NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 50,000 Max Cl/Cd: 26.38 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63010a-il-50000-n5.txt Download as CSV file: xf-n63010a-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A010 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.7052 0.09375 0.08675 -0.0112 1.0000 0.0473
-10.250 -0.7169 0.08745 0.08044 -0.0154 1.0000 0.0468
-10.000 -0.7318 0.08164 0.07458 -0.0190 1.0000 0.0464
-9.750 -0.7481 0.07658 0.06944 -0.0212 1.0000 0.0460
-9.500 -0.7639 0.07222 0.06496 -0.0217 1.0000 0.0457
-9.250 -0.7754 0.06787 0.06041 -0.0217 1.0000 0.0455
-9.000 -0.7824 0.06362 0.05589 -0.0212 1.0000 0.0454
-8.750 -0.7851 0.05947 0.05142 -0.0205 1.0000 0.0453
-8.500 -0.7838 0.05547 0.04704 -0.0194 1.0000 0.0454
-8.250 -0.7786 0.05166 0.04275 -0.0182 1.0000 0.0456
-8.000 -0.7700 0.04817 0.03875 -0.0169 1.0000 0.0466
-7.750 -0.7585 0.04512 0.03507 -0.0154 1.0000 0.0483
-7.500 -0.7430 0.04200 0.03170 -0.0145 1.0000 0.0502
-7.250 -0.7242 0.03937 0.02882 -0.0136 1.0000 0.0518
-7.000 -0.7029 0.03687 0.02600 -0.0127 1.0000 0.0533
-6.750 -0.6796 0.03459 0.02343 -0.0119 1.0000 0.0553
-6.500 -0.6545 0.03256 0.02112 -0.0111 1.0000 0.0581
-6.250 -0.6300 0.03081 0.01915 -0.0103 1.0000 0.0630
-6.000 -0.6069 0.02931 0.01762 -0.0094 1.0000 0.0690
-5.750 -0.5834 0.02790 0.01599 -0.0082 1.0000 0.0746
-5.500 -0.5641 0.02654 0.01465 -0.0069 1.0000 0.0847
-5.250 -0.5455 0.02520 0.01330 -0.0056 1.0000 0.0981
-5.000 -0.5284 0.02373 0.01191 -0.0041 1.0000 0.1192
-4.750 -0.5141 0.02186 0.01052 -0.0026 1.0000 0.1757
-4.500 -0.5083 0.01946 0.00962 0.0003 1.0000 0.3869
-4.250 -0.4973 0.01885 0.00975 0.0045 1.0000 0.5597
-4.000 -0.4824 0.01885 0.00992 0.0085 1.0000 0.6455
-3.750 -0.4678 0.01918 0.01035 0.0132 1.0000 0.7156
-3.500 -0.4497 0.01956 0.01067 0.0174 1.0000 0.7643
-3.250 -0.4272 0.01969 0.01066 0.0199 1.0000 0.7949
-3.000 -0.4042 0.01961 0.01041 0.0214 1.0000 0.8178
-2.750 -0.3783 0.01949 0.01012 0.0222 1.0000 0.8363
-2.500 -0.3507 0.01935 0.00978 0.0223 1.0000 0.8530
-2.250 -0.3222 0.01920 0.00948 0.0222 1.0000 0.8692
-2.000 -0.2922 0.01907 0.00920 0.0215 1.0000 0.8852
-1.750 -0.2604 0.01894 0.00892 0.0204 1.0000 0.9009
-1.500 -0.2263 0.01883 0.00869 0.0186 1.0000 0.9163
-1.250 -0.1899 0.01874 0.00850 0.0163 1.0000 0.9314
-1.000 -0.1509 0.01866 0.00831 0.0133 1.0000 0.9458
-0.750 -0.1094 0.01857 0.00816 0.0096 1.0000 0.9596
