M6 (65%) (m665-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: M6 (65%) (m665-il) Reynolds number: 50,000 Max Cl/Cd: 33.07 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m665-il-50000-n5.txt Download as CSV file: xf-m665-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: M6 (65%)
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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5425 0.11174 0.10453 -0.0007 1.0000 0.1106
-9.250 -0.5445 0.10820 0.10107 -0.0029 1.0000 0.1130
-8.750 -0.5574 0.09462 0.08758 -0.0150 1.0000 0.0596
-8.500 -0.5526 0.09074 0.08372 -0.0151 1.0000 0.0584
-8.250 -0.5552 0.08667 0.07970 -0.0164 1.0000 0.0577
-7.750 -0.5583 0.07843 0.07144 -0.0186 1.0000 0.0552
-7.500 -0.5592 0.07404 0.06701 -0.0198 1.0000 0.0538
-7.250 -0.5599 0.06940 0.06228 -0.0208 1.0000 0.0522
-6.750 -0.5607 0.05911 0.05126 -0.0212 1.0000 0.0482
-6.500 -0.5530 0.05531 0.04720 -0.0201 1.0000 0.0480
-6.000 -0.5326 0.04894 0.04060 -0.0178 1.0000 0.0505
-5.750 -0.5207 0.04609 0.03752 -0.0163 1.0000 0.0522
-5.500 -0.5081 0.04297 0.03404 -0.0146 1.0000 0.0530
-5.250 -0.4941 0.04008 0.03077 -0.0127 1.0000 0.0546
-5.000 -0.4788 0.03739 0.02760 -0.0108 1.0000 0.0571
-4.750 -0.4615 0.03473 0.02444 -0.0088 1.0000 0.0584
-4.500 -0.4421 0.03226 0.02142 -0.0069 1.0000 0.0594
-4.250 -0.4208 0.03003 0.01879 -0.0054 1.0000 0.0608
-4.000 -0.3989 0.02812 0.01674 -0.0043 1.0000 0.0631
-3.750 -0.3763 0.02670 0.01517 -0.0032 1.0000 0.0677
-3.500 -0.3515 0.02541 0.01351 -0.0022 1.0000 0.0734
-3.250 -0.3273 0.02414 0.01231 -0.0016 1.0000 0.0813
-3.000 -0.2992 0.02299 0.01091 -0.0012 1.0000 0.0887
-2.750 -0.2737 0.02202 0.00990 -0.0007 1.0000 0.0991
-2.500 -0.2509 0.02117 0.00903 0.0002 1.0000 0.1150
-2.250 -0.0707 0.01668 0.00747 -0.0267 1.0000 1.0000
-2.000 -0.0539 0.01659 0.00712 -0.0249 1.0000 1.0000
-1.750 -0.0381 0.01654 0.00687 -0.0230 1.0000 1.0000
-1.500 -0.0236 0.01654 0.00669 -0.0210 1.0000 1.0000
-1.250 -0.0104 0.01659 0.00661 -0.0187 1.0000 1.0000
-1.000 0.0006 0.01671 0.00661 -0.0162 1.0000 1.0000
-0.750 0.0091 0.01690 0.00671 -0.0134 1.0000 1.0000
-0.500 0.0213 0.01714 0.00685 -0.0115 0.9973 1.0000
-0.250 0.0677 0.01727 0.00682 -0.0160 0.9785 1.0000
0.000 0.1133 0.01738 0.00681 -0.0201 0.9607 1.0000
0.250 0.1586 0.01745 0.00681 -0.0241 0.9434 1.0000
0.500 0.2036 0.01750 0.00681 -0.0278 0.9265 1.0000
0.750 0.2468 0.01753 0.00682 -0.0310 0.9091 1.0000
1.000 0.2858 0.01757 0.00686 -0.0333 0.8903 1.0000
1.250 0.3209 0.01764 0.00694 -0.0347 0.8707 1.0000
1.500 0.3542 0.01772 0.00703 -0.0356 0.8519 1.0000
1.750 0.3835 0.01785 0.00721 -0.0356 0.8324 1.0000
2.000 0.4101 0.01802 0.00740 -0.0351 0.8128 1.0000
2.250 0.4361 0.01819 0.00762 -0.0344 0.7941 1.0000
2.500 0.4608 0.01839 0.00786 -0.0334 0.7757 1.0000
2.750 0.4836 0.01864 0.00821 -0.0321 0.7562 1.0000
3.000 0.5067 0.01887 0.00851 -0.0308 0.7372 1.0000
3.250 0.5298 0.01908 0.00879 -0.0293 0.7183 1.0000
3.500 0.5511 0.01935 0.00917 -0.0277 0.6968 1.0000
3.750 0.5732 0.01956 0.00952 -0.0260 0.6758 1.0000
4.000 0.5948 0.01982 0.00991 -0.0243 0.6535 1.0000
4.250 0.6170 0.02005 0.01027 -0.0226 0.6313 1.0000
4.500 0.6377 0.02027 0.01063 -0.0206 0.6044 1.0000
4.750 0.6554 0.02033 0.01077 -0.0177 0.5642 1.0000
5.000 0.6696 0.02034 0.01064 -0.0140 0.5017 1.0000
5.250 0.6825 0.02064 0.01065 -0.0105 0.4163 1.0000
5.500 0.6937 0.02157 0.01099 -0.0073 0.2920 1.0000
6.000 0.7122 0.02579 0.01377 -0.0027 0.1041 1.0000
6.250 0.7251 0.02746 0.01532 -0.0006 0.0867 1.0000
6.500 0.7395 0.02892 0.01682 0.0013 0.0753 1.0000
6.750 0.7542 0.03051 0.01844 0.0034 0.0693 1.0000
7.000 0.7727 0.03195 0.02012 0.0052 0.0643 1.0000
7.250 0.7906 0.03352 0.02172 0.0066 0.0594 1.0000
7.500 0.8110 0.03524 0.02360 0.0080 0.0554 1.0000
7.750 0.8328 0.03715 0.02579 0.0092 0.0529 1.0000
8.000 0.8534 0.03935 0.02827 0.0105 0.0511 1.0000
8.250 0.8719 0.04176 0.03102 0.0118 0.0498 1.0000
8.500 0.8877 0.04438 0.03396 0.0132 0.0487 1.0000
8.750 0.9007 0.04722 0.03713 0.0147 0.0480 1.0000
9.000 0.9108 0.05017 0.04037 0.0163 0.0471 1.0000
9.500 0.9182 0.05680 0.04759 0.0198 0.0453 1.0000
9.750 0.9129 0.06013 0.05135 0.0219 0.0447 1.0000
10.000 0.9036 0.06366 0.05522 0.0238 0.0443 1.0000
10.250 0.8904 0.06722 0.05903 0.0255 0.0441 1.0000
10.500 0.8735 0.07074 0.06272 0.0270 0.0442 1.0000
10.750 0.8551 0.07481 0.06693 0.0270 0.0443 1.0000
11.000 0.8359 0.07968 0.07192 0.0253 0.0445 1.0000
11.250 0.8176 0.08529 0.07761 0.0224 0.0448 1.0000
11.500 0.8015 0.09143 0.08379 0.0189 0.0451 1.0000
11.750 0.7882 0.09785 0.09023 0.0152 0.0455 1.0000
12.000 0.7753 0.10486 0.09723 0.0111 0.0457 1.0000
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Polar data table (+)
Polar graphs
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