M6 (85%) (m685-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: M6 (85%) (m685-il) Reynolds number: 50,000 Max Cl/Cd: 33.18 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m685-il-50000.txt Download as CSV file: xf-m685-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: M6 (85%)
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 -0.4839 0.12284 0.11553 -0.0063 1.0000 0.2201
-10.000 -0.4641 0.11768 0.11036 -0.0052 1.0000 0.2308
-9.750 -0.4894 0.11723 0.11005 -0.0078 1.0000 0.2368
-9.500 -0.4651 0.11182 0.10462 -0.0062 1.0000 0.2506
-9.250 -0.4559 0.10773 0.10056 -0.0059 1.0000 0.2600
-9.000 -0.4847 0.10725 0.10025 -0.0079 1.0000 0.2692
-8.750 -0.4635 0.10244 0.09542 -0.0064 1.0000 0.2840
-8.500 -0.4498 0.09838 0.09136 -0.0055 1.0000 0.2972
-8.250 -0.4449 0.09504 0.08807 -0.0048 1.0000 0.3113
-8.000 -0.4428 0.09201 0.08511 -0.0040 1.0000 0.3285
-7.750 -0.4478 0.08953 0.08272 -0.0030 1.0000 0.3478
-7.500 -0.4339 0.08597 0.07919 -0.0010 1.0000 0.3720
-7.250 -0.4441 0.08418 0.07751 0.0011 1.0000 0.3977
-7.000 -0.4358 0.08142 0.07480 0.0040 1.0000 0.4304
-6.750 -0.4094 0.07770 0.07105 0.0066 1.0000 0.4667
-6.500 -0.3889 0.07487 0.06823 0.0100 1.0000 0.5147
-5.250 -0.5229 0.04971 0.04221 -0.0079 1.0000 0.1813
-5.000 -0.5242 0.04617 0.03777 -0.0042 1.0000 0.1577
-4.750 -0.5138 0.04327 0.03479 -0.0018 1.0000 0.1535
-4.500 -0.5064 0.04082 0.03209 0.0012 1.0000 0.1507
-4.250 -0.4983 0.03864 0.02956 0.0042 1.0000 0.1493
-4.000 -0.4885 0.03667 0.02725 0.0070 1.0000 0.1489
-3.750 -0.4765 0.03486 0.02510 0.0095 1.0000 0.1494
-3.500 -0.4629 0.03330 0.02319 0.0118 1.0000 0.1518
-3.250 -0.4474 0.03188 0.02138 0.0139 1.0000 0.1548
-3.000 -0.4297 0.03058 0.01965 0.0158 1.0000 0.1572
-2.750 -0.0689 0.02009 0.01201 -0.0350 1.0000 1.0000
-2.500 -0.0665 0.02028 0.01195 -0.0315 1.0000 1.0000
-2.250 -0.0694 0.02060 0.01209 -0.0273 1.0000 1.0000
-2.000 -0.0736 0.02100 0.01233 -0.0231 1.0000 1.0000
-1.750 -0.0764 0.02142 0.01259 -0.0193 1.0000 1.0000
-1.500 -0.0770 0.02185 0.01284 -0.0158 1.0000 1.0000
-1.250 -0.0756 0.02227 0.01309 -0.0127 1.0000 1.0000
-1.000 -0.0724 0.02272 0.01337 -0.0099 1.0000 1.0000
-0.750 -0.0676 0.02318 0.01367 -0.0074 1.0000 1.0000
-0.500 -0.0616 0.02366 0.01399 -0.0051 1.0000 1.0000
-0.250 -0.0117 0.02438 0.01449 -0.0112 0.9872 1.0000
0.000 0.0411 0.02508 0.01500 -0.0175 0.9727 1.0000
0.250 0.0884 0.02568 0.01547 -0.0226 0.9575 1.0000
0.500 0.1341 0.02627 0.01594 -0.0272 0.9427 1.0000
0.750 0.1795 0.02683 0.01642 -0.0316 0.9281 1.0000
