NACA 64-012A AIRFOIL (n64012a-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-012A AIRFOIL (n64012a-il) Reynolds number: 200,000 Max Cl/Cd: 47.82 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012a-il-200000.txt Download as CSV file: xf-n64012a-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-012A 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5835 0.08250 0.07916 -0.0263 1.0000 0.0782
-10.500 -0.6174 0.07344 0.07006 -0.0314 1.0000 0.0777
-10.250 -0.6497 0.06625 0.06282 -0.0345 1.0000 0.0767
-10.000 -0.6833 0.06032 0.05679 -0.0359 1.0000 0.0758
-9.750 -0.7130 0.05575 0.05211 -0.0353 1.0000 0.0755
-9.500 -0.7374 0.05214 0.04836 -0.0329 1.0000 0.0763
-9.250 -0.7629 0.04915 0.04505 -0.0296 1.0000 0.0792
-9.000 -0.8229 0.05520 0.05073 -0.0240 1.0000 0.0769
-8.750 -0.8492 0.04113 0.03508 -0.0171 1.0000 0.0429
-8.500 -0.8403 0.03691 0.03053 -0.0151 1.0000 0.0411
-8.250 -0.8281 0.03324 0.02643 -0.0129 1.0000 0.0396
-8.000 -0.8119 0.03035 0.02316 -0.0111 1.0000 0.0391
-7.750 -0.7928 0.02813 0.02065 -0.0097 1.0000 0.0393
-7.500 -0.7727 0.02659 0.01890 -0.0084 1.0000 0.0406
-7.250 -0.7520 0.02522 0.01735 -0.0072 1.0000 0.0422
-7.000 -0.7305 0.02379 0.01575 -0.0059 1.0000 0.0431
-6.750 -0.7092 0.02261 0.01442 -0.0046 1.0000 0.0441
-6.500 -0.6884 0.02088 0.01265 -0.0034 1.0000 0.0455
-6.250 -0.6709 0.01960 0.01141 -0.0018 1.0000 0.0478
-6.000 -0.6545 0.01888 0.01071 0.0001 1.0000 0.0511
-5.750 -0.6387 0.01824 0.01003 0.0022 1.0000 0.0545
-5.500 -0.6279 0.01720 0.00901 0.0050 1.0000 0.0580
-5.250 -0.6161 0.01652 0.00837 0.0076 1.0000 0.0630
-5.000 -0.6043 0.01587 0.00769 0.0101 1.0000 0.0703
-4.750 -0.5913 0.01525 0.00710 0.0122 1.0000 0.0826
-4.500 -0.5576 0.01348 0.00597 0.0096 0.9943 0.1761
-4.250 -0.5342 0.01117 0.00540 0.0086 0.9865 0.5208
-4.000 -0.4973 0.01103 0.00544 0.0067 0.9795 0.5967
-3.750 -0.4587 0.01104 0.00546 0.0045 0.9726 0.6375
-3.500 -0.4212 0.01108 0.00551 0.0027 0.9656 0.6660
-3.250 -0.3827 0.01114 0.00556 0.0007 0.9591 0.6895
-2.750 -0.3132 0.01137 0.00584 -0.0011 0.9443 0.7363
-2.500 -0.2802 0.01143 0.00589 -0.0017 0.9365 0.7526
-2.250 -0.2476 0.01137 0.00580 -0.0024 0.9281 0.7635
-2.000 -0.2183 0.01131 0.00568 -0.0026 0.9183 0.7742
-1.750 -0.1866 0.01123 0.00558 -0.0031 0.9108 0.7822
-1.500 -0.1601 0.01115 0.00547 -0.0027 0.8997 0.7910
-1.250 -0.1324 0.01110 0.00539 -0.0024 0.8902 0.7992
-1.000 -0.1050 0.01102 0.00528 -0.0021 0.8812 0.8075
