NACA M6 AIRFOIL (m6-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M6 AIRFOIL (m6-il) Reynolds number: 100,000 Max Cl/Cd: 48.82 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m6-il-100000.txt Download as CSV file: xf-m6-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M6 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5092 0.09825 0.09335 -0.0077 1.0000 0.1366
-8.750 -0.5586 0.09302 0.08820 -0.0187 1.0000 0.1405
-8.500 -0.5592 0.08779 0.08301 -0.0193 1.0000 0.1426
-8.250 -0.5256 0.08487 0.08012 -0.0163 1.0000 0.1463
-8.000 -0.5267 0.08146 0.07673 -0.0169 1.0000 0.1506
-7.750 -0.5753 0.07729 0.07232 -0.0223 1.0000 0.1579
-7.500 -0.5463 0.07294 0.06815 -0.0206 1.0000 0.1605
-7.250 -0.5337 0.07002 0.06527 -0.0198 1.0000 0.1653
-7.000 -0.5495 0.06634 0.06136 -0.0211 1.0000 0.1753
-6.750 -0.5313 0.06317 0.05831 -0.0198 1.0000 0.1790
-6.500 -0.5413 0.06087 0.05579 -0.0188 1.0000 0.1914
-6.250 -0.5281 0.05808 0.05319 -0.0166 1.0000 0.1956
-6.000 -0.5533 0.05779 0.05268 -0.0116 1.0000 0.2074
-5.750 -0.5527 0.05527 0.05035 -0.0081 1.0000 0.2100
-5.500 -0.5563 0.05388 0.04900 -0.0044 1.0000 0.2148
-5.250 -0.5208 0.03911 0.03211 -0.0097 0.9900 0.1062
-5.000 -0.4820 0.03409 0.02608 -0.0117 0.9826 0.0946
-4.750 -0.4428 0.03126 0.02287 -0.0143 0.9735 0.0934
-4.500 -0.4004 0.02913 0.02027 -0.0172 0.9655 0.0945
-4.250 -0.3553 0.02712 0.01788 -0.0205 0.9580 0.0951
-4.000 -0.3108 0.02545 0.01588 -0.0235 0.9499 0.0960
-3.750 -0.2657 0.02361 0.01397 -0.0268 0.9424 0.0984
-3.500 -0.2264 0.02250 0.01289 -0.0290 0.9322 0.1032
-3.250 -0.1837 0.02161 0.01185 -0.0315 0.9242 0.1107
-3.000 -0.1529 0.02059 0.01103 -0.0320 0.9118 0.1189
-2.750 -0.1240 0.01977 0.01030 -0.0318 0.9001 0.1322
-2.500 -0.0977 0.01876 0.00953 -0.0310 0.8899 0.1740
-2.250 -0.0830 0.01674 0.00889 -0.0287 0.8782 0.3964
-2.000 0.1271 0.01720 0.01063 -0.0570 0.8853 1.0000
-1.750 0.1474 0.01719 0.01044 -0.0556 0.8700 1.0000
-1.500 0.1679 0.01719 0.01027 -0.0542 0.8563 1.0000
-1.250 0.1882 0.01719 0.01012 -0.0528 0.8435 1.0000
-1.000 0.2105 0.01722 0.01002 -0.0519 0.8301 1.0000
-0.750 0.2329 0.01727 0.00995 -0.0510 0.8177 1.0000
-0.500 0.2546 0.01730 0.00986 -0.0499 0.8066 1.0000
-0.250 0.2768 0.01734 0.00978 -0.0488 0.7955 1.0000
0.000 0.3003 0.01743 0.00979 -0.0482 0.7837 1.0000
0.250 0.3229 0.01750 0.00976 -0.0471 0.7737 1.0000
0.500 0.3459 0.01757 0.00975 -0.0462 0.7630 1.0000
0.750 0.3698 0.01771 0.00984 -0.0457 0.7519 1.0000
1.000 0.3924 0.01777 0.00979 -0.0444 0.7433 1.0000
1.250 0.4165 0.01793 0.00994 -0.0440 0.7317 1.0000
1.500 0.4402 0.01810 0.01007 -0.0432 0.7217 1.0000
1.750 0.4632 0.01819 0.01008 -0.0421 0.7125 1.0000
2.000 0.4869 0.01843 0.01033 -0.0416 0.7014 1.0000
2.250 0.5099 0.01853 0.01037 -0.0404 0.6927 1.0000
2.500 0.5334 0.01875 0.01059 -0.0398 0.6817 1.0000
