NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 50,000 Max Cl/Cd: 28.75 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64110-il-50000-n5.txt Download as CSV file: xf-n64110-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-110 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6197 0.09249 0.08548 -0.0172 1.0000 0.0504
-9.750 -0.6345 0.08481 0.07783 -0.0233 1.0000 0.0491
-9.500 -0.6580 0.07798 0.07097 -0.0286 1.0000 0.0478
-9.250 -0.6882 0.07275 0.06557 -0.0310 1.0000 0.0466
-8.750 -0.7067 0.06467 0.05688 -0.0314 1.0000 0.0457
-8.250 -0.6999 0.05684 0.04876 -0.0310 1.0000 0.0472
-8.000 -0.6905 0.05390 0.04571 -0.0306 1.0000 0.0491
-7.750 -0.6822 0.05065 0.04214 -0.0299 1.0000 0.0504
-7.500 -0.6719 0.04725 0.03835 -0.0291 1.0000 0.0513
-7.250 -0.6587 0.04393 0.03459 -0.0282 1.0000 0.0516
-7.000 -0.6427 0.04086 0.03109 -0.0273 1.0000 0.0523
-6.750 -0.6241 0.03808 0.02788 -0.0263 1.0000 0.0532
-6.500 -0.6035 0.03565 0.02498 -0.0253 1.0000 0.0550
-6.250 -0.5817 0.03374 0.02254 -0.0242 1.0000 0.0582
-6.000 -0.5595 0.03159 0.02043 -0.0236 1.0000 0.0615
-5.750 -0.5356 0.02987 0.01857 -0.0226 1.0000 0.0643
-5.500 -0.5115 0.02839 0.01692 -0.0214 1.0000 0.0678
-5.250 -0.4900 0.02710 0.01552 -0.0201 1.0000 0.0738
-5.000 -0.4707 0.02597 0.01436 -0.0186 1.0000 0.0809
-4.750 -0.4530 0.02485 0.01315 -0.0168 1.0000 0.0868
-4.500 -0.4372 0.02380 0.01207 -0.0151 1.0000 0.0975
-4.250 -0.4223 0.02267 0.01104 -0.0134 1.0000 0.1147
-4.000 -0.4083 0.02128 0.00990 -0.0118 1.0000 0.1466
-3.750 -0.4031 0.01876 0.00889 -0.0096 1.0000 0.3587
-3.500 -0.3970 0.01829 0.00934 -0.0045 1.0000 0.5879
-3.250 -0.3882 0.01851 0.00974 0.0005 1.0000 0.6750
-3.000 -0.3861 0.01919 0.01065 0.0082 1.0000 0.7526
-2.750 -0.3820 0.01982 0.01133 0.0156 1.0000 0.8081
-2.500 -0.3679 0.01993 0.01130 0.0191 1.0000 0.8366
-2.250 -0.3423 0.01984 0.01095 0.0192 0.9968 0.8520
-2.000 -0.3028 0.01983 0.01066 0.0164 0.9873 0.8656
-1.750 -0.2627 0.01980 0.01040 0.0134 0.9781 0.8779
-1.500 -0.2203 0.01979 0.01016 0.0100 0.9695 0.8888
-1.250 -0.1760 0.01980 0.00999 0.0061 0.9617 0.8997
-1.000 -0.1332 0.01978 0.00983 0.0026 0.9530 0.9109
-0.750 -0.0897 0.01978 0.00969 -0.0012 0.9448 0.9224
-0.500 -0.0399 0.01979 0.00961 -0.0061 0.9379 0.9315
-0.250 0.0075 0.01979 0.00954 -0.0106 0.9301 0.9414
0.000 0.0562 0.01979 0.00949 -0.0154 0.9229 0.9513
0.250 0.1050 0.01977 0.00947 -0.0203 0.9149 0.9603
0.500 0.1561 0.01974 0.00947 -0.0255 0.9078 0.9688
0.750 0.2015 0.01973 0.00951 -0.0298 0.8987 0.9792
1.000 0.2504 0.01969 0.00956 -0.0346 0.8913 0.9890
1.250 0.2907 0.01972 0.00969 -0.0380 0.8813 1.0000
1.500 0.3113 0.01987 0.00990 -0.0375 0.8689 1.0000
1.750 0.3312 0.02006 0.01019 -0.0368 0.8573 1.0000
2.000 0.3529 0.02025 0.01046 -0.0363 0.8467 1.0000
2.250 0.3723 0.02050 0.01080 -0.0353 0.8355 1.0000
2.500 0.3897 0.02080 0.01120 -0.0339 0.8235 1.0000
2.750 0.4095 0.02109 0.01161 -0.0328 0.8116 1.0000
3.000 0.4316 0.02135 0.01203 -0.0320 0.7999 1.0000
3.250 0.4561 0.02157 0.01242 -0.0313 0.7884 1.0000
3.500 0.4807 0.02175 0.01280 -0.0304 0.7759 1.0000
3.750 0.5014 0.02175 0.01299 -0.0284 0.7545 1.0000
4.000 0.5242 0.02051 0.01184 -0.0234 0.7059 1.0000
4.250 0.5377 0.01944 0.01059 -0.0171 0.6250 1.0000
4.500 0.5531 0.01924 0.01021 -0.0132 0.5312 1.0000
4.750 0.5596 0.02047 0.01002 -0.0087 0.2710 1.0000
5.000 0.5664 0.02305 0.01151 -0.0067 0.1408 1.0000
5.250 0.5814 0.02469 0.01289 -0.0054 0.1074 1.0000
5.500 0.5988 0.02608 0.01430 -0.0041 0.0933 1.0000
5.750 0.6175 0.02747 0.01571 -0.0028 0.0840 1.0000
6.000 0.6382 0.02881 0.01705 -0.0019 0.0746 1.0000
6.250 0.6639 0.03025 0.01866 -0.0011 0.0690 1.0000
6.500 0.6913 0.03184 0.02037 -0.0005 0.0641 1.0000
6.750 0.7180 0.03375 0.02234 -0.0003 0.0594 1.0000
7.000 0.7447 0.03575 0.02471 0.0002 0.0557 1.0000
7.250 0.7699 0.03818 0.02746 0.0008 0.0538 1.0000
7.500 0.7924 0.04084 0.03046 0.0015 0.0523 1.0000
7.750 0.8120 0.04356 0.03349 0.0021 0.0509 1.0000
8.000 0.8290 0.04652 0.03666 0.0026 0.0488 1.0000
8.250 0.8406 0.04980 0.04041 0.0038 0.0473 1.0000
8.500 0.8485 0.05323 0.04439 0.0051 0.0463 1.0000
8.750 0.8534 0.05698 0.04859 0.0062 0.0460 1.0000
9.000 0.8545 0.06088 0.05286 0.0073 0.0459 1.0000
9.250 0.8512 0.06491 0.05722 0.0082 0.0459 1.0000
9.500 0.8439 0.06897 0.06155 0.0088 0.0460 1.0000
9.750 0.8321 0.07292 0.06570 0.0094 0.0462 1.0000
10.000 0.8164 0.07702 0.06993 0.0094 0.0465 1.0000
10.250 0.8004 0.08161 0.07463 0.0080 0.0467 1.0000
10.500 0.7846 0.08688 0.07996 0.0054 0.0470 1.0000
10.750 0.7707 0.09271 0.08584 0.0020 0.0474 1.0000
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Polar data table (+)
Polar graphs
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