N-22 (n22-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: N-22 (n22-il) Reynolds number: 50,000 Max Cl/Cd: 36.02 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n22-il-50000-n5.txt Download as CSV file: xf-n22-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: N-22
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3602 0.09712 0.08991 -0.0420 0.9999 0.0862
-10.500 -0.3645 0.09422 0.08709 -0.0415 0.9999 0.0854
-10.250 -0.3745 0.09146 0.08444 -0.0408 0.9999 0.0848
-10.000 -0.3894 0.08891 0.08202 -0.0397 0.9999 0.0845
-9.750 -0.4042 0.08602 0.07925 -0.0392 0.9999 0.0843
-9.500 -0.4181 0.08285 0.07616 -0.0392 0.9999 0.0841
-9.250 -0.4311 0.07941 0.07277 -0.0395 0.9999 0.0840
-9.000 -0.4424 0.07568 0.06906 -0.0400 0.9999 0.0839
-8.750 -0.4511 0.07167 0.06502 -0.0408 0.9999 0.0838
-8.500 -0.4563 0.06748 0.06074 -0.0417 0.9999 0.0836
-8.250 -0.4428 0.06160 0.05458 -0.0468 0.9951 0.0832
-8.000 -0.4184 0.05536 0.04790 -0.0531 0.9874 0.0833
-7.750 -0.3911 0.04991 0.04188 -0.0585 0.9802 0.0848
-7.500 -0.3608 0.04502 0.03621 -0.0631 0.9734 0.0873
-7.250 -0.3308 0.04202 0.03279 -0.0657 0.9663 0.0889
-7.000 -0.2972 0.03978 0.03027 -0.0682 0.9601 0.0906
-6.750 -0.2662 0.03787 0.02806 -0.0698 0.9527 0.0928
-6.500 -0.2311 0.03610 0.02592 -0.0721 0.9461 0.0970
-6.250 -0.1975 0.03428 0.02349 -0.0738 0.9390 0.1014
-6.000 -0.1640 0.03303 0.02219 -0.0753 0.9318 0.1047
-5.750 -0.1284 0.03193 0.02091 -0.0771 0.9253 0.1097
-5.500 -0.0959 0.03094 0.01958 -0.0780 0.9169 0.1169
-5.250 -0.0560 0.03004 0.01870 -0.0804 0.9116 0.1248
-5.000 -0.0278 0.02933 0.01781 -0.0805 0.9016 0.1337
-4.750 0.0134 0.02857 0.01695 -0.0828 0.8960 0.1482
-4.500 0.0414 0.02804 0.01641 -0.0829 0.8858 0.1631
-4.250 0.0820 0.02729 0.01573 -0.0851 0.8799 0.1867
-4.000 0.1095 0.02684 0.01531 -0.0851 0.8689 0.2144
-3.750 0.1439 0.02620 0.01485 -0.0863 0.8609 0.2542
-3.500 0.1748 0.02559 0.01451 -0.0868 0.8513 0.3108
-3.250 0.2032 0.02490 0.01427 -0.0869 0.8412 0.3997
-3.000 0.2317 0.02381 0.01402 -0.0861 0.8333 0.5749
-2.750 0.2741 0.02283 0.01384 -0.0866 0.8231 0.9773
-2.500 0.3097 0.02274 0.01339 -0.0877 0.8136 1.0001
-2.250 0.3391 0.02275 0.01314 -0.0877 0.8021 1.0001
-2.000 0.3660 0.02284 0.01300 -0.0873 0.7899 1.0001
-1.750 0.3963 0.02283 0.01277 -0.0874 0.7792 1.0001
-1.500 0.4269 0.02280 0.01254 -0.0875 0.7686 1.0001
-1.250 0.4524 0.02294 0.01253 -0.0868 0.7557 1.0001
-1.000 0.4800 0.02302 0.01245 -0.0865 0.7439 1.0001
-0.750 0.5122 0.02292 0.01220 -0.0866 0.7342 1.0001
-0.500 0.5367 0.02312 0.01229 -0.0859 0.7206 1.0001
-0.250 0.5623 0.02329 0.01236 -0.0853 0.7076 1.0001
0.000 0.5894 0.02340 0.01237 -0.0848 0.6953 1.0001
0.250 0.6195 0.02340 0.01224 -0.0847 0.6845 1.0001
