NACA 25112 (naca25112-jf) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 25112 (naca25112-jf) Reynolds number: 50,000 Max Cl/Cd: 17.68 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca25112-jf-50000.txt Download as CSV file: xf-naca25112-jf-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 25112
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4517 0.09133 0.08486 0.0130 1.0000 0.3743
-7.500 -0.4336 0.08811 0.08166 0.0156 1.0000 0.4064
-7.250 -0.4327 0.08599 0.07962 0.0184 1.0000 0.4427
-7.000 -0.3976 0.08218 0.07579 0.0212 1.0000 0.4835
-6.750 -0.3720 0.07941 0.07302 0.0244 1.0000 0.5356
-6.500 -0.3417 0.07660 0.07020 0.0277 1.0000 0.5993
-6.250 -0.3125 0.07376 0.06737 0.0297 1.0000 0.6583
-5.500 -0.4830 0.04731 0.03948 -0.0107 1.0000 0.1825
-5.250 -0.4739 0.04424 0.03611 -0.0087 1.0000 0.1769
-5.000 -0.4661 0.04169 0.03305 -0.0062 1.0000 0.1730
-4.750 -0.4567 0.03998 0.03079 -0.0036 1.0000 0.1707
-4.500 -0.4460 0.03797 0.02853 -0.0014 1.0000 0.1705
-4.250 -0.4349 0.03599 0.02654 0.0002 1.0000 0.1735
-4.000 -0.4233 0.03464 0.02511 0.0018 1.0000 0.1775
-3.750 -0.4098 0.03340 0.02363 0.0032 1.0000 0.1805
-3.500 -0.3947 0.03240 0.02230 0.0042 1.0000 0.1849
-3.250 -0.3783 0.03136 0.02111 0.0049 1.0000 0.1911
-3.000 -0.3207 0.02987 0.01930 -0.0013 0.9854 0.2062
-2.750 -0.2555 0.02829 0.01775 -0.0084 0.9691 0.2298
-2.500 -0.1894 0.02688 0.01642 -0.0153 0.9524 0.2669
-2.250 0.0152 0.02311 0.01531 -0.0393 0.9647 1.0000
-2.000 0.0972 0.02291 0.01457 -0.0501 0.9426 1.0000
-1.750 0.1591 0.02274 0.01406 -0.0569 0.9180 1.0000
-1.500 0.2046 0.02269 0.01376 -0.0603 0.8936 1.0000
-1.250 0.2378 0.02278 0.01364 -0.0612 0.8706 1.0000
-1.000 0.2621 0.02305 0.01374 -0.0606 0.8484 1.0000
-0.750 0.2851 0.02333 0.01387 -0.0597 0.8284 1.0000
-0.500 0.0507 0.00674 0.00137 0.0051 0.4333 0.6298
-0.250 0.3280 0.02405 0.01432 -0.0574 0.7933 1.0000
0.000 0.3487 0.02448 0.01464 -0.0561 0.7773 1.0000
0.250 0.3692 0.02497 0.01502 -0.0549 0.7624 1.0000
0.500 0.3895 0.02545 0.01541 -0.0536 0.7483 1.0000
0.750 0.4105 0.02586 0.01570 -0.0521 0.7357 1.0000
1.000 0.4299 0.02651 0.01631 -0.0510 0.7220 1.0000
1.250 0.4486 0.02725 0.01701 -0.0499 0.7085 1.0000
1.500 0.4680 0.02795 0.01765 -0.0487 0.6965 1.0000
1.750 0.4892 0.02841 0.01803 -0.0471 0.6860 1.0000
2.000 0.5060 0.02942 0.01906 -0.0461 0.6730 1.0000
2.250 0.5238 0.03035 0.01997 -0.0449 0.6617 1.0000
2.500 0.5453 0.03085 0.02040 -0.0433 0.6521 1.0000
2.750 0.5593 0.03217 0.02176 -0.0423 0.6397 1.0000
3.000 0.5765 0.03320 0.02280 -0.0410 0.6296 1.0000
3.250 0.5958 0.03400 0.02358 -0.0396 0.6195 1.0000
3.500 0.6067 0.03564 0.02527 -0.0384 0.6082 1.0000
3.750 0.6311 0.03603 0.02561 -0.0367 0.6001 1.0000
4.000 0.6374 0.03809 0.02777 -0.0355 0.5882 1.0000
4.250 0.6487 0.03973 0.02944 -0.0341 0.5783 1.0000
4.500 0.6679 0.04066 0.03038 -0.0327 0.5690 1.0000
4.750 0.6670 0.04335 0.03313 -0.0312 0.5581 1.0000
5.000 0.6992 0.04326 0.03305 -0.0297 0.5504 1.0000
5.250 0.6829 0.04731 0.03718 -0.0283 0.5390 1.0000
5.500 0.6974 0.04883 0.03873 -0.0268 0.5299 1.0000
5.750 0.6967 0.05166 0.04160 -0.0256 0.5202 1.0000
6.000 0.6888 0.05520 0.04516 -0.0245 0.5111 1.0000
6.250 0.7108 0.05622 0.04624 -0.0233 0.5017 1.0000
6.500 0.6759 0.06204 0.05201 -0.0227 0.4943 1.0000
6.750 0.7328 0.06034 0.05048 -0.0211 0.4833 1.0000
7.000 0.6711 0.06831 0.05831 -0.0211 0.4788 1.0000
7.250 0.6597 0.07226 0.06226 -0.0208 0.4726 1.0000
7.500 0.6741 0.07441 0.06449 -0.0202 0.4632 1.0000
7.750 0.6567 0.07910 0.06918 -0.0206 0.4599 1.0000
8.000 0.6824 0.08049 0.07065 -0.0197 0.4479 1.0000
8.250 0.6646 0.08541 0.07558 -0.0205 0.4460 1.0000
8.500 0.6568 0.08991 0.08012 -0.0214 0.4458 1.0000
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Polar data table (+)
Polar graphs
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