EPPLER 426 AIRFOIL (e426-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 426 AIRFOIL (e426-il) Reynolds number: 50,000 Max Cl/Cd: 22.85 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e426-il-50000.txt Download as CSV file: xf-e426-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 426 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4738 0.09579 0.08792 0.0240 1.0000 0.4383
-8.000 -0.4642 0.09286 0.08497 0.0245 1.0000 0.4553
-7.500 -0.6187 0.06977 0.06233 -0.0053 1.0000 0.2828
-7.250 -0.6143 0.06463 0.05711 -0.0073 1.0000 0.2804
-7.000 -0.6133 0.05904 0.05142 -0.0099 1.0000 0.2771
-6.750 -0.6156 0.05297 0.04514 -0.0132 1.0000 0.2754
-6.500 -0.6108 0.04787 0.03974 -0.0155 1.0000 0.2761
-6.250 -0.6004 0.04345 0.03496 -0.0170 1.0000 0.2769
-6.000 -0.5872 0.03967 0.03072 -0.0181 1.0000 0.2805
-5.750 -0.5680 0.03722 0.02810 -0.0179 1.0000 0.2855
-5.500 -0.5471 0.03512 0.02582 -0.0176 1.0000 0.2905
-5.250 -0.5266 0.03276 0.02310 -0.0177 1.0000 0.2956
-5.000 -0.5048 0.03091 0.02108 -0.0173 1.0000 0.3025
-4.750 -0.4823 0.02946 0.01950 -0.0167 1.0000 0.3118
-4.500 -0.4595 0.02799 0.01793 -0.0161 1.0000 0.3208
-4.250 -0.4362 0.02658 0.01629 -0.0157 1.0000 0.3311
-4.000 -0.4131 0.02547 0.01526 -0.0148 1.0000 0.3422
-3.750 -0.3903 0.02446 0.01425 -0.0139 1.0000 0.3564
-3.500 -0.3675 0.02353 0.01336 -0.0130 1.0000 0.3721
-3.250 -0.3450 0.02271 0.01259 -0.0120 1.0000 0.3899
-3.000 -0.3229 0.02196 0.01191 -0.0110 1.0000 0.4108
-2.750 -0.3013 0.02128 0.01137 -0.0098 1.0000 0.4351
-2.500 -0.2805 0.02066 0.01091 -0.0085 1.0000 0.4639
-2.250 -0.2611 0.02007 0.01053 -0.0068 1.0000 0.4987
-2.000 -0.2435 0.01953 0.01033 -0.0047 1.0000 0.5426
-1.750 -0.2284 0.01901 0.01021 -0.0018 1.0000 0.6057
-1.500 -0.2187 0.01864 0.01026 0.0026 1.0000 0.6834
-1.250 -0.2130 0.01837 0.01034 0.0081 1.0000 0.7605
-1.000 -0.2058 0.01820 0.01042 0.0133 1.0000 0.8386
-0.750 -0.1419 0.01840 0.01069 0.0089 1.0000 0.9401
-0.500 -0.0396 0.01843 0.01049 -0.0061 1.0000 1.0000
-0.250 -0.0714 0.01825 0.01022 -0.0010 1.0000 1.0000
0.000 -0.0689 0.01849 0.01028 0.0001 1.0000 1.0000
0.250 -0.0142 0.01926 0.01091 -0.0072 0.9832 1.0000
0.500 0.0780 0.01981 0.01142 -0.0197 0.9467 1.0000
0.750 0.1821 0.01973 0.01139 -0.0327 0.9123 1.0000
1.000 0.2548 0.01942 0.01112 -0.0391 0.8733 1.0000
1.250 0.3034 0.01919 0.01085 -0.0408 0.8315 1.0000
1.500 0.3294 0.01923 0.01077 -0.0386 0.7873 1.0000
1.750 0.3515 0.01931 0.01067 -0.0354 0.7431 1.0000
2.000 0.3703 0.01953 0.01069 -0.0319 0.6954 1.0000
2.250 0.3899 0.01989 0.01079 -0.0288 0.6459 1.0000
2.500 0.4104 0.02036 0.01100 -0.0263 0.5992 1.0000
2.750 0.4325 0.02089 0.01124 -0.0242 0.5602 1.0000
3.000 0.4556 0.02155 0.01171 -0.0228 0.5273 1.0000
3.250 0.4791 0.02228 0.01230 -0.0216 0.4995 1.0000
3.500 0.5033 0.02298 0.01283 -0.0205 0.4764 1.0000
3.750 0.5280 0.02374 0.01344 -0.0195 0.4569 1.0000
4.000 0.5521 0.02461 0.01429 -0.0187 0.4379 1.0000
4.250 0.5763 0.02549 0.01513 -0.0179 0.4207 1.0000
4.500 0.6009 0.02641 0.01601 -0.0172 0.4055 1.0000
4.750 0.6253 0.02737 0.01694 -0.0164 0.3909 1.0000
5.000 0.6486 0.02855 0.01823 -0.0158 0.3769 1.0000
5.250 0.6716 0.02982 0.01963 -0.0152 0.3637 1.0000
5.500 0.6946 0.03122 0.02111 -0.0146 0.3524 1.0000
5.750 0.7196 0.03236 0.02215 -0.0138 0.3413 1.0000
6.000 0.7390 0.03420 0.02431 -0.0133 0.3295 1.0000
6.250 0.7594 0.03614 0.02641 -0.0127 0.3203 1.0000
6.500 0.7819 0.03769 0.02801 -0.0120 0.3104 1.0000
6.750 0.7957 0.04049 0.03118 -0.0117 0.3013 1.0000
7.000 0.8200 0.04200 0.03264 -0.0108 0.2931 1.0000
7.250 0.8243 0.04614 0.03729 -0.0109 0.2860 1.0000
7.500 0.8458 0.04786 0.03902 -0.0100 0.2775 1.0000
7.750 0.8450 0.05280 0.04435 -0.0104 0.2730 1.0000
8.000 0.8280 0.05965 0.05155 -0.0122 0.2706 1.0000
8.250 0.7995 0.06792 0.06000 -0.0155 0.2714 1.0000
8.500 0.7756 0.07582 0.06794 -0.0192 0.2738 1.0000
8.750 0.7693 0.08193 0.07405 -0.0214 0.2756 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 426 AIRFOIL (e426-il)