M6 (65%) (m665-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: M6 (65%) (m665-il) Reynolds number: 50,000 Max Cl/Cd: 33.8 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m665-il-50000.txt Download as CSV file: xf-m665-il-50000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5301 0.09891 0.09206 0.0043 1.0000 0.2580
-8.000 -0.5302 0.09578 0.08901 0.0045 1.0000 0.2732
-7.750 -0.5361 0.09301 0.08633 0.0046 1.0000 0.2891
-7.500 -0.5202 0.08871 0.08205 0.0066 1.0000 0.3100
-7.250 -0.5204 0.08583 0.07925 0.0077 1.0000 0.3310
-7.000 -0.5145 0.08240 0.07587 0.0094 1.0000 0.3537
-6.750 -0.5338 0.08074 0.07436 0.0112 1.0000 0.3780
-6.500 -0.5007 0.07640 0.06997 0.0149 1.0000 0.4186
-6.250 -0.5047 0.07448 0.06816 0.0198 1.0000 0.4661
-6.000 -0.4369 0.06959 0.06304 0.0251 1.0000 0.5637
-4.750 -0.4779 0.04334 0.03483 -0.0104 1.0000 0.1600
-4.500 -0.4598 0.03981 0.03105 -0.0088 1.0000 0.1511
-4.250 -0.4412 0.03666 0.02726 -0.0068 1.0000 0.1410
-4.000 -0.4221 0.03382 0.02402 -0.0050 1.0000 0.1371
-3.750 -0.4024 0.03156 0.02131 -0.0031 1.0000 0.1390
-3.500 -0.3813 0.02960 0.01887 -0.0013 1.0000 0.1427
-3.250 -0.3579 0.02760 0.01648 0.0002 1.0000 0.1449
-3.000 -0.3335 0.02570 0.01453 0.0012 1.0000 0.1517
-2.750 -0.1073 0.01694 0.00865 -0.0300 1.0000 1.0000
-2.500 -0.0879 0.01680 0.00794 -0.0284 1.0000 1.0000
-2.250 -0.0702 0.01668 0.00745 -0.0267 1.0000 1.0000
-2.000 -0.0535 0.01660 0.00712 -0.0250 1.0000 1.0000
-1.750 -0.0379 0.01655 0.00687 -0.0230 1.0000 1.0000
-1.500 -0.0234 0.01655 0.00671 -0.0210 1.0000 1.0000
-1.250 -0.0104 0.01661 0.00663 -0.0187 1.0000 1.0000
-1.000 0.0005 0.01673 0.00663 -0.0162 1.0000 1.0000
-0.750 0.0089 0.01692 0.00673 -0.0133 1.0000 1.0000
-0.500 0.0151 0.01720 0.00691 -0.0103 1.0000 1.0000
-0.250 0.0207 0.01755 0.00716 -0.0073 1.0000 1.0000
0.000 0.0265 0.01794 0.00746 -0.0045 1.0000 1.0000
0.250 0.0332 0.01838 0.00780 -0.0020 1.0000 1.0000
0.500 0.0410 0.01885 0.00818 0.0003 1.0000 1.0000
0.750 0.0496 0.01936 0.00861 0.0022 1.0000 1.0000
1.000 0.0590 0.01991 0.00909 0.0040 1.0000 1.0000
1.250 0.0691 0.02049 0.00962 0.0055 1.0000 1.0000
1.500 0.0936 0.02123 0.01033 0.0041 0.9953 1.0000
1.750 0.1534 0.02219 0.01133 -0.0040 0.9763 1.0000
2.000 0.2128 0.02308 0.01229 -0.0115 0.9576 1.0000
2.250 0.2662 0.02381 0.01314 -0.0177 0.9381 1.0000
2.500 0.3188 0.02444 0.01392 -0.0234 0.9178 1.0000
2.750 0.3822 0.02492 0.01466 -0.0306 0.8985 1.0000
3.000 0.4305 0.02535 0.01530 -0.0348 0.8766 1.0000
3.250 0.4942 0.02538 0.01564 -0.0408 0.8549 1.0000
3.500 0.5329 0.02562 0.01616 -0.0422 0.8300 1.0000
3.750 0.5731 0.02570 0.01650 -0.0431 0.8053 1.0000
4.000 0.6127 0.02557 0.01665 -0.0432 0.7810 1.0000
4.250 0.6365 0.02583 0.01717 -0.0411 0.7543 1.0000
4.500 0.6629 0.02558 0.01714 -0.0380 0.7240 1.0000
4.750 0.6843 0.02452 0.01618 -0.0322 0.6810 1.0000
5.000 0.7004 0.02319 0.01486 -0.0253 0.6280 1.0000
5.250 0.7140 0.02208 0.01375 -0.0189 0.5676 1.0000
5.500 0.7195 0.02129 0.01272 -0.0115 0.4631 1.0000
5.750 0.7096 0.02384 0.01324 -0.0039 0.2331 1.0000
6.000 0.7223 0.02645 0.01519 -0.0011 0.1732 1.0000
6.250 0.7428 0.02851 0.01702 0.0007 0.1469 1.0000
6.500 0.7684 0.03061 0.01909 0.0019 0.1326 1.0000
6.750 0.7933 0.03284 0.02126 0.0028 0.1213 1.0000
7.000 0.8170 0.03516 0.02404 0.0042 0.1160 1.0000
7.250 0.8394 0.03787 0.02710 0.0056 0.1132 1.0000
7.500 0.8589 0.04090 0.03051 0.0071 0.1119 1.0000
7.750 0.8759 0.04407 0.03401 0.0087 0.1103 1.0000
8.000 0.8919 0.04786 0.03791 0.0099 0.1081 1.0000
8.250 0.9007 0.05135 0.04190 0.0119 0.1077 1.0000
8.500 0.9057 0.05502 0.04613 0.0141 0.1090 1.0000
8.750 0.8905 0.05981 0.05175 0.0168 0.1136 1.0000
9.000 0.8816 0.06473 0.05702 0.0182 0.1171 1.0000
9.250 0.8830 0.06962 0.06206 0.0191 0.1203 1.0000
9.500 0.8458 0.07495 0.06771 0.0194 0.1251 1.0000
9.750 0.8063 0.08077 0.07357 0.0184 0.1287 1.0000
10.000 0.8129 0.08634 0.07918 0.0181 0.1359 1.0000
10.250 0.7612 0.09784 0.09058 0.0079 0.1501 1.0000
10.500 0.6775 0.12250 0.11489 -0.0192 0.3049 1.0000
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Polar data table (+)
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