NACA M12 AIRFOIL (m12-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M12 AIRFOIL (m12-il) Reynolds number: 100,000 Max Cl/Cd: 49.03 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m12-il-100000.txt Download as CSV file: xf-m12-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M12 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 -0.4873 0.09552 0.09037 -0.0197 1.0000 0.1480
-8.750 -0.5268 0.09031 0.08529 -0.0288 1.0000 0.1545
-8.500 -0.5504 0.08503 0.08006 -0.0314 1.0000 0.1568
-8.250 -0.5075 0.08284 0.07787 -0.0267 1.0000 0.1608
-8.000 -0.5136 0.07935 0.07442 -0.0275 1.0000 0.1657
-7.750 -0.5628 0.07426 0.06926 -0.0316 1.0000 0.1729
-7.500 -0.5365 0.07135 0.06645 -0.0291 1.0000 0.1763
-7.250 -0.5558 0.06883 0.06382 -0.0289 1.0000 0.1865
-7.000 -0.5533 0.06499 0.06004 -0.0271 1.0000 0.1909
-6.750 -0.5438 0.06280 0.05790 -0.0248 1.0000 0.1974
-6.500 -0.5544 0.06005 0.05509 -0.0224 1.0000 0.2079
-6.250 -0.5916 0.04449 0.03772 -0.0208 1.0000 0.1110
-6.000 -0.5809 0.04151 0.03463 -0.0186 1.0000 0.1051
-5.750 -0.5772 0.03752 0.02958 -0.0145 1.0000 0.0950
-5.500 -0.5643 0.03485 0.02673 -0.0126 1.0000 0.0940
-5.250 -0.5501 0.03270 0.02432 -0.0107 1.0000 0.0934
-5.000 -0.5345 0.03108 0.02235 -0.0089 1.0000 0.0941
-4.750 -0.5175 0.02979 0.02071 -0.0073 1.0000 0.0950
-4.500 -0.4986 0.02819 0.01886 -0.0060 1.0000 0.0957
-4.250 -0.4779 0.02662 0.01715 -0.0052 1.0000 0.0967
-4.000 -0.4565 0.02549 0.01594 -0.0045 1.0000 0.0984
-3.750 -0.4345 0.02462 0.01501 -0.0039 0.9998 0.1007
-3.500 -0.3883 0.02375 0.01399 -0.0076 0.9936 0.1071
-3.250 -0.3453 0.02262 0.01289 -0.0109 0.9866 0.1137
-3.000 -0.3011 0.02184 0.01207 -0.0141 0.9796 0.1241
-2.750 -0.2593 0.02075 0.01125 -0.0172 0.9723 0.1453
-2.500 -0.2225 0.01937 0.01037 -0.0193 0.9642 0.2087
-2.250 -0.1943 0.01671 0.01008 -0.0196 0.9585 0.6305
-2.000 -0.1009 0.01781 0.01177 -0.0265 0.9641 0.9619
-1.750 0.0241 0.01806 0.01162 -0.0449 0.9718 1.0000
-1.500 0.0740 0.01787 0.01123 -0.0501 0.9609 1.0000
-1.250 0.1271 0.01763 0.01083 -0.0557 0.9519 1.0000
-1.000 0.1778 0.01735 0.01042 -0.0606 0.9420 1.0000
-0.750 0.2191 0.01713 0.01010 -0.0636 0.9289 1.0000
-0.500 0.2570 0.01692 0.00981 -0.0658 0.9154 1.0000
-0.250 0.2906 0.01675 0.00956 -0.0669 0.9014 1.0000
0.000 0.3201 0.01661 0.00935 -0.0671 0.8870 1.0000
0.250 0.3464 0.01650 0.00918 -0.0666 0.8724 1.0000
0.500 0.3708 0.01641 0.00903 -0.0655 0.8580 1.0000
0.750 0.3934 0.01637 0.00893 -0.0642 0.8432 1.0000
1.000 0.4149 0.01638 0.00888 -0.0627 0.8278 1.0000
1.250 0.4363 0.01640 0.00887 -0.0612 0.8127 1.0000
1.500 0.4577 0.01644 0.00886 -0.0597 0.7978 1.0000
1.750 0.4794 0.01649 0.00886 -0.0582 0.7831 1.0000
2.000 0.5012 0.01653 0.00887 -0.0567 0.7685 1.0000
2.250 0.5231 0.01657 0.00886 -0.0552 0.7539 1.0000
