EPPLER 561 AIRFOIL (e561-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 561 AIRFOIL (e561-il) Reynolds number: 1,000,000 Max Cl/Cd: 134.35 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e561-il-1000000.txt Download as CSV file: xf-e561-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-18.250 -0.5830 0.14519 0.14300 -0.0386 1.0000 0.0125
-18.000 -0.6175 0.13373 0.13138 -0.0443 1.0000 0.0124
-17.750 -0.6396 0.12530 0.12283 -0.0486 1.0000 0.0124
-17.500 -0.6583 0.11796 0.11537 -0.0523 1.0000 0.0124
-17.250 -0.6713 0.11188 0.10921 -0.0553 1.0000 0.0124
-17.000 -0.6819 0.10637 0.10361 -0.0581 1.0000 0.0124
-16.750 -0.6904 0.10146 0.09863 -0.0604 1.0000 0.0124
-16.500 -0.6996 0.09639 0.09348 -0.0629 1.0000 0.0124
-16.250 -0.7057 0.09204 0.08907 -0.0649 1.0000 0.0124
-16.000 -0.7128 0.08767 0.08463 -0.0669 1.0000 0.0124
-15.750 -0.7205 0.08314 0.08003 -0.0690 1.0000 0.0125
-15.500 -0.7274 0.07923 0.07605 -0.0705 1.0000 0.0125
-15.250 -0.7289 0.07604 0.07283 -0.0717 1.0000 0.0125
-15.000 -0.7367 0.07230 0.06902 -0.0729 1.0000 0.0125
-14.750 -0.7402 0.06904 0.06571 -0.0740 1.0000 0.0125
-14.500 -0.7486 0.06573 0.06235 -0.0746 1.0000 0.0127
-14.250 -0.7544 0.06269 0.05927 -0.0752 1.0000 0.0127
-14.000 -0.7619 0.05953 0.05607 -0.0758 1.0000 0.0127
-13.750 -0.7732 0.05642 0.05292 -0.0760 1.0000 0.0128
-13.500 -0.7700 0.05273 0.04917 -0.0798 0.9988 0.0128
-13.250 -0.7633 0.04857 0.04495 -0.0853 0.9965 0.0130
-13.000 -0.7529 0.04458 0.04089 -0.0912 0.9942 0.0130
-12.750 -0.7462 0.04052 0.03677 -0.0966 0.9907 0.0131
-12.500 -0.7367 0.03634 0.03251 -0.1028 0.9863 0.0133
-12.250 -0.7204 0.03223 0.02831 -0.1102 0.9833 0.0133
-12.000 -0.7109 0.02790 0.02389 -0.1163 0.9750 0.0135
-11.750 -0.6764 0.02335 0.01917 -0.1278 0.9720 0.0137
-11.500 -0.6590 0.02143 0.01715 -0.1297 0.9633 0.0138
-11.250 -0.6267 0.01995 0.01560 -0.1332 0.9603 0.0141
-11.000 -0.6077 0.01882 0.01440 -0.1333 0.9504 0.0142
-10.750 -0.5817 0.01775 0.01325 -0.1344 0.9430 0.0145
-10.500 -0.5514 0.01677 0.01220 -0.1361 0.9356 0.0148
-10.250 -0.5144 0.01588 0.01124 -0.1389 0.9296 0.0150
-10.000 -0.4715 0.01515 0.01043 -0.1427 0.9219 0.0153
-9.750 -0.4203 0.01379 0.00895 -0.1491 0.9144 0.0161
-9.250 -0.3269 0.01254 0.00746 -0.1581 0.8803 0.0174
-9.000 -0.2971 0.01219 0.00696 -0.1588 0.8569 0.0179
-8.750 -0.2709 0.01193 0.00656 -0.1587 0.8350 0.0184
-8.500 -0.2467 0.01154 0.00604 -0.1583 0.8148 0.0193
-8.250 -0.2221 0.01124 0.00563 -0.1578 0.7967 0.0205
-8.000 -0.1969 0.01101 0.00530 -0.1574 0.7804 0.0218
-7.750 -0.1717 0.01068 0.00490 -0.1570 0.7647 0.0252
-7.500 -0.1467 0.01018 0.00442 -0.1568 0.7500 0.0407
