EPPLER 407 AIRFOIL (e407-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 407 AIRFOIL (e407-il) Reynolds number: 50,000 Max Cl/Cd: 30.92 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e407-il-50000-n5.txt Download as CSV file: xf-e407-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 407 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4847 0.11275 0.10605 -0.0481 1.0000 0.0382
-10.500 -0.5186 0.10101 0.09431 -0.0545 1.0000 0.0360
-10.250 -0.5338 0.09608 0.08941 -0.0556 1.0000 0.0358
-10.000 -0.5502 0.09156 0.08496 -0.0563 1.0000 0.0355
-9.750 -0.5721 0.08691 0.08031 -0.0570 1.0000 0.0353
-9.500 -0.5931 0.08301 0.07641 -0.0568 1.0000 0.0350
-9.250 -0.6159 0.07963 0.07300 -0.0559 1.0000 0.0348
-9.000 -0.6416 0.07665 0.06999 -0.0541 1.0000 0.0346
-8.750 -0.6660 0.07380 0.06709 -0.0522 1.0000 0.0344
-8.500 -0.6839 0.07043 0.06359 -0.0514 1.0000 0.0342
-8.250 -0.6982 0.06673 0.05966 -0.0504 1.0000 0.0343
-8.000 -0.7062 0.06298 0.05562 -0.0497 1.0000 0.0343
-7.750 -0.7085 0.05913 0.05141 -0.0492 1.0000 0.0346
-7.500 -0.7047 0.05539 0.04725 -0.0488 1.0000 0.0348
-7.250 -0.6956 0.05195 0.04329 -0.0485 1.0000 0.0352
-7.000 -0.6822 0.04874 0.03956 -0.0480 1.0000 0.0359
-6.750 -0.6671 0.04582 0.03643 -0.0474 1.0000 0.0371
-6.500 -0.6509 0.04386 0.03434 -0.0469 1.0000 0.0396
-6.250 -0.6327 0.04206 0.03228 -0.0461 1.0000 0.0432
-6.000 -0.6132 0.04035 0.03019 -0.0447 1.0000 0.0467
-5.750 -0.5909 0.03883 0.02865 -0.0436 0.9982 0.0503
-5.500 -0.5651 0.03780 0.02740 -0.0432 0.9955 0.0589
-5.250 -0.5412 0.03667 0.02630 -0.0429 0.9928 0.0677
-5.000 -0.5167 0.03551 0.02506 -0.0430 0.9897 0.0821
-4.750 -0.4901 0.03414 0.02378 -0.0442 0.9866 0.1053
-4.500 -0.4594 0.03227 0.02229 -0.0472 0.9842 0.1574
-4.250 -0.4279 0.02965 0.02133 -0.0515 0.9824 0.3571
-4.000 -0.4162 0.03100 0.02335 -0.0458 0.9773 0.5084
-3.750 -0.3879 0.03234 0.02434 -0.0452 0.9731 0.5767
-3.500 -0.3670 0.03377 0.02551 -0.0422 0.9687 0.6107
-3.250 -0.3483 0.03469 0.02621 -0.0394 0.9637 0.6362
-3.000 -0.3265 0.03555 0.02685 -0.0373 0.9592 0.6575
-2.750 -0.3037 0.03606 0.02713 -0.0361 0.9548 0.6765
-2.500 -0.2841 0.03640 0.02729 -0.0341 0.9496 0.6915
-2.250 -0.2584 0.03676 0.02740 -0.0336 0.9452 0.7056
-2.000 -0.2384 0.03687 0.02735 -0.0322 0.9399 0.7168
-1.750 -0.2150 0.03696 0.02727 -0.0315 0.9347 0.7264
-1.500 -0.1820 0.03708 0.02718 -0.0334 0.9307 0.7356
-1.250 -0.1614 0.03695 0.02688 -0.0327 0.9244 0.7420
-1.000 -0.1279 0.03698 0.02672 -0.0350 0.9196 0.7489
-0.750 -0.1025 0.03697 0.02658 -0.0352 0.9141 0.7538
-0.500 -0.0735 0.03697 0.02644 -0.0366 0.9082 0.7596
-0.250 -0.0383 0.03710 0.02641 -0.0389 0.9038 0.7644
0.000 -0.0170 0.03704 0.02628 -0.0386 0.8965 0.7690
0.250 0.0181 0.03717 0.02630 -0.0411 0.8914 0.7739
0.500 0.0451 0.03721 0.02626 -0.0420 0.8845 0.7784
0.750 0.0754 0.03731 0.02630 -0.0432 0.8784 0.7825
1.000 0.1047 0.03741 0.02635 -0.0445 0.8715 0.7870
