EPPLER 266 AIRFOIL (e266-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 266 AIRFOIL (e266-il) Reynolds number: 1,000,000 Max Cl/Cd: 124.91 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e266-il-1000000-n5.txt Download as CSV file: xf-e266-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.000 -0.5510 0.06527 0.06171 -0.1052 0.6781 0.0060
-14.750 -0.5767 0.05733 0.05362 -0.1102 0.6750 0.0058
-14.500 -0.6048 0.05053 0.04662 -0.1135 0.6716 0.0058
-14.250 -0.6286 0.04547 0.04137 -0.1146 0.6683 0.0058
-14.000 -0.6422 0.04201 0.03775 -0.1145 0.6652 0.0059
-13.750 -0.6574 0.03876 0.03433 -0.1135 0.6617 0.0058
-13.500 -0.6614 0.03665 0.03207 -0.1124 0.6580 0.0059
-13.250 -0.6747 0.03387 0.02907 -0.1102 0.6543 0.0059
-13.000 -0.6743 0.03227 0.02735 -0.1086 0.6513 0.0060
-12.750 -0.6723 0.03080 0.02574 -0.1068 0.6481 0.0061
-12.500 -0.6742 0.02899 0.02377 -0.1045 0.6451 0.0061
-12.250 -0.6672 0.02794 0.02260 -0.1028 0.6422 0.0062
-12.000 -0.6628 0.02663 0.02115 -0.1007 0.6393 0.0062
-11.500 -0.6473 0.02442 0.01868 -0.0969 0.6348 0.0063
-11.250 -0.6366 0.02350 0.01765 -0.0951 0.6327 0.0063
-11.000 -0.6235 0.02282 0.01689 -0.0936 0.6305 0.0064
-10.750 -0.6110 0.02200 0.01596 -0.0919 0.6284 0.0064
-10.500 -0.5979 0.02117 0.01503 -0.0902 0.6263 0.0064
-10.250 -0.5864 0.02001 0.01375 -0.0883 0.6243 0.0065
-10.000 -0.5731 0.01916 0.01280 -0.0866 0.6222 0.0067
-9.750 -0.5573 0.01858 0.01217 -0.0852 0.6203 0.0068
-9.500 -0.5415 0.01788 0.01139 -0.0836 0.6190 0.0068
-9.250 -0.5250 0.01732 0.01079 -0.0822 0.6175 0.0069
-9.000 -0.5076 0.01688 0.01031 -0.0808 0.6159 0.0071
-8.750 -0.4901 0.01647 0.00985 -0.0794 0.6141 0.0072
-8.500 -0.4737 0.01597 0.00929 -0.0777 0.6124 0.0072
-8.250 -0.4563 0.01557 0.00884 -0.0761 0.6106 0.0073
-8.000 -0.4373 0.01524 0.00847 -0.0749 0.6089 0.0075
-7.750 -0.4181 0.01480 0.00798 -0.0736 0.6074 0.0076
-7.500 -0.3979 0.01443 0.00755 -0.0726 0.6059 0.0078
-7.000 -0.3561 0.01371 0.00675 -0.0706 0.6034 0.0081
-6.750 -0.3342 0.01338 0.00639 -0.0698 0.6024 0.0083
-6.500 -0.3117 0.01309 0.00607 -0.0691 0.6013 0.0085
-6.250 -0.2889 0.01281 0.00577 -0.0684 0.6001 0.0087
-6.000 -0.2657 0.01256 0.00549 -0.0679 0.5989 0.0088
-5.750 -0.2442 0.01217 0.00507 -0.0669 0.5977 0.0091
-5.500 -0.2217 0.01186 0.00473 -0.0662 0.5963 0.0094
-5.250 -0.1982 0.01160 0.00446 -0.0656 0.5949 0.0097
