NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-212 AIRFOIL (n63212-il) Reynolds number: 50,000 Max Cl/Cd: 33.46 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63212-il-50000.txt Download as CSV file: xf-n63212-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4822 0.10997 0.10290 -0.0018 1.0000 0.3363
-9.250 -0.4630 0.10475 0.09767 -0.0023 1.0000 0.3392
-9.000 -0.5882 0.08212 0.07542 -0.0339 1.0000 0.1705
-8.750 -0.5953 0.07455 0.06785 -0.0372 1.0000 0.1512
-8.500 -0.6355 0.06726 0.06038 -0.0413 1.0000 0.1385
-8.250 -0.6627 0.06099 0.05367 -0.0428 1.0000 0.1286
-8.000 -0.6596 0.05687 0.04928 -0.0424 1.0000 0.1251
-7.750 -0.6542 0.05285 0.04505 -0.0417 1.0000 0.1230
-7.500 -0.6492 0.04901 0.04086 -0.0407 1.0000 0.1205
-7.250 -0.6428 0.04534 0.03674 -0.0393 1.0000 0.1177
-7.000 -0.6336 0.04209 0.03295 -0.0376 1.0000 0.1156
-6.750 -0.6204 0.03927 0.02970 -0.0358 1.0000 0.1152
-6.500 -0.6045 0.03677 0.02693 -0.0342 1.0000 0.1161
-6.250 -0.5879 0.03469 0.02467 -0.0326 1.0000 0.1199
-6.000 -0.5712 0.03286 0.02247 -0.0308 1.0000 0.1253
-5.750 -0.5526 0.03084 0.02041 -0.0291 1.0000 0.1314
-5.500 -0.5332 0.02927 0.01866 -0.0273 1.0000 0.1407
-5.250 -0.5148 0.02770 0.01717 -0.0254 1.0000 0.1554
-5.000 -0.4983 0.02612 0.01585 -0.0233 1.0000 0.1775
-4.750 -0.4838 0.02451 0.01451 -0.0212 1.0000 0.2126
-4.500 -0.4750 0.02162 0.01289 -0.0189 1.0000 0.3211
-4.250 -0.4838 0.02190 0.01482 -0.0085 1.0000 0.6065
-4.000 -0.4819 0.02333 0.01622 -0.0003 1.0000 0.6696
-3.750 -0.4787 0.02453 0.01738 0.0076 1.0000 0.7108
-3.500 -0.4758 0.02543 0.01817 0.0150 1.0000 0.7504
-3.250 -0.4718 0.02636 0.01903 0.0230 1.0000 0.7880
-3.000 -0.4638 0.02697 0.01954 0.0297 1.0000 0.8279
-2.750 -0.4325 0.02768 0.02000 0.0326 1.0000 0.8696
-2.500 -0.3735 0.02796 0.01988 0.0283 1.0000 0.9024
-2.250 -0.2948 0.02786 0.01937 0.0186 1.0000 0.9245
-2.000 -0.2427 0.02751 0.01873 0.0125 1.0000 0.9435
-1.750 -0.1863 0.02714 0.01812 0.0052 1.0000 0.9603
-1.500 -0.1303 0.02677 0.01755 -0.0025 1.0000 0.9762
-1.250 -0.0766 0.02646 0.01707 -0.0100 1.0000 0.9920
-1.000 -0.0532 0.02628 0.01681 -0.0126 1.0000 1.0000
-0.750 -0.0627 0.02611 0.01663 -0.0095 1.0000 1.0000
-0.500 -0.0728 0.02587 0.01635 -0.0063 1.0000 1.0000
-0.250 -0.0836 0.02555 0.01601 -0.0030 1.0000 1.0000
0.000 -0.0949 0.02514 0.01558 0.0005 1.0000 1.0000
0.250 -0.0765 0.02525 0.01562 -0.0012 0.9931 1.0000
0.500 -0.0461 0.02567 0.01596 -0.0047 0.9822 1.0000
