NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: NACA 63-212 AIRFOIL (n63212-il) Reynolds number: 200,000 Max Cl/Cd: 65.64 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63212-il-200000.txt Download as CSV file: xf-n63212-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6188 0.07649 0.07291 -0.0428 1.0000 0.0706
-9.750 -0.6488 0.07203 0.06836 -0.0456 1.0000 0.0708
-9.500 -0.6830 0.06955 0.06566 -0.0455 1.0000 0.0712
-9.250 -0.7000 0.06768 0.06347 -0.0448 1.0000 0.0716
-8.000 -0.6962 0.03595 0.02970 -0.0383 1.0000 0.0386
-7.750 -0.6850 0.03261 0.02623 -0.0365 1.0000 0.0376
-7.500 -0.6777 0.03021 0.02362 -0.0338 1.0000 0.0371
-7.250 -0.6705 0.02846 0.02166 -0.0309 1.0000 0.0370
-7.000 -0.6616 0.02715 0.02012 -0.0283 1.0000 0.0374
-6.750 -0.6319 0.02574 0.01839 -0.0292 0.9966 0.0382
-6.500 -0.5946 0.02372 0.01613 -0.0315 0.9918 0.0383
-6.250 -0.5586 0.02134 0.01362 -0.0334 0.9872 0.0390
-6.000 -0.5211 0.01947 0.01177 -0.0357 0.9833 0.0406
-5.750 -0.4857 0.01831 0.01062 -0.0377 0.9766 0.0430
-5.500 -0.4458 0.01727 0.00953 -0.0404 0.9721 0.0469
-5.250 -0.4109 0.01606 0.00837 -0.0425 0.9655 0.0543
-5.000 -0.3738 0.01490 0.00721 -0.0449 0.9593 0.0712
-4.750 -0.3419 0.01349 0.00616 -0.0467 0.9515 0.1273
-4.500 -0.3163 0.01131 0.00541 -0.0484 0.9433 0.4125
-4.250 -0.2886 0.01102 0.00546 -0.0483 0.9336 0.5200
-4.000 -0.2564 0.01099 0.00543 -0.0489 0.9268 0.5643
-3.750 -0.2302 0.01101 0.00543 -0.0482 0.9162 0.5910
-3.500 -0.2026 0.01105 0.00544 -0.0477 0.9075 0.6126
-3.250 -0.1763 0.01115 0.00552 -0.0469 0.8986 0.6364
-3.000 -0.1515 0.01131 0.00568 -0.0457 0.8889 0.6591
-2.750 -0.1258 0.01143 0.00577 -0.0446 0.8813 0.6770
-2.500 -0.1008 0.01153 0.00585 -0.0436 0.8711 0.6921
-2.250 -0.0749 0.01158 0.00585 -0.0428 0.8631 0.7052
-2.000 -0.0488 0.01159 0.00580 -0.0422 0.8540 0.7161
-1.750 -0.0229 0.01160 0.00579 -0.0415 0.8456 0.7249
-1.500 0.0038 0.01158 0.00571 -0.0410 0.8374 0.7351
-1.250 0.0299 0.01159 0.00571 -0.0404 0.8287 0.7436
-1.000 0.0565 0.01157 0.00563 -0.0399 0.8211 0.7526
-0.750 0.0831 0.01158 0.00563 -0.0396 0.8122 0.7618
-0.500 0.1095 0.01156 0.00557 -0.0390 0.8050 0.7698
0.000 0.1625 0.01157 0.00555 -0.0380 0.7891 0.7874
0.250 0.1893 0.01158 0.00558 -0.0378 0.7799 0.7970
0.500 0.2155 0.01159 0.00556 -0.0371 0.7733 0.8056
0.750 0.2415 0.01161 0.00562 -0.0367 0.7640 0.8148
1.000 0.2684 0.01162 0.00560 -0.0362 0.7575 0.8249
1.250 0.2933 0.01165 0.00570 -0.0355 0.7481 0.8335
