NACA 65-410 (naca65410-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: NACA 65-410 (naca65410-il) Reynolds number: 1,000,000 Max Cl/Cd: 114.39 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca65410-il-1000000.txt Download as CSV file: xf-naca65410-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4383 0.08550 0.08388 -0.0435 1.0000 0.0117
-9.250 -0.4410 0.08099 0.07940 -0.0456 1.0000 0.0117
-9.000 -0.4457 0.07644 0.07487 -0.0478 1.0000 0.0117
-7.500 -0.4181 0.03516 0.03265 -0.0848 0.9661 0.0106
-7.250 -0.4131 0.02386 0.02047 -0.0851 0.9514 0.0100
-7.000 -0.3984 0.01790 0.01378 -0.0839 0.9403 0.0104
-6.750 -0.3765 0.01573 0.01127 -0.0832 0.9309 0.0110
-6.500 -0.3510 0.01504 0.01041 -0.0829 0.9230 0.0115
-6.250 -0.3286 0.01314 0.00824 -0.0824 0.9145 0.0126
-6.000 -0.3028 0.01273 0.00778 -0.0822 0.9065 0.0132
-5.750 -0.2768 0.01231 0.00726 -0.0820 0.8991 0.0139
-5.500 -0.2510 0.01173 0.00660 -0.0818 0.8914 0.0147
-5.000 -0.1982 0.01094 0.00565 -0.0814 0.8765 0.0162
-4.750 -0.1740 0.00989 0.00447 -0.0810 0.8693 0.0177
-4.500 -0.1474 0.00953 0.00410 -0.0810 0.8618 0.0188
-4.250 -0.1207 0.00924 0.00374 -0.0809 0.8552 0.0199
-4.000 -0.0936 0.00892 0.00339 -0.0809 0.8479 0.0212
-3.750 -0.0664 0.00876 0.00316 -0.0809 0.8413 0.0224
-3.500 -0.0394 0.00833 0.00266 -0.0809 0.8340 0.0238
-3.000 0.0155 0.00776 0.00199 -0.0810 0.8207 0.0287
-2.750 0.0432 0.00763 0.00180 -0.0811 0.8143 0.0311
-2.500 0.0711 0.00742 0.00155 -0.0812 0.8077 0.0359
-2.250 0.0990 0.00727 0.00138 -0.0813 0.8011 0.0430
-2.000 0.1264 0.00665 0.00118 -0.0817 0.7948 0.1782
-1.750 0.1527 0.00534 0.00096 -0.0826 0.7883 0.5285
-1.500 0.1806 0.00515 0.00096 -0.0828 0.7823 0.6072
-1.250 0.2088 0.00508 0.00095 -0.0829 0.7758 0.6416
-1.000 0.2365 0.00504 0.00098 -0.0830 0.7699 0.6818
-0.750 0.2646 0.00501 0.00102 -0.0830 0.7636 0.7088
-0.500 0.2926 0.00503 0.00104 -0.0831 0.7576 0.7253
-0.250 0.3207 0.00505 0.00107 -0.0831 0.7516 0.7414
0.000 0.3489 0.00506 0.00108 -0.0833 0.7455 0.7488
0.250 0.3771 0.00509 0.00110 -0.0834 0.7398 0.7556
0.500 0.4055 0.00511 0.00111 -0.0836 0.7336 0.7623
0.750 0.4335 0.00515 0.00115 -0.0836 0.7279 0.7690
1.000 0.4619 0.00516 0.00118 -0.0838 0.7207 0.7760
1.250 0.4894 0.00520 0.00121 -0.0838 0.7110 0.7828
1.500 0.5170 0.00526 0.00123 -0.0837 0.6986 0.7900
1.750 0.5446 0.00527 0.00129 -0.0837 0.6882 0.7969
2.000 0.5716 0.00536 0.00133 -0.0836 0.6708 0.8044
2.250 0.5983 0.00543 0.00138 -0.0834 0.6480 0.8115