2.500 0.5454 0.01661 0.00886 -0.0537 0.7395 1.0000
2.750 0.5680 0.01665 0.00886 -0.0523 0.7255 1.0000
3.000 0.5911 0.01669 0.00884 -0.0509 0.7120 1.0000
3.250 0.6132 0.01683 0.00897 -0.0496 0.6972 1.0000
3.500 0.6354 0.01699 0.00914 -0.0484 0.6825 1.0000
3.750 0.6578 0.01715 0.00930 -0.0471 0.6680 1.0000
4.000 0.6803 0.01732 0.00946 -0.0458 0.6536 1.0000
4.250 0.7029 0.01748 0.00963 -0.0446 0.6391 1.0000
4.500 0.7257 0.01765 0.00979 -0.0433 0.6246 1.0000
4.750 0.7485 0.01781 0.00995 -0.0420 0.6096 1.0000
5.000 0.7712 0.01798 0.01011 -0.0407 0.5941 1.0000
5.250 0.7935 0.01814 0.01029 -0.0394 0.5774 1.0000
5.500 0.8154 0.01829 0.01043 -0.0379 0.5590 1.0000
5.750 0.8372 0.01838 0.01049 -0.0363 0.5392 1.0000
6.000 0.8595 0.01847 0.01050 -0.0348 0.5189 1.0000
6.250 0.8790 0.01864 0.01070 -0.0330 0.4959 1.0000
6.500 0.8992 0.01877 0.01078 -0.0312 0.4725 1.0000
6.750 0.9179 0.01896 0.01097 -0.0293 0.4476 1.0000
7.000 0.9368 0.01917 0.01113 -0.0274 0.4228 1.0000
7.250 0.9541 0.01946 0.01145 -0.0254 0.3954 1.0000
7.500 0.9708 0.01980 0.01177 -0.0234 0.3666 1.0000
7.750 0.9861 0.02024 0.01217 -0.0211 0.3343 1.0000
8.000 0.9987 0.02085 0.01263 -0.0186 0.2968 1.0000
8.250 1.0073 0.02182 0.01336 -0.0156 0.2502 1.0000
8.500 1.0119 0.02326 0.01447 -0.0122 0.2000 1.0000
8.750 1.0170 0.02484 0.01574 -0.0090 0.1644 1.0000
9.000 1.0254 0.02636 0.01706 -0.0063 0.1430 1.0000
9.250 1.0369 0.02779 0.01837 -0.0041 0.1285 1.0000
9.500 1.0512 0.02931 0.01973 -0.0023 0.1180 1.0000
9.750 1.0679 0.03065 0.02111 -0.0008 0.1094 1.0000
10.000 1.0875 0.03232 0.02276 0.0001 0.1023 1.0000
10.250 1.1072 0.03382 0.02430 0.0011 0.0963 1.0000
10.500 1.1311 0.03588 0.02634 0.0012 0.0910 1.0000
10.750 1.1477 0.03752 0.02821 0.0026 0.0866 1.0000
11.000 1.1695 0.03944 0.03012 0.0030 0.0826 1.0000
11.250 1.1885 0.04207 0.03291 0.0037 0.0793 1.0000
11.500 1.1969 0.04411 0.03530 0.0059 0.0767 1.0000
11.750 1.2068 0.04646 0.03792 0.0076 0.0745 1.0000
12.000 1.2179 0.04873 0.04034 0.0091 0.0722 1.0000
12.250 1.2358 0.05224 0.04384 0.0091 0.0696 1.0000
12.500 1.2280 0.05502 0.04699 0.0123 0.0687 1.0000
12.750 1.2140 0.05756 0.04989 0.0162 0.0681 1.0000
13.000 1.1985 0.06062 0.05326 0.0193 0.0678 1.0000
13.250 1.1810 0.06413 0.05707 0.0216 0.0677 1.0000
13.500 1.1610 0.06810 0.06132 0.0230 0.0677 1.0000
13.750 1.1384 0.07263 0.06610 0.0235 0.0678 1.0000
14.000 1.1138 0.07779 0.07149 0.0231 0.0680 1.0000
14.250 1.0876 0.08365 0.07755 0.0216 0.0683 1.0000
14.500 1.0611 0.09026 0.08434 0.0193 0.0687 1.0000
14.750 1.0355 0.09755 0.09177 0.0162 0.0691 1.0000
15.000 0.8811 0.13343 0.12800 -0.0082 0.0804 1.0000
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Polar data table (+)
Polar graphs
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