NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M2 AIRFOIL (m2-il) Reynolds number: 500,000 Max Cl/Cd: 54.37 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m2-il-500000-n5.txt Download as CSV file: xf-m2-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M2 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
------ -------- --------- --------- -------- -------- --------
-13.000 -1.1224 0.05222 0.04966 -0.0173 1.0000 0.0057
-12.750 -1.1497 0.04597 0.04313 -0.0186 1.0000 0.0057
-12.500 -1.1677 0.04185 0.03875 -0.0169 1.0000 0.0057
-12.250 -1.1749 0.03847 0.03508 -0.0149 1.0000 0.0058
-12.000 -1.1738 0.03547 0.03180 -0.0134 1.0000 0.0059
-11.750 -1.1673 0.03291 0.02896 -0.0120 1.0000 0.0061
-11.500 -1.1565 0.03074 0.02652 -0.0108 1.0000 0.0063
-11.250 -1.1424 0.02892 0.02440 -0.0098 1.0000 0.0065
-11.000 -1.1283 0.02699 0.02224 -0.0087 1.0000 0.0067
-10.750 -1.1116 0.02541 0.02048 -0.0078 1.0000 0.0071
-10.500 -1.0919 0.02423 0.01915 -0.0072 1.0000 0.0074
-10.250 -1.0713 0.02313 0.01790 -0.0066 1.0000 0.0078
-10.000 -1.0502 0.02204 0.01665 -0.0060 1.0000 0.0082
-9.750 -1.0286 0.02098 0.01543 -0.0054 1.0000 0.0086
-9.500 -1.0063 0.02000 0.01427 -0.0049 1.0000 0.0090
-9.250 -0.9835 0.01907 0.01319 -0.0044 1.0000 0.0095
-9.000 -0.9607 0.01813 0.01213 -0.0039 1.0000 0.0104
-8.750 -0.9360 0.01758 0.01153 -0.0036 1.0000 0.0113
-8.500 -0.9111 0.01706 0.01094 -0.0034 1.0000 0.0123
-8.250 -0.8863 0.01647 0.01025 -0.0031 1.0000 0.0133
-8.000 -0.8615 0.01584 0.00958 -0.0028 1.0000 0.0147
-7.750 -0.8355 0.01552 0.00925 -0.0026 1.0000 0.0161
-7.500 -0.8095 0.01520 0.00888 -0.0025 1.0000 0.0178
-7.250 -0.7831 0.01493 0.00853 -0.0024 1.0000 0.0193
-7.000 -0.7578 0.01442 0.00795 -0.0021 1.0000 0.0206
-6.750 -0.7323 0.01395 0.00745 -0.0019 1.0000 0.0223
-6.500 -0.7062 0.01364 0.00712 -0.0017 1.0000 0.0240
-6.250 -0.6801 0.01331 0.00674 -0.0016 1.0000 0.0257
-6.000 -0.6542 0.01297 0.00633 -0.0013 1.0000 0.0271
-5.750 -0.6281 0.01269 0.00600 -0.0011 1.0000 0.0281
-5.500 -0.6022 0.01242 0.00569 -0.0008 1.0000 0.0287
-5.250 -0.5777 0.01186 0.00508 -0.0003 1.0000 0.0301
-5.000 -0.5533 0.01137 0.00454 0.0002 1.0000 0.0318
-4.750 -0.5288 0.01101 0.00416 0.0008 1.0000 0.0333
-4.500 -0.5045 0.01071 0.00384 0.0014 1.0000 0.0349
-4.250 -0.4803 0.01045 0.00355 0.0021 1.0000 0.0361
-4.000 -0.4560 0.01022 0.00328 0.0027 1.0000 0.0373
-3.750 -0.4248 0.01000 0.00302 0.0019 0.9978 0.0385
-3.500 -0.3880 0.00977 0.00278 -0.0002 0.9946 0.0405
-3.250 -0.3532 0.00951 0.00254 -0.0018 0.9893 0.0462
-3.000 -0.3189 0.00921 0.00238 -0.0034 0.9847 0.0687
-2.750 -0.2872 0.00904 0.00227 -0.0043 0.9786 0.0859
-2.500 -0.2553 0.00889 0.00215 -0.0052 0.9705 0.0949
-2.250 -0.2255 0.00879 0.00206 -0.0056 0.9606 0.1030
-2.000 -0.1981 0.00865 0.00196 -0.0054 0.9500 0.1126
-1.750 -0.1716 0.00854 0.00186 -0.0050 0.9386 0.1193
-1.500 -0.1457 0.00840 0.00176 -0.0045 0.9237 0.1302
-1.250 -0.1199 0.00823 0.00166 -0.0039 0.9050 0.1494
