NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 50,000 Max Cl/Cd: 34.77 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m11-il-50000.txt Download as CSV file: xf-m11-il-50000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5205 0.12495 0.11764 -0.0006 1.0000 0.1890
-10.000 -0.5045 0.11980 0.11248 0.0002 1.0000 0.1969
-9.750 -0.5182 0.11863 0.11142 -0.0022 1.0000 0.2043
-9.500 -0.5013 0.11364 0.10642 -0.0011 1.0000 0.2164
-9.250 -0.4971 0.11000 0.10282 -0.0014 1.0000 0.2246
-9.000 -0.5105 0.10862 0.10155 -0.0032 1.0000 0.2340
-8.750 -0.4942 0.10392 0.09684 -0.0021 1.0000 0.2466
-8.500 -0.4867 0.10019 0.09315 -0.0017 1.0000 0.2577
-8.000 -0.4858 0.09402 0.08711 -0.0016 1.0000 0.2833
-7.500 -0.4822 0.08832 0.08151 0.0003 1.0000 0.3214
-7.250 -0.4702 0.08485 0.07807 0.0025 1.0000 0.3478
-7.000 -0.4868 0.08323 0.07660 0.0039 1.0000 0.3680
-6.750 -0.4598 0.07904 0.07237 0.0068 1.0000 0.4030
-6.500 -0.4494 0.07627 0.06963 0.0098 1.0000 0.4394
-6.250 -0.4322 0.07330 0.06667 0.0132 1.0000 0.4839
-6.000 -0.4169 0.07091 0.06428 0.0183 1.0000 0.5478
-5.750 -0.3491 0.06617 0.05935 0.0212 1.0000 0.6542
-5.000 -0.2581 0.05551 0.04859 0.0195 1.0000 0.7878
-4.500 -0.4395 0.04358 0.03553 -0.0144 1.0000 0.2020
-4.250 -0.4224 0.04021 0.03172 -0.0133 1.0000 0.1837
-4.000 -0.4048 0.03732 0.02830 -0.0117 1.0000 0.1695
-3.750 -0.3863 0.03492 0.02521 -0.0098 1.0000 0.1587
-3.500 -0.3681 0.03268 0.02272 -0.0082 1.0000 0.1548
-3.250 -0.3488 0.03073 0.02038 -0.0064 1.0000 0.1514
-3.000 -0.3286 0.02907 0.01838 -0.0048 1.0000 0.1505
-2.750 -0.3087 0.02774 0.01682 -0.0033 1.0000 0.1543
-2.500 -0.2876 0.02656 0.01534 -0.0019 1.0000 0.1576
-2.250 -0.2648 0.02545 0.01392 -0.0007 1.0000 0.1597
-2.000 -0.2394 0.02421 0.01266 -0.0001 1.0000 0.1638
-1.750 -0.2105 0.02327 0.01162 -0.0001 1.0000 0.1716
-1.500 -0.0511 0.01720 0.00853 -0.0211 1.0000 1.0000
-1.250 -0.0416 0.01740 0.00833 -0.0182 1.0000 1.0000
-1.000 -0.0326 0.01764 0.00830 -0.0154 1.0000 1.0000
-0.750 -0.0233 0.01792 0.00834 -0.0129 1.0000 1.0000
-0.500 -0.0136 0.01825 0.00847 -0.0105 1.0000 1.0000
-0.250 -0.0035 0.01862 0.00867 -0.0084 1.0000 1.0000
0.000 0.0072 0.01903 0.00893 -0.0064 1.0000 1.0000
0.250 0.0183 0.01949 0.00924 -0.0046 1.0000 1.0000
0.500 0.0299 0.01999 0.00962 -0.0031 1.0000 1.0000
0.750 0.0420 0.02053 0.01005 -0.0017 1.0000 1.0000
1.000 0.0544 0.02112 0.01054 -0.0005 1.0000 1.0000
1.250 0.0785 0.02188 0.01122 -0.0018 0.9958 1.0000
1.500 0.1366 0.02297 0.01226 -0.0094 0.9774 1.0000
1.750 0.1925 0.02394 0.01322 -0.0162 0.9588 1.0000
2.000 0.2498 0.02483 0.01414 -0.0230 0.9407 1.0000
2.250 0.2934 0.02552 0.01489 -0.0270 0.9209 1.0000
2.500 0.3421 0.02616 0.01563 -0.0316 0.9015 1.0000
2.750 0.3974 0.02668 0.01631 -0.0371 0.8829 1.0000
3.000 0.4390 0.02719 0.01695 -0.0400 0.8628 1.0000
3.250 0.4878 0.02752 0.01748 -0.0438 0.8426 1.0000
3.500 0.5412 0.02756 0.01778 -0.0475 0.8234 1.0000
3.750 0.5701 0.02801 0.01839 -0.0474 0.8012 1.0000
4.000 0.6147 0.02786 0.01849 -0.0486 0.7799 1.0000
4.250 0.6422 0.02805 0.01888 -0.0473 0.7551 1.0000
4.500 0.6755 0.02773 0.01878 -0.0458 0.7294 1.0000
4.750 0.7057 0.02719 0.01841 -0.0432 0.7007 1.0000
5.000 0.7316 0.02643 0.01783 -0.0394 0.6671 1.0000
5.250 0.7573 0.02524 0.01669 -0.0349 0.6287 1.0000
5.500 0.7781 0.02432 0.01579 -0.0303 0.5852 1.0000
5.750 0.7950 0.02367 0.01512 -0.0255 0.5326 1.0000
6.000 0.8067 0.02320 0.01443 -0.0199 0.4557 1.0000
6.250 0.8074 0.02416 0.01450 -0.0136 0.3289 1.0000
6.500 0.8153 0.02643 0.01590 -0.0097 0.2424 1.0000
6.750 0.8323 0.02854 0.01760 -0.0076 0.1989 1.0000
7.000 0.8539 0.03068 0.01957 -0.0061 0.1712 1.0000
7.250 0.8788 0.03319 0.02192 -0.0052 0.1525 1.0000
7.500 0.9006 0.03563 0.02450 -0.0040 0.1383 1.0000
7.750 0.9216 0.03832 0.02755 -0.0024 0.1297 1.0000
8.000 0.9390 0.04133 0.03086 -0.0009 0.1227 1.0000
8.250 0.9529 0.04433 0.03430 0.0011 0.1172 1.0000
8.500 0.9666 0.04787 0.03819 0.0028 0.1150 1.0000
8.750 0.9819 0.05206 0.04247 0.0039 0.1123 1.0000
9.000 0.9835 0.05585 0.04678 0.0064 0.1114 1.0000
9.250 0.9809 0.05985 0.05125 0.0087 0.1110 1.0000
9.500 0.9749 0.06412 0.05590 0.0108 0.1109 1.0000
9.750 0.9721 0.06881 0.06081 0.0122 0.1117 1.0000
10.000 0.9120 0.07430 0.06692 0.0144 0.1189 1.0000
10.250 0.8855 0.07948 0.07217 0.0147 0.1220 1.0000
10.500 0.8690 0.08501 0.07772 0.0136 0.1241 1.0000
10.750 0.8602 0.09088 0.08363 0.0120 0.1264 1.0000
11.000 0.7629 0.11271 0.10516 -0.0093 0.1610 1.0000
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Polar data table (+)
Polar graphs
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