EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Reynolds number: 50,000 Max Cl/Cd: 25.62 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea61012-il-50000.txt Download as CSV file: xf-ea61012-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6871 0.08506 0.07754 -0.0134 1.0000 0.2131
-9.250 -0.7071 0.07654 0.06901 -0.0173 1.0000 0.1993
-9.000 -0.7409 0.06972 0.06211 -0.0191 1.0000 0.1910
-8.750 -0.8066 0.06145 0.05321 -0.0197 1.0000 0.1798
-8.500 -0.8032 0.05725 0.04880 -0.0192 1.0000 0.1790
-8.250 -0.8002 0.05337 0.04461 -0.0184 1.0000 0.1787
-8.000 -0.7943 0.04977 0.04066 -0.0174 1.0000 0.1786
-7.750 -0.7843 0.04647 0.03699 -0.0164 1.0000 0.1786
-7.500 -0.7708 0.04345 0.03362 -0.0154 1.0000 0.1787
-7.250 -0.7565 0.04085 0.03057 -0.0142 1.0000 0.1799
-7.000 -0.7365 0.03836 0.02805 -0.0136 1.0000 0.1834
-6.750 -0.7158 0.03635 0.02594 -0.0128 1.0000 0.1873
-6.500 -0.6956 0.03439 0.02374 -0.0118 1.0000 0.1907
-6.250 -0.6754 0.03261 0.02164 -0.0107 1.0000 0.1943
-6.000 -0.6542 0.03093 0.01994 -0.0097 1.0000 0.2001
-5.750 -0.6338 0.02959 0.01858 -0.0085 1.0000 0.2079
-5.500 -0.6126 0.02824 0.01719 -0.0072 1.0000 0.2157
-5.250 -0.5913 0.02705 0.01604 -0.0059 1.0000 0.2256
-5.000 -0.5716 0.02595 0.01509 -0.0043 1.0000 0.2392
-4.750 -0.5549 0.02492 0.01420 -0.0023 1.0000 0.2592
-4.500 -0.5419 0.02375 0.01339 0.0002 1.0000 0.2877
-4.250 -0.5357 0.02217 0.01247 0.0036 1.0000 0.3501
-4.000 -0.5435 0.02121 0.01351 0.0127 1.0000 0.5819
-3.750 -0.5422 0.02286 0.01524 0.0215 1.0000 0.6841
-3.500 -0.5336 0.02414 0.01641 0.0283 1.0000 0.7331
-3.250 -0.5250 0.02486 0.01701 0.0341 1.0000 0.7727
-3.000 -0.5023 0.02578 0.01780 0.0384 1.0000 0.8089
-2.750 -0.2099 0.02956 0.02072 0.0073 1.0000 0.9147
-2.500 -0.1312 0.02848 0.01940 -0.0023 1.0000 0.9454
-2.250 -0.0787 0.02754 0.01833 -0.0085 1.0000 0.9684
-2.000 -0.0191 0.02629 0.01697 -0.0163 1.0000 0.9869
-1.750 0.0244 0.02542 0.01603 -0.0216 1.0000 1.0000
-1.500 0.0193 0.02558 0.01618 -0.0184 1.0000 1.0000
-1.250 0.0143 0.02572 0.01632 -0.0151 1.0000 1.0000
-1.000 0.0102 0.02584 0.01642 -0.0119 1.0000 1.0000
-0.750 0.0070 0.02592 0.01649 -0.0088 1.0000 1.0000
-0.500 0.0043 0.02597 0.01654 -0.0059 1.0000 1.0000
-0.250 0.0021 0.02600 0.01657 -0.0029 1.0000 1.0000
0.000 0.0000 0.02601 0.01657 0.0000 1.0000 1.0000
0.250 -0.0021 0.02600 0.01656 0.0029 1.0000 1.0000
0.500 -0.0043 0.02597 0.01653 0.0059 1.0000 1.0000
0.750 -0.0070 0.02591 0.01649 0.0088 1.0000 1.0000
1.000 -0.0102 0.02583 0.01641 0.0119 1.0000 1.0000
1.250 -0.0143 0.02572 0.01631 0.0151 1.0000 1.0000
1.500 -0.0193 0.02557 0.01617 0.0184 1.0000 1.0000
1.750 -0.0244 0.02541 0.01602 0.0216 1.0000 1.0000
2.000 0.0189 0.02627 0.01695 0.0164 0.9869 1.0000
2.250 0.0781 0.02751 0.01831 0.0086 0.9685 1.0000
2.500 0.1311 0.02846 0.01939 0.0023 0.9455 1.0000
2.750 0.2105 0.02955 0.02071 -0.0074 0.9144 1.0000
3.000 0.5022 0.02578 0.01780 -0.0384 0.8090 1.0000
3.250 0.5250 0.02486 0.01701 -0.0340 0.7727 1.0000
3.500 0.5335 0.02413 0.01640 -0.0282 0.7330 1.0000
3.750 0.5422 0.02287 0.01524 -0.0216 0.6843 1.0000
4.000 0.5436 0.02122 0.01352 -0.0128 0.5833 1.0000
4.250 0.5356 0.02217 0.01247 -0.0036 0.3501 1.0000
4.500 0.5419 0.02375 0.01339 -0.0002 0.2878 1.0000
4.750 0.5549 0.02491 0.01419 0.0023 0.2593 1.0000
5.000 0.5715 0.02595 0.01509 0.0043 0.2392 1.0000
5.250 0.5912 0.02705 0.01604 0.0059 0.2256 1.0000
5.500 0.6126 0.02824 0.01719 0.0072 0.2156 1.0000
5.750 0.6337 0.02959 0.01858 0.0085 0.2079 1.0000
6.000 0.6542 0.03093 0.01994 0.0097 0.2002 1.0000
6.250 0.6754 0.03261 0.02164 0.0107 0.1943 1.0000
6.500 0.6956 0.03439 0.02374 0.0118 0.1907 1.0000
6.750 0.7158 0.03635 0.02594 0.0128 0.1874 1.0000
7.000 0.7365 0.03836 0.02805 0.0136 0.1834 1.0000
7.250 0.7566 0.04085 0.03057 0.0142 0.1799 1.0000
7.500 0.7708 0.04346 0.03363 0.0154 0.1787 1.0000
7.750 0.7844 0.04646 0.03699 0.0164 0.1786 1.0000
8.000 0.7942 0.04979 0.04068 0.0174 0.1786 1.0000
8.250 0.8002 0.05337 0.04462 0.0184 0.1787 1.0000
8.500 0.8031 0.05727 0.04882 0.0192 0.1789 1.0000
8.750 0.8065 0.06146 0.05322 0.0197 0.1798 1.0000
9.000 0.7410 0.06974 0.06213 0.0190 0.1909 1.0000
9.250 0.7050 0.07672 0.06919 0.0171 0.1994 1.0000
9.500 0.4267 0.10811 0.10085 -0.0112 0.3935 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il)