NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 16-012 (naca16012-il) Reynolds number: 50,000 Max Cl/Cd: 22.09 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca16012-il-50000-n5.txt Download as CSV file: xf-naca16012-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 16-012
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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6613 0.09230 0.08497 -0.0522 1.0000 0.0431
-11.000 -0.6789 0.08799 0.08064 -0.0519 1.0000 0.0429
-10.750 -0.6985 0.08413 0.07673 -0.0506 1.0000 0.0427
-10.500 -0.7192 0.08073 0.07328 -0.0481 1.0000 0.0426
-10.250 -0.7410 0.07777 0.07025 -0.0445 1.0000 0.0424
-10.000 -0.7633 0.07522 0.06763 -0.0396 1.0000 0.0423
-9.750 -0.7844 0.07271 0.06501 -0.0343 1.0000 0.0423
-9.500 -0.8021 0.06985 0.06199 -0.0294 1.0000 0.0422
-9.250 -0.8176 0.06697 0.05891 -0.0243 1.0000 0.0423
-9.000 -0.8314 0.06408 0.05574 -0.0190 1.0000 0.0425
-8.750 -0.8432 0.06124 0.05254 -0.0137 1.0000 0.0429
-8.500 -0.8503 0.05818 0.04924 -0.0091 1.0000 0.0437
-8.250 -0.8491 0.05568 0.04658 -0.0058 1.0000 0.0454
-8.000 -0.8469 0.05337 0.04407 -0.0023 1.0000 0.0470
-7.750 -0.8446 0.05074 0.04110 0.0016 1.0000 0.0484
-7.500 -0.8394 0.04792 0.03787 0.0053 1.0000 0.0495
-7.250 -0.8306 0.04520 0.03470 0.0086 1.0000 0.0512
-7.000 -0.8190 0.04284 0.03175 0.0117 1.0000 0.0544
-6.750 -0.8028 0.04041 0.02903 0.0136 1.0000 0.0578
-6.500 -0.7820 0.03831 0.02670 0.0149 1.0000 0.0611
-6.250 -0.7597 0.03658 0.02459 0.0161 1.0000 0.0671
-6.000 -0.7327 0.03477 0.02265 0.0162 1.0000 0.0733
-5.750 -0.7012 0.03330 0.02086 0.0158 1.0000 0.0806
-5.500 -0.6722 0.03206 0.01957 0.0155 1.0000 0.0908
-5.250 -0.6404 0.03086 0.01817 0.0151 1.0000 0.0993
-5.000 -0.6184 0.02984 0.01702 0.0162 1.0000 0.1097
-4.750 -0.6009 0.02883 0.01600 0.0179 1.0000 0.1265
-4.500 -0.5862 0.02766 0.01506 0.0200 1.0000 0.1577
-4.250 -0.5760 0.02609 0.01406 0.0227 1.0000 0.2328
-4.000 -0.4895 0.02660 0.01719 0.0174 1.0000 0.8241
-3.750 -0.3486 0.03088 0.02056 0.0014 1.0000 0.9281
-3.500 -0.2900 0.03108 0.02032 -0.0049 1.0000 0.9519
-3.250 -0.2487 0.03082 0.01976 -0.0084 1.0000 0.9645
-3.000 -0.2100 0.03045 0.01913 -0.0117 1.0000 0.9742
-2.750 -0.1731 0.03006 0.01853 -0.0147 1.0000 0.9830
-2.500 -0.1367 0.02967 0.01796 -0.0176 1.0000 0.9911
-2.250 -0.1001 0.02929 0.01742 -0.0206 1.0000 0.9986
-2.000 -0.0845 0.02907 0.01710 -0.0194 1.0000 1.0000
-1.750 -0.0740 0.02889 0.01686 -0.0170 1.0000 1.0000
-1.500 -0.0635 0.02874 0.01664 -0.0146 1.0000 1.0000
-1.250 -0.0529 0.02861 0.01646 -0.0122 1.0000 1.0000
-1.000 -0.0424 0.02851 0.01632 -0.0098 1.0000 1.0000