-1.000 -0.0961 0.00764 0.00145 -0.0032 0.8783 0.2706
-0.750 -0.0740 0.00705 0.00129 -0.0021 0.8371 0.4333
-0.500 -0.0499 0.00688 0.00120 -0.0012 0.7899 0.5057
-0.250 -0.0246 0.00661 0.00116 -0.0007 0.7516 0.6024
0.000 0.0000 0.00643 0.00116 0.0000 0.6973 0.6972
0.250 0.0246 0.00661 0.00116 0.0007 0.6024 0.7517
0.500 0.0500 0.00687 0.00120 0.0012 0.5069 0.7898
0.750 0.0741 0.00705 0.00129 0.0021 0.4332 0.8372
1.000 0.0961 0.00764 0.00145 0.0032 0.2691 0.8785
1.250 0.1199 0.00823 0.00166 0.0039 0.1498 0.9051
1.500 0.1457 0.00840 0.00176 0.0045 0.1304 0.9238
1.750 0.1717 0.00854 0.00186 0.0050 0.1194 0.9387
2.000 0.1982 0.00865 0.00196 0.0054 0.1126 0.9499
2.250 0.2256 0.00879 0.00206 0.0056 0.1030 0.9606
2.500 0.2553 0.00889 0.00215 0.0052 0.0949 0.9707
2.750 0.2873 0.00904 0.00227 0.0043 0.0860 0.9788
3.000 0.3190 0.00921 0.00238 0.0034 0.0689 0.9847
3.250 0.3532 0.00951 0.00254 0.0018 0.0462 0.9893
3.500 0.3880 0.00977 0.00278 0.0002 0.0405 0.9946
3.750 0.4248 0.01000 0.00302 -0.0019 0.0385 0.9978
4.000 0.4559 0.01022 0.00328 -0.0027 0.0373 1.0000
4.250 0.4802 0.01045 0.00355 -0.0021 0.0361 1.0000
4.500 0.5044 0.01071 0.00384 -0.0014 0.0349 1.0000
4.750 0.5287 0.01101 0.00416 -0.0008 0.0333 1.0000
5.000 0.5533 0.01137 0.00454 -0.0002 0.0318 1.0000
5.250 0.5777 0.01187 0.00508 0.0003 0.0300 1.0000
5.500 0.6022 0.01242 0.00569 0.0008 0.0287 1.0000
5.750 0.6282 0.01269 0.00600 0.0011 0.0281 1.0000
6.000 0.6542 0.01297 0.00633 0.0013 0.0271 1.0000
6.250 0.6802 0.01331 0.00674 0.0015 0.0257 1.0000
6.500 0.7062 0.01364 0.00712 0.0017 0.0240 1.0000
6.750 0.7324 0.01395 0.00745 0.0019 0.0223 1.0000
7.000 0.7578 0.01443 0.00795 0.0021 0.0206 1.0000
7.250 0.7832 0.01493 0.00853 0.0023 0.0193 1.0000
7.500 0.8095 0.01521 0.00889 0.0025 0.0178 1.0000
7.750 0.8357 0.01552 0.00925 0.0026 0.0161 1.0000
8.000 0.8617 0.01585 0.00958 0.0027 0.0147 1.0000
8.250 0.8864 0.01648 0.01025 0.0030 0.0133 1.0000
8.500 0.9112 0.01706 0.01094 0.0034 0.0123 1.0000
8.750 0.9361 0.01759 0.01154 0.0036 0.0113 1.0000
9.000 0.9608 0.01813 0.01213 0.0038 0.0104 1.0000
9.250 0.9837 0.01907 0.01319 0.0043 0.0094 1.0000
9.500 1.0065 0.02000 0.01427 0.0048 0.0090 1.0000
9.750 1.0288 0.02099 0.01543 0.0054 0.0086 1.0000
10.000 1.0504 0.02204 0.01665 0.0059 0.0082 1.0000
10.250 1.0715 0.02313 0.01790 0.0065 0.0078 1.0000
10.500 1.0921 0.02424 0.01916 0.0071 0.0074 1.0000
10.750 1.1118 0.02541 0.02048 0.0078 0.0071 1.0000
11.000 1.1285 0.02700 0.02225 0.0087 0.0067 1.0000
11.250 1.1427 0.02892 0.02441 0.0097 0.0065 1.0000
11.500 1.1568 0.03075 0.02652 0.0107 0.0063 1.0000
11.750 1.1677 0.03292 0.02896 0.0119 0.0061 1.0000
12.000 1.1742 0.03548 0.03180 0.0133 0.0059 1.0000
12.250 1.1753 0.03849 0.03511 0.0148 0.0058 1.0000
12.500 1.1682 0.04188 0.03878 0.0168 0.0057 1.0000
12.750 1.1502 0.04601 0.04317 0.0185 0.0057 1.0000
13.000 1.1226 0.05233 0.04977 0.0171 0.0057 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M2 AIRFOIL (m2-il)