EPPLER 266 AIRFOIL (e266-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 266 AIRFOIL (e266-il) Reynolds number: 50,000 Max Cl/Cd: 13.4 at α=14° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e266-il-50000-n5.txt Download as CSV file: xf-e266-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 266 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.3244 0.12375 0.11711 -0.0618 1.0000 0.0505
-12.250 -0.3230 0.12044 0.11389 -0.0622 1.0000 0.0500
-12.000 -0.3249 0.11716 0.11072 -0.0627 1.0000 0.0496
-11.750 -0.3313 0.11390 0.10763 -0.0630 1.0000 0.0493
-11.500 -0.3265 0.10867 0.10245 -0.0671 0.9829 0.0488
-11.250 -0.3216 0.10238 0.09618 -0.0728 0.9637 0.0485
-11.000 -0.3185 0.09564 0.08943 -0.0790 0.9461 0.0479
-10.750 -0.3203 0.08794 0.08168 -0.0862 0.9310 0.0469
-10.500 -0.3311 0.08027 0.07389 -0.0935 0.9172 0.0459
-10.250 -0.3486 0.07348 0.06691 -0.0990 0.9039 0.0450
-10.000 -0.3654 0.06825 0.06139 -0.1019 0.8909 0.0443
-9.750 -0.3800 0.06413 0.05696 -0.1026 0.8791 0.0440
-9.500 -0.3927 0.06100 0.05351 -0.1013 0.8678 0.0438
-9.250 -0.4014 0.05844 0.05066 -0.0992 0.8574 0.0439
-9.000 -0.4041 0.05568 0.04756 -0.0975 0.8500 0.0441
-8.750 -0.4070 0.05341 0.04497 -0.0950 0.8415 0.0442
-8.500 -0.4037 0.05093 0.04208 -0.0930 0.8353 0.0443
-8.250 -0.3988 0.04869 0.03947 -0.0908 0.8289 0.0444
-8.000 -0.3900 0.04657 0.03700 -0.0889 0.8228 0.0446
-7.750 -0.3743 0.04439 0.03442 -0.0878 0.8182 0.0451
-7.500 -0.3587 0.04258 0.03228 -0.0864 0.8128 0.0458
-7.250 -0.3388 0.04088 0.03027 -0.0854 0.8077 0.0466
-7.000 -0.3122 0.03915 0.02816 -0.0853 0.8038 0.0489
-6.750 -0.2784 0.03748 0.02619 -0.0862 0.8009 0.0519
-6.500 -0.2503 0.03639 0.02506 -0.0865 0.7968 0.0553
-6.250 -0.2109 0.03523 0.02369 -0.0878 0.7933 0.0592
-6.000 -0.1765 0.03419 0.02252 -0.0886 0.7900 0.0642
-5.750 -0.1512 0.03340 0.02160 -0.0882 0.7866 0.0729
-5.500 -0.1315 0.03258 0.02078 -0.0872 0.7835 0.0850
-5.250 -0.1246 0.03212 0.02046 -0.0844 0.7786 0.1023
-5.000 -0.1198 0.03128 0.01999 -0.0814 0.7746 0.1438
-4.750 -0.0761 0.03071 0.02243 -0.0791 0.7732 0.6161
-4.500 -0.0736 0.03192 0.02345 -0.0733 0.7699 0.6902
-4.000 -0.0448 0.03493 0.02597 -0.0653 0.7625 0.7764
-3.750 0.0025 0.03607 0.02673 -0.0668 0.7598 0.8081
-3.500 0.0483 0.03653 0.02682 -0.0693 0.7573 0.8297
-3.250 0.0842 0.03669 0.02670 -0.0708 0.7550 0.8457
-3.000 0.1195 0.03667 0.02641 -0.0725 0.7529 0.8583
-2.750 0.1453 0.03676 0.02627 -0.0728 0.7502 0.8698
-2.500 0.1532 0.03729 0.02672 -0.0704 0.7457 0.8820
-2.250 0.1850 0.03735 0.02660 -0.0720 0.7425 0.8902
-2.000 0.2045 0.03752 0.02661 -0.0715 0.7396 0.8995
-1.750 0.2354 0.03749 0.02641 -0.0729 0.7373 0.9068
-1.500 0.2606 0.03748 0.02624 -0.0732 0.7351 0.9141
-1.250 0.2636 0.03831 0.02705 -0.0706 0.7301 0.9227
-1.000 0.2794 0.03878 0.02744 -0.0699 0.7263 0.9294
-0.750 0.2938 0.03914 0.02771 -0.0686 0.7231 0.9364
-0.500 0.3260 0.03916 0.02761 -0.0704 0.7207 0.9407
-0.250 0.3545 0.03920 0.02753 -0.0713 0.7186 0.9455
0.000 0.3471 0.04058 0.02896 -0.0677 0.7126 0.9524
0.250 0.3588 0.04124 0.02958 -0.0664 0.7085 0.9579
0.500 0.3814 0.04156 0.02984 -0.0666 0.7054 0.9620