1.000 0.2257 0.02735 0.01689 -0.0360 0.9139 1.0000
1.250 0.2757 0.02782 0.01733 -0.0408 0.9002 1.0000
1.500 0.3158 0.02829 0.01781 -0.0438 0.8854 1.0000
1.750 0.3533 0.02878 0.01831 -0.0461 0.8706 1.0000
2.000 0.3895 0.02929 0.01885 -0.0482 0.8560 1.0000
2.250 0.4240 0.02980 0.01941 -0.0498 0.8412 1.0000
2.500 0.4549 0.03034 0.02001 -0.0506 0.8263 1.0000
2.750 0.4835 0.03091 0.02064 -0.0510 0.8114 1.0000
3.000 0.5111 0.03148 0.02128 -0.0511 0.7963 1.0000
3.250 0.5382 0.03203 0.02191 -0.0509 0.7808 1.0000
3.500 0.5642 0.03256 0.02253 -0.0503 0.7651 1.0000
3.750 0.5893 0.03307 0.02315 -0.0495 0.7490 1.0000
4.000 0.6130 0.03361 0.02378 -0.0483 0.7328 1.0000
4.250 0.6360 0.03413 0.02440 -0.0470 0.7163 1.0000
4.500 0.6587 0.03462 0.02503 -0.0454 0.6996 1.0000
4.750 0.6810 0.03510 0.02563 -0.0437 0.6828 1.0000
5.000 0.7040 0.03549 0.02614 -0.0420 0.6657 1.0000
5.250 0.7272 0.03582 0.02663 -0.0401 0.6485 1.0000
5.500 0.7520 0.03598 0.02695 -0.0381 0.6311 1.0000
5.750 0.7807 0.03580 0.02693 -0.0362 0.6132 1.0000
6.000 0.7942 0.03630 0.02758 -0.0330 0.5909 1.0000
6.250 0.8357 0.03385 0.02524 -0.0300 0.5633 1.0000
6.500 0.8670 0.03191 0.02338 -0.0264 0.5327 1.0000
6.750 0.8899 0.03069 0.02223 -0.0227 0.4998 1.0000
7.000 0.9111 0.02908 0.02059 -0.0185 0.4582 1.0000
7.250 0.9241 0.02795 0.01932 -0.0135 0.4049 1.0000
7.500 0.9255 0.02789 0.01879 -0.0074 0.3300 1.0000
7.750 0.9199 0.02953 0.01964 -0.0015 0.2531 1.0000
8.000 0.9220 0.03169 0.02129 0.0026 0.2036 1.0000
8.250 0.9335 0.03381 0.02312 0.0053 0.1732 1.0000
8.500 0.9541 0.03598 0.02500 0.0067 0.1531 1.0000
8.750 0.9743 0.03811 0.02713 0.0080 0.1391 1.0000
9.000 0.9940 0.04061 0.02990 0.0095 0.1307 1.0000
9.250 1.0161 0.04323 0.03247 0.0103 0.1232 1.0000
9.500 1.0245 0.04579 0.03551 0.0129 0.1187 1.0000
9.750 1.0362 0.04857 0.03858 0.0148 0.1154 1.0000
10.000 1.0473 0.05168 0.04193 0.0166 0.1136 1.0000
10.250 1.0549 0.05503 0.04552 0.0186 0.1124 1.0000
10.500 1.0607 0.05881 0.04949 0.0204 0.1112 1.0000
10.750 1.0544 0.06249 0.05346 0.0232 0.1107 1.0000
11.000 1.0464 0.06621 0.05746 0.0258 0.1108 1.0000
11.250 1.0319 0.06992 0.06142 0.0285 0.1109 1.0000
11.500 0.9358 0.07466 0.06670 0.0336 0.1169 1.0000
11.750 0.8931 0.08138 0.07355 0.0319 0.1193 1.0000
12.000 0.8585 0.08912 0.08131 0.0285 0.1217 1.0000
12.250 0.8389 0.09644 0.08864 0.0256 0.1234 1.0000
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Polar data table (+)
Polar graphs
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