MH 22 7.2% (mh22-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: MH 22 7.2% (mh22-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.73 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-mh22-il-1000000.txt Download as CSV file: xf-mh22-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: MH 22 7.2%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4520 0.08126 0.07986 0.0025 1.0000 0.0048
-8.000 -0.4528 0.07659 0.07520 0.0007 1.0000 0.0049
-7.750 -0.4548 0.07198 0.07061 -0.0012 1.0000 0.0049
-7.500 -0.4576 0.06752 0.06617 -0.0031 1.0000 0.0050
-7.250 -0.4642 0.06313 0.06180 -0.0051 1.0000 0.0050
-7.000 -0.4735 0.05837 0.05706 -0.0086 1.0000 0.0049
-6.750 -0.4737 0.05257 0.05124 -0.0133 1.0000 0.0049
-6.500 -0.4705 0.04622 0.04484 -0.0177 1.0000 0.0049
-6.250 -0.4623 0.04015 0.03867 -0.0212 1.0000 0.0050
-6.000 -0.4514 0.03416 0.03255 -0.0236 1.0000 0.0051
-5.750 -0.4375 0.02865 0.02688 -0.0251 1.0000 0.0054
-5.500 -0.4187 0.02398 0.02204 -0.0260 0.9933 0.0057
-5.250 -0.3883 0.02156 0.01942 -0.0270 0.9316 0.0064
-5.000 -0.3762 0.01814 0.01561 -0.0253 0.8816 0.0064
-4.250 -0.3268 0.01281 0.00801 -0.0229 0.8512 0.0032
-3.750 -0.2781 0.01022 0.00477 -0.0206 0.7991 0.0031
-3.500 -0.2527 0.00949 0.00385 -0.0199 0.7796 0.0037
-3.250 -0.2264 0.00909 0.00333 -0.0194 0.7627 0.0045
-3.000 -0.1998 0.00873 0.00285 -0.0190 0.7484 0.0053
-2.750 -0.1730 0.00839 0.00238 -0.0186 0.7359 0.0058
-2.500 -0.1459 0.00816 0.00207 -0.0183 0.7248 0.0071
-2.250 -0.1186 0.00785 0.00155 -0.0179 0.7150 0.0089
-2.000 -0.0930 0.00718 0.00125 -0.0175 0.7057 0.1086
-1.750 -0.0666 0.00684 0.00113 -0.0173 0.6971 0.1834
-1.500 -0.0403 0.00651 0.00101 -0.0171 0.6890 0.2653
-1.250 -0.0144 0.00610 0.00093 -0.0169 0.6808 0.3758
-1.000 0.0106 0.00563 0.00086 -0.0165 0.6734 0.5174
-0.750 0.0346 0.00515 0.00082 -0.0158 0.6658 0.6642
-0.500 0.0573 0.00470 0.00079 -0.0146 0.6587 0.7974
-0.250 0.0930 0.00437 0.00082 -0.0158 0.6512 0.9519
0.000 0.1388 0.00443 0.00083 -0.0197 0.6429 0.9828
0.250 0.1801 0.00450 0.00083 -0.0226 0.6348 0.9930
0.500 0.2240 0.00453 0.00080 -0.0262 0.6257 0.9985
0.750 0.2561 0.00455 0.00078 -0.0272 0.6166 1.0000
1.000 0.2820 0.00458 0.00078 -0.0267 0.6075 1.0000
1.250 0.3079 0.00463 0.00077 -0.0263 0.5974 1.0000
1.500 0.3340 0.00467 0.00078 -0.0259 0.5859 1.0000
1.750 0.3601 0.00472 0.00079 -0.0254 0.5726 1.0000
2.000 0.3861 0.00478 0.00081 -0.0250 0.5575 1.0000
2.250 0.4122 0.00486 0.00086 -0.0246 0.5386 1.0000
2.500 0.4381 0.00497 0.00089 -0.0242 0.5159 1.0000
2.750 0.4639 0.00510 0.00095 -0.0238 0.4903 1.0000
3.000 0.4896 0.00528 0.00102 -0.0234 0.4609 1.0000
3.500 0.5386 0.00607 0.00133 -0.0224 0.3355 1.0000
3.750 0.5629 0.00650 0.00151 -0.0220 0.2739 1.0000
4.000 0.5861 0.00717 0.00177 -0.0214 0.1865 1.0000
4.250 0.6069 0.00831 0.00229 -0.0207 0.0576 1.0000
4.500 0.6299 0.00906 0.00278 -0.0199 0.0032 1.0000
4.750 0.6555 0.00933 0.00314 -0.0194 0.0023 1.0000
5.000 0.6805 0.00972 0.00365 -0.0187 0.0022 1.0000
5.250 0.7049 0.01024 0.00430 -0.0179 0.0022 1.0000
5.500 0.7289 0.01083 0.00499 -0.0171 0.0023 1.0000
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