NACA 63-412 AIRFOIL (n63412-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-412 AIRFOIL (n63412-il) Reynolds number: 100,000 Max Cl/Cd: 56.64 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63412-il-100000.txt Download as CSV file: xf-n63412-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-412 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4645 0.09508 0.09024 -0.0421 1.0000 0.1434
-9.250 -0.4258 0.09255 0.08760 -0.0366 1.0000 0.1486
-9.000 -0.4586 0.08890 0.08412 -0.0421 1.0000 0.1568
-8.750 -0.4385 0.08534 0.08055 -0.0393 1.0000 0.1611
-8.500 -0.4366 0.08262 0.07788 -0.0390 1.0000 0.1683
-8.250 -0.5188 0.07808 0.07362 -0.0468 1.0000 0.1722
-8.000 -0.4524 0.07664 0.07211 -0.0383 1.0000 0.1827
-7.750 -0.5190 0.07267 0.06836 -0.0423 1.0000 0.1865
-7.500 -0.4863 0.07128 0.06703 -0.0359 1.0000 0.1947
-6.750 -0.5919 0.04422 0.03794 -0.0449 1.0000 0.0835
-6.500 -0.5809 0.04042 0.03396 -0.0440 1.0000 0.0803
-6.250 -0.5672 0.03727 0.03041 -0.0434 1.0000 0.0793
-6.000 -0.5501 0.03439 0.02709 -0.0428 1.0000 0.0783
-5.750 -0.5300 0.03160 0.02377 -0.0423 1.0000 0.0765
-5.500 -0.5084 0.02948 0.02124 -0.0417 1.0000 0.0764
-5.250 -0.4687 0.02769 0.01912 -0.0445 0.9951 0.0801
-5.000 -0.4289 0.02617 0.01719 -0.0468 0.9893 0.0838
-4.750 -0.3894 0.02434 0.01526 -0.0492 0.9844 0.0875
-4.500 -0.3531 0.02332 0.01423 -0.0512 0.9780 0.0959
-4.250 -0.3148 0.02209 0.01310 -0.0534 0.9724 0.1048
-4.000 -0.2796 0.02114 0.01225 -0.0553 0.9658 0.1207
-3.750 -0.2422 0.01996 0.01128 -0.0577 0.9594 0.1573
-3.500 -0.2109 0.01769 0.01123 -0.0594 0.9543 0.5682
-3.250 -0.1834 0.01819 0.01177 -0.0584 0.9453 0.6331
-3.000 -0.1516 0.01877 0.01229 -0.0582 0.9379 0.6752
-2.750 -0.1258 0.01933 0.01286 -0.0566 0.9292 0.7085
-2.500 -0.1030 0.01994 0.01348 -0.0540 0.9211 0.7399
-2.250 -0.0788 0.02040 0.01390 -0.0518 0.9132 0.7683
-2.000 -0.0585 0.02071 0.01418 -0.0492 0.9047 0.7889
-1.750 -0.0282 0.02079 0.01419 -0.0490 0.8978 0.8063
-1.500 -0.0028 0.02084 0.01414 -0.0484 0.8897 0.8205
-1.250 0.0294 0.02079 0.01401 -0.0490 0.8829 0.8342
-1.000 0.0574 0.02076 0.01391 -0.0490 0.8753 0.8474
-0.750 0.0864 0.02069 0.01377 -0.0491 0.8679 0.8606
-0.500 0.1149 0.02062 0.01364 -0.0491 0.8608 0.8741
-0.250 0.1406 0.02054 0.01352 -0.0487 0.8527 0.8878
0.000 0.1715 0.02042 0.01335 -0.0491 0.8466 0.9015
0.250 0.1953 0.02038 0.01329 -0.0485 0.8378 0.9162
0.500 0.2383 0.02011 0.01299 -0.0508 0.8339 0.9288
0.750 0.2647 0.02023 0.01312 -0.0512 0.8237 0.9449
1.000 0.3188 0.01999 0.01286 -0.0559 0.8198 0.9557
