NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 200,000 Max Cl/Cd: 43.46 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012-il-200000.txt Download as CSV file: xf-n64012-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY 64-012 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6820 0.10002 0.09657 -0.0066 1.0000 0.0689
-11.000 -0.6784 0.09648 0.09303 -0.0076 1.0000 0.0708
-10.750 -0.6957 0.08791 0.08449 -0.0140 1.0000 0.0720
-10.500 -0.7302 0.07866 0.07519 -0.0214 1.0000 0.0718
-10.250 -0.7741 0.07129 0.06769 -0.0259 1.0000 0.0701
-10.000 -0.7897 0.06776 0.06408 -0.0261 1.0000 0.0712
-9.750 -0.8077 0.06453 0.06076 -0.0249 1.0000 0.0721
-9.500 -0.8204 0.06154 0.05759 -0.0236 1.0000 0.0749
-9.250 -0.8518 0.06341 0.05870 -0.0193 1.0000 0.0788
-8.750 -0.8539 0.03981 0.03371 -0.0155 1.0000 0.0420
-8.500 -0.8411 0.03576 0.02934 -0.0143 1.0000 0.0408
-8.250 -0.8246 0.03259 0.02582 -0.0132 1.0000 0.0403
-8.000 -0.8051 0.03003 0.02294 -0.0122 1.0000 0.0404
-7.750 -0.7838 0.02830 0.02094 -0.0113 1.0000 0.0416
-7.500 -0.7612 0.02676 0.01913 -0.0106 1.0000 0.0425
-7.250 -0.7375 0.02525 0.01740 -0.0099 1.0000 0.0430
-7.000 -0.7127 0.02270 0.01476 -0.0096 1.0000 0.0442
-6.750 -0.6897 0.02111 0.01320 -0.0090 1.0000 0.0459
-6.500 -0.6691 0.02017 0.01229 -0.0081 1.0000 0.0485
-6.250 -0.6523 0.01941 0.01150 -0.0064 1.0000 0.0511
-6.000 -0.6415 0.01883 0.01086 -0.0037 1.0000 0.0532
-5.750 -0.6368 0.01793 0.00997 -0.0002 1.0000 0.0552
-5.500 -0.6009 0.01676 0.00887 -0.0027 0.9926 0.0611
-5.250 -0.5630 0.01569 0.00780 -0.0056 0.9851 0.0711
-5.000 -0.5255 0.01467 0.00683 -0.0083 0.9775 0.0889
-4.750 -0.5004 0.01200 0.00543 -0.0102 0.9676 0.3006
-4.500 -0.4737 0.01090 0.00532 -0.0107 0.9578 0.5358
-4.250 -0.4428 0.01091 0.00537 -0.0111 0.9473 0.5820
-4.000 -0.4128 0.01098 0.00541 -0.0112 0.9372 0.6109
-3.750 -0.3838 0.01108 0.00548 -0.0110 0.9279 0.6319
-3.500 -0.3590 0.01120 0.00555 -0.0100 0.9163 0.6516
-3.250 -0.3352 0.01141 0.00578 -0.0085 0.9054 0.6737
-3.000 -0.3116 0.01172 0.00611 -0.0067 0.8960 0.6955
-2.750 -0.2878 0.01190 0.00628 -0.0052 0.8860 0.7111
-2.500 -0.2627 0.01191 0.00624 -0.0044 0.8759 0.7203
-2.250 -0.2374 0.01188 0.00613 -0.0036 0.8674 0.7280
-2.000 -0.2115 0.01184 0.00603 -0.0031 0.8575 0.7350
-1.750 -0.1856 0.01183 0.00597 -0.0025 0.8490 0.7429
-1.500 -0.1596 0.01178 0.00588 -0.0020 0.8406 0.7500
-1.250 -0.1331 0.01176 0.00584 -0.0016 0.8317 0.7567
-1.000 -0.1069 0.01171 0.00571 -0.0012 0.8241 0.7640
-0.750 -0.0802 0.01170 0.00571 -0.0009 0.8150 0.7705
