NACA 22112 (naca22112-jf) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 22112 (naca22112-jf) Reynolds number: 50,000 Max Cl/Cd: 23.87 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca22112-jf-50000.txt Download as CSV file: xf-naca22112-jf-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 22112
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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.6740 0.06513 0.05795 -0.0169 1.0000 0.1816
-7.750 -0.6816 0.05884 0.05123 -0.0171 1.0000 0.1651
-7.500 -0.6868 0.05369 0.04537 -0.0157 1.0000 0.1526
-7.250 -0.6742 0.04979 0.04128 -0.0144 1.0000 0.1485
-7.000 -0.6675 0.04593 0.03671 -0.0122 1.0000 0.1415
-6.750 -0.6526 0.04312 0.03336 -0.0102 1.0000 0.1376
-6.500 -0.6335 0.04014 0.03014 -0.0087 1.0000 0.1349
-6.250 -0.6137 0.03753 0.02718 -0.0071 1.0000 0.1328
-6.000 -0.5928 0.03532 0.02465 -0.0055 1.0000 0.1333
-5.750 -0.5704 0.03337 0.02241 -0.0041 1.0000 0.1349
-5.500 -0.5459 0.03155 0.02032 -0.0028 1.0000 0.1362
-5.250 -0.5193 0.02989 0.01845 -0.0018 1.0000 0.1375
-5.000 -0.4915 0.02822 0.01680 -0.0010 1.0000 0.1419
-4.750 -0.4645 0.02686 0.01550 -0.0002 1.0000 0.1485
-4.500 -0.4378 0.02576 0.01433 0.0010 1.0000 0.1545
-4.250 -0.4154 0.02455 0.01328 0.0023 1.0000 0.1647
-4.000 -0.3957 0.02347 0.01232 0.0040 1.0000 0.1792
-3.750 -0.3775 0.02230 0.01138 0.0058 1.0000 0.2024
-3.500 -0.3672 0.02025 0.01057 0.0084 1.0000 0.3321
-3.250 -0.3645 0.01897 0.01113 0.0159 1.0000 0.6711
-3.000 -0.3380 0.02039 0.01289 0.0237 1.0000 0.8420
-2.750 -0.0291 0.02367 0.01455 -0.0152 1.0000 0.9891
-2.500 0.0115 0.02267 0.01340 -0.0200 1.0000 1.0000
-2.250 0.0206 0.02207 0.01277 -0.0189 1.0000 1.0000
-2.000 0.0274 0.02158 0.01226 -0.0173 1.0000 1.0000
-1.750 0.0305 0.02121 0.01190 -0.0150 1.0000 1.0000
-1.500 0.0289 0.02097 0.01168 -0.0119 1.0000 1.0000
-1.250 0.0220 0.02087 0.01159 -0.0080 1.0000 1.0000
-1.000 0.0097 0.02091 0.01163 -0.0032 1.0000 1.0000
-0.750 -0.0051 0.02106 0.01178 0.0018 1.0000 1.0000
-0.500 -0.0186 0.02133 0.01201 0.0065 1.0000 1.0000
-0.250 -0.0280 0.02170 0.01233 0.0104 1.0000 1.0000
0.000 -0.0329 0.02217 0.01274 0.0135 1.0000 1.0000
0.250 0.0339 0.02304 0.01355 0.0041 0.9805 1.0000
0.500 0.1282 0.02362 0.01414 -0.0093 0.9489 1.0000
0.750 0.2357 0.02352 0.01412 -0.0237 0.9144 1.0000
1.000 0.3280 0.02269 0.01339 -0.0335 0.8763 1.0000
1.250 0.3742 0.02214 0.01284 -0.0347 0.8352 1.0000
1.500 0.4043 0.02177 0.01242 -0.0329 0.7954 1.0000
2.000 0.4460 0.02154 0.01196 -0.0265 0.7152 1.0000
2.250 0.4658 0.02158 0.01181 -0.0234 0.6750 1.0000
2.500 0.4845 0.02181 0.01185 -0.0204 0.6320 1.0000
2.750 0.5041 0.02210 0.01190 -0.0176 0.5908 1.0000
3.000 0.5233 0.02259 0.01218 -0.0152 0.5488 1.0000
3.250 0.5432 0.02316 0.01250 -0.0131 0.5109 1.0000
3.500 0.5637 0.02385 0.01301 -0.0113 0.4777 1.0000
3.750 0.5844 0.02463 0.01363 -0.0097 0.4493 1.0000
4.000 0.6058 0.02541 0.01424 -0.0082 0.4255 1.0000
4.250 0.6279 0.02631 0.01497 -0.0070 0.4063 1.0000
4.500 0.6491 0.02730 0.01596 -0.0058 0.3890 1.0000
4.750 0.6707 0.02828 0.01688 -0.0047 0.3737 1.0000
5.000 0.6908 0.02944 0.01815 -0.0035 0.3599 1.0000
5.250 0.7105 0.03081 0.01965 -0.0024 0.3486 1.0000
5.500 0.7316 0.03203 0.02087 -0.0014 0.3384 1.0000
5.750 0.7509 0.03337 0.02233 -0.0002 0.3281 1.0000
6.000 0.7685 0.03509 0.02423 0.0009 0.3201 1.0000
6.250 0.7869 0.03672 0.02601 0.0020 0.3123 1.0000
6.500 0.8034 0.03860 0.02804 0.0032 0.3051 1.0000
6.750 0.8168 0.04071 0.03040 0.0045 0.2976 1.0000
7.000 0.8397 0.04228 0.03190 0.0052 0.2919 1.0000
7.250 0.8396 0.04593 0.03608 0.0067 0.2873 1.0000
7.500 0.8452 0.04901 0.03944 0.0079 0.2819 1.0000
7.750 0.8704 0.05035 0.04069 0.0087 0.2757 1.0000
8.000 0.8556 0.05575 0.04654 0.0098 0.2741 1.0000
8.250 0.8338 0.06209 0.05317 0.0100 0.2737 1.0000
8.500 0.8071 0.06922 0.06045 0.0090 0.2746 1.0000
8.750 0.7825 0.07651 0.06781 0.0072 0.2765 1.0000
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Polar data table (+)
Polar graphs
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