-0.500 -0.0688 0.01850 0.00804 0.0059 1.0000 0.9742
-0.250 -0.0270 0.01842 0.00795 0.0018 1.0000 0.9885
0.000 0.0000 0.01838 0.00791 0.0000 1.0000 1.0000
0.250 0.0270 0.01842 0.00795 -0.0018 0.9885 1.0000
0.500 0.0688 0.01850 0.00804 -0.0059 0.9742 1.0000
0.750 0.1094 0.01857 0.00816 -0.0096 0.9596 1.0000
1.000 0.1508 0.01865 0.00831 -0.0133 0.9458 1.0000
1.250 0.1899 0.01874 0.00849 -0.0163 0.9314 1.0000
1.500 0.2263 0.01883 0.00869 -0.0186 0.9163 1.0000
1.750 0.2603 0.01894 0.00892 -0.0204 0.9009 1.0000
2.000 0.2922 0.01906 0.00920 -0.0215 0.8852 1.0000
2.250 0.3222 0.01920 0.00948 -0.0221 0.8692 1.0000
2.500 0.3507 0.01935 0.00978 -0.0223 0.8530 1.0000
2.750 0.3782 0.01949 0.01012 -0.0222 0.8363 1.0000
3.000 0.4042 0.01961 0.01041 -0.0214 0.8178 1.0000
3.250 0.4271 0.01969 0.01066 -0.0199 0.7949 1.0000
3.500 0.4496 0.01956 0.01067 -0.0174 0.7643 1.0000
3.750 0.4678 0.01918 0.01035 -0.0132 0.7157 1.0000
4.000 0.4824 0.01885 0.00992 -0.0084 0.6456 1.0000
4.250 0.4972 0.01885 0.00975 -0.0045 0.5597 1.0000
4.500 0.5083 0.01946 0.00962 -0.0003 0.3869 1.0000
4.750 0.5141 0.02185 0.01052 0.0026 0.1757 1.0000
5.000 0.5284 0.02373 0.01191 0.0041 0.1192 1.0000
5.250 0.5456 0.02520 0.01330 0.0056 0.0981 1.0000
5.500 0.5641 0.02654 0.01465 0.0069 0.0847 1.0000
5.750 0.5834 0.02790 0.01599 0.0082 0.0746 1.0000
6.000 0.6070 0.02931 0.01762 0.0094 0.0690 1.0000
6.250 0.6300 0.03081 0.01915 0.0103 0.0630 1.0000
6.500 0.6545 0.03256 0.02112 0.0111 0.0581 1.0000
6.750 0.6796 0.03459 0.02343 0.0119 0.0553 1.0000
7.000 0.7030 0.03687 0.02600 0.0127 0.0533 1.0000
7.250 0.7242 0.03937 0.02882 0.0136 0.0518 1.0000
7.500 0.7431 0.04200 0.03170 0.0145 0.0502 1.0000
7.750 0.7585 0.04512 0.03507 0.0154 0.0483 1.0000
8.000 0.7701 0.04817 0.03875 0.0169 0.0466 1.0000
8.250 0.7787 0.05166 0.04275 0.0182 0.0456 1.0000
8.500 0.7839 0.05548 0.04705 0.0194 0.0454 1.0000
8.750 0.7852 0.05948 0.05143 0.0204 0.0453 1.0000
9.000 0.7825 0.06362 0.05590 0.0212 0.0454 1.0000
9.250 0.7756 0.06788 0.06042 0.0216 0.0455 1.0000
9.500 0.7642 0.07224 0.06498 0.0216 0.0457 1.0000
9.750 0.7484 0.07661 0.06947 0.0212 0.0460 1.0000
10.000 0.7322 0.08167 0.07462 0.0189 0.0464 1.0000
10.250 0.7174 0.08750 0.08049 0.0153 0.0468 1.0000
10.500 0.7057 0.09381 0.08681 0.0111 0.0473 1.0000
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Polar data table (+)
Polar graphs
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