NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 66-210 (naca66210-il) Reynolds number: 100,000 Max Cl/Cd: 36.72 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca66210-il-100000-n5.txt Download as CSV file: xf-naca66210-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66-210
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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5075 0.08989 0.08486 -0.0420 1.0000 0.0362
-9.500 -0.5145 0.08429 0.07932 -0.0454 1.0000 0.0341
-9.250 -0.5291 0.07791 0.07299 -0.0502 1.0000 0.0325
-8.750 -0.5780 0.06947 0.06434 -0.0495 1.0000 0.0281
-8.500 -0.5920 0.06652 0.06140 -0.0467 1.0000 0.0278
-8.250 -0.6074 0.06405 0.05888 -0.0430 1.0000 0.0275
-8.000 -0.6218 0.06167 0.05643 -0.0390 1.0000 0.0273
-7.750 -0.6342 0.05912 0.05378 -0.0351 1.0000 0.0271
-7.500 -0.6303 0.05509 0.04949 -0.0348 0.9961 0.0269
-7.250 -0.6177 0.05049 0.04451 -0.0358 0.9900 0.0267
-7.000 -0.6013 0.04612 0.03970 -0.0366 0.9849 0.0268
-6.750 -0.5820 0.04305 0.03602 -0.0365 0.9791 0.0281
-6.500 -0.5640 0.03875 0.03124 -0.0371 0.9748 0.0296
-6.250 -0.5444 0.03573 0.02784 -0.0369 0.9697 0.0302
-6.000 -0.5195 0.03295 0.02464 -0.0372 0.9663 0.0308
-5.750 -0.4925 0.03060 0.02190 -0.0377 0.9635 0.0316
-5.500 -0.4698 0.02895 0.01995 -0.0372 0.9590 0.0333
-5.250 -0.4420 0.02748 0.01818 -0.0375 0.9557 0.0361
-5.000 -0.4111 0.02579 0.01618 -0.0381 0.9533 0.0374
-4.750 -0.3785 0.02425 0.01440 -0.0390 0.9515 0.0387
-4.500 -0.3549 0.02319 0.01315 -0.0382 0.9472 0.0399
-4.250 -0.3294 0.02203 0.01193 -0.0380 0.9439 0.0424
-4.000 -0.3027 0.02109 0.01104 -0.0383 0.9410 0.0468
-3.750 -0.2739 0.02029 0.01016 -0.0388 0.9386 0.0500
-3.500 -0.2492 0.01970 0.00943 -0.0385 0.9351 0.0533
-3.250 -0.2282 0.01913 0.00882 -0.0376 0.9303 0.0595
-3.000 -0.2015 0.01865 0.00831 -0.0377 0.9272 0.0707
-2.750 -0.1738 0.01803 0.00780 -0.0380 0.9248 0.1001
-2.500 -0.1755 0.01550 0.00796 -0.0329 0.9201 0.6311
-2.250 -0.1646 0.01601 0.00876 -0.0275 0.9152 0.7868
-2.000 -0.1466 0.01667 0.00938 -0.0234 0.9120 0.8415
-1.750 -0.1207 0.01726 0.00987 -0.0212 0.9102 0.8742
-1.500 -0.0844 0.01778 0.01027 -0.0212 0.9095 0.9013
-1.250 -0.0658 0.01784 0.01023 -0.0196 0.9045 0.9114
-1.000 -0.0343 0.01787 0.01012 -0.0207 0.9015 0.9161
-0.750 -0.0060 0.01786 0.01002 -0.0213 0.8987 0.9218
-0.500 0.0255 0.01785 0.00992 -0.0225 0.8964 0.9258
-0.250 0.0620 0.01785 0.00984 -0.0247 0.8948 0.9284
0.000 0.0802 0.01792 0.00989 -0.0234 0.8891 0.9341