0.750 0.4116 0.04173 0.02993 -0.0680 0.7031 0.9653
1.000 0.4028 0.04318 0.03141 -0.0641 0.6971 0.9716
1.250 0.4117 0.04408 0.03231 -0.0628 0.6922 0.9759
1.500 0.4353 0.04451 0.03271 -0.0633 0.6890 0.9792
1.750 0.4629 0.04479 0.03295 -0.0641 0.6866 0.9822
2.250 0.4570 0.04738 0.03559 -0.0584 0.6741 0.9919
2.500 0.4832 0.04777 0.03597 -0.0591 0.6712 0.9945
2.750 0.5144 0.04803 0.03622 -0.0604 0.6689 0.9968
3.000 0.4790 0.05014 0.03838 -0.0533 0.6590 1.0000
3.250 0.4933 0.05055 0.03879 -0.0518 0.6553 1.0000
3.500 0.5150 0.05081 0.03904 -0.0511 0.6526 1.0000
4.000 0.4916 0.05291 0.04117 -0.0416 0.6388 1.0000
4.250 0.4804 0.05394 0.04222 -0.0370 0.6319 1.0000
4.500 0.4789 0.05467 0.04296 -0.0336 0.6255 1.0000
4.750 0.4974 0.05492 0.04323 -0.0324 0.6219 1.0000
5.000 0.4780 0.05605 0.04438 -0.0270 0.6132 1.0000
5.250 0.4876 0.05644 0.04479 -0.0248 0.6078 1.0000
5.500 0.5096 0.05659 0.04498 -0.0239 0.6046 1.0000
5.750 0.4843 0.05778 0.04619 -0.0181 0.5941 1.0000
6.000 0.5026 0.05799 0.04642 -0.0168 0.5900 1.0000
6.250 0.4902 0.05900 0.04745 -0.0126 0.5806 1.0000
6.500 0.5065 0.05934 0.04783 -0.0114 0.5753 1.0000
6.750 0.5079 0.06028 0.04880 -0.0091 0.5671 1.0000
7.000 0.5234 0.06084 0.04942 -0.0081 0.5606 1.0000
7.500 0.5486 0.06241 0.05111 -0.0061 0.5456 1.0000
8.000 0.5785 0.06399 0.05286 -0.0049 0.5303 1.0000
8.250 0.5872 0.06516 0.05410 -0.0041 0.5209 1.0000
8.500 0.6116 0.06549 0.05454 -0.0040 0.5147 1.0000
8.750 0.6178 0.06684 0.05598 -0.0032 0.5041 1.0000
9.000 0.6482 0.06676 0.05604 -0.0034 0.4989 1.0000
9.250 0.6527 0.06828 0.05765 -0.0026 0.4872 1.0000
9.500 0.6641 0.06940 0.05888 -0.0021 0.4772 1.0000
9.750 0.6915 0.06933 0.05898 -0.0021 0.4705 1.0000
10.000 0.6982 0.07079 0.06055 -0.0015 0.4586 1.0000
10.250 0.7118 0.07177 0.06165 -0.0011 0.4484 1.0000
10.500 0.7390 0.07152 0.06157 -0.0009 0.4411 1.0000
10.750 0.7469 0.07295 0.06313 -0.0004 0.4286 1.0000
11.000 0.7596 0.07397 0.06431 0.0000 0.4174 1.0000
11.250 0.7913 0.07299 0.06352 0.0004 0.4105 1.0000
11.500 0.7998 0.07435 0.06503 0.0008 0.3974 1.0000
11.750 0.8111 0.07547 0.06630 0.0013 0.3850 1.0000
12.000 0.8499 0.07315 0.06423 0.0021 0.3794 1.0000
12.250 0.8600 0.07421 0.06545 0.0026 0.3657 1.0000
12.750 0.8881 0.07542 0.06700 0.0038 0.3390 1.0000
13.000 0.9078 0.07525 0.06702 0.0046 0.3257 1.0000
13.250 0.9295 0.07477 0.06671 0.0054 0.3118 1.0000
13.500 0.9503 0.07442 0.06650 0.0062 0.2962 1.0000
13.750 0.9693 0.07430 0.06650 0.0070 0.2786 1.0000
14.000 0.9899 0.07389 0.06611 0.0080 0.2594 1.0000
14.250 0.9979 0.07537 0.06760 0.0085 0.2399 1.0000
14.500 1.0018 0.07749 0.06974 0.0087 0.2204 1.0000
14.750 1.0063 0.07951 0.07170 0.0090 0.2011 1.0000
15.000 1.0088 0.08188 0.07397 0.0091 0.1830 1.0000
15.250 1.0055 0.08527 0.07738 0.0087 0.1654 1.0000
15.500 1.0016 0.08881 0.08090 0.0081 0.1486 1.0000
15.750 0.9958 0.09272 0.08475 0.0073 0.1321 1.0000
16.000 0.9899 0.09678 0.08875 0.0062 0.1173 1.0000
16.250 0.9844 0.10092 0.09286 0.0050 0.1041 1.0000
16.500 0.9794 0.10514 0.09705 0.0037 0.0922 1.0000
16.750 0.9761 0.10919 0.10111 0.0024 0.0823 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 266 AIRFOIL (e266-il)