-7.250 -0.1215 0.00963 0.00397 -0.1567 0.7359 0.0671
-7.000 -0.0955 0.00929 0.00366 -0.1566 0.7227 0.0880
-6.750 -0.0691 0.00907 0.00343 -0.1564 0.7100 0.1046
-6.500 -0.0418 0.00884 0.00323 -0.1564 0.6981 0.1207
-6.250 -0.0146 0.00870 0.00307 -0.1563 0.6864 0.1341
-6.000 0.0123 0.00859 0.00293 -0.1561 0.6741 0.1473
-5.750 0.0402 0.00849 0.00282 -0.1560 0.6629 0.1587
-5.500 0.0678 0.00839 0.00271 -0.1559 0.6526 0.1711
-5.250 0.0955 0.00832 0.00261 -0.1558 0.6420 0.1806
-5.000 0.1236 0.00828 0.00254 -0.1558 0.6322 0.1902
-4.500 0.1795 0.00819 0.00240 -0.1556 0.6122 0.2092
-4.250 0.2071 0.00818 0.00235 -0.1555 0.6026 0.2190
-4.000 0.2355 0.00814 0.00230 -0.1555 0.5939 0.2291
-3.750 0.2633 0.00816 0.00227 -0.1553 0.5851 0.2371
-3.500 0.2917 0.00812 0.00222 -0.1554 0.5766 0.2459
-3.250 0.3194 0.00815 0.00220 -0.1552 0.5678 0.2528
-3.000 0.3479 0.00813 0.00216 -0.1552 0.5597 0.2602
-2.750 0.3756 0.00816 0.00215 -0.1551 0.5516 0.2675
-2.500 0.4042 0.00816 0.00213 -0.1551 0.5446 0.2740
-2.250 0.4321 0.00818 0.00213 -0.1550 0.5368 0.2822
-2.000 0.4603 0.00821 0.00213 -0.1549 0.5296 0.2888
-1.750 0.4884 0.00822 0.00213 -0.1549 0.5223 0.2969
-1.500 0.5161 0.00828 0.00215 -0.1547 0.5155 0.3041
-1.250 0.5446 0.00829 0.00216 -0.1548 0.5094 0.3114
-1.000 0.5723 0.00834 0.00218 -0.1546 0.5025 0.3188
-0.750 0.6002 0.00839 0.00221 -0.1545 0.4964 0.3250
-0.500 0.6284 0.00841 0.00224 -0.1545 0.4904 0.3335
-0.250 0.6558 0.00849 0.00228 -0.1543 0.4843 0.3403
0.000 0.6838 0.00853 0.00232 -0.1543 0.4791 0.3473
0.250 0.7119 0.00857 0.00236 -0.1542 0.4736 0.3551
0.500 0.7391 0.00866 0.00242 -0.1540 0.4677 0.3618
0.750 0.7668 0.00870 0.00248 -0.1539 0.4628 0.3695
1.000 0.7948 0.00876 0.00253 -0.1538 0.4582 0.3767
1.250 0.8221 0.00882 0.00259 -0.1537 0.4531 0.3846
1.500 0.8488 0.00893 0.00268 -0.1534 0.4476 0.3923
1.750 0.8769 0.00897 0.00274 -0.1533 0.4439 0.3996
2.000 0.9045 0.00902 0.00282 -0.1532 0.4396 0.4084
2.250 0.9312 0.00913 0.00290 -0.1529 0.4349 0.4161
2.500 0.9579 0.00922 0.00301 -0.1527 0.4302 0.4252
3.000 1.0128 0.00933 0.00318 -0.1524 0.4227 0.4451
3.250 1.0393 0.00944 0.00328 -0.1521 0.4184 0.4551
3.500 1.0651 0.00957 0.00342 -0.1517 0.4137 0.4672
3.750 1.0927 0.00960 0.00352 -0.1516 0.4108 0.4808
4.000 1.1196 0.00967 0.00363 -0.1514 0.4070 0.4950
4.250 1.1457 0.00977 0.00375 -0.1511 0.4029 0.5114
4.500 1.1710 0.00990 0.00391 -0.1506 0.3984 0.5308
4.750 1.1976 0.00997 0.00405 -0.1504 0.3953 0.5537
5.250 1.2502 0.01009 0.00434 -0.1498 0.3879 0.6114
5.500 1.2751 0.01020 0.00453 -0.1493 0.3838 0.6497
5.750 1.2999 0.01030 0.00473 -0.1488 0.3798 0.6953
