EPPLER 266 AIRFOIL (e266-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 266 AIRFOIL (e266-il) Reynolds number: 200,000 Max Cl/Cd: 61.92 at α=10° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e266-il-200000.txt Download as CSV file: xf-e266-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.2029 0.09890 0.09488 -0.0951 0.8256 0.0585
-11.500 -0.2014 0.09515 0.09103 -0.0971 0.8153 0.0609
-11.250 -0.2063 0.09060 0.08644 -0.1000 0.8068 0.0626
-11.000 -0.1655 0.07925 0.07521 -0.0953 0.7785 0.0677
-10.750 -0.1601 0.07623 0.07217 -0.0956 0.7745 0.0689
-10.500 -0.1602 0.07249 0.06838 -0.0966 0.7712 0.0710
-10.250 -0.1737 0.06650 0.06240 -0.0999 0.7686 0.0727
-10.000 -0.2055 0.05861 0.05450 -0.1051 0.7664 0.0732
-9.750 -0.2364 0.05313 0.04895 -0.1068 0.7634 0.0732
-9.500 -0.2652 0.04904 0.04477 -0.1063 0.7603 0.0737
-9.250 -0.2916 0.04601 0.04160 -0.1043 0.7572 0.0746
-9.000 -0.3239 0.04409 0.03948 -0.0996 0.7543 0.0761
-7.750 -0.3818 0.03516 0.02819 -0.0850 0.7448 0.0419
-7.500 -0.3673 0.03232 0.02500 -0.0832 0.7421 0.0401
-7.250 -0.3516 0.02973 0.02161 -0.0803 0.7398 0.0365
-7.000 -0.3266 0.02833 0.01997 -0.0798 0.7374 0.0358
-6.750 -0.2995 0.02629 0.01778 -0.0798 0.7346 0.0354
-6.500 -0.2713 0.02480 0.01614 -0.0799 0.7317 0.0354
-6.250 -0.2422 0.02361 0.01482 -0.0801 0.7292 0.0357
-6.000 -0.2120 0.02246 0.01360 -0.0805 0.7269 0.0362
-5.750 -0.1833 0.02118 0.01236 -0.0809 0.7249 0.0375
-5.500 -0.1605 0.02056 0.01175 -0.0803 0.7230 0.0399
-5.250 -0.1387 0.02016 0.01129 -0.0795 0.7212 0.0430
-5.000 -0.1218 0.01952 0.01069 -0.0778 0.7189 0.0455
-4.750 -0.1058 0.01902 0.01022 -0.0761 0.7163 0.0492
-4.500 -0.0894 0.01855 0.00974 -0.0743 0.7138 0.0562
-4.250 -0.0796 0.01762 0.00916 -0.0714 0.7116 0.1071
-4.000 -0.0884 0.01587 0.00868 -0.0662 0.7097 0.3566
-3.750 -0.0891 0.01506 0.00938 -0.0600 0.7080 0.6498
-3.500 -0.0632 0.01605 0.01039 -0.0582 0.7065 0.7309
-3.250 -0.0393 0.01729 0.01155 -0.0560 0.7047 0.7661
-3.000 -0.0092 0.01897 0.01320 -0.0540 0.7024 0.7886
-2.750 0.0251 0.02026 0.01444 -0.0532 0.7001 0.8041
-2.500 0.0541 0.02108 0.01517 -0.0524 0.6982 0.8179
-2.250 0.1159 0.02213 0.01609 -0.0568 0.6966 0.8268
-2.000 0.1576 0.02255 0.01639 -0.0587 0.6948 0.8369
-1.750 0.1765 0.02278 0.01655 -0.0570 0.6928 0.8489
-1.500 0.2354 0.02284 0.01646 -0.0628 0.6912 0.8517
-1.250 0.2779 0.02292 0.01643 -0.0658 0.6898 0.8563
-1.000 0.2886 0.02310 0.01661 -0.0630 0.6880 0.8654
-0.750 0.3279 0.02320 0.01669 -0.0657 0.6857 0.8690
-0.500 0.3232 0.02344 0.01697 -0.0601 0.6829 0.8791
-0.250 0.3679 0.02340 0.01688 -0.0638 0.6808 0.8814
0.000 0.4033 0.02342 0.01686 -0.0658 0.6789 0.8849
0.250 0.4170 0.02352 0.01693 -0.0636 0.6771 0.8911
0.500 0.4387 0.02350 0.01687 -0.0629 0.6753 0.8955
0.750 0.4770 0.02346 0.01678 -0.0654 0.6737 0.8979
1.000 0.4947 0.02378 0.01714 -0.0643 0.6712 0.9021
1.250 0.4764 0.02429 0.01775 -0.0565 0.6675 0.9091
1.500 0.5085 0.02440 0.01787 -0.0581 0.6646 0.9113
1.750 0.5352 0.02447 0.01795 -0.0586 0.6624 0.9139
2.000 0.5549 0.02449 0.01796 -0.0575 0.6604 0.9171
2.250 0.5458 0.02452 0.01796 -0.0507 0.6585 0.9222
2.500 0.5859 0.02454 0.01794 -0.0536 0.6570 0.9233
2.750 0.5615 0.02574 0.01935 -0.0458 0.6499 0.9282
3.000 0.5209 0.02589 0.01952 -0.0335 0.6458 0.9350
3.250 0.5589 0.02578 0.01940 -0.0357 0.6441 0.9360
3.500 0.5997 0.02557 0.01917 -0.0384 0.6427 0.9368