-0.500 -0.0536 0.01166 0.00560 -0.0006 0.8084 0.7782
-0.250 -0.0267 0.01167 0.00565 -0.0003 0.7997 0.7845
0.000 -0.0001 0.01165 0.00557 0.0000 0.7932 0.7932
0.250 0.0266 0.01167 0.00565 0.0003 0.7845 0.7997
0.500 0.0535 0.01166 0.00560 0.0006 0.7782 0.8084
0.750 0.0801 0.01170 0.00571 0.0009 0.7705 0.8150
1.000 0.1068 0.01171 0.00571 0.0012 0.7640 0.8241
1.250 0.1330 0.01176 0.00583 0.0017 0.7567 0.8317
1.500 0.1595 0.01178 0.00588 0.0020 0.7500 0.8406
1.750 0.1855 0.01183 0.00597 0.0025 0.7429 0.8490
2.000 0.2114 0.01183 0.00603 0.0031 0.7350 0.8575
2.250 0.2373 0.01188 0.00613 0.0036 0.7280 0.8675
2.500 0.2626 0.01191 0.00624 0.0044 0.7203 0.8759
2.750 0.2877 0.01190 0.00628 0.0052 0.7111 0.8860
3.000 0.3115 0.01172 0.00611 0.0067 0.6956 0.8960
3.250 0.3351 0.01142 0.00578 0.0085 0.6738 0.9054
3.500 0.3589 0.01120 0.00555 0.0100 0.6516 0.9163
3.750 0.3837 0.01107 0.00548 0.0110 0.6318 0.9279
4.000 0.4127 0.01098 0.00540 0.0113 0.6108 0.9372
4.250 0.4428 0.01091 0.00537 0.0111 0.5820 0.9473
4.500 0.4737 0.01090 0.00532 0.0107 0.5357 0.9578
4.750 0.5005 0.01198 0.00542 0.0102 0.3030 0.9676
5.000 0.5256 0.01467 0.00683 0.0082 0.0891 0.9775
5.250 0.5630 0.01569 0.00780 0.0055 0.0711 0.9851
5.500 0.6010 0.01675 0.00885 0.0027 0.0612 0.9926
5.750 0.6366 0.01793 0.00997 0.0002 0.0552 1.0000
6.000 0.6414 0.01883 0.01086 0.0037 0.0533 1.0000
6.250 0.6523 0.01941 0.01149 0.0064 0.0512 1.0000
6.500 0.6691 0.02016 0.01227 0.0081 0.0484 1.0000
6.750 0.6897 0.02111 0.01321 0.0090 0.0459 1.0000
7.000 0.7126 0.02267 0.01474 0.0096 0.0442 1.0000
7.250 0.7375 0.02526 0.01742 0.0099 0.0430 1.0000
7.500 0.7611 0.02675 0.01912 0.0106 0.0425 1.0000
7.750 0.7836 0.02823 0.02087 0.0114 0.0415 1.0000
8.000 0.8050 0.03000 0.02291 0.0122 0.0403 1.0000
8.250 0.8245 0.03258 0.02581 0.0132 0.0403 1.0000
8.500 0.8410 0.03575 0.02933 0.0144 0.0409 1.0000
8.750 0.8538 0.03974 0.03364 0.0155 0.0420 1.0000
9.250 0.8516 0.06340 0.05868 0.0194 0.0788 1.0000
9.500 0.8204 0.06154 0.05758 0.0237 0.0749 1.0000
9.750 0.8069 0.06462 0.06085 0.0249 0.0724 1.0000
10.000 0.7900 0.06771 0.06404 0.0261 0.0712 1.0000
10.250 0.7636 0.07219 0.06862 0.0253 0.0712 1.0000
10.500 0.7367 0.07780 0.07431 0.0223 0.0710 1.0000
10.750 0.6953 0.08806 0.08463 0.0139 0.0722 1.0000
11.000 0.6800 0.09615 0.09270 0.0079 0.0706 1.0000
11.250 0.6804 0.10054 0.09708 0.0061 0.0690 1.0000
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Polar data table (+)
Polar graphs
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