N-22 (n22-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: N-22 (n22-il) Reynolds number: 50,000 Max Cl/Cd: 24.26 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n22-il-50000.txt Download as CSV file: xf-n22-il-50000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.3307 0.11470 0.10756 -0.0276 0.9999 0.2401
-10.750 -0.3082 0.10951 0.10237 -0.0263 0.9999 0.2478
-10.500 -0.3246 0.10905 0.10204 -0.0264 0.9999 0.2567
-10.250 -0.3082 0.10459 0.09760 -0.0252 0.9999 0.2635
-10.000 -0.3216 0.10374 0.09688 -0.0241 0.9999 0.2730
-9.750 -0.3152 0.10029 0.09350 -0.0228 0.9999 0.2792
-9.500 -0.3243 0.09901 0.09233 -0.0208 0.9999 0.2891
-9.250 -0.3307 0.09671 0.09015 -0.0187 0.9999 0.2953
-9.000 -0.3380 0.09527 0.08882 -0.0158 0.9999 0.3056
-8.750 -0.3940 0.09720 0.09102 -0.0113 0.9999 0.3087
-8.500 -0.3650 0.09260 0.08640 -0.0088 0.9999 0.3227
-8.250 -0.3711 0.09049 0.08438 -0.0056 0.9999 0.3316
-8.000 -0.4034 0.09054 0.08460 -0.0023 0.9999 0.3416
-7.750 -0.3952 0.08769 0.08179 0.0011 0.9999 0.3550
-7.500 -0.4029 0.08577 0.07997 0.0042 0.9999 0.3673
-7.250 -0.4171 0.08422 0.07851 0.0069 0.9999 0.3812
-7.000 -0.4263 0.08249 0.07686 0.0098 0.9999 0.3975
-6.750 -0.4300 0.08051 0.07494 0.0129 0.9999 0.4150
-6.500 -0.4309 0.07845 0.07294 0.0162 0.9999 0.4341
-6.250 -0.4332 0.07659 0.07114 0.0197 0.9999 0.4566
-6.000 -0.4373 0.07466 0.06929 0.0236 0.9999 0.4839
-5.750 -0.3479 0.04991 0.04230 -0.0397 0.9999 0.1923
-5.500 -0.3233 0.04686 0.03887 -0.0415 0.9999 0.1879
-5.250 -0.2986 0.04403 0.03571 -0.0430 0.9999 0.1846
-5.000 -0.2722 0.04153 0.03276 -0.0446 0.9999 0.1832
-4.750 -0.2464 0.03974 0.03052 -0.0458 0.9999 0.1867
-4.500 -0.2195 0.03823 0.02841 -0.0468 0.9999 0.1904
-4.250 -0.1969 0.03695 0.02709 -0.0471 0.9999 0.1946
-4.000 -0.1738 0.03619 0.02614 -0.0474 0.9999 0.2028
-3.750 -0.1493 0.03545 0.02524 -0.0479 0.9995 0.2127
-3.500 -0.0929 0.03483 0.02441 -0.0538 0.9880 0.2345
-3.250 -0.0407 0.03432 0.02386 -0.0587 0.9760 0.2669
-3.000 0.0097 0.03388 0.02351 -0.0631 0.9636 0.3197
-2.750 0.0603 0.03327 0.02335 -0.0673 0.9514 0.4144
-2.500 0.0973 0.03075 0.02310 -0.0666 0.9399 1.0001
-2.250 0.1456 0.03155 0.02306 -0.0710 0.9244 1.0001
-2.000 0.1873 0.03227 0.02339 -0.0741 0.9088 1.0001
-1.750 0.2273 0.03294 0.02375 -0.0768 0.8930 1.0001
-1.500 0.2655 0.03358 0.02414 -0.0791 0.8772 1.0001
-1.250 0.3024 0.03418 0.02455 -0.0811 0.8615 1.0001