2.750 0.5566 0.01897 0.01082 -0.0390 0.6714 1.0000
3.000 0.5798 0.01905 0.01085 -0.0377 0.6625 1.0000
3.250 0.6028 0.01932 0.01117 -0.0370 0.6507 1.0000
3.500 0.6259 0.01948 0.01131 -0.0360 0.6409 1.0000
3.750 0.6490 0.01959 0.01141 -0.0348 0.6306 1.0000
4.000 0.6716 0.01984 0.01172 -0.0339 0.6189 1.0000
4.250 0.6947 0.01995 0.01181 -0.0328 0.6090 1.0000
4.500 0.7178 0.02008 0.01196 -0.0317 0.5981 1.0000
4.750 0.7401 0.02035 0.01230 -0.0307 0.5863 1.0000
5.000 0.7633 0.02048 0.01243 -0.0295 0.5760 1.0000
5.250 0.7867 0.02058 0.01256 -0.0284 0.5652 1.0000
5.500 0.8086 0.02087 0.01294 -0.0274 0.5528 1.0000
5.750 0.8312 0.02104 0.01315 -0.0262 0.5409 1.0000
6.000 0.8545 0.02102 0.01313 -0.0249 0.5283 1.0000
6.250 0.8779 0.02093 0.01302 -0.0235 0.5146 1.0000
6.500 0.9004 0.02086 0.01296 -0.0221 0.4990 1.0000
6.750 0.9217 0.02084 0.01298 -0.0206 0.4814 1.0000
7.000 0.9422 0.02075 0.01296 -0.0190 0.4613 1.0000
7.250 0.9629 0.02063 0.01282 -0.0173 0.4405 1.0000
7.500 0.9834 0.02058 0.01272 -0.0156 0.4197 1.0000
7.750 1.0017 0.02075 0.01299 -0.0139 0.3967 1.0000
8.000 1.0201 0.02091 0.01313 -0.0121 0.3735 1.0000
8.250 1.0360 0.02122 0.01351 -0.0101 0.3454 1.0000
8.500 1.0496 0.02166 0.01390 -0.0078 0.3120 1.0000
8.750 1.0580 0.02245 0.01454 -0.0049 0.2669 1.0000
9.000 1.0576 0.02393 0.01566 -0.0011 0.2091 1.0000
9.250 1.0527 0.02588 0.01720 0.0032 0.1648 1.0000
9.500 1.0515 0.02764 0.01872 0.0069 0.1397 1.0000
9.750 1.0532 0.02921 0.02014 0.0102 0.1246 1.0000
10.000 1.0564 0.03070 0.02155 0.0133 0.1142 1.0000
10.250 1.0622 0.03228 0.02293 0.0159 0.1063 1.0000
10.500 1.0724 0.03367 0.02440 0.0181 0.0991 1.0000
10.750 1.0898 0.03539 0.02599 0.0192 0.0926 1.0000
11.000 1.1056 0.03694 0.02766 0.0206 0.0875 1.0000
11.250 1.1396 0.03928 0.02979 0.0195 0.0814 1.0000
11.500 1.1491 0.04090 0.03169 0.0215 0.0783 1.0000
11.750 1.1656 0.04288 0.03386 0.0226 0.0751 1.0000
12.000 1.1929 0.04537 0.03629 0.0220 0.0716 1.0000
12.250 1.2041 0.04823 0.03938 0.0232 0.0693 1.0000
12.500 1.2002 0.05031 0.04178 0.0263 0.0680 1.0000
12.750 1.1969 0.05284 0.04461 0.0287 0.0670 1.0000
13.000 1.1917 0.05570 0.04776 0.0307 0.0662 1.0000
13.250 1.1830 0.05884 0.05117 0.0325 0.0656 1.0000
13.500 1.1710 0.06228 0.05489 0.0337 0.0652 1.0000
13.750 1.1556 0.06612 0.05900 0.0345 0.0649 1.0000
14.000 1.1360 0.07054 0.06368 0.0345 0.0648 1.0000
14.250 1.1115 0.07575 0.06916 0.0336 0.0649 1.0000
14.500 1.0817 0.08202 0.07570 0.0317 0.0655 1.0000
14.750 1.0470 0.08957 0.08349 0.0284 0.0664 1.0000
15.000 1.0094 0.09843 0.09256 0.0238 0.0675 1.0000
15.250 0.9720 0.10841 0.10269 0.0183 0.0687 1.0000
15.500 0.9407 0.11870 0.11307 0.0128 0.0698 1.0000
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Polar data table (+)
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