0.750 -0.0174 0.02615 0.01635 -0.0076 0.9713 1.0000
1.000 0.0070 0.02657 0.01670 -0.0096 0.9606 1.0000
1.250 0.0369 0.02720 0.01726 -0.0123 0.9503 1.0000
1.500 0.0739 0.02807 0.01807 -0.0160 0.9403 1.0000
1.750 0.1043 0.02880 0.01876 -0.0185 0.9296 1.0000
2.000 0.1338 0.02960 0.01952 -0.0207 0.9191 1.0000
2.250 0.1692 0.03056 0.02047 -0.0237 0.9087 1.0000
2.500 0.2036 0.03151 0.02144 -0.0264 0.8977 1.0000
2.750 0.2289 0.03237 0.02231 -0.0276 0.8862 1.0000
3.000 0.2583 0.03334 0.02332 -0.0294 0.8747 1.0000
3.250 0.2922 0.03438 0.02443 -0.0317 0.8627 1.0000
3.500 0.3305 0.03545 0.02559 -0.0346 0.8502 1.0000
3.750 0.3557 0.03641 0.02662 -0.0354 0.8366 1.0000
4.000 0.3829 0.03741 0.02774 -0.0364 0.8223 1.0000
4.250 0.4130 0.03841 0.02886 -0.0376 0.8070 1.0000
4.500 0.4464 0.03934 0.02995 -0.0391 0.7907 1.0000
4.750 0.4890 0.04009 0.03093 -0.0412 0.7730 1.0000
5.000 0.5088 0.04097 0.03195 -0.0405 0.7529 1.0000
5.250 0.5487 0.04130 0.03253 -0.0414 0.7304 1.0000
5.500 0.5900 0.04120 0.03275 -0.0417 0.7052 1.0000
5.750 0.6551 0.03896 0.03096 -0.0416 0.6729 1.0000
6.000 0.7400 0.03185 0.02440 -0.0368 0.6265 1.0000
6.250 0.7874 0.02666 0.01947 -0.0303 0.5698 1.0000
6.500 0.8006 0.02393 0.01651 -0.0221 0.4429 1.0000
6.750 0.7898 0.02631 0.01691 -0.0153 0.2619 1.0000
7.000 0.7998 0.02913 0.01891 -0.0128 0.2017 1.0000
7.250 0.8263 0.03143 0.02091 -0.0121 0.1687 1.0000
7.500 0.8596 0.03374 0.02314 -0.0121 0.1489 1.0000
7.750 0.8907 0.03618 0.02562 -0.0121 0.1355 1.0000
8.000 0.9229 0.03919 0.02859 -0.0124 0.1278 1.0000
8.250 0.9437 0.04176 0.03167 -0.0114 0.1218 1.0000
8.500 0.9680 0.04471 0.03470 -0.0111 0.1165 1.0000
8.750 0.9902 0.04877 0.03897 -0.0108 0.1147 1.0000
9.000 1.0031 0.05238 0.04309 -0.0095 0.1143 1.0000
9.250 1.0108 0.05613 0.04733 -0.0080 0.1138 1.0000
9.500 1.0140 0.06004 0.05170 -0.0065 0.1135 1.0000
9.750 1.0135 0.06421 0.05626 -0.0051 0.1136 1.0000
10.000 1.0107 0.06866 0.06102 -0.0040 0.1140 1.0000
10.250 0.9800 0.07280 0.06572 -0.0020 0.1172 1.0000
10.500 0.9270 0.07837 0.07161 -0.0012 0.1222 1.0000
10.750 0.8986 0.08447 0.07782 -0.0030 0.1250 1.0000
11.000 0.8850 0.09065 0.08404 -0.0050 0.1272 1.0000
11.250 0.7821 0.11226 0.10556 -0.0242 0.1491 1.0000
11.500 0.8074 0.11671 0.11008 -0.0225 0.1572 1.0000
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Polar data table (+)
Polar graphs
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