NACA 65-410 (naca65410-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 65-410 (naca65410-il) Reynolds number: 500,000 Max Cl/Cd: 87.51 at α=2.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca65410-il-500000-n5.txt Download as CSV file: xf-naca65410-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65-410
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4574 0.08633 0.08402 -0.0422 1.0000 0.0069
-9.500 -0.4645 0.08114 0.07888 -0.0445 1.0000 0.0069
-9.250 -0.4726 0.07614 0.07392 -0.0470 1.0000 0.0069
-9.000 -0.4867 0.07053 0.06836 -0.0502 1.0000 0.0069
-8.750 -0.4994 0.06306 0.06090 -0.0593 0.9911 0.0067
-8.500 -0.4981 0.05329 0.05091 -0.0701 0.9783 0.0067
-8.000 -0.5154 0.02055 0.01579 -0.0794 0.9496 0.0081
-7.750 -0.4873 0.02035 0.01556 -0.0802 0.9424 0.0085
-7.500 -0.4576 0.02039 0.01558 -0.0811 0.9364 0.0089
-7.250 -0.4328 0.01913 0.01408 -0.0812 0.9280 0.0096
-7.000 -0.4085 0.01735 0.01195 -0.0810 0.9203 0.0104
-6.750 -0.3837 0.01604 0.01036 -0.0808 0.9121 0.0110
-6.500 -0.3593 0.01485 0.00898 -0.0805 0.9044 0.0117
-6.250 -0.3334 0.01434 0.00838 -0.0804 0.8967 0.0123
-6.000 -0.3071 0.01407 0.00806 -0.0804 0.8893 0.0132
-5.750 -0.2808 0.01366 0.00754 -0.0803 0.8819 0.0143
-5.500 -0.2550 0.01312 0.00688 -0.0801 0.8744 0.0152
-5.250 -0.2289 0.01263 0.00627 -0.0799 0.8673 0.0159
-5.000 -0.2041 0.01179 0.00535 -0.0796 0.8601 0.0170
-4.750 -0.1777 0.01146 0.00496 -0.0796 0.8531 0.0182
-4.500 -0.1511 0.01116 0.00462 -0.0795 0.8460 0.0195
-4.250 -0.1247 0.01078 0.00416 -0.0794 0.8390 0.0206
-4.000 -0.0979 0.01044 0.00376 -0.0794 0.8324 0.0217
-3.750 -0.0709 0.01019 0.00343 -0.0794 0.8258 0.0227
-3.500 -0.0440 0.00979 0.00296 -0.0794 0.8194 0.0247
-3.250 -0.0169 0.00950 0.00262 -0.0794 0.8126 0.0268
-3.000 0.0106 0.00930 0.00236 -0.0795 0.8063 0.0291
-2.750 0.0383 0.00913 0.00214 -0.0796 0.7999 0.0318
-2.500 0.0659 0.00898 0.00193 -0.0796 0.7942 0.0360
-2.250 0.0937 0.00880 0.00177 -0.0798 0.7878 0.0455
-2.000 0.1212 0.00849 0.00161 -0.0800 0.7816 0.0998
-1.750 0.1481 0.00769 0.00140 -0.0806 0.7756 0.2903
-1.500 0.1742 0.00666 0.00132 -0.0812 0.7696 0.5808
-1.250 0.2012 0.00650 0.00139 -0.0811 0.7642 0.6637
-1.000 0.2289 0.00647 0.00141 -0.0810 0.7580 0.6913
-0.750 0.2565 0.00648 0.00143 -0.0810 0.7524 0.7128
-0.500 0.2845 0.00649 0.00144 -0.0810 0.7467 0.7270
-0.250 0.3126 0.00650 0.00144 -0.0811 0.7408 0.7336
0.000 0.3408 0.00653 0.00144 -0.0812 0.7354 0.7400
0.250 0.3688 0.00655 0.00146 -0.0813 0.7295 0.7463
0.750 0.4249 0.00661 0.00153 -0.0815 0.7182 0.7590
1.000 0.4530 0.00664 0.00157 -0.0816 0.7124 0.7659
1.250 0.4808 0.00668 0.00163 -0.0816 0.7070 0.7721