6.000 1.3259 0.01029 0.00490 -0.1484 0.3767 0.7565
6.250 1.3481 0.01026 0.00509 -0.1472 0.3729 0.8450
6.500 1.3636 0.01018 0.00520 -0.1446 0.3691 1.0000
6.750 1.3855 0.01042 0.00541 -0.1435 0.3644 1.0000
7.000 1.4108 0.01053 0.00555 -0.1430 0.3614 1.0000
7.250 1.4349 0.01068 0.00571 -0.1423 0.3574 1.0000
7.500 1.4573 0.01088 0.00590 -0.1414 0.3530 1.0000
7.750 1.4779 0.01115 0.00614 -0.1401 0.3479 1.0000
8.000 1.5025 0.01128 0.00630 -0.1395 0.3442 1.0000
8.250 1.5253 0.01147 0.00650 -0.1386 0.3394 1.0000
8.500 1.5457 0.01174 0.00675 -0.1374 0.3339 1.0000
8.750 1.5679 0.01195 0.00698 -0.1364 0.3293 1.0000
9.000 1.5898 0.01216 0.00720 -0.1354 0.3245 1.0000
9.250 1.6090 0.01248 0.00749 -0.1340 0.3184 1.0000
9.500 1.6300 0.01273 0.00776 -0.1329 0.3130 1.0000
9.750 1.6501 0.01302 0.00805 -0.1317 0.3062 1.0000
10.000 1.6674 0.01340 0.00841 -0.1300 0.2992 1.0000
10.250 1.6867 0.01371 0.00874 -0.1287 0.2917 1.0000
10.500 1.7027 0.01417 0.00916 -0.1269 0.2841 1.0000
10.750 1.7197 0.01458 0.00957 -0.1253 0.2750 1.0000
11.000 1.7345 0.01509 0.01006 -0.1234 0.2663 1.0000
11.250 1.7482 0.01565 0.01060 -0.1213 0.2572 1.0000
11.500 1.7625 0.01621 0.01116 -0.1194 0.2490 1.0000
11.750 1.7725 0.01697 0.01187 -0.1170 0.2394 1.0000
12.000 1.7850 0.01764 0.01255 -0.1150 0.2303 1.0000
12.250 1.7930 0.01855 0.01342 -0.1125 0.2194 1.0000
12.500 1.7991 0.01960 0.01444 -0.1099 0.2094 1.0000
12.750 1.8063 0.02065 0.01547 -0.1076 0.1989 1.0000
13.000 1.8111 0.02188 0.01669 -0.1051 0.1887 1.0000
13.250 1.8134 0.02334 0.01813 -0.1026 0.1789 1.0000
13.500 1.8140 0.02500 0.01976 -0.1001 0.1681 1.0000
13.750 1.8138 0.02681 0.02156 -0.0978 0.1573 1.0000
14.000 1.8121 0.02886 0.02360 -0.0956 0.1473 1.0000
14.250 1.8079 0.03124 0.02596 -0.0935 0.1379 1.0000
14.500 1.8034 0.03379 0.02850 -0.0917 0.1285 1.0000
14.750 1.7994 0.03644 0.03116 -0.0902 0.1201 1.0000
15.000 1.7914 0.03961 0.03433 -0.0888 0.1118 1.0000
15.250 1.7853 0.04274 0.03748 -0.0878 0.1046 1.0000
15.500 1.7789 0.04605 0.04082 -0.0870 0.0984 1.0000
15.750 1.7696 0.04983 0.04462 -0.0863 0.0920 1.0000
16.000 1.7629 0.05344 0.04828 -0.0860 0.0867 1.0000
16.250 1.7532 0.05755 0.05242 -0.0858 0.0812 1.0000
16.500 1.7458 0.06151 0.05643 -0.0858 0.0765 1.0000
16.750 1.7346 0.06605 0.06100 -0.0860 0.0714 1.0000
17.000 1.7262 0.07038 0.06539 -0.0864 0.0671 1.0000
17.250 1.7156 0.07511 0.07016 -0.0870 0.0627 1.0000
17.500 1.7060 0.07980 0.07490 -0.0878 0.0586 1.0000
17.750 1.6960 0.08465 0.07980 -0.0887 0.0546 1.0000
18.000 1.6858 0.08965 0.08484 -0.0898 0.0508 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 561 AIRFOIL (e561-il)