NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 100,000 Max Cl/Cd: 39.85 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012-il-100000.txt Download as CSV file: xf-n64012-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6344 0.09389 0.08907 -0.0022 1.0000 0.1863
-9.500 -0.6503 0.08864 0.08393 -0.0052 1.0000 0.1966
-8.750 -0.8370 0.05259 0.04573 -0.0190 1.0000 0.0785
-8.500 -0.8234 0.04726 0.04030 -0.0187 1.0000 0.0764
-8.250 -0.8112 0.04310 0.03583 -0.0176 1.0000 0.0736
-8.000 -0.7984 0.03911 0.03130 -0.0160 1.0000 0.0705
-7.750 -0.7825 0.03581 0.02739 -0.0143 1.0000 0.0687
-7.500 -0.7617 0.03320 0.02441 -0.0133 1.0000 0.0688
-7.250 -0.7392 0.03123 0.02235 -0.0127 1.0000 0.0715
-7.000 -0.7163 0.02940 0.02029 -0.0119 1.0000 0.0742
-6.750 -0.6917 0.02753 0.01820 -0.0111 1.0000 0.0762
-6.500 -0.6670 0.02603 0.01645 -0.0103 1.0000 0.0785
-6.250 -0.6431 0.02405 0.01465 -0.0097 1.0000 0.0834
-6.000 -0.6228 0.02295 0.01354 -0.0085 1.0000 0.0900
-5.750 -0.6059 0.02161 0.01231 -0.0066 1.0000 0.0956
-5.500 -0.5941 0.02072 0.01150 -0.0041 1.0000 0.1039
-5.250 -0.5870 0.01978 0.01068 -0.0009 1.0000 0.1150
-5.000 -0.5810 0.01886 0.00986 0.0023 1.0000 0.1300
-4.750 -0.5771 0.01754 0.00890 0.0056 1.0000 0.1695
-4.500 -0.5895 0.01480 0.00836 0.0112 1.0000 0.5149
-4.250 -0.5801 0.01515 0.00888 0.0153 1.0000 0.6104
-4.000 -0.5669 0.01564 0.00939 0.0186 1.0000 0.6503
-3.750 -0.5486 0.01616 0.00985 0.0208 0.9984 0.6798
-3.500 -0.5115 0.01705 0.01070 0.0202 0.9888 0.7131
-3.250 -0.4779 0.01831 0.01198 0.0213 0.9798 0.7477
-3.000 -0.4428 0.01958 0.01322 0.0226 0.9718 0.7759
-2.750 -0.4084 0.02006 0.01360 0.0222 0.9628 0.7928
-2.500 -0.3609 0.02037 0.01378 0.0190 0.9571 0.8007
-2.250 -0.3273 0.02025 0.01353 0.0172 0.9473 0.8102
-2.000 -0.2807 0.02035 0.01352 0.0136 0.9416 0.8174
-1.750 -0.2454 0.02027 0.01334 0.0117 0.9325 0.8261
-1.500 -0.2014 0.02031 0.01330 0.0085 0.9262 0.8331
-1.250 -0.1686 0.02020 0.01313 0.0071 0.9173 0.8411
-1.000 -0.1304 0.02021 0.01308 0.0049 0.9101 0.8477
-0.750 -0.0981 0.02010 0.01294 0.0036 0.9016 0.8557
-0.500 -0.0639 0.02012 0.01293 0.0022 0.8935 0.8623
-0.250 -0.0322 0.02004 0.01283 0.0011 0.8860 0.8706
0.000 0.0000 0.02010 0.01289 0.0000 0.8778 0.8779
0.250 0.0322 0.02004 0.01282 -0.0011 0.8706 0.8859
0.500 0.0639 0.02012 0.01293 -0.0022 0.8623 0.8935
0.750 0.0981 0.02010 0.01293 -0.0036 0.8557 0.9016
1.000 0.1304 0.02021 0.01308 -0.0049 0.8476 0.9101
1.250 0.1686 0.02020 0.01313 -0.0070 0.8410 0.9173
1.500 0.2013 0.02031 0.01330 -0.0085 0.8330 0.9263
1.750 0.2454 0.02027 0.01334 -0.0117 0.8261 0.9325
2.000 0.2805 0.02035 0.01352 -0.0136 0.8174 0.9416
2.250 0.3274 0.02025 0.01352 -0.0173 0.8102 0.9473
2.500 0.3610 0.02037 0.01378 -0.0190 0.8007 0.9571
2.750 0.4086 0.02005 0.01359 -0.0222 0.7928 0.9628
3.000 0.4430 0.01958 0.01321 -0.0226 0.7759 0.9719
3.250 0.4779 0.01831 0.01197 -0.0213 0.7477 0.9799
3.500 0.5115 0.01705 0.01070 -0.0202 0.7132 0.9889
3.750 0.5486 0.01616 0.00985 -0.0208 0.6799 0.9985
4.000 0.5668 0.01564 0.00938 -0.0186 0.6503 1.0000
4.250 0.5800 0.01515 0.00889 -0.0153 0.6107 1.0000
4.500 0.5894 0.01479 0.00836 -0.0112 0.5167 1.0000
4.750 0.5771 0.01752 0.00889 -0.0056 0.1703 1.0000
5.000 0.5809 0.01885 0.00985 -0.0023 0.1302 1.0000
5.250 0.5869 0.01977 0.01067 0.0010 0.1150 1.0000
5.500 0.5941 0.02073 0.01150 0.0041 0.1045 1.0000
5.750 0.6059 0.02161 0.01231 0.0066 0.0956 1.0000
6.000 0.6228 0.02295 0.01353 0.0085 0.0900 1.0000
6.250 0.6432 0.02404 0.01465 0.0097 0.0836 1.0000
6.500 0.6670 0.02602 0.01645 0.0103 0.0785 1.0000
6.750 0.6917 0.02754 0.01821 0.0111 0.0761 1.0000
7.000 0.7163 0.02939 0.02029 0.0119 0.0743 1.0000
7.250 0.7392 0.03124 0.02236 0.0127 0.0716 1.0000
7.500 0.7617 0.03320 0.02442 0.0133 0.0688 1.0000
7.750 0.7825 0.03581 0.02738 0.0143 0.0686 1.0000
8.000 0.7984 0.03909 0.03129 0.0160 0.0705 1.0000
8.250 0.8111 0.04310 0.03583 0.0176 0.0736 1.0000
8.500 0.8234 0.04726 0.04030 0.0187 0.0764 1.0000
8.750 0.8363 0.05229 0.04546 0.0192 0.0783 1.0000
9.250 0.6891 0.08208 0.07750 0.0112 0.2094 1.0000
9.500 0.6511 0.08859 0.08386 0.0053 0.1965 1.0000
9.750 0.6349 0.09386 0.08904 0.0021 0.1862 1.0000
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Polar data table (+)
Polar graphs
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