USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=500,000 Ncrit=5
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Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 500,000 Max Cl/Cd: 96.54 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa22-il-500000-n5.txt Download as CSV file: xf-usa22-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 22 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 -0.3712 0.08551 0.08346 -0.0193 1.0000 0.0068
-8.000 -0.3716 0.08133 0.07932 -0.0214 1.0000 0.0068
-7.500 -0.3911 0.02455 0.02138 -0.0778 0.9614 0.0064
-7.250 -0.3634 0.02021 0.01636 -0.0805 0.9488 0.0066
-7.000 -0.3343 0.01799 0.01372 -0.0818 0.9356 0.0069
-6.750 -0.3061 0.01654 0.01193 -0.0823 0.9210 0.0071
-6.500 -0.2801 0.01498 0.01004 -0.0823 0.9050 0.0075
-6.250 -0.2540 0.01412 0.00897 -0.0820 0.8884 0.0080
-6.000 -0.2278 0.01352 0.00820 -0.0817 0.8711 0.0085
-5.750 -0.2015 0.01302 0.00754 -0.0813 0.8527 0.0093
-5.500 -0.1753 0.01253 0.00684 -0.0808 0.8301 0.0101
-5.250 -0.1493 0.01207 0.00613 -0.0803 0.8031 0.0106
-5.000 -0.1236 0.01140 0.00522 -0.0797 0.7749 0.0120
-4.500 -0.0706 0.01084 0.00431 -0.0788 0.7288 0.0152
-4.250 -0.0437 0.01057 0.00391 -0.0785 0.7090 0.0187
-4.000 -0.0166 0.01043 0.00366 -0.0782 0.6908 0.0231
-3.750 0.0107 0.01030 0.00345 -0.0780 0.6739 0.0274
-3.500 0.0383 0.01027 0.00333 -0.0778 0.6579 0.0320
-3.250 0.0657 0.01016 0.00311 -0.0775 0.6410 0.0349
-3.000 0.0931 0.01008 0.00295 -0.0774 0.6234 0.0388
-2.750 0.1205 0.01004 0.00281 -0.0772 0.6060 0.0421
-2.500 0.1480 0.01004 0.00270 -0.0770 0.5891 0.0445
-2.250 0.1753 0.00984 0.00238 -0.0767 0.5717 0.0481
-2.000 0.2027 0.00975 0.00219 -0.0765 0.5541 0.0509
-1.750 0.2302 0.00969 0.00204 -0.0764 0.5375 0.0539
-1.500 0.2577 0.00967 0.00191 -0.0762 0.5216 0.0569
-1.250 0.2852 0.00964 0.00179 -0.0760 0.5067 0.0622
-1.000 0.3127 0.00958 0.00169 -0.0759 0.4934 0.0725
-0.750 0.3397 0.00939 0.00163 -0.0758 0.4813 0.1304
-0.500 0.3670 0.00933 0.00161 -0.0757 0.4699 0.1666
-0.250 0.3945 0.00928 0.00162 -0.0756 0.4593 0.2027
0.000 0.4220 0.00930 0.00164 -0.0755 0.4491 0.2336
0.250 0.4495 0.00935 0.00168 -0.0754 0.4385 0.2591
0.500 0.4769 0.00941 0.00172 -0.0752 0.4272 0.2796
0.750 0.5046 0.00946 0.00176 -0.0751 0.4177 0.2928
1.000 0.5321 0.00953 0.00179 -0.0750 0.4103 0.3052
1.250 0.5598 0.00955 0.00185 -0.0750 0.4039 0.3231
1.500 0.5873 0.00961 0.00189 -0.0749 0.3971 0.3352
1.750 0.6150 0.00965 0.00194 -0.0748 0.3906 0.3457
2.000 0.6426 0.00970 0.00200 -0.0747 0.3841 0.3598
2.250 0.6701 0.00971 0.00208 -0.0747 0.3786 0.3822
