EPPLER 472 AIRFOIL (e472-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: EPPLER 472 AIRFOIL (e472-il) Reynolds number: 100,000 Max Cl/Cd: 26.55 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e472-il-100000.txt Download as CSV file: xf-e472-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 472 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.7600 0.07873 0.07271 -0.0218 1.0000 0.1647
-10.250 -0.7733 0.07285 0.06681 -0.0232 1.0000 0.1654
-10.000 -0.7886 0.06800 0.06195 -0.0233 1.0000 0.1663
-9.750 -0.8009 0.06426 0.05818 -0.0221 1.0000 0.1677
-9.500 -0.8126 0.06048 0.05433 -0.0210 1.0000 0.1697
-9.250 -0.8491 0.05379 0.04738 -0.0195 1.0000 0.1708
-9.000 -0.8628 0.04945 0.04277 -0.0174 1.0000 0.1736
-8.750 -0.8774 0.04482 0.03771 -0.0146 1.0000 0.1767
-8.500 -0.8906 0.04063 0.03284 -0.0110 1.0000 0.1801
-8.250 -0.8635 0.03990 0.03236 -0.0106 1.0000 0.1837
-8.000 -0.8443 0.03898 0.03142 -0.0093 1.0000 0.1879
-7.750 -0.8366 0.03668 0.02878 -0.0070 1.0000 0.1924
-7.500 -0.8300 0.03425 0.02589 -0.0044 1.0000 0.1967
-7.250 -0.8063 0.03345 0.02525 -0.0036 1.0000 0.2006
-7.000 -0.7875 0.03240 0.02412 -0.0021 1.0000 0.2055
-6.750 -0.7744 0.03076 0.02207 0.0000 1.0000 0.2105
-6.500 -0.7557 0.02937 0.02061 0.0014 1.0000 0.2147
-6.250 -0.7335 0.02863 0.01991 0.0025 1.0000 0.2196
-6.000 -0.7150 0.02757 0.01863 0.0042 1.0000 0.2252
-5.750 -0.6963 0.02633 0.01718 0.0057 1.0000 0.2300
-5.500 -0.6739 0.02556 0.01648 0.0068 1.0000 0.2349
-5.250 -0.6530 0.02478 0.01558 0.0082 1.0000 0.2410
-5.000 -0.6326 0.02383 0.01446 0.0096 1.0000 0.2465
-4.750 -0.6099 0.02311 0.01380 0.0106 1.0000 0.2516
-4.500 -0.5882 0.02244 0.01303 0.0119 1.0000 0.2579
-4.250 -0.5663 0.02168 0.01221 0.0131 1.0000 0.2640
-4.000 -0.5437 0.02110 0.01169 0.0142 1.0000 0.2702
-3.750 -0.5217 0.02054 0.01098 0.0155 1.0000 0.2770
-3.500 -0.4988 0.01988 0.01044 0.0165 1.0000 0.2833
-3.250 -0.4764 0.01941 0.00996 0.0177 1.0000 0.2911
-3.000 -0.4539 0.01886 0.00944 0.0188 1.0000 0.2984
-2.750 -0.4313 0.01841 0.00906 0.0199 1.0000 0.3062
-2.500 -0.4090 0.01796 0.00864 0.0210 1.0000 0.3148
-2.250 -0.3869 0.01758 0.00836 0.0222 1.0000 0.3243
-2.000 -0.3650 0.01718 0.00806 0.0233 1.0000 0.3340
-1.750 -0.3435 0.01689 0.00780 0.0245 1.0000 0.3454
-1.500 -0.3224 0.01655 0.00767 0.0257 1.0000 0.3573
-1.250 -0.3022 0.01628 0.00757 0.0270 1.0000 0.3713
-1.000 -0.2830 0.01605 0.00752 0.0284 1.0000 0.3874
-0.750 -0.2260 0.01569 0.00751 0.0224 0.9837 0.4201
