EPPLER 654 AIRFOIL (e654-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 654 AIRFOIL (e654-il) Reynolds number: 500,000 Max Cl/Cd: 118 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e654-il-500000-n5.txt Download as CSV file: xf-e654-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 654 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.3461 0.06921 0.06500 -0.1251 0.7519 0.0058
-13.750 -0.3832 0.05870 0.05427 -0.1317 0.7488 0.0058
-13.500 -0.4036 0.05170 0.04710 -0.1364 0.7449 0.0058
-13.250 -0.4165 0.04658 0.04182 -0.1397 0.7411 0.0058
-13.000 -0.4349 0.04152 0.03653 -0.1425 0.7375 0.0058
-12.750 -0.4384 0.03867 0.03357 -0.1436 0.7342 0.0057
-12.500 -0.4488 0.03539 0.03011 -0.1443 0.7308 0.0058
-12.250 -0.4503 0.03321 0.02780 -0.1443 0.7273 0.0058
-12.000 -0.4513 0.03122 0.02567 -0.1439 0.7241 0.0058
-11.750 -0.4515 0.02941 0.02370 -0.1431 0.7210 0.0058
-11.500 -0.4492 0.02783 0.02198 -0.1421 0.7181 0.0058
-11.250 -0.4464 0.02631 0.02032 -0.1408 0.7152 0.0059
-11.000 -0.4409 0.02507 0.01897 -0.1394 0.7122 0.0059
-10.750 -0.4344 0.02401 0.01781 -0.1378 0.7095 0.0059
-10.500 -0.4295 0.02305 0.01672 -0.1357 0.7069 0.0059
-10.250 -0.4228 0.02210 0.01565 -0.1338 0.7046 0.0060
-9.750 -0.3964 0.02028 0.01361 -0.1313 0.7005 0.0061
-9.500 -0.3792 0.01952 0.01278 -0.1305 0.6985 0.0062
-9.250 -0.3606 0.01880 0.01200 -0.1298 0.6962 0.0064
-9.000 -0.3403 0.01820 0.01132 -0.1292 0.6940 0.0065
-8.750 -0.3192 0.01762 0.01067 -0.1287 0.6918 0.0067
-8.500 -0.2972 0.01708 0.01005 -0.1282 0.6898 0.0069
-8.250 -0.2745 0.01658 0.00947 -0.1279 0.6880 0.0071
-8.000 -0.2512 0.01611 0.00891 -0.1276 0.6862 0.0074
-7.750 -0.2271 0.01567 0.00839 -0.1273 0.6846 0.0077
-7.500 -0.2025 0.01524 0.00791 -0.1272 0.6830 0.0080
-7.250 -0.1780 0.01477 0.00741 -0.1270 0.6811 0.0084
-7.000 -0.1526 0.01440 0.00700 -0.1270 0.6792 0.0090
-6.750 -0.1267 0.01405 0.00661 -0.1270 0.6772 0.0098
-6.500 -0.1006 0.01372 0.00624 -0.1270 0.6755 0.0107
-6.250 -0.0743 0.01341 0.00590 -0.1270 0.6738 0.0118
-6.000 -0.0476 0.01313 0.00558 -0.1271 0.6722 0.0135
-5.750 -0.0206 0.01287 0.00529 -0.1273 0.6707 0.0160
-5.500 0.0066 0.01262 0.00503 -0.1274 0.6690 0.0198
-5.250 0.0339 0.01238 0.00480 -0.1277 0.6674 0.0255
-5.000 0.0613 0.01212 0.00458 -0.1279 0.6657 0.0339
-4.750 0.0889 0.01187 0.00439 -0.1282 0.6641 0.0449
-4.500 0.1167 0.01164 0.00420 -0.1285 0.6625 0.0572