2.500 0.6964 0.00936 0.00217 -0.0746 0.3725 0.5647
2.750 0.7261 0.00855 0.00229 -0.0748 0.3664 1.0000
3.000 0.7535 0.00865 0.00238 -0.0747 0.3600 1.0000
3.250 0.7806 0.00880 0.00247 -0.0746 0.3530 1.0000
3.500 0.8079 0.00892 0.00259 -0.0744 0.3455 1.0000
3.750 0.8348 0.00908 0.00270 -0.0743 0.3372 1.0000
4.000 0.8619 0.00922 0.00283 -0.0741 0.3288 1.0000
4.250 0.8888 0.00939 0.00297 -0.0740 0.3206 1.0000
4.500 0.9156 0.00956 0.00313 -0.0738 0.3114 1.0000
4.750 0.9422 0.00976 0.00330 -0.0736 0.2992 1.0000
5.000 0.9679 0.01006 0.00349 -0.0733 0.2781 1.0000
5.250 0.9936 0.01036 0.00373 -0.0731 0.2591 1.0000
5.750 1.0437 0.01113 0.00430 -0.0724 0.2180 1.0000
6.000 1.0679 0.01161 0.00465 -0.0720 0.1950 1.0000
6.250 1.0913 0.01220 0.00509 -0.0715 0.1644 1.0000
6.500 1.1086 0.01357 0.00594 -0.0703 0.0845 1.0000
6.750 1.1255 0.01500 0.00708 -0.0690 0.0170 1.0000
7.000 1.1494 0.01547 0.00760 -0.0685 0.0137 1.0000
7.250 1.1725 0.01604 0.00826 -0.0678 0.0107 1.0000
7.500 1.1961 0.01650 0.00881 -0.0672 0.0099 1.0000
7.750 1.2190 0.01703 0.00943 -0.0666 0.0090 1.0000
8.000 1.2412 0.01763 0.01010 -0.0659 0.0082 1.0000
8.250 1.2613 0.01844 0.01098 -0.0649 0.0073 1.0000
8.500 1.2815 0.01918 0.01181 -0.0640 0.0067 1.0000
8.750 1.3012 0.01994 0.01267 -0.0630 0.0063 1.0000
9.000 1.3196 0.02078 0.01360 -0.0618 0.0059 1.0000
9.250 1.3367 0.02168 0.01462 -0.0605 0.0056 1.0000
9.500 1.3524 0.02264 0.01566 -0.0591 0.0053 1.0000
9.750 1.3658 0.02371 0.01682 -0.0575 0.0051 1.0000
10.000 1.3739 0.02505 0.01825 -0.0552 0.0049 1.0000
10.250 1.3760 0.02657 0.01989 -0.0522 0.0047 1.0000
10.500 1.3828 0.02785 0.02128 -0.0501 0.0045 1.0000
10.750 1.3873 0.02939 0.02293 -0.0482 0.0043 1.0000
11.000 1.3909 0.03116 0.02482 -0.0466 0.0041 1.0000
11.250 1.3913 0.03341 0.02720 -0.0454 0.0040 1.0000
11.500 1.3904 0.03605 0.02996 -0.0447 0.0039 1.0000
11.750 1.3894 0.03898 0.03301 -0.0445 0.0038 1.0000
12.000 1.3876 0.04222 0.03637 -0.0447 0.0037 1.0000
12.250 1.3852 0.04568 0.03995 -0.0451 0.0036 1.0000
12.500 1.3815 0.04938 0.04379 -0.0457 0.0035 1.0000
12.750 1.3778 0.05316 0.04769 -0.0464 0.0035 1.0000
13.000 1.3732 0.05709 0.05172 -0.0472 0.0034 1.0000
13.250 1.3677 0.06118 0.05591 -0.0480 0.0034 1.0000
13.500 1.3623 0.06527 0.06011 -0.0490 0.0033 1.0000
13.750 1.3559 0.06949 0.06443 -0.0498 0.0033 1.0000
14.000 1.3496 0.07377 0.06880 -0.0508 0.0032 1.0000
14.250 1.3429 0.07812 0.07324 -0.0518 0.0032 1.0000
14.500 1.3365 0.08247 0.07768 -0.0527 0.0031 1.0000
14.750 1.3299 0.08683 0.08213 -0.0536 0.0031 1.0000
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