-0.750 -0.0792 0.01099 0.00523 -0.0015 0.8707 0.8166
-0.500 -0.0522 0.01096 0.00519 -0.0011 0.8619 0.8246
-0.250 -0.0266 0.01092 0.00513 -0.0005 0.8521 0.8347
0.000 0.0000 0.01094 0.00517 0.0000 0.8426 0.8426
0.250 0.0265 0.01092 0.00513 0.0005 0.8347 0.8521
0.500 0.0521 0.01096 0.00519 0.0011 0.8246 0.8620
0.750 0.0791 0.01099 0.00523 0.0016 0.8166 0.8707
1.000 0.1049 0.01102 0.00528 0.0021 0.8075 0.8812
1.250 0.1324 0.01110 0.00539 0.0024 0.7992 0.8902
1.500 0.1600 0.01115 0.00547 0.0027 0.7910 0.8997
1.750 0.1865 0.01123 0.00558 0.0031 0.7822 0.9108
2.000 0.2182 0.01131 0.00568 0.0026 0.7742 0.9183
2.250 0.2475 0.01137 0.00580 0.0025 0.7635 0.9281
2.500 0.2801 0.01143 0.00589 0.0017 0.7526 0.9366
2.750 0.3131 0.01137 0.00584 0.0011 0.7363 0.9444
3.250 0.3827 0.01113 0.00556 -0.0007 0.6895 0.9592
3.500 0.4211 0.01108 0.00550 -0.0026 0.6660 0.9657
3.750 0.4586 0.01104 0.00546 -0.0045 0.6376 0.9726
4.000 0.4973 0.01103 0.00544 -0.0067 0.5968 0.9796
4.250 0.5342 0.01117 0.00540 -0.0086 0.5209 0.9865
4.500 0.5576 0.01348 0.00597 -0.0096 0.1758 0.9944
4.750 0.5911 0.01524 0.00709 -0.0122 0.0827 1.0000
5.000 0.6040 0.01586 0.00768 -0.0100 0.0703 1.0000
5.250 0.6159 0.01652 0.00836 -0.0075 0.0630 1.0000
5.500 0.6277 0.01719 0.00901 -0.0050 0.0580 1.0000
5.750 0.6386 0.01824 0.01002 -0.0022 0.0545 1.0000
6.000 0.6543 0.01887 0.01071 -0.0001 0.0511 1.0000
6.250 0.6708 0.01959 0.01141 0.0018 0.0478 1.0000
6.500 0.6883 0.02088 0.01264 0.0034 0.0455 1.0000
6.750 0.7091 0.02260 0.01442 0.0047 0.0441 1.0000
7.000 0.7304 0.02379 0.01575 0.0059 0.0431 1.0000
7.250 0.7520 0.02521 0.01735 0.0072 0.0422 1.0000
7.500 0.7726 0.02658 0.01890 0.0084 0.0406 1.0000
7.750 0.7928 0.02814 0.02065 0.0097 0.0393 1.0000
8.000 0.8119 0.03036 0.02316 0.0111 0.0391 1.0000
8.250 0.8282 0.03325 0.02643 0.0129 0.0396 1.0000
8.500 0.8404 0.03692 0.03054 0.0151 0.0411 1.0000
8.750 0.8493 0.04115 0.03510 0.0171 0.0429 1.0000
9.000 0.8231 0.05521 0.05074 0.0239 0.0769 1.0000
9.250 0.7628 0.04907 0.04496 0.0295 0.0793 1.0000
9.500 0.7370 0.05206 0.04829 0.0329 0.0763 1.0000
9.750 0.7127 0.05568 0.05204 0.0353 0.0755 1.0000
10.000 0.6829 0.06025 0.05673 0.0358 0.0757 1.0000
10.250 0.6497 0.06617 0.06275 0.0344 0.0767 1.0000
10.500 0.6173 0.07338 0.07000 0.0313 0.0777 1.0000
10.750 0.5831 0.08246 0.07912 0.0262 0.0782 1.0000
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Polar data table (+)
Polar graphs
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