NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il) Reynolds number: 100,000 Max Cl/Cd: 45.72 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca651212a06-il-100000-n5.txt Download as CSV file: xf-naca651212a06-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65(1)-212 a=0.6
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5196 0.09394 0.08895 -0.0366 1.0000 0.0340
-10.500 -0.5255 0.08758 0.08266 -0.0407 1.0000 0.0337
-10.250 -0.5349 0.08049 0.07559 -0.0459 1.0000 0.0330
-10.000 -0.5507 0.07381 0.06889 -0.0509 1.0000 0.0323
-9.750 -0.5709 0.06824 0.06326 -0.0543 1.0000 0.0323
-9.500 -0.5899 0.06401 0.05895 -0.0552 1.0000 0.0318
-9.250 -0.6077 0.06068 0.05553 -0.0540 1.0000 0.0317
-9.000 -0.6218 0.05740 0.05211 -0.0523 1.0000 0.0316
-8.750 -0.6384 0.05502 0.04961 -0.0486 1.0000 0.0316
-8.500 -0.6330 0.05052 0.04473 -0.0498 0.9859 0.0315
-8.250 -0.6193 0.04598 0.03971 -0.0515 0.9751 0.0315
-8.000 -0.6029 0.04203 0.03526 -0.0524 0.9655 0.0317
-7.750 -0.5838 0.03866 0.03137 -0.0529 0.9568 0.0321
-7.500 -0.5608 0.03595 0.02813 -0.0535 0.9497 0.0338
-7.250 -0.5396 0.03383 0.02552 -0.0531 0.9409 0.0350
-7.000 -0.5137 0.03167 0.02289 -0.0534 0.9349 0.0356
-6.750 -0.4909 0.02990 0.02079 -0.0528 0.9267 0.0361
-6.500 -0.4642 0.02758 0.01825 -0.0531 0.9211 0.0371
-6.250 -0.4398 0.02613 0.01670 -0.0529 0.9143 0.0385
-6.000 -0.4146 0.02509 0.01557 -0.0527 0.9079 0.0410
-5.750 -0.3887 0.02404 0.01438 -0.0525 0.9022 0.0433
-5.500 -0.3647 0.02303 0.01324 -0.0518 0.8954 0.0448
-5.250 -0.3398 0.02211 0.01221 -0.0512 0.8902 0.0464
-5.000 -0.3190 0.02121 0.01130 -0.0502 0.8835 0.0485
-4.750 -0.2974 0.02052 0.01061 -0.0494 0.8776 0.0528
-4.500 -0.2748 0.01995 0.00992 -0.0486 0.8724 0.0572
-4.250 -0.2542 0.01937 0.00926 -0.0475 0.8658 0.0616
-4.000 -0.2316 0.01883 0.00866 -0.0466 0.8609 0.0702
-3.750 -0.2102 0.01823 0.00813 -0.0457 0.8556 0.0892
-3.500 -0.1944 0.01703 0.00757 -0.0443 0.8494 0.2066
-3.250 -0.1860 0.01543 0.00747 -0.0411 0.8445 0.5311
-3.000 -0.1673 0.01543 0.00772 -0.0388 0.8391 0.6371
-2.750 -0.1491 0.01567 0.00815 -0.0358 0.8339 0.7065
-2.500 -0.1308 0.01598 0.00849 -0.0325 0.8297 0.7593
-2.250 -0.1080 0.01599 0.00842 -0.0310 0.8252 0.7826
-2.000 -0.0826 0.01593 0.00825 -0.0307 0.8200 0.7920
-1.750 -0.0558 0.01583 0.00804 -0.0306 0.8161 0.8006
-1.500 -0.0272 0.01576 0.00787 -0.0306 0.8129 0.8079
-1.250 -0.0009 0.01580 0.00786 -0.0304 0.8076 0.8151
-1.000 0.0265 0.01583 0.00785 -0.0304 0.8031 0.8225
-0.750 0.0563 0.01588 0.00786 -0.0306 0.7997 0.8284
-0.500 0.0862 0.01591 0.00784 -0.0308 0.7969 0.8354
-0.250 0.1140 0.01614 0.00808 -0.0309 0.7913 0.8414
0.000 0.1432 0.01627 0.00821 -0.0311 0.7871 0.8485
0.250 0.1735 0.01638 0.00829 -0.0315 0.7838 0.8560