3.750 0.6372 0.02543 0.01900 -0.0404 0.6412 0.9376
4.000 0.5342 0.02720 0.02095 -0.0192 0.6292 0.9482
4.250 0.5738 0.02690 0.02065 -0.0214 0.6279 0.9487
4.500 0.6178 0.02650 0.02025 -0.0243 0.6268 0.9490
4.750 0.6633 0.02615 0.01989 -0.0274 0.6257 0.9494
5.000 0.5904 0.02827 0.02214 -0.0128 0.6125 0.9586
5.250 0.6359 0.02769 0.02158 -0.0156 0.6116 0.9588
5.500 0.6838 0.02709 0.02100 -0.0188 0.6108 0.9588
5.750 0.7338 0.02647 0.02038 -0.0224 0.6098 0.9588
6.000 0.6314 0.03161 0.02572 -0.0076 0.5868 0.9710
6.250 0.6935 0.02991 0.02405 -0.0115 0.5886 0.9699
6.500 0.7624 0.02835 0.02253 -0.0170 0.5900 0.9681
6.750 0.8408 0.02655 0.02074 -0.0241 0.5921 0.9658
7.000 0.8226 0.02780 0.02211 -0.0183 0.5819 0.9716
7.250 0.8846 0.02640 0.02077 -0.0231 0.5810 0.9706
7.500 0.9468 0.02493 0.01933 -0.0281 0.5795 0.9696
7.750 0.9428 0.02598 0.02054 -0.0245 0.5698 0.9744
8.000 0.9960 0.02471 0.01932 -0.0281 0.5670 0.9743
8.250 1.0555 0.02313 0.01778 -0.0326 0.5643 0.9738
8.500 1.0739 0.02324 0.01805 -0.0319 0.5555 0.9768
8.750 1.1243 0.02188 0.01675 -0.0352 0.5502 0.9771
9.000 1.1465 0.02165 0.01667 -0.0347 0.5418 0.9801
9.250 1.1904 0.02034 0.01539 -0.0369 0.5336 0.9811
9.500 1.2024 0.02014 0.01534 -0.0348 0.5208 0.9860
9.750 1.2182 0.01973 0.01503 -0.0333 0.5051 0.9918
10.000 1.2192 0.01969 0.01504 -0.0294 0.4882 1.0000
10.250 1.1974 0.02008 0.01544 -0.0219 0.4761 1.0000
10.500 1.1928 0.02064 0.01598 -0.0179 0.4547 1.0000
10.750 1.1970 0.02124 0.01647 -0.0153 0.4286 1.0000
11.000 1.1940 0.02250 0.01755 -0.0124 0.3959 1.0000
11.250 1.1876 0.02427 0.01913 -0.0097 0.3636 1.0000
11.500 1.1785 0.02647 0.02114 -0.0072 0.3327 1.0000
11.750 1.1700 0.02882 0.02335 -0.0051 0.3050 1.0000
12.000 1.1611 0.03134 0.02571 -0.0032 0.2783 1.0000
12.250 1.1518 0.03403 0.02825 -0.0016 0.2511 1.0000
12.500 1.1457 0.03660 0.03070 -0.0003 0.2262 1.0000
12.750 1.1401 0.03925 0.03324 0.0008 0.2033 1.0000
13.000 1.1367 0.04183 0.03572 0.0017 0.1812 1.0000
13.250 1.1329 0.04454 0.03832 0.0024 0.1603 1.0000
13.500 1.1307 0.04718 0.04088 0.0030 0.1396 1.0000
13.750 1.1275 0.05001 0.04360 0.0035 0.1204 1.0000
14.000 1.1219 0.05315 0.04660 0.0040 0.0963 1.0000
14.250 1.1145 0.05656 0.04984 0.0044 0.0728 1.0000
14.500 1.1029 0.06050 0.05361 0.0049 0.0525 1.0000
14.750 1.0948 0.06425 0.05728 0.0052 0.0427 1.0000
15.000 1.0914 0.06759 0.06065 0.0053 0.0369 1.0000
15.250 1.0868 0.07109 0.06415 0.0052 0.0343 1.0000
15.500 1.0879 0.07403 0.06717 0.0051 0.0320 1.0000
15.750 1.0900 0.07686 0.07007 0.0048 0.0304 1.0000
16.000 1.0903 0.07994 0.07319 0.0045 0.0285 1.0000
16.250 1.0927 0.08265 0.07592 0.0043 0.0279 1.0000
16.500 1.0973 0.08497 0.07825 0.0045 0.0269 1.0000
16.750 1.1042 0.08723 0.08062 0.0044 0.0259 1.0000
17.000 1.1122 0.08927 0.08276 0.0044 0.0252 1.0000
17.250 1.1206 0.09128 0.08485 0.0044 0.0244 1.0000
17.500 1.1299 0.09313 0.08676 0.0044 0.0238 1.0000
17.750 1.1392 0.09498 0.08865 0.0043 0.0232 1.0000
18.000 1.1526 0.09606 0.08967 0.0048 0.0221 1.0000
18.250 1.1654 0.09751 0.09122 0.0053 0.0218 1.0000
18.500 1.1666 0.10079 0.09471 0.0044 0.0215 1.0000
18.750 1.1660 0.10437 0.09849 0.0034 0.0211 1.0000
19.000 1.1668 0.10775 0.10208 0.0026 0.0210 1.0000
19.250 1.1652 0.11156 0.10608 0.0014 0.0209 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 266 AIRFOIL (e266-il)