1.250 0.3655 0.02010 0.01299 -0.0602 0.8116 0.9680
1.500 0.4240 0.01990 0.01280 -0.0662 0.8064 0.9769
1.750 0.4820 0.01969 0.01262 -0.0720 0.8013 0.9859
2.000 0.5319 0.01974 0.01272 -0.0772 0.7925 0.9989
2.250 0.5439 0.01996 0.01295 -0.0752 0.7833 1.0000
2.500 0.5566 0.02010 0.01310 -0.0731 0.7746 1.0000
2.750 0.5649 0.02058 0.01362 -0.0708 0.7637 1.0000
3.000 0.5983 0.02044 0.01349 -0.0716 0.7579 1.0000
3.250 0.6188 0.02093 0.01403 -0.0713 0.7467 1.0000
3.500 0.6514 0.02092 0.01405 -0.0720 0.7395 1.0000
3.750 0.6796 0.02109 0.01430 -0.0722 0.7296 1.0000
4.000 0.7066 0.02131 0.01459 -0.0722 0.7189 1.0000
4.250 0.7443 0.02079 0.01412 -0.0728 0.7114 1.0000
4.500 0.7712 0.02080 0.01422 -0.0723 0.6983 1.0000
4.750 0.8004 0.02054 0.01403 -0.0718 0.6845 1.0000
5.000 0.8311 0.02002 0.01358 -0.0711 0.6693 1.0000
5.250 0.8632 0.01923 0.01282 -0.0703 0.6518 1.0000
5.500 0.8893 0.01864 0.01228 -0.0687 0.6278 1.0000
5.750 0.9169 0.01797 0.01160 -0.0673 0.6019 1.0000
6.000 0.9420 0.01758 0.01125 -0.0658 0.5737 1.0000
6.250 0.9639 0.01740 0.01114 -0.0641 0.5387 1.0000
6.500 0.9833 0.01736 0.01106 -0.0619 0.4889 1.0000
6.750 0.9923 0.01791 0.01113 -0.0582 0.3732 1.0000
7.000 0.9792 0.02086 0.01259 -0.0527 0.1912 1.0000
7.250 0.9772 0.02342 0.01448 -0.0490 0.1325 1.0000
7.500 0.9860 0.02514 0.01604 -0.0465 0.1096 1.0000
7.750 0.9983 0.02674 0.01752 -0.0445 0.0960 1.0000
8.000 1.0160 0.02846 0.01916 -0.0432 0.0867 1.0000
8.250 1.0403 0.03050 0.02098 -0.0429 0.0784 1.0000
8.500 1.0656 0.03205 0.02272 -0.0425 0.0725 1.0000
8.750 1.0980 0.03436 0.02493 -0.0432 0.0681 1.0000
9.000 1.1293 0.03714 0.02788 -0.0436 0.0650 1.0000
9.250 1.1513 0.03920 0.03026 -0.0428 0.0620 1.0000
9.500 1.1735 0.04176 0.03310 -0.0422 0.0601 1.0000
9.750 1.1928 0.04474 0.03643 -0.0412 0.0593 1.0000
10.000 1.2068 0.04805 0.04015 -0.0397 0.0591 1.0000
10.250 1.2141 0.05176 0.04432 -0.0377 0.0596 1.0000
10.500 1.2159 0.05558 0.04859 -0.0353 0.0600 1.0000
10.750 1.2124 0.05932 0.05273 -0.0328 0.0602 1.0000
11.000 1.2026 0.06310 0.05686 -0.0300 0.0606 1.0000
11.250 1.1867 0.06674 0.06080 -0.0270 0.0611 1.0000
11.500 1.1676 0.07071 0.06504 -0.0246 0.0618 1.0000
11.750 1.1465 0.07507 0.06963 -0.0233 0.0626 1.0000
12.000 1.1248 0.07994 0.07471 -0.0229 0.0634 1.0000
12.250 1.1032 0.08535 0.08030 -0.0236 0.0642 1.0000
12.500 1.0857 0.09127 0.08634 -0.0250 0.0652 1.0000
12.750 0.9692 0.09557 0.09108 -0.0209 0.0669 1.0000
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Polar data table (+)
Polar graphs
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