0.500 0.2028 0.01648 0.00838 -0.0317 0.7804 0.8654
0.750 0.2172 0.01671 0.00861 -0.0298 0.7740 0.8890
1.000 0.2389 0.01677 0.00865 -0.0288 0.7698 0.9056
1.250 0.2683 0.01674 0.00863 -0.0291 0.7668 0.9163
1.500 0.2976 0.01689 0.00883 -0.0299 0.7616 0.9264
1.750 0.3307 0.01698 0.00899 -0.0313 0.7571 0.9355
2.000 0.3660 0.01701 0.00906 -0.0329 0.7536 0.9443
2.250 0.4059 0.01701 0.00912 -0.0354 0.7510 0.9504
2.500 0.4403 0.01727 0.00951 -0.0375 0.7449 0.9599
2.750 0.4803 0.01737 0.00972 -0.0403 0.7406 0.9667
3.000 0.5194 0.01738 0.00983 -0.0427 0.7370 0.9743
3.250 0.5569 0.01753 0.01015 -0.0452 0.7309 0.9829
3.500 0.5949 0.01753 0.01030 -0.0475 0.7242 0.9914
3.750 0.6318 0.01744 0.01035 -0.0493 0.7171 1.0000
4.000 0.6503 0.01701 0.01000 -0.0468 0.7005 1.0000
4.250 0.6675 0.01595 0.00888 -0.0429 0.6638 1.0000
4.500 0.6824 0.01532 0.00814 -0.0393 0.6127 1.0000
4.750 0.6973 0.01525 0.00801 -0.0364 0.5600 1.0000
5.000 0.7069 0.01555 0.00765 -0.0324 0.4276 1.0000
5.250 0.6987 0.01729 0.00833 -0.0267 0.2573 1.0000
5.500 0.6955 0.01904 0.00927 -0.0224 0.1295 1.0000
5.750 0.7016 0.02031 0.01014 -0.0193 0.0795 1.0000
6.000 0.7130 0.02126 0.01098 -0.0169 0.0641 1.0000
6.250 0.7268 0.02207 0.01184 -0.0148 0.0567 1.0000
6.750 0.7525 0.02386 0.01373 -0.0107 0.0475 1.0000
7.000 0.7658 0.02476 0.01467 -0.0088 0.0440 1.0000
7.250 0.7775 0.02580 0.01573 -0.0067 0.0419 1.0000
7.500 0.7904 0.02715 0.01705 -0.0048 0.0402 1.0000
7.750 0.8104 0.02827 0.01827 -0.0037 0.0388 1.0000
8.000 0.8326 0.02944 0.01958 -0.0031 0.0367 1.0000
8.250 0.8552 0.03068 0.02090 -0.0026 0.0345 1.0000
8.500 0.8798 0.03215 0.02245 -0.0024 0.0329 1.0000
8.750 0.9060 0.03391 0.02429 -0.0025 0.0319 1.0000
9.000 0.9331 0.03616 0.02665 -0.0028 0.0310 1.0000
9.250 0.9587 0.03905 0.02975 -0.0029 0.0304 1.0000
9.500 0.9759 0.04125 0.03229 -0.0018 0.0300 1.0000
9.750 0.9886 0.04353 0.03494 -0.0002 0.0296 1.0000
10.000 0.9971 0.04584 0.03763 0.0017 0.0290 1.0000
10.250 1.0015 0.04823 0.04040 0.0040 0.0282 1.0000
10.500 1.0013 0.05066 0.04316 0.0066 0.0275 1.0000
10.750 0.9969 0.05335 0.04615 0.0093 0.0272 1.0000
11.000 0.9886 0.05630 0.04939 0.0119 0.0270 1.0000
11.250 0.9780 0.05947 0.05283 0.0141 0.0270 1.0000
11.500 0.9641 0.06291 0.05653 0.0158 0.0270 1.0000
11.750 0.9485 0.06670 0.06057 0.0168 0.0270 1.0000
12.000 0.9303 0.07097 0.06506 0.0171 0.0271 1.0000
12.250 0.9099 0.07578 0.07008 0.0166 0.0272 1.0000
12.500 0.8878 0.08127 0.07575 0.0150 0.0273 1.0000
12.750 0.8643 0.08759 0.08224 0.0123 0.0275 1.0000
13.000 0.8386 0.09520 0.09000 0.0081 0.0278 1.0000
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Polar data table (+)
Polar graphs
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