1.250 0.1381 0.03755 0.02643 -0.0467 0.8649 0.7916
1.500 0.1647 0.03763 0.02649 -0.0471 0.8576 0.7954
1.750 0.1966 0.03773 0.02659 -0.0487 0.8506 0.7996
2.000 0.2263 0.03789 0.02673 -0.0502 0.8425 0.8041
2.250 0.2593 0.03795 0.02682 -0.0517 0.8358 0.8081
2.750 0.3227 0.03811 0.02705 -0.0545 0.8206 0.8166
3.000 0.3470 0.03824 0.02724 -0.0548 0.8100 0.8208
3.500 0.4099 0.03824 0.02739 -0.0570 0.7934 0.8295
3.750 0.4364 0.03838 0.02764 -0.0577 0.7821 0.8341
4.000 0.4619 0.03838 0.02774 -0.0576 0.7715 0.8385
4.250 0.5008 0.03810 0.02757 -0.0595 0.7644 0.8433
4.500 0.5275 0.03816 0.02779 -0.0600 0.7521 0.8482
4.750 0.5516 0.03811 0.02788 -0.0596 0.7397 0.8530
5.000 0.5794 0.03803 0.02794 -0.0599 0.7275 0.8583
5.250 0.6088 0.03785 0.02796 -0.0603 0.7154 0.8637
5.500 0.6388 0.03748 0.02777 -0.0604 0.7038 0.8692
5.750 0.6765 0.03677 0.02726 -0.0615 0.6936 0.8752
6.000 0.7033 0.03632 0.02705 -0.0609 0.6797 0.8813
6.250 0.7331 0.03576 0.02670 -0.0607 0.6654 0.8882
6.500 0.7628 0.03504 0.02620 -0.0603 0.6506 0.8953
6.750 0.7860 0.03483 0.02620 -0.0594 0.6311 0.9033
7.000 0.8146 0.03411 0.02572 -0.0586 0.6126 0.9122
7.250 0.8500 0.03299 0.02486 -0.0585 0.5931 0.9221
7.500 0.8771 0.03248 0.02454 -0.0577 0.5671 0.9344
7.750 0.9103 0.03171 0.02392 -0.0575 0.5370 0.9511
8.000 0.9388 0.03126 0.02355 -0.0570 0.5009 1.0000
8.250 0.9689 0.03141 0.02371 -0.0572 0.4592 1.0000
8.500 0.9919 0.03208 0.02429 -0.0569 0.4153 1.0000
8.750 1.0096 0.03314 0.02522 -0.0561 0.3720 1.0000
9.000 1.0238 0.03449 0.02643 -0.0552 0.3304 1.0000
9.250 1.0356 0.03608 0.02790 -0.0542 0.2909 1.0000
9.500 1.0456 0.03787 0.02956 -0.0532 0.2538 1.0000
9.750 1.0549 0.03981 0.03138 -0.0523 0.2204 1.0000
10.000 1.0632 0.04191 0.03341 -0.0514 0.1905 1.0000
10.250 1.0718 0.04410 0.03553 -0.0506 0.1639 1.0000
10.500 1.0804 0.04639 0.03776 -0.0499 0.1412 1.0000
10.750 1.0885 0.04876 0.04009 -0.0492 0.1224 1.0000
11.000 1.0975 0.05118 0.04248 -0.0486 0.1074 1.0000
11.250 1.1078 0.05359 0.04493 -0.0480 0.0945 1.0000
11.500 1.1183 0.05603 0.04748 -0.0475 0.0831 1.0000
11.750 1.1326 0.05844 0.05004 -0.0470 0.0744 1.0000
12.000 1.1413 0.06103 0.05257 -0.0466 0.0676 1.0000
12.250 1.1541 0.06374 0.05565 -0.0461 0.0609 1.0000
12.500 1.1665 0.06641 0.05842 -0.0457 0.0565 1.0000
12.750 1.1753 0.06975 0.06209 -0.0452 0.0524 1.0000
13.000 1.1767 0.07333 0.06599 -0.0449 0.0488 1.0000
13.250 1.1778 0.07683 0.06965 -0.0448 0.0462 1.0000
13.500 1.1843 0.08040 0.07329 -0.0447 0.0442 1.0000
13.750 1.1770 0.08534 0.07860 -0.0449 0.0432 1.0000
14.000 1.1637 0.09088 0.08454 -0.0458 0.0424 1.0000
14.250 1.1476 0.09696 0.09096 -0.0475 0.0419 1.0000
14.500 1.1293 0.10371 0.09803 -0.0502 0.0415 1.0000
14.750 1.1093 0.11124 0.10582 -0.0539 0.0415 1.0000
15.000 1.0878 0.11968 0.11449 -0.0588 0.0418 1.0000
15.250 1.0659 0.12909 0.12407 -0.0648 0.0422 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 407 AIRFOIL (e407-il)