USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 100,000 Max Cl/Cd: 53.22 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa40-il-100000-n5.txt Download as CSV file: xf-usa40-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 40 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3249 0.09602 0.09061 -0.0440 1.0000 0.0531
-9.250 -0.3289 0.09250 0.08716 -0.0447 1.0000 0.0530
-9.000 -0.3363 0.08891 0.08365 -0.0451 1.0000 0.0530
-8.750 -0.3463 0.08559 0.08041 -0.0451 1.0000 0.0530
-8.500 -0.3593 0.08274 0.07766 -0.0441 1.0000 0.0529
-8.250 -0.3774 0.08041 0.07544 -0.0420 1.0000 0.0527
-8.000 -0.3935 0.07741 0.07254 -0.0416 0.9982 0.0525
-7.750 -0.3800 0.06806 0.06313 -0.0548 0.9870 0.0519
-7.500 -0.3767 0.04981 0.04426 -0.0801 0.9714 0.0508
-7.250 -0.3572 0.04219 0.03599 -0.0889 0.9608 0.0509
-7.000 -0.3283 0.03731 0.03048 -0.0944 0.9535 0.0514
-6.750 -0.2977 0.03370 0.02623 -0.0982 0.9457 0.0526
-6.500 -0.2601 0.03127 0.02342 -0.1018 0.9413 0.0539
-6.250 -0.2309 0.02988 0.02190 -0.1030 0.9321 0.0549
-6.000 -0.1940 0.02822 0.02002 -0.1055 0.9264 0.0559
-5.750 -0.1625 0.02677 0.01835 -0.1066 0.9166 0.0570
-5.500 -0.1249 0.02528 0.01661 -0.1086 0.9089 0.0583
-5.250 -0.0937 0.02408 0.01518 -0.1093 0.8973 0.0598
-5.000 -0.0570 0.02298 0.01398 -0.1109 0.8894 0.0620
-4.750 -0.0285 0.02226 0.01324 -0.1110 0.8766 0.0643
-4.500 0.0033 0.02145 0.01233 -0.1116 0.8661 0.0671
-4.250 0.0349 0.02064 0.01138 -0.1120 0.8551 0.0698
-4.000 0.0635 0.01992 0.01072 -0.1121 0.8423 0.0726
-3.750 0.0944 0.01929 0.01003 -0.1124 0.8307 0.0773
-3.500 0.1239 0.01867 0.00939 -0.1126 0.8179 0.0833
-3.250 0.1524 0.01818 0.00884 -0.1125 0.8038 0.0912
-3.000 0.1814 0.01762 0.00831 -0.1125 0.7904 0.1013
-2.750 0.2109 0.01713 0.00784 -0.1126 0.7770 0.1171
-2.500 0.2394 0.01669 0.00754 -0.1127 0.7627 0.1438
-2.250 0.2674 0.01634 0.00734 -0.1126 0.7477 0.1900
-2.000 0.2955 0.01603 0.00711 -0.1125 0.7328 0.2391
-1.750 0.3234 0.01571 0.00693 -0.1124 0.7181 0.2878
-1.500 0.3511 0.01546 0.00679 -0.1121 0.7035 0.3425
-1.250 0.3786 0.01525 0.00667 -0.1118 0.6885 0.4034
-1.000 0.4049 0.01506 0.00666 -0.1112 0.6730 0.4721
-0.750 0.4297 0.01492 0.00678 -0.1101 0.6579 0.5623
-0.500 0.4540 0.01494 0.00691 -0.1087 0.6431 0.6392
-0.250 0.4786 0.01502 0.00698 -0.1074 0.6284 0.6902
0.000 0.5037 0.01512 0.00698 -0.1062 0.6139 0.7251
0.250 0.5287 0.01522 0.00701 -0.1052 0.5984 0.7534
0.500 0.5533 0.01530 0.00702 -0.1041 0.5833 0.7785
0.750 0.5782 0.01538 0.00701 -0.1032 0.5685 0.8022
1.000 0.6025 0.01544 0.00699 -0.1020 0.5542 0.8242
1.500 0.6490 0.01544 0.00692 -0.0993 0.5260 0.8856
1.750 0.6798 0.01538 0.00685 -0.0996 0.5117 1.0000
2.000 0.7085 0.01564 0.00695 -0.0999 0.4987 1.0000
