NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 100,000 Max Cl/Cd: 37.68 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63010a-il-100000.txt Download as CSV file: xf-n63010a-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5333 0.09589 0.09126 -0.0031 1.0000 0.1582
-9.500 -0.5678 0.08998 0.08545 -0.0083 1.0000 0.1640
-9.250 -0.7092 0.09003 0.08540 -0.0081 1.0000 0.1499
-9.000 -0.6843 0.08783 0.08316 -0.0023 1.0000 0.1560
-8.750 -0.7103 0.08144 0.07684 -0.0094 1.0000 0.1625
-7.500 -0.7593 0.04430 0.03682 -0.0146 1.0000 0.0698
-7.250 -0.7406 0.04136 0.03354 -0.0132 1.0000 0.0682
-7.000 -0.7239 0.03769 0.02954 -0.0119 1.0000 0.0677
-6.750 -0.7050 0.03416 0.02569 -0.0107 1.0000 0.0662
-6.500 -0.6840 0.03122 0.02236 -0.0093 1.0000 0.0652
-6.250 -0.6612 0.02874 0.01949 -0.0081 1.0000 0.0653
-6.000 -0.6370 0.02666 0.01711 -0.0070 1.0000 0.0665
-5.750 -0.6129 0.02480 0.01502 -0.0060 1.0000 0.0702
-5.500 -0.5888 0.02298 0.01327 -0.0053 1.0000 0.0747
-5.250 -0.5648 0.02153 0.01176 -0.0041 1.0000 0.0796
-5.000 -0.5441 0.01997 0.01029 -0.0027 1.0000 0.0869
-4.750 -0.5257 0.01873 0.00916 -0.0010 1.0000 0.1011
-4.500 -0.5097 0.01737 0.00797 0.0011 1.0000 0.1256
-4.250 -0.5073 0.01422 0.00655 0.0046 1.0000 0.3556
-4.000 -0.5020 0.01332 0.00688 0.0102 1.0000 0.6258
-3.750 -0.4854 0.01342 0.00706 0.0136 1.0000 0.6908
-3.500 -0.4684 0.01359 0.00721 0.0168 1.0000 0.7354
-3.250 -0.4516 0.01380 0.00744 0.0204 1.0000 0.7717
-3.000 -0.4367 0.01406 0.00771 0.0245 1.0000 0.8070
-2.750 -0.4225 0.01435 0.00800 0.0289 1.0000 0.8400
-2.500 -0.4055 0.01458 0.00816 0.0324 1.0000 0.8682
-2.250 -0.3785 0.01479 0.00828 0.0337 1.0000 0.8894
-2.000 -0.3487 0.01485 0.00822 0.0335 1.0000 0.9091
-1.750 -0.3135 0.01489 0.00813 0.0319 1.0000 0.9264
-1.500 -0.2686 0.01495 0.00805 0.0282 1.0000 0.9404
-1.250 -0.2145 0.01500 0.00798 0.0226 1.0000 0.9512
-1.000 -0.1590 0.01500 0.00786 0.0164 1.0000 0.9616
-0.750 -0.1070 0.01494 0.00774 0.0106 1.0000 0.9731
-0.500 -0.0550 0.01485 0.00762 0.0047 1.0000 0.9847
-0.250 -0.0016 0.01474 0.00748 -0.0018 1.0000 0.9957
0.000 0.0000 0.01472 0.00746 0.0000 1.0000 1.0000
0.250 0.0016 0.01474 0.00748 0.0018 0.9958 1.0000
0.500 0.0550 0.01485 0.00762 -0.0047 0.9847 1.0000
0.750 0.1069 0.01494 0.00774 -0.0106 0.9731 1.0000
1.000 0.1590 0.01499 0.00786 -0.0164 0.9616 1.0000
1.250 0.2145 0.01500 0.00798 -0.0226 0.9513 1.0000
1.500 0.2686 0.01494 0.00804 -0.0282 0.9404 1.0000
1.750 0.3135 0.01488 0.00813 -0.0319 0.9264 1.0000
2.000 0.3486 0.01485 0.00822 -0.0335 0.9091 1.0000
2.250 0.3784 0.01479 0.00827 -0.0336 0.8894 1.0000
2.500 0.4054 0.01458 0.00816 -0.0324 0.8682 1.0000
2.750 0.4224 0.01435 0.00799 -0.0288 0.8400 1.0000
3.000 0.4366 0.01406 0.00771 -0.0244 0.8071 1.0000
3.250 0.4514 0.01380 0.00744 -0.0203 0.7717 1.0000
3.500 0.4682 0.01359 0.00721 -0.0168 0.7355 1.0000
3.750 0.4853 0.01342 0.00706 -0.0135 0.6910 1.0000
4.000 0.5019 0.01332 0.00688 -0.0101 0.6260 1.0000
4.250 0.5073 0.01421 0.00655 -0.0046 0.3576 1.0000
4.500 0.5095 0.01736 0.00796 -0.0011 0.1257 1.0000
4.750 0.5256 0.01873 0.00916 0.0010 0.1011 1.0000
5.000 0.5440 0.01997 0.01029 0.0027 0.0870 1.0000
5.250 0.5647 0.02153 0.01175 0.0042 0.0796 1.0000
5.500 0.5887 0.02297 0.01327 0.0053 0.0747 1.0000
5.750 0.6128 0.02479 0.01501 0.0060 0.0702 1.0000
6.000 0.6369 0.02666 0.01710 0.0070 0.0664 1.0000
6.250 0.6611 0.02874 0.01949 0.0081 0.0652 1.0000
6.500 0.6840 0.03122 0.02236 0.0093 0.0652 1.0000
6.750 0.7050 0.03417 0.02569 0.0107 0.0662 1.0000
7.000 0.7239 0.03770 0.02954 0.0119 0.0677 1.0000
7.250 0.7406 0.04137 0.03356 0.0132 0.0682 1.0000
7.500 0.7594 0.04430 0.03683 0.0146 0.0698 1.0000
9.000 0.6854 0.08781 0.08313 0.0023 0.1559 1.0000
9.250 0.7087 0.09015 0.08552 0.0077 0.1498 1.0000
9.500 0.6706 0.09782 0.09303 -0.0036 0.1433 1.0000
9.750 0.7220 0.09830 0.09363 0.0086 0.1367 1.0000
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