1.500 0.5087 0.00672 0.00169 -0.0817 0.7000 0.7789
1.750 0.5357 0.00677 0.00174 -0.0815 0.6859 0.7854
2.000 0.5624 0.00685 0.00180 -0.0813 0.6698 0.7927
2.250 0.5893 0.00692 0.00188 -0.0811 0.6565 0.7997
2.500 0.6152 0.00703 0.00195 -0.0807 0.6317 0.8069
2.750 0.6385 0.00730 0.00203 -0.0798 0.5789 0.8141
3.000 0.6586 0.00786 0.00221 -0.0784 0.4828 0.8224
3.250 0.6742 0.00898 0.00268 -0.0766 0.3350 0.8307
3.500 0.6917 0.01005 0.00319 -0.0752 0.2058 0.8397
3.750 0.7093 0.01104 0.00372 -0.0738 0.0939 0.8482
4.000 0.7308 0.01166 0.00411 -0.0729 0.0445 0.8570
4.250 0.7546 0.01198 0.00443 -0.0722 0.0337 0.8665
4.500 0.7783 0.01224 0.00474 -0.0715 0.0298 0.8762
4.750 0.8012 0.01259 0.00513 -0.0707 0.0255 0.8866
5.250 0.8463 0.01321 0.00590 -0.0688 0.0215 0.9101
5.500 0.8676 0.01353 0.00629 -0.0676 0.0198 0.9240
6.250 0.9327 0.01497 0.00792 -0.0647 0.0159 1.0000
6.500 0.9556 0.01553 0.00853 -0.0642 0.0154 1.0000
6.750 0.9778 0.01613 0.00919 -0.0635 0.0148 1.0000
7.000 0.9993 0.01679 0.00992 -0.0627 0.0143 1.0000
7.250 1.0205 0.01747 0.01068 -0.0618 0.0138 1.0000
7.500 1.0414 0.01819 0.01145 -0.0609 0.0133 1.0000
7.750 1.0623 0.01886 0.01217 -0.0601 0.0128 1.0000
8.000 1.0832 0.01951 0.01285 -0.0593 0.0122 1.0000
8.250 1.1012 0.02055 0.01395 -0.0581 0.0116 1.0000
8.500 1.1184 0.02198 0.01552 -0.0568 0.0112 1.0000
8.750 1.1384 0.02292 0.01657 -0.0559 0.0109 1.0000
9.000 1.1582 0.02388 0.01765 -0.0549 0.0106 1.0000
9.250 1.1775 0.02484 0.01873 -0.0539 0.0102 1.0000
9.500 1.1958 0.02581 0.01983 -0.0528 0.0097 1.0000
9.750 1.2125 0.02670 0.02084 -0.0515 0.0093 1.0000
10.000 1.2280 0.02773 0.02199 -0.0501 0.0089 1.0000
10.250 1.2420 0.02874 0.02310 -0.0484 0.0087 1.0000
10.500 1.2548 0.02968 0.02417 -0.0467 0.0084 1.0000
10.750 1.2661 0.03081 0.02540 -0.0449 0.0082 1.0000
11.000 1.2749 0.03222 0.02694 -0.0430 0.0079 1.0000
11.250 1.2784 0.03473 0.02970 -0.0406 0.0076 1.0000
11.500 1.2860 0.03628 0.03146 -0.0388 0.0074 1.0000
11.750 1.2902 0.03832 0.03375 -0.0368 0.0071 1.0000
12.000 1.2914 0.04067 0.03634 -0.0348 0.0068 1.0000
12.250 1.2888 0.04343 0.03935 -0.0328 0.0065 1.0000
12.500 1.2854 0.04619 0.04234 -0.0312 0.0063 1.0000
12.750 1.2771 0.04963 0.04602 -0.0297 0.0061 1.0000
13.000 1.2661 0.05348 0.05011 -0.0287 0.0060 1.0000
13.250 1.2542 0.05760 0.05444 -0.0283 0.0059 1.0000
13.500 1.2386 0.06247 0.05952 -0.0286 0.0058 1.0000
13.750 1.2195 0.06822 0.06549 -0.0298 0.0058 1.0000
14.000 1.1869 0.07678 0.07433 -0.0327 0.0058 1.0000
14.250 1.1356 0.09042 0.08829 -0.0398 0.0059 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 65-410 (naca65410-il)