EPPLER 561 AIRFOIL (e561-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 561 AIRFOIL (e561-il) Reynolds number: 200,000 Max Cl/Cd: 69.96 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e561-il-200000.txt Download as CSV file: xf-e561-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 561 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
------ -------- --------- --------- -------- -------- --------
-13.000 -0.5578 0.07712 0.07291 -0.0768 1.0000 0.0349
-12.750 -0.5846 0.07088 0.06658 -0.0800 1.0000 0.0347
-12.500 -0.6111 0.06529 0.06089 -0.0825 1.0000 0.0346
-12.250 -0.6325 0.06092 0.05644 -0.0839 1.0000 0.0344
-12.000 -0.6575 0.05685 0.05228 -0.0847 1.0000 0.0343
-11.750 -0.6817 0.05363 0.04901 -0.0843 1.0000 0.0341
-11.500 -0.7083 0.05114 0.04648 -0.0826 1.0000 0.0340
-11.250 -0.7431 0.04942 0.04475 -0.0793 1.0000 0.0339
-11.000 -0.7263 0.04384 0.03877 -0.0912 0.9910 0.0340
-10.750 -0.6991 0.03988 0.03443 -0.0986 0.9837 0.0343
-10.500 -0.6720 0.03686 0.03108 -0.1033 0.9755 0.0347
-10.250 -0.6395 0.03431 0.02830 -0.1072 0.9700 0.0352
-9.750 -0.5758 0.03066 0.02465 -0.1120 0.9575 0.0370
-9.500 -0.5440 0.02906 0.02297 -0.1143 0.9500 0.0381
-9.250 -0.5093 0.02750 0.02128 -0.1170 0.9442 0.0396
-9.000 -0.4771 0.02593 0.01966 -0.1192 0.9366 0.0411
-8.750 -0.4369 0.02438 0.01815 -0.1231 0.9330 0.0441
-8.500 -0.4075 0.02314 0.01678 -0.1244 0.9232 0.0469
-8.250 -0.3674 0.02143 0.01510 -0.1282 0.9192 0.0520
-8.000 -0.3348 0.01988 0.01353 -0.1304 0.9108 0.0608
-7.750 -0.2947 0.01775 0.01160 -0.1347 0.9054 0.0956
-7.500 -0.2473 0.01675 0.01064 -0.1391 0.9020 0.1319
-7.250 -0.2113 0.01629 0.01016 -0.1410 0.8916 0.1518
-7.000 -0.1628 0.01583 0.00959 -0.1451 0.8861 0.1706
-6.750 -0.1253 0.01556 0.00924 -0.1470 0.8744 0.1860
-6.500 -0.0817 0.01542 0.00897 -0.1501 0.8646 0.2038
-6.250 -0.0422 0.01535 0.00882 -0.1523 0.8526 0.2194
-6.000 -0.0089 0.01527 0.00871 -0.1532 0.8383 0.2315
-5.750 0.0240 0.01515 0.00845 -0.1541 0.8243 0.2426
-5.500 0.0566 0.01512 0.00826 -0.1549 0.8109 0.2535
-5.000 0.1162 0.01498 0.00787 -0.1553 0.7839 0.2726
-4.750 0.1436 0.01491 0.00781 -0.1550 0.7704 0.2804
-4.500 0.1728 0.01488 0.00756 -0.1551 0.7580 0.2901
-4.250 0.2012 0.01484 0.00750 -0.1550 0.7463 0.2979
-4.000 0.2281 0.01482 0.00731 -0.1547 0.7337 0.3071
-3.750 0.2557 0.01478 0.00727 -0.1545 0.7225 0.3146
-3.500 0.2841 0.01481 0.00709 -0.1544 0.7114 0.3240
-3.250 0.3102 0.01475 0.00707 -0.1539 0.6999 0.3313
-3.000 0.3390 0.01480 0.00694 -0.1540 0.6903 0.3407
-2.750 0.3652 0.01474 0.00690 -0.1535 0.6795 0.3481
-2.500 0.3932 0.01480 0.00684 -0.1534 0.6700 0.3572
-2.250 0.4202 0.01476 0.00675 -0.1532 0.6601 0.3650
-2.000 0.4475 0.01481 0.00675 -0.1529 0.6508 0.3739
-1.750 0.4752 0.01481 0.00667 -0.1528 0.6421 0.3825
-1.500 0.5024 0.01486 0.00669 -0.1526 0.6335 0.3910
-1.250 0.5299 0.01488 0.00662 -0.1524 0.6250 0.4000
-1.000 0.5572 0.01494 0.00667 -0.1522 0.6169 0.4089
-0.750 0.5844 0.01496 0.00663 -0.1520 0.6087 0.4180
-0.500 0.6126 0.01506 0.00668 -0.1520 0.6019 0.4271
-0.250 0.6391 0.01509 0.00671 -0.1516 0.5938 0.4367
0.000 0.6676 0.01518 0.00674 -0.1517 0.5873 0.4464
0.250 0.6937 0.01524 0.00682 -0.1513 0.5798 0.4560
0.500 0.7212 0.01532 0.00688 -0.1512 0.5731 0.4667
0.750 0.7496 0.01544 0.00696 -0.1512 0.5673 0.4766
1.000 0.7757 0.01553 0.00707 -0.1509 0.5603 0.4882
1.250 0.8031 0.01559 0.00715 -0.1507 0.5544 0.4989
1.500 0.8307 0.01573 0.00728 -0.1507 0.5486 0.5107
