NACA 642-015 AIRFOIL (n64015-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 642-015 AIRFOIL (n64015-il) Reynolds number: 100,000 Max Cl/Cd: 40.33 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64015-il-100000.txt Download as CSV file: xf-n64015-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 642-015 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.4667 0.12727 0.12255 -0.0018 1.0000 0.1709
-12.250 -0.4982 0.12219 0.11753 -0.0061 1.0000 0.1775
-12.000 -0.9138 0.07275 0.06689 -0.0348 1.0000 0.0774
-11.750 -0.9063 0.06823 0.06226 -0.0348 1.0000 0.0755
-11.500 -0.9147 0.06393 0.05782 -0.0341 1.0000 0.0743
-11.250 -0.9272 0.06021 0.05392 -0.0324 1.0000 0.0733
-11.000 -0.9385 0.05665 0.05011 -0.0302 1.0000 0.0723
-10.750 -0.9455 0.05299 0.04614 -0.0281 1.0000 0.0713
-10.500 -0.9480 0.04950 0.04228 -0.0260 1.0000 0.0706
-10.250 -0.9442 0.04627 0.03869 -0.0242 1.0000 0.0702
-10.000 -0.9346 0.04357 0.03569 -0.0228 1.0000 0.0710
-9.750 -0.9223 0.04124 0.03307 -0.0215 1.0000 0.0726
-9.500 -0.9077 0.03897 0.03046 -0.0202 1.0000 0.0741
-9.250 -0.8895 0.03677 0.02792 -0.0192 1.0000 0.0753
-9.000 -0.8699 0.03493 0.02575 -0.0182 1.0000 0.0766
-8.750 -0.8385 0.03215 0.02308 -0.0191 1.0000 0.0800
-8.500 -0.8146 0.03071 0.02161 -0.0189 1.0000 0.0843
-8.250 -0.7905 0.02936 0.02008 -0.0183 1.0000 0.0885
-8.000 -0.7604 0.02752 0.01843 -0.0188 1.0000 0.0940
-7.750 -0.7419 0.02652 0.01742 -0.0176 1.0000 0.1010
-7.500 -0.7275 0.02534 0.01642 -0.0158 1.0000 0.1077
-7.250 -0.7281 0.02478 0.01590 -0.0116 1.0000 0.1137
-7.000 -0.7348 0.02421 0.01541 -0.0066 1.0000 0.1197
-6.750 -0.7408 0.02368 0.01495 -0.0019 1.0000 0.1275
-6.500 -0.7459 0.02295 0.01433 0.0026 1.0000 0.1379
-6.250 -0.7336 0.02138 0.01310 0.0034 0.9945 0.1757
-6.000 -0.7287 0.01808 0.01191 0.0042 0.9816 0.4552
-5.750 -0.6919 0.01885 0.01271 0.0030 0.9716 0.5527
-5.500 -0.6538 0.01991 0.01373 0.0022 0.9625 0.5883
-5.250 -0.6109 0.02098 0.01467 0.0005 0.9553 0.6174
-5.000 -0.5659 0.02271 0.01643 0.0000 0.9483 0.6388
-4.750 -0.5114 0.02463 0.01830 -0.0017 0.9443 0.6590
-4.500 -0.4658 0.02627 0.01988 -0.0019 0.9378 0.6746
-4.250 -0.4159 0.02748 0.02097 -0.0035 0.9321 0.6881
-4.000 -0.3828 0.02757 0.02094 -0.0044 0.9237 0.7008
-3.750 -0.3317 0.02834 0.02161 -0.0066 0.9181 0.7064
-3.500 -0.3139 0.02776 0.02091 -0.0063 0.9077 0.7173
-3.250 -0.2706 0.02803 0.02110 -0.0079 0.9009 0.7215
-3.000 -0.2542 0.02759 0.02057 -0.0070 0.8905 0.7298
-2.750 -0.2266 0.02728 0.02017 -0.0073 0.8827 0.7356
-2.500 -0.2009 0.02728 0.02012 -0.0070 0.8739 0.7405
-2.250 -0.1912 0.02653 0.01928 -0.0055 0.8656 0.7498
-2.000 -0.1641 0.02667 0.01939 -0.0052 0.8570 0.7535
-1.750 -0.1367 0.02651 0.01917 -0.0052 0.8505 0.7589
-1.500 -0.1301 0.02610 0.01871 -0.0032 0.8406 0.7673
-1.250 -0.0985 0.02610 0.01868 -0.0035 0.8349 0.7712
