NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 50,000 Max Cl/Cd: 35.16 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m11-il-50000-n5.txt Download as CSV file: xf-m11-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M11 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5050 0.09919 0.09217 -0.0129 1.0000 0.1329
-8.500 -0.5217 0.08998 0.08299 -0.0226 1.0000 0.0768
-8.250 -0.5208 0.08608 0.07912 -0.0231 1.0000 0.0756
-8.000 -0.5223 0.08198 0.07503 -0.0246 1.0000 0.0750
-7.750 -0.5229 0.07783 0.07088 -0.0259 1.0000 0.0742
-7.500 -0.5227 0.07363 0.06665 -0.0269 1.0000 0.0729
-7.250 -0.5224 0.06919 0.06214 -0.0278 1.0000 0.0709
-7.000 -0.5222 0.06448 0.05728 -0.0285 1.0000 0.0687
-6.500 -0.5199 0.05486 0.04682 -0.0280 1.0000 0.0648
-6.250 -0.5075 0.05242 0.04449 -0.0268 1.0000 0.0673
-6.000 -0.4970 0.04977 0.04168 -0.0255 1.0000 0.0696
-5.750 -0.4870 0.04672 0.03834 -0.0239 1.0000 0.0705
-5.500 -0.4761 0.04357 0.03482 -0.0221 1.0000 0.0701
-5.250 -0.4635 0.04069 0.03153 -0.0203 1.0000 0.0700
-5.000 -0.4494 0.03809 0.02855 -0.0183 1.0000 0.0702
-4.750 -0.4341 0.03585 0.02595 -0.0165 1.0000 0.0712
-4.500 -0.4178 0.03400 0.02373 -0.0146 1.0000 0.0741
-4.250 -0.4003 0.03219 0.02149 -0.0128 1.0000 0.0762
-4.000 -0.3814 0.03045 0.01939 -0.0111 1.0000 0.0768
-3.750 -0.3614 0.02890 0.01748 -0.0096 1.0000 0.0774
-3.500 -0.3407 0.02754 0.01582 -0.0082 1.0000 0.0781
-3.250 -0.3196 0.02623 0.01440 -0.0070 1.0000 0.0793
-3.000 -0.2982 0.02515 0.01324 -0.0059 1.0000 0.0810
-2.750 -0.2761 0.02427 0.01225 -0.0049 1.0000 0.0831
-2.500 -0.2531 0.02353 0.01139 -0.0040 1.0000 0.0860
-2.250 -0.2305 0.02301 0.01068 -0.0032 1.0000 0.0921
-2.000 -0.2098 0.02247 0.01017 -0.0023 1.0000 0.0987
-1.750 -0.1809 0.02197 0.00951 -0.0029 0.9955 0.1066
-1.500 -0.1408 0.02134 0.00887 -0.0057 0.9853 0.1206
-1.000 -0.0053 0.01761 0.00808 -0.0209 0.9885 1.0000
-0.750 0.0389 0.01779 0.00793 -0.0247 0.9752 1.0000
-0.500 0.0817 0.01793 0.00783 -0.0281 0.9613 1.0000
-0.250 0.1246 0.01805 0.00777 -0.0316 0.9476 1.0000
0.000 0.1688 0.01813 0.00770 -0.0351 0.9347 1.0000
0.250 0.2106 0.01819 0.00766 -0.0381 0.9205 1.0000
0.500 0.2500 0.01823 0.00763 -0.0404 0.9055 1.0000
0.750 0.2860 0.01828 0.00762 -0.0420 0.8893 1.0000
1.000 0.3197 0.01835 0.00766 -0.0431 0.8726 1.0000
1.250 0.3519 0.01842 0.00771 -0.0438 0.8555 1.0000
1.500 0.3826 0.01850 0.00778 -0.0441 0.8386 1.0000
1.750 0.4120 0.01858 0.00786 -0.0440 0.8217 1.0000
2.000 0.4400 0.01869 0.00799 -0.0436 0.8049 1.0000
2.250 0.4654 0.01885 0.00816 -0.0428 0.7874 1.0000
2.500 0.4896 0.01904 0.00839 -0.0417 0.7694 1.0000
2.750 0.5138 0.01923 0.00864 -0.0406 0.7518 1.0000
3.000 0.5378 0.01943 0.00889 -0.0394 0.7345 1.0000
3.250 0.5619 0.01963 0.00915 -0.0383 0.7172 1.0000
3.500 0.5859 0.01985 0.00946 -0.0371 0.6996 1.0000
3.750 0.6083 0.02013 0.00984 -0.0357 0.6801 1.0000
4.000 0.6313 0.02037 0.01018 -0.0343 0.6607 1.0000
4.250 0.6543 0.02056 0.01048 -0.0327 0.6401 1.0000
4.500 0.6755 0.02077 0.01079 -0.0309 0.6156 1.0000
4.750 0.6962 0.02094 0.01103 -0.0288 0.5882 1.0000
5.000 0.7160 0.02107 0.01119 -0.0265 0.5564 1.0000
5.250 0.7342 0.02123 0.01139 -0.0240 0.5183 1.0000
5.500 0.7514 0.02147 0.01160 -0.0214 0.4744 1.0000
5.750 0.7678 0.02184 0.01191 -0.0188 0.4253 1.0000
6.000 0.7830 0.02239 0.01233 -0.0163 0.3681 1.0000
6.250 0.7958 0.02326 0.01293 -0.0137 0.2974 1.0000
6.500 0.8061 0.02462 0.01388 -0.0111 0.2213 1.0000
6.750 0.8164 0.02628 0.01515 -0.0088 0.1708 1.0000
7.000 0.8280 0.02797 0.01664 -0.0067 0.1415 1.0000
7.250 0.8402 0.02957 0.01813 -0.0047 0.1202 1.0000
7.500 0.8543 0.03118 0.01974 -0.0027 0.1063 1.0000
7.750 0.8689 0.03276 0.02135 -0.0010 0.0929 1.0000
8.000 0.8856 0.03442 0.02319 0.0007 0.0822 1.0000
8.250 0.9023 0.03620 0.02508 0.0022 0.0729 1.0000
8.500 0.9199 0.03812 0.02703 0.0035 0.0660 1.0000
8.750 0.9395 0.04045 0.02967 0.0047 0.0604 1.0000
9.000 0.9544 0.04256 0.03200 0.0062 0.0550 1.0000
9.250 0.9709 0.04531 0.03478 0.0071 0.0518 1.0000
9.500 0.9811 0.04839 0.03837 0.0090 0.0498 1.0000
9.750 0.9859 0.05146 0.04192 0.0111 0.0475 1.0000
10.000 0.9874 0.05446 0.04527 0.0131 0.0454 1.0000
10.250 0.9873 0.05733 0.04845 0.0150 0.0437 1.0000
10.500 0.9831 0.06041 0.05176 0.0170 0.0427 1.0000
10.750 0.9729 0.06366 0.05523 0.0191 0.0423 1.0000
11.000 0.9592 0.06722 0.05900 0.0206 0.0420 1.0000
11.250 0.9420 0.07128 0.06327 0.0210 0.0419 1.0000
11.500 0.9202 0.07632 0.06852 0.0198 0.0422 1.0000
11.750 0.8915 0.08302 0.07543 0.0164 0.0428 1.0000
12.000 0.8549 0.09272 0.08523 0.0097 0.0443 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M11 AIRFOIL (m11-il)