-4.250 0.1446 0.01139 0.00403 -0.1288 0.6609 0.0754
-4.000 0.1726 0.01113 0.00386 -0.1292 0.6592 0.1010
-3.750 0.2007 0.01085 0.00370 -0.1297 0.6575 0.1338
-3.500 0.2291 0.01060 0.00355 -0.1302 0.6557 0.1700
-3.250 0.2577 0.01037 0.00342 -0.1307 0.6541 0.2095
-3.000 0.2867 0.01013 0.00331 -0.1313 0.6527 0.2558
-2.750 0.3158 0.00986 0.00321 -0.1321 0.6513 0.3158
-2.500 0.3450 0.00954 0.00312 -0.1328 0.6498 0.3851
-2.250 0.3744 0.00924 0.00305 -0.1336 0.6480 0.4548
-2.000 0.4038 0.00900 0.00303 -0.1343 0.6462 0.5231
-1.750 0.4331 0.00889 0.00307 -0.1347 0.6444 0.5796
-1.500 0.4622 0.00887 0.00312 -0.1351 0.6427 0.6145
-1.250 0.4914 0.00890 0.00316 -0.1354 0.6410 0.6341
-1.000 0.5206 0.00894 0.00318 -0.1358 0.6392 0.6492
-0.750 0.5497 0.00899 0.00321 -0.1361 0.6375 0.6617
-0.500 0.5790 0.00905 0.00326 -0.1365 0.6359 0.6723
-0.250 0.6081 0.00912 0.00330 -0.1369 0.6341 0.6819
0.000 0.6366 0.00916 0.00337 -0.1371 0.6321 0.6894
0.250 0.6653 0.00921 0.00343 -0.1374 0.6299 0.6972
0.500 0.6940 0.00926 0.00347 -0.1377 0.6275 0.7038
0.750 0.7223 0.00930 0.00353 -0.1379 0.6251 0.7094
1.000 0.7510 0.00936 0.00356 -0.1382 0.6227 0.7153
1.250 0.7800 0.00942 0.00358 -0.1385 0.6203 0.7205
1.500 0.8081 0.00948 0.00365 -0.1387 0.6179 0.7246
1.750 0.8359 0.00952 0.00374 -0.1388 0.6150 0.7290
2.000 0.8641 0.00957 0.00380 -0.1391 0.6119 0.7333
2.250 0.8923 0.00962 0.00384 -0.1393 0.6088 0.7370
2.500 0.9202 0.00966 0.00388 -0.1395 0.6057 0.7400
2.750 0.9480 0.00972 0.00393 -0.1396 0.6027 0.7430
3.000 0.9753 0.00977 0.00404 -0.1397 0.5990 0.7460
3.250 1.0027 0.00983 0.00412 -0.1398 0.5950 0.7493
3.500 1.0301 0.00989 0.00417 -0.1400 0.5911 0.7524
3.750 1.0575 0.00996 0.00422 -0.1401 0.5874 0.7554
4.000 1.0841 0.01001 0.00435 -0.1400 0.5829 0.7580
4.250 1.1104 0.01008 0.00445 -0.1399 0.5779 0.7606
4.500 1.1364 0.01016 0.00453 -0.1398 0.5729 0.7633
4.750 1.1624 0.01024 0.00466 -0.1397 0.5674 0.7663
5.000 1.1879 0.01033 0.00477 -0.1395 0.5609 0.7695
5.250 1.2128 0.01045 0.00488 -0.1392 0.5542 0.7725
5.500 1.2369 0.01056 0.00502 -0.1387 0.5458 0.7750
5.750 1.2599 0.01069 0.00517 -0.1380 0.5371 0.7776
6.000 1.2815 0.01086 0.00535 -0.1370 0.5266 0.7808
6.250 1.3030 0.01105 0.00554 -0.1361 0.5157 0.7843
6.500 1.3222 0.01129 0.00576 -0.1347 0.5030 0.7878
6.750 1.3384 0.01158 0.00603 -0.1328 0.4887 0.7910
7.000 1.3494 0.01191 0.00633 -0.1299 0.4732 0.7942
7.250 1.3589 0.01234 0.00673 -0.1268 0.4568 0.7980
