NACA 64-108 AIRFOIL (n64108-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-108 AIRFOIL (n64108-il) Reynolds number: 500,000 Max Cl/Cd: 60.03 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64108-il-500000.txt Download as CSV file: xf-n64108-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6169 0.12434 0.12201 0.0106 1.0000 0.0140
-11.000 -0.6139 0.11977 0.11746 0.0084 1.0000 0.0140
-10.750 -0.6110 0.11530 0.11300 0.0063 1.0000 0.0140
-5.750 -0.5354 0.02604 0.02132 -0.0225 1.0000 0.0210
-5.500 -0.5180 0.02096 0.01577 -0.0212 1.0000 0.0194
-5.250 -0.4962 0.01773 0.01216 -0.0198 1.0000 0.0184
-5.000 -0.4739 0.01595 0.01014 -0.0186 1.0000 0.0188
-4.750 -0.4523 0.01472 0.00878 -0.0173 1.0000 0.0195
-4.500 -0.4322 0.01383 0.00780 -0.0159 1.0000 0.0204
-4.250 -0.4022 0.01371 0.00760 -0.0165 0.9977 0.0216
-4.000 -0.3699 0.01132 0.00511 -0.0178 0.9931 0.0236
-3.750 -0.3348 0.01053 0.00427 -0.0196 0.9878 0.0260
-3.500 -0.2976 0.00995 0.00364 -0.0218 0.9836 0.0291
-3.250 -0.2629 0.00947 0.00311 -0.0234 0.9767 0.0336
-3.000 -0.2273 0.00893 0.00253 -0.0252 0.9701 0.0442
-2.750 -0.1970 0.00794 0.00201 -0.0263 0.9600 0.1658
-2.500 -0.1731 0.00625 0.00167 -0.0267 0.9482 0.5279
-2.250 -0.1468 0.00595 0.00163 -0.0263 0.9368 0.6139
-2.000 -0.1210 0.00583 0.00160 -0.0257 0.9253 0.6576
-1.500 -0.0705 0.00569 0.00156 -0.0241 0.9016 0.7272
-1.250 -0.0457 0.00566 0.00160 -0.0231 0.8900 0.7648
-1.000 -0.0202 0.00565 0.00162 -0.0224 0.8787 0.7880
-0.750 0.0060 0.00564 0.00159 -0.0218 0.8676 0.8020
-0.500 0.0324 0.00563 0.00157 -0.0213 0.8567 0.8137
-0.250 0.0589 0.00563 0.00155 -0.0209 0.8455 0.8251
0.000 0.0856 0.00562 0.00155 -0.0205 0.8343 0.8369
0.250 0.1122 0.00563 0.00156 -0.0202 0.8232 0.8487
0.500 0.1387 0.00565 0.00157 -0.0197 0.8123 0.8607
0.750 0.1649 0.00567 0.00159 -0.0192 0.8017 0.8730
1.000 0.1907 0.00568 0.00162 -0.0186 0.7894 0.8851
1.250 0.2161 0.00569 0.00164 -0.0179 0.7756 0.8973
1.500 0.2408 0.00570 0.00166 -0.0170 0.7545 0.9100
1.750 0.2640 0.00575 0.00161 -0.0156 0.7186 0.9237
2.000 0.2878 0.00580 0.00159 -0.0144 0.6839 0.9383
2.250 0.3134 0.00585 0.00162 -0.0138 0.6576 0.9532
2.500 0.3430 0.00595 0.00165 -0.0141 0.6248 0.9673
2.750 0.3758 0.00626 0.00171 -0.0152 0.5453 0.9800
3.000 0.4061 0.00760 0.00201 -0.0168 0.2832 0.9947
3.250 0.4277 0.00918 0.00259 -0.0168 0.0575 1.0000
3.500 0.4530 0.00966 0.00301 -0.0166 0.0396 1.0000
3.750 0.4786 0.01020 0.00360 -0.0163 0.0315 1.0000
4.000 0.5046 0.01068 0.00411 -0.0160 0.0276 1.0000
4.250 0.5289 0.01152 0.00499 -0.0156 0.0245 1.0000
4.500 0.5513 0.01288 0.00645 -0.0147 0.0228 1.0000
4.750 0.5776 0.01336 0.00697 -0.0145 0.0217 1.0000
5.000 0.6030 0.01413 0.00782 -0.0141 0.0204 1.0000
5.250 0.6280 0.01518 0.00895 -0.0136 0.0195 1.0000
5.500 0.6533 0.01636 0.01023 -0.0131 0.0185 1.0000
5.750 0.6784 0.01793 0.01194 -0.0125 0.0179 1.0000
6.000 0.7033 0.01912 0.01324 -0.0121 0.0168 1.0000
6.250 0.7256 0.02166 0.01595 -0.0116 0.0155 1.0000
6.500 0.7477 0.02449 0.01912 -0.0107 0.0155 1.0000
6.750 0.7692 0.02719 0.02214 -0.0097 0.0156 1.0000
7.000 0.7353 0.03007 0.02650 -0.0051 0.0230 1.0000
7.250 0.7904 0.04121 0.03764 -0.0045 0.0208 1.0000
7.500 0.8042 0.04472 0.04144 -0.0034 0.0193 1.0000
7.750 0.8147 0.04828 0.04523 -0.0026 0.0182 1.0000
8.000 0.8237 0.05137 0.04849 -0.0021 0.0173 1.0000
8.250 0.8315 0.05394 0.05116 -0.0017 0.0166 1.0000
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Polar data table (+)
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