-5.000 -0.1741 0.01138 0.00422 -0.0651 0.5936 0.0101
-4.750 -0.1498 0.01119 0.00400 -0.0647 0.5922 0.0105
-4.500 -0.1252 0.01100 0.00379 -0.0643 0.5910 0.0109
-4.250 -0.1003 0.01085 0.00361 -0.0640 0.5898 0.0113
-4.000 -0.0751 0.01070 0.00343 -0.0637 0.5884 0.0116
-3.750 -0.0498 0.01052 0.00325 -0.0634 0.5877 0.0125
-3.500 -0.0242 0.01036 0.00309 -0.0632 0.5868 0.0138
-3.250 0.0019 0.01023 0.00296 -0.0631 0.5858 0.0151
-2.750 0.0503 0.00969 0.00264 -0.0622 0.5835 0.0648
-2.500 0.0742 0.00941 0.00251 -0.0618 0.5824 0.1074
-2.250 0.0975 0.00908 0.00238 -0.0613 0.5814 0.1641
-1.750 0.1369 0.00797 0.00204 -0.0592 0.5791 0.3882
-1.500 0.1530 0.00720 0.00180 -0.0574 0.5779 0.5423
-1.250 0.1694 0.00662 0.00173 -0.0554 0.5766 0.6886
-1.000 0.1958 0.00663 0.00182 -0.0552 0.5754 0.7285
-0.750 0.2235 0.00670 0.00191 -0.0552 0.5742 0.7479
-0.500 0.2516 0.00679 0.00196 -0.0554 0.5728 0.7580
-0.250 0.2801 0.00686 0.00203 -0.0556 0.5719 0.7646
0.000 0.3087 0.00691 0.00207 -0.0559 0.5710 0.7701
0.250 0.3373 0.00695 0.00210 -0.0562 0.5699 0.7736
0.500 0.3660 0.00698 0.00214 -0.0565 0.5688 0.7759
0.750 0.3947 0.00701 0.00217 -0.0569 0.5676 0.7776
1.000 0.4234 0.00704 0.00218 -0.0573 0.5663 0.7789
1.250 0.4520 0.00707 0.00220 -0.0576 0.5650 0.7802
1.500 0.4805 0.00710 0.00222 -0.0580 0.5638 0.7815
1.750 0.5090 0.00713 0.00224 -0.0583 0.5623 0.7826
2.000 0.5373 0.00716 0.00227 -0.0587 0.5609 0.7837
2.250 0.5656 0.00721 0.00229 -0.0590 0.5595 0.7848
2.500 0.5937 0.00727 0.00233 -0.0593 0.5577 0.7860
2.750 0.6220 0.00730 0.00237 -0.0597 0.5562 0.7869
3.000 0.6503 0.00731 0.00240 -0.0600 0.5541 0.7878
3.250 0.6784 0.00732 0.00244 -0.0603 0.5515 0.7888
3.500 0.7062 0.00733 0.00247 -0.0605 0.5486 0.7897
3.750 0.7337 0.00737 0.00251 -0.0607 0.5458 0.7907
4.000 0.7607 0.00742 0.00256 -0.0608 0.5426 0.7916
4.250 0.7886 0.00745 0.00263 -0.0611 0.5397 0.7927
4.500 0.8163 0.00748 0.00269 -0.0613 0.5365 0.7938
4.750 0.8436 0.00752 0.00275 -0.0615 0.5328 0.7949
5.000 0.8699 0.00758 0.00281 -0.0614 0.5284 0.7960
5.250 0.8969 0.00762 0.00288 -0.0615 0.5231 0.7970
5.500 0.9233 0.00767 0.00295 -0.0615 0.5169 0.7981
5.750 0.9490 0.00775 0.00303 -0.0614 0.5106 0.7991
6.000 0.9747 0.00783 0.00312 -0.0612 0.5020 0.8002
6.250 0.9968 0.00798 0.00322 -0.0604 0.4844 0.8013
6.500 1.0167 0.00822 0.00339 -0.0592 0.4616 0.8023
6.750 1.0290 0.00870 0.00369 -0.0565 0.4251 0.8035