-0.750 -0.0318 0.02843 0.01621 -0.0073 1.0000 1.0000
-0.500 -0.0212 0.02837 0.01613 -0.0049 1.0000 1.0000
-0.250 -0.0106 0.02834 0.01609 -0.0025 1.0000 1.0000
0.000 0.0000 0.02833 0.01607 0.0000 1.0000 1.0000
0.250 0.0106 0.02834 0.01608 0.0025 1.0000 1.0000
0.500 0.0212 0.02837 0.01613 0.0049 1.0000 1.0000
0.750 0.0318 0.02843 0.01621 0.0073 1.0000 1.0000
1.000 0.0424 0.02850 0.01632 0.0098 1.0000 1.0000
1.250 0.0529 0.02860 0.01646 0.0122 1.0000 1.0000
1.500 0.0635 0.02873 0.01663 0.0146 1.0000 1.0000
1.750 0.0740 0.02888 0.01685 0.0170 1.0000 1.0000
2.000 0.0845 0.02906 0.01709 0.0194 1.0000 1.0000
2.250 0.1000 0.02928 0.01740 0.0207 0.9987 1.0000
2.500 0.1366 0.02966 0.01795 0.0176 0.9911 1.0000
2.750 0.1730 0.03005 0.01852 0.0147 0.9831 1.0000
3.000 0.2099 0.03043 0.01911 0.0117 0.9743 1.0000
3.250 0.2486 0.03080 0.01974 0.0085 0.9645 1.0000
3.500 0.2898 0.03106 0.02030 0.0049 0.9520 1.0000
3.750 0.3483 0.03087 0.02054 -0.0013 0.9282 1.0000
4.000 0.4896 0.02659 0.01718 -0.0174 0.8241 1.0000
4.250 0.5760 0.02608 0.01405 -0.0227 0.2329 1.0000
4.500 0.5861 0.02766 0.01505 -0.0200 0.1578 1.0000
4.750 0.6009 0.02882 0.01600 -0.0179 0.1265 1.0000
5.000 0.6184 0.02983 0.01701 -0.0162 0.1098 1.0000
5.250 0.6404 0.03085 0.01816 -0.0150 0.0994 1.0000
5.500 0.6721 0.03205 0.01957 -0.0155 0.0908 1.0000
5.750 0.7011 0.03329 0.02085 -0.0158 0.0806 1.0000
6.000 0.7326 0.03477 0.02264 -0.0162 0.0733 1.0000
6.250 0.7596 0.03657 0.02458 -0.0161 0.0671 1.0000
6.500 0.7819 0.03830 0.02669 -0.0149 0.0611 1.0000
6.750 0.8027 0.04040 0.02902 -0.0136 0.0578 1.0000
7.000 0.8190 0.04283 0.03174 -0.0117 0.0545 1.0000
7.250 0.8305 0.04520 0.03469 -0.0086 0.0512 1.0000
7.500 0.8394 0.04791 0.03786 -0.0053 0.0495 1.0000
7.750 0.8445 0.05073 0.04109 -0.0016 0.0484 1.0000
8.000 0.8469 0.05337 0.04407 0.0023 0.0471 1.0000
8.250 0.8492 0.05566 0.04656 0.0058 0.0454 1.0000
8.500 0.8503 0.05817 0.04923 0.0091 0.0438 1.0000
8.750 0.8432 0.06123 0.05253 0.0137 0.0429 1.0000
9.000 0.8315 0.06407 0.05573 0.0190 0.0425 1.0000
9.250 0.8177 0.06697 0.05890 0.0243 0.0423 1.0000
9.500 0.8022 0.06985 0.06199 0.0293 0.0422 1.0000
9.750 0.7846 0.07270 0.06500 0.0343 0.0423 1.0000
10.000 0.7635 0.07523 0.06763 0.0396 0.0423 1.0000
10.250 0.7411 0.07779 0.07027 0.0444 0.0424 1.0000
10.500 0.7195 0.08074 0.07329 0.0481 0.0426 1.0000
10.750 0.6988 0.08414 0.07674 0.0506 0.0427 1.0000
11.000 0.6793 0.08800 0.08064 0.0519 0.0429 1.0000
11.250 0.6618 0.09230 0.08497 0.0521 0.0431 1.0000
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Polar data table (+)
Polar graphs
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