7.500 1.3660 0.01290 0.00724 -0.1234 0.4387 0.8025
8.000 1.3753 0.01433 0.00857 -0.1163 0.4033 0.8116
8.250 1.3790 0.01519 0.00940 -0.1130 0.3869 0.8165
8.500 1.3823 0.01617 0.01035 -0.1098 0.3710 0.8218
8.750 1.3852 0.01726 0.01142 -0.1068 0.3554 0.8268
9.000 1.3890 0.01840 0.01256 -0.1040 0.3412 0.8326
9.250 1.3917 0.01972 0.01385 -0.1014 0.3258 0.8393
9.500 1.3947 0.02106 0.01519 -0.0989 0.3116 0.8462
9.750 1.3977 0.02249 0.01661 -0.0965 0.2976 0.8542
10.000 1.4012 0.02393 0.01804 -0.0943 0.2852 0.8624
10.500 1.4078 0.02690 0.02103 -0.0901 0.2588 0.8851
11.000 1.4145 0.02979 0.02401 -0.0862 0.2349 1.0000
11.250 1.4191 0.03154 0.02571 -0.0848 0.2221 1.0000
11.500 1.4244 0.03327 0.02741 -0.0835 0.2104 1.0000
11.750 1.4315 0.03490 0.02902 -0.0824 0.1989 1.0000
12.000 1.4375 0.03666 0.03076 -0.0813 0.1881 1.0000
12.250 1.4435 0.03846 0.03252 -0.0803 0.1776 1.0000
12.500 1.4491 0.04033 0.03437 -0.0793 0.1669 1.0000
12.750 1.4565 0.04208 0.03612 -0.0785 0.1573 1.0000
13.000 1.4621 0.04402 0.03804 -0.0776 0.1481 1.0000
13.250 1.4682 0.04597 0.03997 -0.0768 0.1390 1.0000
13.500 1.4750 0.04788 0.04188 -0.0762 0.1301 1.0000
13.750 1.4803 0.04996 0.04396 -0.0755 0.1218 1.0000
14.000 1.4859 0.05206 0.04605 -0.0749 0.1133 1.0000
14.250 1.4922 0.05415 0.04815 -0.0744 0.1062 1.0000
14.500 1.4970 0.05641 0.05040 -0.0739 0.0989 1.0000
14.750 1.5032 0.05856 0.05258 -0.0735 0.0926 1.0000
15.000 1.5078 0.06092 0.05494 -0.0732 0.0866 1.0000
15.250 1.5132 0.06324 0.05729 -0.0729 0.0807 1.0000
15.500 1.5174 0.06574 0.05981 -0.0727 0.0755 1.0000
15.750 1.5223 0.06819 0.06229 -0.0725 0.0703 1.0000
16.000 1.5255 0.07088 0.06501 -0.0724 0.0656 1.0000
16.250 1.5294 0.07351 0.06767 -0.0724 0.0610 1.0000
16.500 1.5319 0.07639 0.07058 -0.0725 0.0570 1.0000
16.750 1.5350 0.07921 0.07345 -0.0726 0.0528 1.0000
17.000 1.5367 0.08228 0.07655 -0.0728 0.0492 1.0000
17.250 1.5392 0.08527 0.07960 -0.0731 0.0458 1.0000
17.500 1.5400 0.08851 0.08288 -0.0735 0.0429 1.0000
17.750 1.5418 0.09167 0.08610 -0.0740 0.0400 1.0000
18.000 1.5415 0.09516 0.08963 -0.0745 0.0374 1.0000
18.250 1.5430 0.09845 0.09301 -0.0752 0.0353 1.0000
18.500 1.5428 0.10198 0.09660 -0.0760 0.0329 1.0000
18.750 1.5420 0.10566 0.10034 -0.0769 0.0309 1.0000
19.000 1.5418 0.10926 0.10402 -0.0778 0.0291 1.0000
19.250 1.5400 0.11312 0.10794 -0.0789 0.0273 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 654 AIRFOIL (e654-il)