USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 500,000 Max Cl/Cd: 104.58 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa41-il-500000-n5.txt Download as CSV file: xf-usa41-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3448 0.08997 0.08777 -0.0261 1.0000 0.0074
-8.000 -0.3504 0.08629 0.08414 -0.0262 1.0000 0.0072
-7.750 -0.3507 0.08461 0.08250 -0.0253 1.0000 0.0075
-7.250 -0.3261 0.07608 0.07400 -0.0347 0.9895 0.0077
-7.000 -0.3047 0.07119 0.06910 -0.0418 0.9803 0.0081
-6.750 -0.2807 0.06512 0.06300 -0.0510 0.9702 0.0090
-6.500 -0.2472 0.05447 0.05227 -0.0672 0.9601 0.0097
-6.250 -0.2081 0.04712 0.04480 -0.0798 0.9529 0.0103
-6.000 -0.1663 0.01687 0.01269 -0.1050 0.9376 0.0138
-5.750 -0.1356 0.01667 0.01241 -0.1058 0.9300 0.0146
-5.500 -0.1036 0.01697 0.01272 -0.1067 0.9232 0.0154
-5.250 -0.0751 0.01622 0.01175 -0.1071 0.9145 0.0167
-5.000 -0.0471 0.01497 0.01012 -0.1072 0.9062 0.0189
-4.750 -0.0187 0.01484 0.00978 -0.1072 0.8973 0.0200
-4.500 0.0066 0.01357 0.00832 -0.1071 0.8881 0.0213
-4.250 0.0334 0.01319 0.00783 -0.1069 0.8790 0.0223
-4.000 0.0599 0.01284 0.00737 -0.1066 0.8690 0.0236
-3.750 0.0862 0.01233 0.00672 -0.1062 0.8584 0.0248
-3.500 0.1126 0.01183 0.00605 -0.1058 0.8470 0.0259
-3.250 0.1391 0.01151 0.00559 -0.1054 0.8345 0.0271
-3.000 0.1654 0.01116 0.00510 -0.1049 0.8203 0.0277
-2.750 0.1915 0.01081 0.00462 -0.1044 0.8041 0.0280
-2.500 0.2176 0.01052 0.00420 -0.1038 0.7861 0.0283
-2.250 0.2434 0.01026 0.00379 -0.1032 0.7671 0.0285
-2.000 0.2693 0.01005 0.00346 -0.1027 0.7473 0.0287
-1.750 0.2944 0.00957 0.00284 -0.1020 0.7269 0.0294
-1.500 0.3199 0.00937 0.00251 -0.1014 0.7068 0.0295
-1.250 0.3455 0.00920 0.00224 -0.1008 0.6854 0.0296
-1.000 0.3710 0.00907 0.00199 -0.1002 0.6641 0.0298
-0.750 0.3967 0.00897 0.00177 -0.0996 0.6425 0.0302
-0.500 0.4223 0.00892 0.00160 -0.0990 0.6206 0.0310
-0.250 0.4481 0.00889 0.00146 -0.0985 0.6000 0.0322
0.000 0.4740 0.00889 0.00137 -0.0980 0.5809 0.0337
0.250 0.5000 0.00893 0.00131 -0.0975 0.5631 0.0352
0.500 0.5261 0.00898 0.00128 -0.0971 0.5475 0.0367
0.750 0.5523 0.00904 0.00126 -0.0967 0.5336 0.0381
1.250 0.6049 0.00910 0.00131 -0.0959 0.5082 0.0596
1.500 0.6307 0.00898 0.00145 -0.0956 0.4973 0.1682
1.750 0.6569 0.00907 0.00154 -0.0952 0.4860 0.1883
2.000 0.6833 0.00915 0.00161 -0.0949 0.4747 0.1962
2.250 0.7097 0.00923 0.00169 -0.0945 0.4640 0.2055
2.500 0.7359 0.00931 0.00178 -0.0942 0.4528 0.2140
2.750 0.7619 0.00940 0.00187 -0.0938 0.4411 0.2243
3.000 0.7877 0.00948 0.00198 -0.0934 0.4280 0.2464
3.500 0.8522 0.00841 0.00240 -0.0959 0.3983 1.0000
4.000 0.9037 0.00876 0.00270 -0.0950 0.3746 1.0000
4.250 0.9294 0.00894 0.00287 -0.0946 0.3652 1.0000
4.500 0.9548 0.00914 0.00305 -0.0941 0.3544 1.0000