7.000 1.0392 0.00921 0.00406 -0.0535 0.3898 0.8051
7.250 1.0404 0.00982 0.00450 -0.0487 0.3514 0.8069
7.500 1.0446 0.01037 0.00494 -0.0445 0.3231 0.8085
7.750 1.0386 0.01124 0.00562 -0.0386 0.2824 0.8104
8.000 1.0391 0.01198 0.00623 -0.0342 0.2544 0.8122
8.250 1.0363 0.01286 0.00697 -0.0294 0.2247 0.8141
8.500 1.0321 0.01388 0.00786 -0.0247 0.1968 0.8160
8.750 1.0338 0.01481 0.00872 -0.0213 0.1775 0.8177
9.000 1.0324 0.01600 0.00982 -0.0178 0.1573 0.8194
9.250 1.0337 0.01721 0.01098 -0.0149 0.1408 0.8213
9.500 1.0367 0.01844 0.01219 -0.0125 0.1263 0.8232
9.750 1.0387 0.01984 0.01354 -0.0102 0.1103 0.8250
10.000 1.0382 0.02147 0.01510 -0.0079 0.0919 0.8268
10.250 1.0405 0.02299 0.01656 -0.0059 0.0779 0.8285
10.500 1.0437 0.02450 0.01803 -0.0042 0.0652 0.8302
10.750 1.0521 0.02573 0.01925 -0.0030 0.0583 0.8318
11.000 1.0527 0.02747 0.02092 -0.0012 0.0440 0.8334
11.250 1.0547 0.02917 0.02256 0.0004 0.0322 0.8349
11.500 1.0573 0.03085 0.02420 0.0018 0.0214 0.8363
11.750 1.0585 0.03268 0.02599 0.0033 0.0114 0.8381
12.000 1.0673 0.03402 0.02736 0.0041 0.0097 0.8398
12.250 1.0763 0.03537 0.02874 0.0049 0.0084 0.8416
12.500 1.0861 0.03669 0.03011 0.0055 0.0078 0.8434
12.750 1.0964 0.03802 0.03148 0.0061 0.0075 0.8451
13.000 1.1069 0.03933 0.03283 0.0066 0.0071 0.8468
13.250 1.1163 0.04079 0.03433 0.0071 0.0068 0.8484
13.500 1.1255 0.04226 0.03584 0.0076 0.0066 0.8498
14.000 1.1438 0.04531 0.03900 0.0084 0.0063 0.8530
14.250 1.1509 0.04705 0.04079 0.0088 0.0060 0.8547
14.500 1.1599 0.04865 0.04245 0.0091 0.0059 0.8566
14.750 1.1686 0.05031 0.04418 0.0093 0.0057 0.8585
15.000 1.1777 0.05193 0.04585 0.0095 0.0057 0.8604
15.250 1.1848 0.05379 0.04777 0.0097 0.0054 0.8623
15.500 1.1931 0.05557 0.04961 0.0098 0.0054 0.8643
16.000 1.2060 0.05957 0.05375 0.0100 0.0052 0.8688
16.250 1.2114 0.06169 0.05594 0.0100 0.0051 0.8713
16.500 1.2184 0.06370 0.05801 0.0099 0.0049 0.8739
16.750 1.2223 0.06608 0.06047 0.0098 0.0048 0.8764
17.000 1.2272 0.06838 0.06284 0.0096 0.0048 0.8791
17.500 1.2330 0.07354 0.06817 0.0091 0.0047 0.8849
17.750 1.2331 0.07651 0.07122 0.0088 0.0045 0.8884
18.000 1.2356 0.07925 0.07405 0.0083 0.0046 0.8923
18.250 1.2267 0.08351 0.07841 0.0077 0.0043 0.8958
18.500 1.2275 0.08658 0.08159 0.0070 0.0042 0.9003
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 266 AIRFOIL (e266-il)