EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Reynolds number: 100,000 Max Cl/Cd: 33.52 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea61012-il-100000.txt Download as CSV file: xf-ea61012-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER EA 6(-1)-012 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5380 0.10183 0.09679 -0.0038 1.0000 0.2132
-10.000 -0.4958 0.09904 0.09394 -0.0011 1.0000 0.2161
-9.750 -0.8636 0.06041 0.05427 -0.0239 1.0000 0.1171
-9.500 -0.8649 0.05602 0.04968 -0.0230 1.0000 0.1156
-9.250 -0.8669 0.05177 0.04515 -0.0217 1.0000 0.1142
-9.000 -0.8654 0.04772 0.04073 -0.0203 1.0000 0.1128
-8.750 -0.8599 0.04388 0.03650 -0.0189 1.0000 0.1113
-8.500 -0.8493 0.04076 0.03300 -0.0175 1.0000 0.1111
-8.250 -0.8345 0.03830 0.03024 -0.0164 1.0000 0.1122
-8.000 -0.8178 0.03604 0.02766 -0.0153 1.0000 0.1134
-7.750 -0.7992 0.03388 0.02519 -0.0142 1.0000 0.1143
-7.500 -0.7791 0.03193 0.02297 -0.0133 1.0000 0.1152
-7.250 -0.7586 0.03039 0.02113 -0.0123 1.0000 0.1171
-7.000 -0.7382 0.02919 0.01962 -0.0111 1.0000 0.1190
-6.750 -0.7153 0.02720 0.01766 -0.0107 1.0000 0.1211
-6.500 -0.6940 0.02590 0.01638 -0.0097 1.0000 0.1233
-6.250 -0.6759 0.02489 0.01535 -0.0083 1.0000 0.1258
-6.000 -0.6619 0.02410 0.01451 -0.0061 1.0000 0.1289
-5.750 -0.6498 0.02345 0.01376 -0.0037 1.0000 0.1321
-5.500 -0.6359 0.02261 0.01290 -0.0015 1.0000 0.1352
-5.250 -0.6214 0.02179 0.01221 0.0004 1.0000 0.1388
-5.000 -0.6069 0.02117 0.01161 0.0024 1.0000 0.1431
-4.750 -0.5917 0.02066 0.01104 0.0043 1.0000 0.1486
-4.500 -0.5777 0.01994 0.01051 0.0062 1.0000 0.1558
-4.250 -0.5625 0.01943 0.01003 0.0080 1.0000 0.1652
-4.000 -0.5484 0.01880 0.00953 0.0099 1.0000 0.1784
-3.750 -0.5272 0.01785 0.00891 0.0104 0.9973 0.2122
-3.500 -0.5159 0.01539 0.00865 0.0123 0.9913 0.5128
-3.250 -0.4895 0.01593 0.00954 0.0137 0.9847 0.6546
-3.000 -0.4595 0.01664 0.01026 0.0143 0.9788 0.6978
-2.750 -0.4319 0.01726 0.01083 0.0151 0.9729 0.7278
-2.500 -0.4002 0.01788 0.01138 0.0151 0.9684 0.7542
-2.250 -0.3783 0.01825 0.01172 0.0169 0.9621 0.7748
-2.000 -0.3511 0.01878 0.01227 0.0184 0.9568 0.7986
-1.750 -0.3249 0.01937 0.01288 0.0205 0.9528 0.8267
-1.500 -0.3075 0.01975 0.01327 0.0239 0.9469 0.8507
-1.250 -0.2742 0.02025 0.01376 0.0245 0.9424 0.8736
-1.000 -0.2189 0.02075 0.01420 0.0205 0.9386 0.8896
-0.750 -0.1686 0.02093 0.01433 0.0165 0.9312 0.8972
-0.500 -0.1177 0.02090 0.01424 0.0120 0.9238 0.9048
-0.250 -0.0393 0.02100 0.01430 0.0026 0.9211 0.9076
0.000 0.0003 0.02106 0.01437 -0.0001 0.9139 0.9139
0.250 0.0399 0.02100 0.01430 -0.0027 0.9076 0.9211
0.500 0.1175 0.02091 0.01425 -0.0119 0.9049 0.9239
0.750 0.1687 0.02093 0.01433 -0.0165 0.8972 0.9312
1.000 0.2188 0.02074 0.01420 -0.0204 0.8895 0.9385
1.250 0.2741 0.02025 0.01376 -0.0244 0.8736 0.9423
1.500 0.3077 0.01975 0.01327 -0.0239 0.8507 0.9470
1.750 0.3248 0.01936 0.01286 -0.0204 0.8263 0.9528
2.000 0.3512 0.01877 0.01226 -0.0184 0.7986 0.9568
2.250 0.3783 0.01824 0.01171 -0.0168 0.7747 0.9621
2.500 0.4004 0.01788 0.01138 -0.0151 0.7544 0.9684
2.750 0.4320 0.01727 0.01084 -0.0151 0.7280 0.9730
3.000 0.4595 0.01664 0.01025 -0.0143 0.6977 0.9788
3.250 0.4895 0.01593 0.00954 -0.0137 0.6539 0.9847
3.500 0.5158 0.01539 0.00863 -0.0122 0.5097 0.9913
3.750 0.5273 0.01784 0.00891 -0.0104 0.2124 0.9973
4.000 0.5483 0.01880 0.00953 -0.0099 0.1786 1.0000
4.250 0.5625 0.01942 0.01002 -0.0080 0.1653 1.0000
4.500 0.5777 0.01993 0.01050 -0.0062 0.1557 1.0000
4.750 0.5916 0.02067 0.01104 -0.0043 0.1487 1.0000
5.000 0.6068 0.02117 0.01161 -0.0024 0.1431 1.0000
5.250 0.6214 0.02179 0.01221 -0.0004 0.1388 1.0000
5.500 0.6358 0.02261 0.01290 0.0015 0.1351 1.0000
5.750 0.6497 0.02344 0.01375 0.0037 0.1321 1.0000
6.000 0.6619 0.02410 0.01450 0.0061 0.1289 1.0000
6.250 0.6758 0.02488 0.01535 0.0083 0.1258 1.0000
6.500 0.6939 0.02590 0.01638 0.0098 0.1233 1.0000
6.750 0.7153 0.02719 0.01765 0.0107 0.1212 1.0000
7.000 0.7381 0.02918 0.01961 0.0112 0.1189 1.0000
7.250 0.7586 0.03038 0.02112 0.0123 0.1171 1.0000
7.500 0.7792 0.03194 0.02297 0.0133 0.1152 1.0000
7.750 0.7992 0.03387 0.02519 0.0142 0.1143 1.0000
8.000 0.8178 0.03604 0.02767 0.0153 0.1134 1.0000
8.250 0.8345 0.03830 0.03024 0.0164 0.1122 1.0000
8.500 0.8495 0.04072 0.03294 0.0175 0.1109 1.0000
8.750 0.8599 0.04389 0.03651 0.0189 0.1113 1.0000
9.000 0.8655 0.04771 0.04073 0.0203 0.1127 1.0000
9.250 0.8669 0.05179 0.04517 0.0217 0.1142 1.0000
9.500 0.8651 0.05603 0.04969 0.0230 0.1156 1.0000
9.750 0.8635 0.06042 0.05428 0.0239 0.1170 1.0000
10.000 0.6146 0.09997 0.09456 0.0000 0.2148 1.0000
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Polar data table (+)
Polar graphs
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