-0.500 -0.1622 0.01489 0.00724 0.0158 0.9560 0.4760
-0.250 -0.1103 0.01376 0.00685 0.0123 0.9181 0.5714
0.000 0.0006 0.01247 0.00676 -0.0001 0.8320 0.8332
0.250 0.1103 0.01376 0.00685 -0.0123 0.5713 0.9181
0.500 0.1624 0.01489 0.00724 -0.0158 0.4758 0.9560
0.750 0.2260 0.01569 0.00751 -0.0225 0.4199 0.9837
1.000 0.2830 0.01605 0.00752 -0.0284 0.3875 1.0000
1.250 0.3022 0.01628 0.00756 -0.0270 0.3713 1.0000
1.500 0.3224 0.01655 0.00767 -0.0257 0.3572 1.0000
1.750 0.3435 0.01689 0.00780 -0.0245 0.3455 1.0000
2.000 0.3650 0.01718 0.00806 -0.0233 0.3340 1.0000
2.250 0.3869 0.01758 0.00836 -0.0222 0.3244 1.0000
2.500 0.4089 0.01796 0.00864 -0.0210 0.3148 1.0000
2.750 0.4313 0.01841 0.00906 -0.0199 0.3062 1.0000
3.000 0.4538 0.01886 0.00944 -0.0188 0.2984 1.0000
3.250 0.4764 0.01941 0.00996 -0.0177 0.2911 1.0000
3.500 0.4987 0.01988 0.01043 -0.0165 0.2833 1.0000
3.750 0.5217 0.02053 0.01097 -0.0155 0.2771 1.0000
4.000 0.5437 0.02110 0.01169 -0.0142 0.2702 1.0000
4.250 0.5663 0.02168 0.01221 -0.0131 0.2640 1.0000
4.500 0.5882 0.02244 0.01303 -0.0119 0.2579 1.0000
4.750 0.6099 0.02310 0.01380 -0.0106 0.2516 1.0000
5.000 0.6326 0.02383 0.01446 -0.0095 0.2465 1.0000
5.250 0.6530 0.02479 0.01559 -0.0082 0.2410 1.0000
5.500 0.6739 0.02556 0.01648 -0.0068 0.2349 1.0000
5.750 0.6963 0.02633 0.01718 -0.0057 0.2300 1.0000
6.000 0.7151 0.02757 0.01863 -0.0042 0.2253 1.0000
6.250 0.7335 0.02863 0.01991 -0.0025 0.2196 1.0000
6.500 0.7557 0.02938 0.02061 -0.0014 0.2147 1.0000
6.750 0.7744 0.03076 0.02207 0.0000 0.2106 1.0000
7.000 0.7875 0.03241 0.02413 0.0021 0.2055 1.0000
7.250 0.8063 0.03346 0.02526 0.0036 0.2007 1.0000
7.500 0.8300 0.03425 0.02590 0.0044 0.1967 1.0000
7.750 0.8367 0.03668 0.02877 0.0070 0.1925 1.0000
8.000 0.8443 0.03897 0.03142 0.0092 0.1879 1.0000
8.250 0.8634 0.03992 0.03238 0.0106 0.1837 1.0000
8.500 0.8905 0.04062 0.03284 0.0110 0.1801 1.0000
8.750 0.8775 0.04481 0.03770 0.0146 0.1767 1.0000
9.000 0.8631 0.04941 0.04273 0.0174 0.1735 1.0000
9.250 0.8487 0.05386 0.04746 0.0195 0.1709 1.0000
9.500 0.8085 0.06093 0.05479 0.0209 0.1698 1.0000
9.750 0.8024 0.06415 0.05806 0.0221 0.1677 1.0000
10.000 0.7876 0.06814 0.06208 0.0232 0.1664 1.0000
10.250 0.7731 0.07293 0.06690 0.0231 0.1654 1.0000
10.500 0.7609 0.07866 0.07265 0.0218 0.1647 1.0000
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Polar data table (+)
Polar graphs
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