0.250 0.1062 0.01794 0.00988 -0.0236 0.8853 0.9391
0.500 0.1426 0.01795 0.00988 -0.0257 0.8830 0.9415
0.750 0.1794 0.01794 0.00988 -0.0279 0.8809 0.9438
1.000 0.1989 0.01806 0.01002 -0.0269 0.8744 0.9500
1.250 0.2310 0.01809 0.01009 -0.0282 0.8707 0.9535
1.500 0.2687 0.01809 0.01016 -0.0306 0.8681 0.9557
1.750 0.2989 0.01817 0.01032 -0.0316 0.8635 0.9595
2.000 0.3256 0.01825 0.01049 -0.0319 0.8578 0.9646
2.250 0.3648 0.01823 0.01061 -0.0344 0.8545 0.9666
2.500 0.3999 0.01826 0.01078 -0.0362 0.8494 0.9696
2.750 0.4324 0.01823 0.01091 -0.0374 0.8423 0.9734
3.000 0.4772 0.01703 0.00989 -0.0384 0.8218 0.9727
3.250 0.5118 0.01559 0.00853 -0.0370 0.7828 0.9744
3.500 0.5387 0.01467 0.00762 -0.0350 0.7148 0.9782
3.750 0.5535 0.01529 0.00656 -0.0307 0.3760 0.9825
4.000 0.5620 0.01780 0.00742 -0.0288 0.0898 0.9899
4.250 0.5879 0.01867 0.00822 -0.0290 0.0653 0.9958
4.500 0.6090 0.01937 0.00896 -0.0283 0.0558 1.0000
4.750 0.6177 0.01990 0.00952 -0.0251 0.0505 1.0000
5.000 0.6244 0.02058 0.01024 -0.0215 0.0476 1.0000
5.250 0.6334 0.02121 0.01098 -0.0183 0.0455 1.0000
5.500 0.6455 0.02200 0.01184 -0.0156 0.0433 1.0000
5.750 0.6618 0.02288 0.01277 -0.0139 0.0400 1.0000
6.000 0.6806 0.02414 0.01398 -0.0128 0.0368 1.0000
6.250 0.7057 0.02579 0.01568 -0.0125 0.0353 1.0000
6.500 0.7324 0.02728 0.01734 -0.0123 0.0343 1.0000
6.750 0.7595 0.02906 0.01934 -0.0121 0.0334 1.0000
7.000 0.7854 0.03111 0.02171 -0.0117 0.0327 1.0000
7.250 0.8084 0.03321 0.02416 -0.0109 0.0317 1.0000
7.500 0.8281 0.03508 0.02635 -0.0099 0.0296 1.0000
7.750 0.8462 0.03724 0.02881 -0.0087 0.0283 1.0000
8.000 0.8617 0.03990 0.03185 -0.0072 0.0277 1.0000
8.250 0.8743 0.04280 0.03514 -0.0054 0.0274 1.0000
8.500 0.8834 0.04595 0.03869 -0.0033 0.0272 1.0000
8.750 0.8886 0.04935 0.04248 -0.0010 0.0271 1.0000
9.000 0.8935 0.05230 0.04572 0.0009 0.0265 1.0000
9.250 0.8955 0.05531 0.04893 0.0028 0.0257 1.0000
9.500 0.8910 0.05916 0.05298 0.0048 0.0250 1.0000
9.750 0.8779 0.06384 0.05786 0.0072 0.0246 1.0000
10.000 0.8634 0.06623 0.06052 0.0106 0.0243 1.0000
10.250 0.8479 0.06956 0.06404 0.0127 0.0242 1.0000
10.500 0.8301 0.07316 0.06784 0.0139 0.0241 1.0000
10.750 0.8132 0.07751 0.07232 0.0139 0.0242 1.0000
11.000 0.7962 0.08237 0.07730 0.0127 0.0243 1.0000
11.250 0.7806 0.08791 0.08291 0.0106 0.0245 1.0000
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Polar data table (+)
Polar graphs
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