NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 66-210 (naca66210-il) Reynolds number: 200,000 Max Cl/Cd: 51.39 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca66210-il-200000.txt Download as CSV file: xf-naca66210-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4029 0.09135 0.08798 -0.0394 1.0000 0.0439
-10.000 -0.4064 0.08692 0.08357 -0.0409 1.0000 0.0450
-9.750 -0.4125 0.08212 0.07880 -0.0429 1.0000 0.0464
-9.500 -0.4205 0.07710 0.07382 -0.0453 1.0000 0.0472
-9.250 -0.4325 0.07160 0.06837 -0.0483 1.0000 0.0477
-9.000 -0.4493 0.06602 0.06282 -0.0515 1.0000 0.0475
-8.750 -0.4705 0.06212 0.05895 -0.0516 1.0000 0.0471
-8.500 -0.4992 0.06023 0.05711 -0.0477 1.0000 0.0465
-8.250 -0.5306 0.05947 0.05640 -0.0420 1.0000 0.0455
-8.000 -0.5613 0.05849 0.05544 -0.0364 1.0000 0.0447
-7.750 -0.5750 0.05483 0.05168 -0.0357 0.9977 0.0455
-7.500 -0.5738 0.04938 0.04592 -0.0384 0.9924 0.0484
-7.250 -0.5677 0.04729 0.04312 -0.0394 0.9853 0.0503
-7.000 -0.6123 0.05711 0.05242 -0.0341 0.9878 0.0503
-6.250 -0.5155 0.02748 0.02288 -0.0427 0.9716 0.0680
-6.000 -0.5413 0.03971 0.03456 -0.0387 0.9734 0.0700
-5.750 -0.5161 0.03682 0.03128 -0.0400 0.9706 0.0806
-5.500 -0.4996 0.03525 0.02935 -0.0390 0.9656 0.0929
-5.250 -0.4764 0.03291 0.02703 -0.0392 0.9623 0.1003
-5.000 -0.4426 0.02789 0.02091 -0.0374 0.9598 0.0661
-4.750 -0.4066 0.02453 0.01703 -0.0370 0.9584 0.0476
-4.500 -0.3729 0.02230 0.01447 -0.0378 0.9570 0.0459
-4.250 -0.3455 0.02091 0.01290 -0.0375 0.9542 0.0461
-4.000 -0.3225 0.02018 0.01202 -0.0366 0.9501 0.0486
-3.750 -0.2938 0.01944 0.01117 -0.0368 0.9475 0.0500
-3.500 -0.2659 0.01796 0.00973 -0.0370 0.9456 0.0524
-3.250 -0.2358 0.01714 0.00892 -0.0378 0.9437 0.0561
-3.000 -0.2022 0.01669 0.00843 -0.0393 0.9422 0.0633
-2.750 -0.1927 0.01622 0.00799 -0.0361 0.9360 0.0692
-2.500 -0.1669 0.01578 0.00755 -0.0361 0.9331 0.0854
-2.250 -0.1624 0.01296 0.00748 -0.0326 0.9302 0.7063
-2.000 -0.1398 0.01352 0.00822 -0.0297 0.9282 0.8161
-1.750 -0.1390 0.01399 0.00872 -0.0234 0.9218 0.8438
-1.500 -0.1300 0.01468 0.00946 -0.0175 0.9180 0.8788
-1.250 -0.1117 0.01526 0.01002 -0.0136 0.9156 0.9065
-1.000 -0.0775 0.01567 0.01035 -0.0137 0.9145 0.9236
-0.750 -0.0247 0.01613 0.01072 -0.0179 0.9147 0.9385
-0.500 -0.0082 0.01622 0.01077 -0.0162 0.9093 0.9464
-0.250 0.0317 0.01628 0.01077 -0.0192 0.9068 0.9493
0.000 0.0715 0.01630 0.01075 -0.0221 0.9046 0.9521
0.250 0.1107 0.01629 0.01072 -0.0249 0.9026 0.9546
0.500 0.1477 0.01627 0.01068 -0.0272 0.9008 0.9572
0.750 0.1832 0.01633 0.01075 -0.0294 0.8975 0.9591
1.000 0.2156 0.01641 0.01085 -0.0310 0.8929 0.9619
1.250 0.2565 0.01639 0.01087 -0.0342 0.8903 0.9635
1.500 0.2992 0.01629 0.01083 -0.0375 0.8881 0.9646
1.750 0.3438 0.01613 0.01074 -0.0410 0.8862 0.9653
2.000 0.3604 0.01628 0.01095 -0.0394 0.8777 0.9713
2.250 0.4199 0.01512 0.00989 -0.0439 0.8675 0.9688
2.500 0.4638 0.01382 0.00866 -0.0449 0.8504 0.9689
2.750 0.4978 0.01247 0.00730 -0.0437 0.8276 0.9704
3.000 0.5280 0.01146 0.00632 -0.0425 0.7911 0.9725
3.250 0.5565 0.01083 0.00565 -0.0414 0.7343 0.9755
3.500 0.5529 0.01223 0.00523 -0.0346 0.3590 0.9828
3.750 0.5627 0.01478 0.00626 -0.0330 0.0751 0.9882
4.000 0.5900 0.01551 0.00700 -0.0334 0.0619 0.9923
4.250 0.6164 0.01645 0.00796 -0.0338 0.0545 0.9963
4.500 0.6462 0.01723 0.00878 -0.0348 0.0493 1.0000
4.750 0.6562 0.01790 0.00945 -0.0318 0.0471 1.0000
5.000 0.6667 0.01891 0.01044 -0.0289 0.0452 1.0000
5.250 0.6869 0.02091 0.01242 -0.0276 0.0437 1.0000
5.500 0.7024 0.02163 0.01328 -0.0253 0.0428 1.0000
5.750 0.7159 0.02224 0.01404 -0.0227 0.0409 1.0000
6.000 0.7346 0.02362 0.01557 -0.0209 0.0404 1.0000
6.250 0.7541 0.02541 0.01756 -0.0191 0.0405 1.0000
6.500 0.7737 0.02777 0.02018 -0.0174 0.0414 1.0000
6.750 0.7982 0.03102 0.02356 -0.0167 0.0446 1.0000
7.000 0.8238 0.03878 0.03296 -0.0104 0.0919 1.0000
7.250 0.8406 0.04090 0.03515 -0.0093 0.0855 1.0000
7.500 0.8472 0.04331 0.03794 -0.0069 0.0759 1.0000
7.750 0.8634 0.04606 0.04070 -0.0062 0.0717 1.0000
8.000 0.8641 0.04951 0.04453 -0.0036 0.0660 1.0000
8.250 0.8715 0.05179 0.04708 -0.0015 0.0611 1.0000
8.500 0.8834 0.05479 0.05010 -0.0008 0.0586 1.0000
8.750 0.8890 0.06485 0.05996 -0.0020 0.0561 1.0000
9.000 0.8751 0.06358 0.05929 0.0032 0.0548 1.0000
9.250 0.8607 0.06551 0.06162 0.0069 0.0512 1.0000
9.500 0.8491 0.06861 0.06487 0.0092 0.0497 1.0000
9.750 0.8344 0.07168 0.06805 0.0114 0.0488 1.0000
10.000 0.8172 0.07528 0.07174 0.0125 0.0483 1.0000
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Polar data table (+)
Polar graphs
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