.0388 0.0146 1.0000
12.750 1.2477 0.07792 0.07164 -0.0397 0.0146 1.0000
13.000 1.2333 0.08328 0.07726 -0.0413 0.0146 1.0000
13.250 1.2172 0.08920 0.08341 -0.0436 0.0146 1.0000
13.500 1.2000 0.09567 0.09011 -0.0468 0.0147 1.0000
13.750 1.1816 0.10284 0.09750 -0.0507 0.0148 1.0000
14.000 1.1620 0.11077 0.10564 -0.0554 0.0149 1.0000
14.250 1.1411 0.11978 0.11484 -0.0612 0.0152 1.0000
14.500 1.1181 0.13032 0.12555 -0.0682 0.0155 1.0000
14.750 1.0915 0.14349 0.13886 -0.0769 0.0163 1.0000
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