NACA 64-108 AIRFOIL (n64108-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-108 AIRFOIL (n64108-il) Reynolds number: 100,000 Max Cl/Cd: 38.97 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64108-il-100000.txt Download as CSV file: xf-n64108-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-108 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6104 0.08831 0.08364 -0.0073 1.0000 0.1066
-8.250 -0.6263 0.08328 0.07869 -0.0137 1.0000 0.1089
-8.000 -0.6541 0.07892 0.07416 -0.0218 1.0000 0.1108
-7.500 -0.6513 0.07102 0.06603 -0.0244 1.0000 0.1246
-7.250 -0.6271 0.06603 0.06135 -0.0212 1.0000 0.1318
-7.000 -0.6263 0.06177 0.05697 -0.0232 1.0000 0.1420
-6.750 -0.6194 0.05813 0.05322 -0.0238 1.0000 0.1554
-6.250 -0.6002 0.05175 0.04668 -0.0234 1.0000 0.1961
-6.000 -0.5865 0.04944 0.04453 -0.0209 1.0000 0.2273
-5.500 -0.5312 0.03583 0.02864 -0.0261 1.0000 0.1089
-5.250 -0.4998 0.03139 0.02288 -0.0236 1.0000 0.0708
-5.000 -0.4766 0.02757 0.01890 -0.0230 1.0000 0.0677
-4.750 -0.4521 0.02501 0.01600 -0.0219 1.0000 0.0660
-4.500 -0.4274 0.02323 0.01390 -0.0208 1.0000 0.0682
-4.250 -0.4030 0.02169 0.01208 -0.0196 1.0000 0.0712
-4.000 -0.3797 0.01966 0.01009 -0.0185 1.0000 0.0736
-3.750 -0.3575 0.01830 0.00874 -0.0172 1.0000 0.0780
-3.500 -0.3368 0.01711 0.00758 -0.0157 1.0000 0.0859
-3.250 -0.3176 0.01602 0.00660 -0.0143 1.0000 0.1003
-3.000 -0.2982 0.01487 0.00557 -0.0128 1.0000 0.1271
-2.750 -0.2983 0.01178 0.00539 -0.0072 1.0000 0.6932
-2.500 -0.2915 0.01192 0.00569 -0.0012 1.0000 0.7827
-2.250 -0.2869 0.01212 0.00595 0.0052 1.0000 0.8400
-2.000 -0.2842 0.01236 0.00619 0.0125 1.0000 0.8966
-1.750 -0.1848 0.01323 0.00670 0.0036 1.0000 0.9632
-1.500 -0.1257 0.01303 0.00626 -0.0033 1.0000 0.9748
-1.250 -0.0720 0.01277 0.00581 -0.0094 1.0000 0.9864
-1.000 -0.0183 0.01249 0.00541 -0.0156 1.0000 0.9984
-0.750 -0.0318 0.01239 0.00532 -0.0108 1.0000 1.0000
-0.500 -0.0609 0.01223 0.00517 -0.0035 1.0000 1.0000
-0.250 -0.0668 0.01218 0.00505 0.0003 1.0000 1.0000
0.000 -0.0526 0.01232 0.00510 0.0011 0.9996 1.0000
0.250 -0.0052 0.01268 0.00539 -0.0040 0.9900 1.0000
0.500 0.0422 0.01306 0.00573 -0.0089 0.9810 1.0000
0.750 0.0875 0.01337 0.00604 -0.0133 0.9712 1.0000
1.000 0.1304 0.01367 0.00637 -0.0172 0.9609 1.0000
1.250 0.1749 0.01395 0.00671 -0.0212 0.9516 1.0000
1.500 0.2220 0.01420 0.00705 -0.0256 0.9426 1.0000
1.750 0.2627 0.01443 0.00742 -0.0287 0.9317 1.0000
2.000 0.3070 0.01461 0.00774 -0.0323 0.9209 1.0000
2.250 0.3508 0.01472 0.00804 -0.0354 0.9091 1.0000
2.500 0.3909 0.01478 0.00834 -0.0375 0.8954 1.0000
2.750 0.4306 0.01346 0.00715 -0.0345 0.8472 1.0000
3.000 0.4435 0.01268 0.00632 -0.0278 0.7930 1.0000
3.250 0.4612 0.01239 0.00604 -0.0238 0.7473 1.0000
3.500 0.4746 0.01218 0.00559 -0.0184 0.6361 1.0000
3.750 0.4709 0.01574 0.00623 -0.0126 0.1267 1.0000
4.000 0.4910 0.01717 0.00753 -0.0113 0.0960 1.0000
4.250 0.5122 0.01853 0.00880 -0.0101 0.0827 1.0000
4.500 0.5362 0.01994 0.01016 -0.0091 0.0757 1.0000
4.750 0.5617 0.02203 0.01214 -0.0084 0.0713 1.0000
5.000 0.5885 0.02332 0.01366 -0.0077 0.0651 1.0000
5.250 0.6159 0.02539 0.01596 -0.0070 0.0637 1.0000
5.500 0.6430 0.02788 0.01884 -0.0061 0.0644 1.0000
5.750 0.6689 0.03098 0.02241 -0.0051 0.0676 1.0000
6.000 0.6925 0.03465 0.02641 -0.0042 0.0704 1.0000
6.250 0.7139 0.03868 0.03075 -0.0035 0.0715 1.0000
7.750 0.7760 0.06989 0.06534 -0.0034 0.1467 1.0000
8.000 0.7926 0.07446 0.06980 -0.0021 0.1384 1.0000
8.250 0.7668 0.07849 0.07406 -0.0058 0.1322 1.0000
8.500 0.7843 0.08295 0.07844 -0.0039 0.1247 1.0000
8.750 0.7542 0.08748 0.08304 -0.0084 0.1234 1.0000
9.000 0.7321 0.09367 0.08915 -0.0156 0.1210 1.0000
9.250 0.7667 0.09623 0.09174 -0.0078 0.1116 1.0000
9.500 0.7424 0.10235 0.09780 -0.0158 0.1112 1.0000
9.750 0.7214 0.10796 0.10330 -0.0235 0.1080 1.0000
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Polar data table (+)
Polar graphs
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