-1.000 -0.0809 0.02602 0.01857 -0.0026 0.8276 0.7776
-0.750 -0.0657 0.02572 0.01825 -0.0015 0.8203 0.7845
-0.500 -0.0338 0.02569 0.01819 -0.0018 0.8161 0.7885
-0.250 -0.0236 0.02571 0.01822 -0.0003 0.8068 0.7962
0.000 0.0001 0.02554 0.01803 0.0000 0.8018 0.8018
0.250 0.0236 0.02571 0.01822 0.0003 0.7962 0.8068
0.500 0.0340 0.02569 0.01819 0.0018 0.7885 0.8161
0.750 0.0658 0.02572 0.01824 0.0015 0.7844 0.8203
1.000 0.0809 0.02602 0.01857 0.0026 0.7777 0.8276
1.250 0.0986 0.02610 0.01868 0.0035 0.7712 0.8350
1.500 0.1301 0.02610 0.01871 0.0032 0.7673 0.8406
1.750 0.1368 0.02652 0.01918 0.0052 0.7589 0.8505
2.000 0.1642 0.02667 0.01939 0.0053 0.7535 0.8570
2.250 0.1915 0.02653 0.01927 0.0055 0.7498 0.8656
2.500 0.2009 0.02728 0.02012 0.0070 0.7405 0.8739
2.750 0.2267 0.02728 0.02017 0.0073 0.7356 0.8827
3.000 0.2538 0.02760 0.02058 0.0071 0.7297 0.8906
3.250 0.2708 0.02802 0.02109 0.0079 0.7215 0.9009
3.500 0.3142 0.02775 0.02090 0.0062 0.7173 0.9077
3.750 0.3318 0.02834 0.02162 0.0066 0.7065 0.9181
4.000 0.3833 0.02754 0.02091 0.0044 0.7008 0.9236
4.250 0.4160 0.02748 0.02097 0.0035 0.6881 0.9322
4.500 0.4657 0.02628 0.01989 0.0019 0.6747 0.9377
4.750 0.5115 0.02462 0.01830 0.0016 0.6590 0.9443
5.000 0.5661 0.02269 0.01642 0.0000 0.6388 0.9483
5.250 0.6109 0.02098 0.01467 -0.0005 0.6174 0.9553
5.500 0.6538 0.01992 0.01374 -0.0023 0.5884 0.9625
5.750 0.6920 0.01886 0.01271 -0.0031 0.5527 0.9717
6.000 0.7287 0.01807 0.01191 -0.0042 0.4546 0.9816
6.250 0.7338 0.02137 0.01309 -0.0035 0.1758 0.9945
6.500 0.7459 0.02294 0.01433 -0.0026 0.1380 1.0000
6.750 0.7407 0.02367 0.01494 0.0019 0.1274 1.0000
7.000 0.7347 0.02420 0.01540 0.0066 0.1196 1.0000
7.250 0.7279 0.02477 0.01589 0.0116 0.1136 1.0000
7.500 0.7274 0.02534 0.01641 0.0158 0.1077 1.0000
7.750 0.7418 0.02652 0.01742 0.0176 0.1008 1.0000
8.000 0.7604 0.02753 0.01843 0.0188 0.0939 1.0000
8.250 0.7904 0.02936 0.02008 0.0184 0.0884 1.0000
8.500 0.8148 0.03071 0.02161 0.0189 0.0843 1.0000
8.750 0.8387 0.03216 0.02309 0.0191 0.0800 1.0000
9.000 0.8701 0.03494 0.02575 0.0182 0.0766 1.0000
9.250 0.8897 0.03677 0.02792 0.0192 0.0753 1.0000
9.500 0.9077 0.03897 0.03045 0.0202 0.0740 1.0000
9.750 0.9224 0.04122 0.03305 0.0215 0.0725 1.0000
10.000 0.9347 0.04358 0.03569 0.0228 0.0709 1.0000
10.250 0.9444 0.04628 0.03870 0.0242 0.0702 1.0000
10.500 0.9483 0.04948 0.04227 0.0260 0.0705 1.0000
10.750 0.9458 0.05300 0.04615 0.0280 0.0713 1.0000
11.000 0.9388 0.05667 0.05013 0.0301 0.0723 1.0000
11.250 0.9267 0.06026 0.05397 0.0324 0.0733 1.0000
11.500 0.9149 0.06399 0.05789 0.0340 0.0744 1.0000
11.750 0.9049 0.06825 0.06228 0.0348 0.0754 1.0000
12.000 0.9141 0.07280 0.06695 0.0347 0.0774 1.0000
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Polar data table (+)
Polar graphs
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