EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Reynolds number: 100,000 Max Cl/Cd: 29.48 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea61009-il-100000.txt Download as CSV file: xf-ea61009-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER EA 6(-1)-009 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6790 0.08929 0.08445 0.0087 1.0000 0.1983
-8.500 -0.6653 0.08531 0.08046 0.0100 1.0000 0.2051
-8.250 -0.7038 0.08013 0.07540 0.0031 1.0000 0.2140
-8.000 -0.6841 0.07633 0.07162 0.0054 1.0000 0.2201
-7.750 -0.7773 0.04921 0.04295 -0.0141 1.0000 0.1151
-7.500 -0.7625 0.04420 0.03759 -0.0138 1.0000 0.1086
-7.250 -0.7523 0.03859 0.03071 -0.0125 1.0000 0.0995
-7.000 -0.7338 0.03476 0.02653 -0.0121 1.0000 0.1009
-6.750 -0.7117 0.03187 0.02334 -0.0116 1.0000 0.1019
-6.500 -0.6879 0.02950 0.02069 -0.0111 1.0000 0.1030
-6.250 -0.6633 0.02775 0.01881 -0.0107 1.0000 0.1061
-6.000 -0.6379 0.02591 0.01668 -0.0102 1.0000 0.1074
-5.750 -0.6117 0.02427 0.01481 -0.0097 1.0000 0.1087
-5.500 -0.5856 0.02302 0.01333 -0.0093 1.0000 0.1117
-5.250 -0.5593 0.02169 0.01190 -0.0088 1.0000 0.1141
-5.000 -0.5335 0.02037 0.01066 -0.0084 1.0000 0.1164
-4.750 -0.5082 0.01936 0.00969 -0.0078 1.0000 0.1193
-4.500 -0.4838 0.01850 0.00884 -0.0071 1.0000 0.1232
-4.250 -0.4605 0.01770 0.00809 -0.0063 1.0000 0.1286
-4.000 -0.4388 0.01696 0.00747 -0.0053 1.0000 0.1348
-3.750 -0.4180 0.01635 0.00688 -0.0041 1.0000 0.1416
-3.500 -0.3991 0.01573 0.00636 -0.0026 1.0000 0.1524
-3.250 -0.3811 0.01506 0.00590 -0.0011 1.0000 0.1749
-3.000 -0.3678 0.01364 0.00533 0.0007 1.0000 0.2982
-2.750 -0.3613 0.01225 0.00543 0.0048 1.0000 0.5915
-2.500 -0.3492 0.01226 0.00579 0.0088 1.0000 0.7003
-2.250 -0.3362 0.01246 0.00606 0.0125 1.0000 0.7592
-2.000 -0.3230 0.01266 0.00628 0.0162 1.0000 0.7971
-1.750 -0.3091 0.01280 0.00640 0.0194 1.0000 0.8267
-1.500 -0.2966 0.01293 0.00653 0.0231 1.0000 0.8579
-1.250 -0.2826 0.01303 0.00663 0.0266 1.0000 0.8904
-1.000 -0.2594 0.01316 0.00675 0.0283 1.0000 0.9236
-0.750 -0.2026 0.01355 0.00709 0.0238 1.0000 0.9555
-0.500 -0.1182 0.01393 0.00735 0.0135 1.0000 0.9738
-0.250 0.0124 0.01395 0.00732 -0.0049 1.0000 1.0000
0.000 0.0000 0.01399 0.00735 0.0000 1.0000 1.0000
0.250 -0.0124 0.01395 0.00732 0.0049 1.0000 1.0000
0.500 0.1194 0.01392 0.00735 -0.0137 0.9736 1.0000
0.750 0.2035 0.01354 0.00707 -0.0239 0.9552 1.0000
1.000 0.2593 0.01315 0.00675 -0.0283 0.9236 1.0000
1.250 0.2823 0.01302 0.00662 -0.0265 0.8901 1.0000
1.500 0.2964 0.01293 0.00653 -0.0231 0.8579 1.0000
1.750 0.3089 0.01280 0.00640 -0.0194 0.8268 1.0000
2.000 0.3229 0.01266 0.00628 -0.0162 0.7973 1.0000
2.250 0.3360 0.01246 0.00606 -0.0125 0.7589 1.0000
2.500 0.3490 0.01226 0.00579 -0.0088 0.7001 1.0000
2.750 0.3611 0.01225 0.00542 -0.0047 0.5916 1.0000
3.000 0.3675 0.01365 0.00533 -0.0007 0.2958 1.0000
3.250 0.3809 0.01506 0.00590 0.0011 0.1749 1.0000
3.500 0.3989 0.01573 0.00636 0.0027 0.1523 1.0000
3.750 0.4179 0.01635 0.00688 0.0041 0.1415 1.0000
4.000 0.4387 0.01696 0.00747 0.0053 0.1347 1.0000
4.250 0.4604 0.01770 0.00808 0.0063 0.1286 1.0000
4.500 0.4837 0.01850 0.00884 0.0072 0.1232 1.0000
4.750 0.5082 0.01936 0.00969 0.0078 0.1193 1.0000
5.000 0.5335 0.02037 0.01065 0.0084 0.1164 1.0000
5.250 0.5593 0.02169 0.01190 0.0088 0.1141 1.0000
5.500 0.5856 0.02302 0.01333 0.0093 0.1117 1.0000
5.750 0.6118 0.02427 0.01482 0.0097 0.1087 1.0000
6.000 0.6379 0.02591 0.01669 0.0102 0.1075 1.0000
6.250 0.6634 0.02774 0.01880 0.0107 0.1061 1.0000
6.500 0.6880 0.02949 0.02068 0.0111 0.1030 1.0000
6.750 0.7118 0.03187 0.02334 0.0116 0.1019 1.0000
7.000 0.7339 0.03475 0.02653 0.0120 0.1010 1.0000
7.250 0.7524 0.03862 0.03074 0.0124 0.0995 1.0000
7.500 0.7470 0.05223 0.04662 0.0124 0.1473 1.0000
7.750 0.7774 0.04927 0.04301 0.0141 0.1151 1.0000
8.000 0.6846 0.07635 0.07164 -0.0054 0.2201 1.0000
8.250 0.7025 0.08022 0.07550 -0.0035 0.2139 1.0000
8.500 0.6673 0.08526 0.08042 -0.0098 0.2047 1.0000
8.750 0.6767 0.08938 0.08453 -0.0093 0.1981 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il)