0.500 0.6441 0.02363 0.01241 -0.0840 0.6707 1.0001
0.750 0.6688 0.02387 0.01259 -0.0833 0.6569 1.0001
1.000 0.6942 0.02410 0.01276 -0.0827 0.6438 1.0001
1.250 0.7208 0.02429 0.01289 -0.0822 0.6313 1.0001
1.500 0.7491 0.02439 0.01290 -0.0819 0.6196 1.0001
1.750 0.7729 0.02473 0.01322 -0.0812 0.6058 1.0001
2.000 0.7970 0.02509 0.01357 -0.0806 0.5925 1.0001
2.250 0.8219 0.02542 0.01387 -0.0800 0.5799 1.0001
2.500 0.8484 0.02568 0.01409 -0.0796 0.5683 1.0001
2.750 0.8743 0.02600 0.01440 -0.0792 0.5566 1.0001
3.000 0.8972 0.02649 0.01490 -0.0786 0.5440 1.0001
3.250 0.9216 0.02693 0.01535 -0.0780 0.5325 1.0001
3.500 0.9490 0.02722 0.01560 -0.0778 0.5225 1.0001
3.750 0.9709 0.02781 0.01624 -0.0770 0.5104 1.0001
4.000 0.9936 0.02837 0.01684 -0.0764 0.4991 1.0001
4.250 1.0190 0.02875 0.01723 -0.0759 0.4885 1.0001
4.500 1.0419 0.02923 0.01773 -0.0752 0.4766 1.0001
4.750 1.0621 0.02984 0.01839 -0.0742 0.4639 1.0001
5.000 1.0834 0.03037 0.01897 -0.0733 0.4516 1.0001
5.250 1.1062 0.03081 0.01941 -0.0725 0.4398 1.0001
5.500 1.1278 0.03131 0.01993 -0.0716 0.4277 1.0001
5.750 1.1451 0.03205 0.02077 -0.0703 0.4152 1.0001
6.000 1.1637 0.03272 0.02153 -0.0692 0.4032 1.0001
6.250 1.1844 0.03328 0.02212 -0.0682 0.3918 1.0001
6.500 1.2027 0.03395 0.02285 -0.0670 0.3800 1.0001
6.750 1.2168 0.03486 0.02390 -0.0654 0.3680 1.0001
7.000 1.2328 0.03566 0.02480 -0.0640 0.3564 1.0001
7.250 1.2516 0.03630 0.02545 -0.0629 0.3452 1.0001
7.500 1.2626 0.03734 0.02665 -0.0610 0.3333 1.0001
7.750 1.2731 0.03843 0.02787 -0.0592 0.3218 1.0001
8.000 1.2864 0.03934 0.02886 -0.0576 0.3108 1.0001
8.250 1.2966 0.04035 0.02996 -0.0556 0.3001 1.0001
8.500 1.3002 0.04182 0.03159 -0.0531 0.2895 1.0001
8.750 1.3102 0.04296 0.03278 -0.0514 0.2797 1.0001
9.000 1.3158 0.04440 0.03435 -0.0494 0.2698 1.0001
9.250 1.3185 0.04620 0.03629 -0.0475 0.2606 1.0001
9.500 1.3316 0.04729 0.03743 -0.0462 0.2526 1.0001
9.750 1.3281 0.04979 0.04017 -0.0443 0.2445 1.0001
10.000 1.3427 0.05082 0.04120 -0.0433 0.2373 1.0001
10.250 1.3338 0.05396 0.04459 -0.0416 0.2302 1.0001
10.500 1.3441 0.05539 0.04608 -0.0406 0.2237 1.0001
10.750 1.3371 0.05864 0.04954 -0.0394 0.2178 1.0001
11.000 1.3343 0.06159 0.05268 -0.0386 0.2123 1.0001
11.250 1.3507 0.06264 0.05376 -0.0378 0.2072 1.0001
11.500 1.3229 0.06857 0.06001 -0.0375 0.2028 1.0001
11.750 1.3100 0.07308 0.06470 -0.0375 0.1982 1.0001
12.000 1.3346 0.07294 0.06456 -0.0365 0.1929 1.0001
12.250 1.2717 0.08462 0.07654 -0.0391 0.1899 1.0001
12.500 1.1434 0.11071 0.10270 -0.0495 0.1839 1.0001
12.750 1.1817 0.10681 0.09894 -0.0466 0.1809 1.0001
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Polar data table (+)
Polar graphs
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