NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63-212 AIRFOIL (n63212-il) Reynolds number: 100,000 Max Cl/Cd: 46.87 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63212-il-100000-n5.txt Download as CSV file: xf-n63212-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-212 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.6133 0.09081 0.08555 -0.0275 1.0000 0.0294
-11.250 -0.6318 0.08213 0.07683 -0.0337 1.0000 0.0292
-11.000 -0.6529 0.07453 0.06916 -0.0393 1.0000 0.0289
-10.750 -0.6702 0.06874 0.06328 -0.0432 1.0000 0.0285
-10.500 -0.6922 0.06336 0.05774 -0.0464 1.0000 0.0286
-10.250 -0.7097 0.05921 0.05345 -0.0476 1.0000 0.0284
-10.000 -0.7266 0.05576 0.04981 -0.0471 1.0000 0.0284
-9.750 -0.7378 0.05231 0.04609 -0.0463 1.0000 0.0286
-9.500 -0.7416 0.04899 0.04249 -0.0454 1.0000 0.0286
-9.250 -0.7412 0.04574 0.03890 -0.0443 1.0000 0.0287
-9.000 -0.7362 0.04266 0.03547 -0.0432 1.0000 0.0288
-8.750 -0.7274 0.03988 0.03233 -0.0421 1.0000 0.0289
-8.500 -0.7158 0.03729 0.02939 -0.0409 1.0000 0.0291
-8.250 -0.7018 0.03500 0.02675 -0.0397 1.0000 0.0293
-8.000 -0.6861 0.03299 0.02443 -0.0384 1.0000 0.0296
-7.750 -0.6696 0.03111 0.02234 -0.0371 1.0000 0.0300
-7.500 -0.6534 0.02931 0.02046 -0.0358 1.0000 0.0306
-7.250 -0.6377 0.02792 0.01901 -0.0342 1.0000 0.0312
-7.000 -0.6233 0.02681 0.01786 -0.0325 1.0000 0.0323
-6.750 -0.6099 0.02590 0.01688 -0.0305 1.0000 0.0337
-6.500 -0.5857 0.02479 0.01566 -0.0306 0.9956 0.0356
-6.250 -0.5519 0.02354 0.01427 -0.0322 0.9874 0.0376
-6.000 -0.5198 0.02217 0.01288 -0.0339 0.9790 0.0397
-5.750 -0.4866 0.02106 0.01173 -0.0357 0.9710 0.0431
-5.500 -0.4524 0.02008 0.01062 -0.0376 0.9635 0.0488
-5.250 -0.4213 0.01913 0.00964 -0.0390 0.9541 0.0583
-5.000 -0.3896 0.01819 0.00881 -0.0405 0.9457 0.0797
-4.750 -0.3608 0.01694 0.00801 -0.0418 0.9365 0.1474
-4.500 -0.3378 0.01528 0.00732 -0.0425 0.9261 0.3235
-4.250 -0.3094 0.01467 0.00720 -0.0429 0.9182 0.4551
-4.000 -0.2827 0.01455 0.00718 -0.0425 0.9079 0.5192
-3.750 -0.2550 0.01457 0.00725 -0.0420 0.8990 0.5708
-3.500 -0.2280 0.01467 0.00737 -0.0413 0.8902 0.6075
-3.250 -0.2018 0.01478 0.00742 -0.0405 0.8807 0.6336
-3.000 -0.1738 0.01486 0.00742 -0.0399 0.8730 0.6539
-2.750 -0.1478 0.01489 0.00738 -0.0393 0.8630 0.6686
-2.500 -0.1200 0.01487 0.00725 -0.0390 0.8551 0.6804
-2.250 -0.0930 0.01484 0.00711 -0.0387 0.8459 0.6921
-2.000 -0.0663 0.01483 0.00704 -0.0381 0.8376 0.7011
-1.750 -0.0392 0.01479 0.00692 -0.0378 0.8291 0.7108
-1.500 -0.0124 0.01477 0.00683 -0.0375 0.8208 0.7203
-1.250 0.0143 0.01474 0.00676 -0.0370 0.8127 0.7287
-1.000 0.0412 0.01472 0.00667 -0.0368 0.8041 0.7388
-0.750 0.0676 0.01471 0.00663 -0.0362 0.7965 0.7465
-0.500 0.0941 0.01471 0.00660 -0.0359 0.7878 0.7560
-0.250 0.1206 0.01470 0.00656 -0.0354 0.7805 0.7645