1.750 0.8569 0.01584 0.00743 -0.1503 0.5424 0.5236
2.000 0.8845 0.01593 0.00755 -0.1502 0.5370 0.5366
2.250 0.9121 0.01609 0.00772 -0.1502 0.5318 0.5506
2.500 0.9374 0.01619 0.00792 -0.1497 0.5258 0.5659
2.750 0.9647 0.01631 0.00807 -0.1496 0.5207 0.5833
3.000 0.9937 0.01651 0.00826 -0.1498 0.5162 0.6030
3.250 1.0172 0.01660 0.00854 -0.1489 0.5105 0.6248
3.500 1.0429 0.01670 0.00874 -0.1485 0.5052 0.6519
3.750 1.0702 0.01683 0.00893 -0.1483 0.5009 0.6855
4.000 1.0936 0.01696 0.00925 -0.1474 0.4962 0.7281
4.250 1.1129 0.01698 0.00952 -0.1455 0.4911 0.7902
4.500 1.1267 0.01675 0.00953 -0.1420 0.4868 0.9415
4.750 1.1596 0.01707 0.00969 -0.1432 0.4826 1.0000
5.000 1.1840 0.01738 0.01005 -0.1428 0.4776 1.0000
5.250 1.2098 0.01765 0.01032 -0.1425 0.4725 1.0000
5.500 1.2381 0.01793 0.01053 -0.1427 0.4681 1.0000
5.750 1.2669 0.01833 0.01086 -0.1431 0.4640 1.0000
6.000 1.2887 0.01863 0.01126 -0.1421 0.4588 1.0000
6.250 1.3141 0.01892 0.01155 -0.1418 0.4540 1.0000
6.500 1.3427 0.01923 0.01180 -0.1420 0.4499 1.0000
6.750 1.3673 0.01964 0.01224 -0.1416 0.4454 1.0000
7.000 1.3888 0.01998 0.01266 -0.1406 0.4403 1.0000
7.250 1.4143 0.02028 0.01296 -0.1402 0.4357 1.0000
7.500 1.4440 0.02064 0.01323 -0.1407 0.4316 1.0000
7.750 1.4628 0.02105 0.01378 -0.1393 0.4267 1.0000
8.000 1.4842 0.02140 0.01420 -0.1382 0.4217 1.0000
8.250 1.5104 0.02169 0.01446 -0.1381 0.4172 1.0000
8.500 1.5361 0.02213 0.01489 -0.1379 0.4127 1.0000
8.750 1.5519 0.02253 0.01545 -0.1359 0.4074 1.0000
9.000 1.5741 0.02283 0.01578 -0.1350 0.4024 1.0000
9.250 1.6034 0.02317 0.01603 -0.1354 0.3978 1.0000
9.500 1.6155 0.02363 0.01668 -0.1329 0.3925 1.0000
9.750 1.6331 0.02396 0.01708 -0.1313 0.3870 1.0000
10.000 1.6591 0.02422 0.01728 -0.1311 0.3820 1.0000
10.250 1.6710 0.02472 0.01792 -0.1286 0.3766 1.0000
10.500 1.6838 0.02508 0.01837 -0.1262 0.3708 1.0000
10.750 1.7064 0.02533 0.01857 -0.1254 0.3656 1.0000
11.000 1.7125 0.02586 0.01924 -0.1219 0.3602 1.0000
11.250 1.7197 0.02630 0.01978 -0.1187 0.3545 1.0000
11.500 1.7394 0.02660 0.02003 -0.1175 0.3491 1.0000
11.750 1.7409 0.02730 0.02089 -0.1136 0.3435 1.0000
12.000 1.7462 0.02788 0.02157 -0.1105 0.3375 1.0000
12.250 1.7650 0.02827 0.02188 -0.1093 0.3316 1.0000
12.500 1.7582 0.02933 0.02316 -0.1048 0.3256 1.0000
12.750 1.7636 0.03008 0.02397 -0.1021 0.3194 1.0000
13.000 1.7688 0.03102 0.02496 -0.0996 0.3131 1.0000
13.250 1.7659 0.03232 0.02639 -0.0965 0.3064 1.0000
13.500 1.7747 0.03321 0.02724 -0.0946 0.2997 1.0000
13.750 1.7658 0.03510 0.02933 -0.0915 0.2926 1.0000
14.000 1.7705 0.03635 0.03052 -0.0897 0.2853 1.0000
14.250 1.7610 0.03869 0.03307 -0.0872 0.2778 1.0000
14.500 1.7616 0.04045 0.03478 -0.0856 0.2701 1.0000
14.750 1.7519 0.04324 0.03776 -0.0838 0.2622 1.0000
15.000 1.7501 0.04545 0.03992 -0.0825 0.2541 1.0000
15.250 1.7387 0.04883 0.04348 -0.0814 0.2457 1.0000
15.500 1.7332 0.05175 0.04639 -0.0805 0.2375 1.0000
15.750 1.7223 0.05546 0.05022 -0.0799 0.2286 1.0000
16.000 1.7130 0.05918 0.05398 -0.0796 0.2201 1.0000
16.250 1.7026 0.06315 0.05797 -0.0795 0.2113 1.0000
16.500 1.6913 0.06751 0.06243 -0.0797 0.2027 1.0000
16.750 1.6808 0.07176 0.06665 -0.0800 0.1942 1.0000
17.000 1.6675 0.07671 0.07171 -0.0807 0.1852 1.0000
17.250 1.6559 0.08148 0.07651 -0.0815 0.1768 1.0000
17.500 1.6430 0.08654 0.08160 -0.0825 0.1683 1.0000
17.750 1.6309 0.09167 0.08681 -0.0837 0.1599 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 561 AIRFOIL (e561-il)