2.500 0.6250 0.00555 0.00145 -0.0832 0.6261 0.8192
2.750 0.6509 0.00569 0.00154 -0.0828 0.5967 0.8268
3.000 0.6760 0.00594 0.00165 -0.0824 0.5527 0.8347
3.250 0.6952 0.00672 0.00194 -0.0810 0.4300 0.8431
3.500 0.7103 0.00802 0.00250 -0.0791 0.2552 0.8520
3.750 0.7272 0.00918 0.00304 -0.0776 0.1120 0.8618
4.000 0.7477 0.00991 0.00347 -0.0765 0.0414 0.8713
4.250 0.7717 0.01026 0.00381 -0.0758 0.0312 0.8806
4.500 0.7968 0.01048 0.00408 -0.0754 0.0281 0.8905
4.750 0.8191 0.01093 0.00458 -0.0744 0.0228 0.9012
5.000 0.8426 0.01117 0.00488 -0.0736 0.0214 0.9123
5.250 0.8654 0.01144 0.00520 -0.0727 0.0199 0.9243
5.500 0.8869 0.01172 0.00554 -0.0715 0.0185 0.9379
5.750 0.9057 0.01208 0.00596 -0.0697 0.0173 0.9568
6.000 0.9242 0.01313 0.00717 -0.0682 0.0158 1.0000
6.250 0.9506 0.01339 0.00744 -0.0682 0.0154 1.0000
6.500 0.9753 0.01381 0.00788 -0.0679 0.0147 1.0000
6.750 0.9984 0.01437 0.00849 -0.0673 0.0141 1.0000
7.000 1.0208 0.01498 0.00915 -0.0666 0.0135 1.0000
7.250 1.0430 0.01561 0.00983 -0.0659 0.0130 1.0000
7.500 1.0648 0.01628 0.01053 -0.0651 0.0126 1.0000
7.750 1.0861 0.01700 0.01129 -0.0643 0.0122 1.0000
8.000 1.1046 0.01821 0.01255 -0.0631 0.0115 1.0000
8.250 1.1228 0.02006 0.01455 -0.0618 0.0109 1.0000
8.500 1.1451 0.02071 0.01529 -0.0611 0.0107 1.0000
8.750 1.1666 0.02160 0.01627 -0.0604 0.0103 1.0000
9.000 1.1874 0.02272 0.01750 -0.0595 0.0099 1.0000
9.250 1.2075 0.02390 0.01881 -0.0586 0.0096 1.0000
9.500 1.2268 0.02494 0.01995 -0.0577 0.0092 1.0000
9.750 1.2452 0.02567 0.02076 -0.0566 0.0088 1.0000
10.000 1.2627 0.02622 0.02136 -0.0554 0.0085 1.0000
10.250 1.2790 0.02674 0.02193 -0.0541 0.0082 1.0000
10.500 1.2883 0.02920 0.02459 -0.0521 0.0078 1.0000
10.750 1.2873 0.03310 0.02892 -0.0487 0.0075 1.0000
11.000 1.2938 0.03430 0.03026 -0.0460 0.0074 1.0000
11.250 1.2966 0.03601 0.03217 -0.0431 0.0072 1.0000
11.500 1.2947 0.03831 0.03469 -0.0400 0.0070 1.0000
11.750 1.2866 0.04130 0.03794 -0.0367 0.0069 1.0000
12.000 1.2712 0.04520 0.04212 -0.0335 0.0067 1.0000
12.250 1.2496 0.04990 0.04711 -0.0307 0.0066 1.0000
12.500 1.2165 0.05639 0.05393 -0.0288 0.0066 1.0000
12.750 1.1748 0.06462 0.06248 -0.0286 0.0067 1.0000
13.000 1.1370 0.07318 0.07128 -0.0305 0.0067 1.0000
13.250 1.1059 0.08164 0.07993 -0.0342 0.0067 1.0000
13.500 1.0738 0.09166 0.09011 -0.0400 0.0069 1.0000
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Polar data table (+)
Polar graphs
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