0.000 0.1466 0.01471 0.00657 -0.0349 0.7718 0.7736
0.250 0.1734 0.01471 0.00654 -0.0345 0.7647 0.7831
0.500 0.1989 0.01474 0.00660 -0.0339 0.7559 0.7918
0.750 0.2257 0.01474 0.00658 -0.0336 0.7489 0.8019
1.000 0.2509 0.01479 0.00667 -0.0329 0.7402 0.8109
1.250 0.2773 0.01479 0.00668 -0.0324 0.7332 0.8206
1.500 0.3030 0.01486 0.00679 -0.0320 0.7244 0.8314
1.750 0.3288 0.01486 0.00682 -0.0312 0.7175 0.8407
2.000 0.3538 0.01493 0.00696 -0.0306 0.7084 0.8514
2.250 0.3801 0.01494 0.00700 -0.0300 0.7015 0.8628
2.500 0.4046 0.01500 0.00716 -0.0293 0.6919 0.8738
2.750 0.4304 0.01501 0.00722 -0.0286 0.6841 0.8852
3.000 0.4559 0.01505 0.00737 -0.0280 0.6745 0.8977
3.250 0.4823 0.01508 0.00748 -0.0276 0.6651 0.9108
3.500 0.5105 0.01506 0.00752 -0.0273 0.6558 0.9243
3.750 0.5396 0.01507 0.00767 -0.0275 0.6434 0.9385
4.000 0.5718 0.01503 0.00770 -0.0282 0.6284 0.9525
4.250 0.6067 0.01492 0.00765 -0.0294 0.6097 0.9670
4.750 0.6682 0.01486 0.00767 -0.0306 0.5535 1.0000
5.000 0.6898 0.01496 0.00768 -0.0294 0.5164 1.0000
5.250 0.7115 0.01518 0.00780 -0.0284 0.4668 1.0000
5.500 0.7320 0.01562 0.00798 -0.0273 0.3991 1.0000
5.750 0.7479 0.01658 0.00838 -0.0258 0.2972 1.0000
6.000 0.7606 0.01808 0.00921 -0.0244 0.1910 1.0000
6.250 0.7756 0.01950 0.01019 -0.0232 0.1230 1.0000
6.500 0.7923 0.02073 0.01118 -0.0222 0.0895 1.0000
6.750 0.8101 0.02184 0.01218 -0.0212 0.0733 1.0000
7.000 0.8277 0.02292 0.01324 -0.0202 0.0633 1.0000
7.250 0.8462 0.02389 0.01429 -0.0192 0.0564 1.0000
7.500 0.8621 0.02508 0.01547 -0.0180 0.0519 1.0000
7.750 0.8798 0.02616 0.01664 -0.0169 0.0480 1.0000
8.000 0.8973 0.02725 0.01777 -0.0159 0.0442 1.0000
8.250 0.9132 0.02860 0.01910 -0.0148 0.0418 1.0000
8.500 0.9318 0.03003 0.02061 -0.0139 0.0402 1.0000
8.750 0.9525 0.03145 0.02217 -0.0132 0.0383 1.0000
9.000 0.9728 0.03287 0.02371 -0.0125 0.0363 1.0000
9.250 0.9914 0.03425 0.02517 -0.0118 0.0344 1.0000
9.500 1.0101 0.03591 0.02686 -0.0113 0.0330 1.0000
9.750 1.0311 0.03816 0.02922 -0.0109 0.0320 1.0000
10.000 1.0492 0.04027 0.03164 -0.0102 0.0313 1.0000
10.250 1.0642 0.04265 0.03433 -0.0092 0.0307 1.0000
10.500 1.0751 0.04521 0.03723 -0.0080 0.0302 1.0000
10.750 1.0805 0.04791 0.04026 -0.0063 0.0298 1.0000
11.000 1.0809 0.05064 0.04330 -0.0044 0.0294 1.0000
11.250 1.0772 0.05346 0.04641 -0.0026 0.0290 1.0000
11.500 1.0704 0.05646 0.04970 -0.0012 0.0285 1.0000
11.750 1.0611 0.05973 0.05323 -0.0002 0.0282 1.0000
12.000 1.0497 0.06335 0.05709 0.0001 0.0279 1.0000
12.250 1.0355 0.06751 0.06149 -0.0003 0.0277 1.0000
12.500 1.0185 0.07228 0.06649 -0.0015 0.0275 1.0000
12.750 0.9982 0.07789 0.07232 -0.0038 0.0274 1.0000
13.000 0.9711 0.08520 0.07988 -0.0077 0.0276 1.0000
13.250 0.9333 0.09556 0.09047 -0.0142 0.0284 1.0000
13.500 0.8914 0.10905 0.10415 -0.0235 0.0293 1.0000
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