-1.000 0.3384 0.03475 0.02495 -0.0828 0.8457 1.0001
-0.750 0.3726 0.03533 0.02538 -0.0842 0.8301 1.0001
-0.500 0.4063 0.03588 0.02581 -0.0853 0.8145 1.0001
-0.250 0.4389 0.03643 0.02625 -0.0863 0.7991 1.0001
0.000 0.4703 0.03700 0.02674 -0.0870 0.7840 1.0001
0.250 0.5017 0.03753 0.02720 -0.0876 0.7691 1.0001
0.500 0.5342 0.03801 0.02763 -0.0882 0.7551 1.0001
0.750 0.5861 0.03763 0.02719 -0.0908 0.7451 1.0001
1.000 0.6088 0.03839 0.02791 -0.0902 0.7301 1.0001
1.250 0.6295 0.03931 0.02881 -0.0894 0.7153 1.0001
1.500 0.6497 0.04031 0.02979 -0.0887 0.7011 1.0001
1.750 0.6721 0.04123 0.03070 -0.0881 0.6881 1.0001
2.000 0.7197 0.04054 0.03001 -0.0893 0.6800 1.0001
2.250 0.7309 0.04218 0.03165 -0.0879 0.6657 1.0001
2.500 0.7410 0.04398 0.03346 -0.0866 0.6521 1.0001
2.750 0.7579 0.04539 0.03489 -0.0858 0.6404 1.0001
3.000 0.7996 0.04492 0.03446 -0.0863 0.6320 1.0001
3.250 0.7986 0.04767 0.03723 -0.0844 0.6183 1.0001
3.500 0.7990 0.05043 0.04002 -0.0828 0.6060 1.0001
3.750 0.8596 0.04830 0.03797 -0.0839 0.5990 1.0001
4.000 0.8456 0.05219 0.04188 -0.0816 0.5850 1.0001
4.250 0.8385 0.05574 0.04546 -0.0800 0.5720 1.0001
4.500 0.9318 0.05019 0.04005 -0.0813 0.5657 1.0001
4.750 0.9155 0.05425 0.04415 -0.0788 0.5513 1.0001
5.000 0.8943 0.05926 0.04918 -0.0770 0.5372 1.0001
5.250 0.8893 0.06277 0.05274 -0.0758 0.5240 1.0001
5.500 0.9075 0.06393 0.05396 -0.0746 0.5118 1.0001
5.750 1.0235 0.05460 0.04485 -0.0744 0.5026 1.0001
6.000 0.9892 0.06061 0.05088 -0.0716 0.4879 1.0001
6.250 0.9655 0.06646 0.05677 -0.0706 0.4729 1.0001
6.500 0.9575 0.07055 0.06091 -0.0696 0.4586 1.0001
6.750 0.9595 0.07348 0.06390 -0.0685 0.4447 1.0001
7.000 0.9774 0.07451 0.06502 -0.0671 0.4317 1.0001
7.250 1.1958 0.05334 0.04415 -0.0673 0.4187 1.0001
7.500 1.2329 0.05267 0.04354 -0.0668 0.4034 1.0001
7.750 1.2693 0.05232 0.04322 -0.0664 0.3885 1.0001
8.000 0.9519 0.09196 0.08268 -0.0656 0.3809 1.0001
8.250 0.9786 0.09213 0.08296 -0.0640 0.3701 1.0001
8.500 0.9363 0.10208 0.09288 -0.0663 0.3621 1.0001
8.750 0.9449 0.10521 0.09608 -0.0660 0.3540 1.0001
9.000 0.9112 0.11409 0.10495 -0.0684 0.3497 1.0001
9.250 0.9441 0.11399 0.10497 -0.0667 0.3396 1.0001
9.500 0.9191 0.12235 0.11333 -0.0693 0.3390 1.0001
9.750 0.9141 0.12854 0.11958 -0.0711 0.3399 1.0001
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Polar data table (+)
Polar graphs
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