1.500 0.3197 0.01166 0.00569 -0.0350 0.7414 0.8435
1.750 0.3447 0.01169 0.00578 -0.0343 0.7320 0.8537
2.000 0.3697 0.01170 0.00581 -0.0333 0.7247 0.8632
2.250 0.3946 0.01170 0.00587 -0.0326 0.7155 0.8740
2.500 0.4191 0.01172 0.00592 -0.0317 0.7071 0.8852
2.750 0.4427 0.01169 0.00594 -0.0304 0.6983 0.8959
3.000 0.4660 0.01168 0.00599 -0.0292 0.6885 0.9077
3.250 0.4899 0.01160 0.00592 -0.0280 0.6785 0.9202
3.500 0.5141 0.01147 0.00581 -0.0268 0.6662 0.9336
3.750 0.5406 0.01133 0.00572 -0.0262 0.6519 0.9474
4.000 0.5720 0.01119 0.00559 -0.0265 0.6356 0.9605
4.250 0.6076 0.01102 0.00546 -0.0279 0.6121 0.9728
4.500 0.6471 0.01093 0.00536 -0.0302 0.5870 0.9840
4.750 0.6885 0.01091 0.00535 -0.0330 0.5593 0.9953
5.000 0.7113 0.01098 0.00542 -0.0325 0.5312 1.0000
5.250 0.7312 0.01114 0.00553 -0.0314 0.4959 1.0000
5.500 0.7526 0.01150 0.00568 -0.0305 0.4327 1.0000
5.750 0.7671 0.01265 0.00611 -0.0290 0.2945 1.0000
6.000 0.7745 0.01493 0.00731 -0.0273 0.1272 1.0000
6.250 0.7909 0.01640 0.00845 -0.0261 0.0822 1.0000
6.500 0.8105 0.01746 0.00944 -0.0253 0.0681 1.0000
6.750 0.8284 0.01866 0.01060 -0.0242 0.0606 1.0000
7.000 0.8491 0.01956 0.01151 -0.0234 0.0548 1.0000
7.250 0.8663 0.02100 0.01290 -0.0222 0.0504 1.0000
7.500 0.8882 0.02197 0.01395 -0.0215 0.0475 1.0000
7.750 0.9100 0.02311 0.01511 -0.0208 0.0450 1.0000
8.000 0.9325 0.02445 0.01642 -0.0203 0.0430 1.0000
8.250 0.9582 0.02693 0.01890 -0.0202 0.0409 1.0000
8.500 0.9810 0.02791 0.02008 -0.0196 0.0394 1.0000
8.750 1.0048 0.02951 0.02185 -0.0191 0.0382 1.0000
9.000 1.0281 0.03145 0.02399 -0.0186 0.0373 1.0000
9.250 1.0496 0.03369 0.02648 -0.0180 0.0368 1.0000
9.500 1.0683 0.03630 0.02941 -0.0170 0.0366 1.0000
9.750 1.0831 0.03928 0.03274 -0.0157 0.0366 1.0000
10.000 1.0944 0.04225 0.03604 -0.0143 0.0363 1.0000
10.250 1.1034 0.04492 0.03899 -0.0128 0.0357 1.0000
10.500 1.1014 0.04909 0.04360 -0.0104 0.0364 1.0000
10.750 1.0917 0.05389 0.04883 -0.0078 0.0378 1.0000
11.000 1.0803 0.05808 0.05327 -0.0054 0.0390 1.0000
11.250 1.0683 0.06261 0.05796 -0.0036 0.0400 1.0000
11.500 0.9641 0.05850 0.05435 0.0021 0.0396 1.0000
11.750 0.9427 0.06425 0.06025 0.0020 0.0402 1.0000
12.000 0.8525 0.07773 0.07435 -0.0026 0.0462 1.0000
12.250 0.8097 0.08774 0.08453 -0.0076 0.0479 1.0000
12.500 0.7834 0.09636 0.09321 -0.0118 0.0495 1.0000
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Polar data table (+)
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