2.500 0.7647 0.01620 0.00723 -0.1001 0.4744 1.0000
2.750 0.7924 0.01651 0.00740 -0.1002 0.4638 1.0000
3.000 0.8198 0.01683 0.00759 -0.1002 0.4536 1.0000
3.250 0.8471 0.01715 0.00784 -0.1001 0.4437 1.0000
3.500 0.8740 0.01751 0.00807 -0.1000 0.4346 1.0000
3.750 0.9009 0.01785 0.00836 -0.0999 0.4251 1.0000
4.000 0.9273 0.01822 0.00862 -0.0997 0.4164 1.0000
4.250 0.9536 0.01858 0.00896 -0.0995 0.4073 1.0000
4.500 0.9796 0.01897 0.00925 -0.0992 0.3990 1.0000
4.750 1.0052 0.01934 0.00962 -0.0989 0.3896 1.0000
5.000 1.0302 0.01972 0.00990 -0.0985 0.3802 1.0000
5.250 1.0548 0.02008 0.01028 -0.0980 0.3695 1.0000
5.500 1.0789 0.02048 0.01058 -0.0974 0.3602 1.0000
5.750 1.1032 0.02085 0.01100 -0.0969 0.3505 1.0000
6.000 1.1270 0.02126 0.01134 -0.0964 0.3424 1.0000
6.250 1.1504 0.02164 0.01179 -0.0957 0.3323 1.0000
6.500 1.1732 0.02207 0.01216 -0.0950 0.3236 1.0000
6.750 1.1964 0.02248 0.01264 -0.0944 0.3153 1.0000
7.000 1.2190 0.02295 0.01308 -0.0937 0.3089 1.0000
7.250 1.2420 0.02339 0.01362 -0.0930 0.3015 1.0000
7.500 1.2632 0.02386 0.01410 -0.0921 0.2937 1.0000
7.750 1.2843 0.02434 0.01465 -0.0912 0.2845 1.0000
8.000 1.3037 0.02486 0.01515 -0.0901 0.2757 1.0000
8.250 1.3240 0.02538 0.01576 -0.0891 0.2670 1.0000
8.500 1.3426 0.02595 0.01635 -0.0879 0.2597 1.0000
8.750 1.3620 0.02652 0.01703 -0.0869 0.2523 1.0000
9.000 1.3791 0.02715 0.01769 -0.0855 0.2446 1.0000
9.250 1.3953 0.02781 0.01842 -0.0841 0.2352 1.0000
9.500 1.4077 0.02855 0.01917 -0.0821 0.2257 1.0000
9.750 1.4208 0.02934 0.02003 -0.0804 0.2155 1.0000
10.000 1.4327 0.03023 0.02097 -0.0786 0.2055 1.0000
10.250 1.4426 0.03125 0.02201 -0.0767 0.1961 1.0000
10.500 1.4537 0.03228 0.02312 -0.0750 0.1873 1.0000
10.750 1.4627 0.03347 0.02438 -0.0733 0.1793 1.0000
11.000 1.4691 0.03488 0.02583 -0.0716 0.1688 1.0000
11.250 1.4745 0.03646 0.02746 -0.0700 0.1576 1.0000
11.500 1.4778 0.03827 0.02931 -0.0684 0.1466 1.0000
11.750 1.4785 0.04043 0.03150 -0.0670 0.1348 1.0000
12.000 1.4760 0.04299 0.03409 -0.0658 0.1229 1.0000
12.250 1.4708 0.04599 0.03713 -0.0649 0.1124 1.0000
12.500 1.4634 0.04943 0.04061 -0.0643 0.1044 1.0000
12.750 1.4550 0.05319 0.04445 -0.0641 0.0987 1.0000
13.000 1.4451 0.05733 0.04869 -0.0642 0.0945 1.0000
13.250 1.4337 0.06187 0.05334 -0.0647 0.0907 1.0000
13.500 1.4193 0.06698 0.05858 -0.0655 0.0876 1.0000
13.750 1.4038 0.07249 0.06423 -0.0667 0.0852 1.0000
14.000 1.3914 0.07776 0.06967 -0.0680 0.0830 1.0000
14.250 1.3779 0.08332 0.07539 -0.0694 0.0808 1.0000
14.500 1.3637 0.08914 0.08136 -0.0711 0.0786 1.0000
14.750 1.3491 0.09511 0.08747 -0.0729 0.0767 1.0000
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Polar data table (+)
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