4.750 0.9799 0.00937 0.00324 -0.0936 0.3393 1.0000
5.000 1.0036 0.00973 0.00348 -0.0929 0.3092 1.0000
5.250 1.0269 0.01013 0.00375 -0.0921 0.2762 1.0000
5.500 1.0495 0.01061 0.00406 -0.0913 0.2393 1.0000
5.750 1.0708 0.01126 0.00446 -0.0903 0.1903 1.0000
6.000 1.0872 0.01246 0.00520 -0.0886 0.1045 1.0000
6.500 1.1246 0.01434 0.00662 -0.0859 0.0210 1.0000
6.750 1.1471 0.01483 0.00712 -0.0850 0.0163 1.0000
7.000 1.1687 0.01541 0.00775 -0.0839 0.0125 1.0000
7.250 1.1909 0.01588 0.00832 -0.0830 0.0113 1.0000
7.500 1.2124 0.01642 0.00893 -0.0819 0.0103 1.0000
7.750 1.2326 0.01707 0.00963 -0.0807 0.0092 1.0000
8.000 1.2506 0.01792 0.01056 -0.0792 0.0082 1.0000
8.250 1.2707 0.01849 0.01121 -0.0780 0.0076 1.0000
8.500 1.2890 0.01922 0.01203 -0.0766 0.0071 1.0000
8.750 1.3063 0.02000 0.01291 -0.0750 0.0067 1.0000
9.000 1.3229 0.02078 0.01376 -0.0733 0.0062 1.0000
9.250 1.3375 0.02168 0.01473 -0.0714 0.0059 1.0000
9.500 1.3474 0.02290 0.01605 -0.0689 0.0057 1.0000
9.750 1.3541 0.02417 0.01743 -0.0658 0.0054 1.0000
10.000 1.3625 0.02516 0.01853 -0.0629 0.0053 1.0000
10.250 1.3705 0.02619 0.01967 -0.0602 0.0050 1.0000
10.500 1.3779 0.02730 0.02089 -0.0575 0.0048 1.0000
10.750 1.3846 0.02850 0.02220 -0.0550 0.0045 1.0000
11.000 1.3895 0.02990 0.02371 -0.0526 0.0044 1.0000
11.250 1.3953 0.03128 0.02522 -0.0505 0.0042 1.0000
11.500 1.3993 0.03290 0.02695 -0.0485 0.0041 1.0000
11.750 1.4009 0.03486 0.02903 -0.0465 0.0040 1.0000
12.000 1.4045 0.03667 0.03094 -0.0450 0.0039 1.0000
12.250 1.4046 0.03897 0.03336 -0.0436 0.0038 1.0000
12.500 1.4040 0.04147 0.03598 -0.0425 0.0037 1.0000
12.750 1.4008 0.04445 0.03910 -0.0415 0.0037 1.0000
13.000 1.3947 0.04794 0.04276 -0.0408 0.0036 1.0000
13.250 1.3902 0.05142 0.04640 -0.0406 0.0035 1.0000
13.500 1.3858 0.05505 0.05022 -0.0406 0.0035 1.0000
13.750 1.3803 0.05898 0.05433 -0.0411 0.0035 1.0000
14.000 1.3729 0.06336 0.05891 -0.0419 0.0034 1.0000
14.250 1.3641 0.06812 0.06385 -0.0431 0.0034 1.0000
14.500 1.3534 0.07341 0.06933 -0.0447 0.0033 1.0000
14.750 1.3424 0.07896 0.07506 -0.0468 0.0033 1.0000
15.000 1.3293 0.08516 0.08146 -0.0493 0.0033 1.0000
15.250 1.3156 0.09167 0.08814 -0.0522 0.0033 1.0000
15.500 1.3001 0.09887 0.09553 -0.0557 0.0033 1.0000
15.750 1.2841 0.10647 0.10331 -0.0596 0.0033 1.0000
16.000 1.2689 0.11420 0.11119 -0.0637 0.0033 1.0000
16.250 1.2524 0.12260 0.11975 -0.0683 0.0033 1.0000
16.500 1.2351 0.13161 0.12893 -0.0735 0.0033 1.0000
16.750 1.2198 0.14039 0.13784 -0.0787 0.0033 1.0000
17.000 1.2027 0.15020 0.14779 -0.0845 0.0033 1.0000
17.250 1.1861 0.16041 0.15812 -0.0906 0.0034 1.0000
17.500 1.1691 0.17136 0.16918 -0.0971 0.0034 1.0000
17.750 1.1452 0.18589 0.18382 -0.1054 0.0035 1.0000
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