NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 100,000 Max Cl/Cd: 41.64 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64110-il-100000.txt Download as CSV file: xf-n64110-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5973 0.12572 0.12064 0.0047 1.0000 0.0995
-11.000 -0.5947 0.12194 0.11687 0.0030 1.0000 0.1042
-10.750 -0.6164 0.11782 0.11287 -0.0040 1.0000 0.1079
-10.500 -0.6061 0.11271 0.10777 -0.0035 1.0000 0.1104
-10.250 -0.5907 0.10951 0.10452 -0.0021 1.0000 0.1155
-10.000 -0.6087 0.10497 0.10010 -0.0079 1.0000 0.1213
-9.750 -0.6448 0.09885 0.09412 -0.0181 1.0000 0.1222
-9.500 -0.5931 0.09708 0.09221 -0.0077 1.0000 0.1308
-9.250 -0.6243 0.09121 0.08648 -0.0164 1.0000 0.1355
-9.000 -0.5984 0.08870 0.08394 -0.0116 1.0000 0.1436
-8.750 -0.6199 0.08300 0.07834 -0.0178 1.0000 0.1483
-8.500 -0.6136 0.07956 0.07494 -0.0170 1.0000 0.1570
-8.250 -0.6372 0.07446 0.06988 -0.0215 1.0000 0.1614
-7.750 -0.6451 0.06678 0.06220 -0.0229 1.0000 0.1836
-6.750 -0.6377 0.04256 0.03594 -0.0282 1.0000 0.1057
-6.500 -0.6185 0.03617 0.02862 -0.0262 1.0000 0.0765
-6.250 -0.6001 0.03232 0.02408 -0.0244 1.0000 0.0701
-6.000 -0.5790 0.03025 0.02145 -0.0226 1.0000 0.0682
-5.750 -0.5594 0.02774 0.01889 -0.0217 1.0000 0.0712
-5.500 -0.5386 0.02591 0.01688 -0.0204 1.0000 0.0731
-5.250 -0.5171 0.02413 0.01487 -0.0189 1.0000 0.0740
-5.000 -0.4961 0.02263 0.01324 -0.0174 1.0000 0.0760
-4.750 -0.4761 0.02152 0.01197 -0.0157 1.0000 0.0799
-4.500 -0.4583 0.02008 0.01067 -0.0141 1.0000 0.0860
-4.250 -0.4410 0.01910 0.00970 -0.0122 1.0000 0.0921
-4.000 -0.4257 0.01802 0.00873 -0.0102 1.0000 0.0999
-3.750 -0.4103 0.01713 0.00791 -0.0084 1.0000 0.1165
-3.500 -0.3946 0.01590 0.00694 -0.0068 1.0000 0.1548
-3.250 -0.3946 0.01340 0.00703 -0.0018 1.0000 0.6427
-3.000 -0.3834 0.01376 0.00745 0.0023 1.0000 0.7092
-2.750 -0.3718 0.01410 0.00781 0.0062 1.0000 0.7476
-2.500 -0.3618 0.01445 0.00815 0.0103 1.0000 0.7819
-2.250 -0.3599 0.01493 0.00869 0.0168 1.0000 0.8209
-2.000 -0.3456 0.01576 0.00958 0.0222 0.9909 0.8698
-1.750 -0.3106 0.01628 0.00998 0.0222 0.9829 0.8975
-1.500 -0.2686 0.01639 0.00993 0.0191 0.9745 0.9109
-1.250 -0.2180 0.01656 0.00992 0.0142 0.9682 0.9217
-1.000 -0.1700 0.01663 0.00986 0.0096 0.9606 0.9321
-0.750 -0.1115 0.01677 0.00988 0.0030 0.9559 0.9406
-0.500 -0.0588 0.01681 0.00983 -0.0026 0.9488 0.9484
-0.250 -0.0003 0.01687 0.00982 -0.0093 0.9437 0.9565
0.000 0.0653 0.01685 0.00978 -0.0174 0.9395 0.9610
0.250 0.1225 0.01680 0.00973 -0.0240 0.9332 0.9683
0.500 0.1863 0.01667 0.00962 -0.0316 0.9288 0.9736
0.750 0.2394 0.01656 0.00957 -0.0374 0.9211 0.9815
1.000 0.2984 0.01635 0.00944 -0.0442 0.9152 0.9877
1.250 0.3502 0.01619 0.00938 -0.0497 0.9066 0.9960
1.500 0.3956 0.01602 0.00934 -0.0537 0.8988 1.0000
1.750 0.4151 0.01612 0.00952 -0.0530 0.8854 1.0000
2.000 0.4329 0.01623 0.00970 -0.0517 0.8718 1.0000
2.250 0.4489 0.01638 0.00992 -0.0501 0.8584 1.0000
2.500 0.4639 0.01656 0.01019 -0.0481 0.8453 1.0000
2.750 0.4797 0.01652 0.01025 -0.0455 0.8297 1.0000
3.000 0.4918 0.01597 0.00972 -0.0406 0.8038 1.0000
3.250 0.5021 0.01497 0.00866 -0.0341 0.7696 1.0000
3.500 0.5131 0.01435 0.00798 -0.0289 0.7342 1.0000
3.750 0.5287 0.01403 0.00769 -0.0254 0.7046 1.0000
4.000 0.5455 0.01373 0.00742 -0.0221 0.6676 1.0000
4.250 0.5588 0.01342 0.00693 -0.0177 0.5763 1.0000
4.500 0.5494 0.01650 0.00746 -0.0114 0.1547 1.0000
4.750 0.5646 0.01811 0.00870 -0.0096 0.1123 1.0000
5.000 0.5829 0.01945 0.00990 -0.0081 0.0973 1.0000
5.250 0.6044 0.02069 0.01108 -0.0070 0.0863 1.0000
5.500 0.6281 0.02229 0.01262 -0.0062 0.0792 1.0000
5.750 0.6548 0.02386 0.01424 -0.0056 0.0748 1.0000
6.000 0.6816 0.02576 0.01612 -0.0053 0.0710 1.0000
6.250 0.7077 0.02803 0.01854 -0.0049 0.0675 1.0000
6.500 0.7334 0.03019 0.02105 -0.0041 0.0663 1.0000
6.750 0.7579 0.03291 0.02416 -0.0033 0.0665 1.0000
7.000 0.7803 0.03625 0.02784 -0.0024 0.0677 1.0000
7.250 0.8020 0.04047 0.03230 -0.0019 0.0694 1.0000
7.500 0.8164 0.04344 0.03636 0.0010 0.0764 1.0000
7.750 0.8320 0.04848 0.04162 0.0017 0.0810 1.0000
9.250 0.7557 0.08724 0.08275 -0.0042 0.1638 1.0000
9.500 0.6159 0.09182 0.08750 -0.0100 0.1780 1.0000
9.750 0.6451 0.09430 0.09001 -0.0063 0.1652 1.0000
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Polar data table (+)
Polar graphs
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