EPPLER 636 AIRFOIL (e636-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 636 AIRFOIL (e636-il) Reynolds number: 100,000 Max Cl/Cd: 44 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e636-il-100000-n5.txt Download as CSV file: xf-e636-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 636 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4489 0.11637 0.11176 -0.0023 1.0000 0.0532
-10.000 -0.4526 0.11274 0.10819 -0.0061 1.0000 0.0535
-9.750 -0.4554 0.10882 0.10433 -0.0097 1.0000 0.0537
-9.500 -0.4592 0.10460 0.10016 -0.0138 1.0000 0.0538
-9.000 -0.4460 0.09513 0.09080 -0.0134 1.0000 0.0549
-8.750 -0.4378 0.09153 0.08724 -0.0129 1.0000 0.0558
-8.500 -0.4364 0.08775 0.08352 -0.0145 1.0000 0.0564
-8.250 -0.4366 0.08408 0.07990 -0.0162 1.0000 0.0573
-8.000 -0.4391 0.08064 0.07650 -0.0172 1.0000 0.0584
-7.750 -0.4437 0.07743 0.07333 -0.0175 1.0000 0.0600
-7.500 -0.4470 0.07394 0.06985 -0.0182 1.0000 0.0607
-7.250 -0.4496 0.07055 0.06642 -0.0189 1.0000 0.0635
-6.750 -0.4327 0.05781 0.05284 -0.0233 0.8602 0.0415
-6.250 -0.4193 0.04857 0.04262 -0.0196 0.7952 0.0331
-6.000 -0.4076 0.04580 0.03961 -0.0180 0.7701 0.0326
-5.750 -0.3953 0.04307 0.03657 -0.0161 0.7490 0.0320
-5.500 -0.3816 0.04035 0.03349 -0.0142 0.7309 0.0314
-5.250 -0.3664 0.03772 0.03047 -0.0122 0.7138 0.0309
-5.000 -0.3495 0.03516 0.02747 -0.0102 0.6980 0.0305
-4.750 -0.3307 0.03285 0.02473 -0.0084 0.6833 0.0302
-4.500 -0.3101 0.03076 0.02219 -0.0067 0.6693 0.0304
-4.250 -0.2879 0.02907 0.02005 -0.0053 0.6560 0.0314
-4.000 -0.2641 0.02752 0.01805 -0.0040 0.6427 0.0326
-3.750 -0.2387 0.02600 0.01617 -0.0030 0.6301 0.0328
-3.500 -0.2121 0.02460 0.01447 -0.0023 0.6181 0.0330
-3.250 -0.1849 0.02340 0.01298 -0.0017 0.6068 0.0333
-3.000 -0.1576 0.02219 0.01162 -0.0013 0.5958 0.0339
-2.750 -0.1310 0.02118 0.01055 -0.0009 0.5847 0.0347
-2.500 -0.1054 0.02040 0.00969 -0.0002 0.5748 0.0358
-2.250 -0.0808 0.01978 0.00898 0.0006 0.5653 0.0381
-2.000 -0.0562 0.01931 0.00839 0.0014 0.5554 0.0414
-1.750 -0.0329 0.01876 0.00774 0.0025 0.5473 0.0444
-1.500 -0.0088 0.01833 0.00723 0.0035 0.5382 0.0481
-1.250 0.0161 0.01800 0.00673 0.0044 0.5303 0.0543
-1.000 0.2378 0.01598 0.00738 -0.0303 0.5091 1.0000
-0.750 0.2603 0.01598 0.00715 -0.0295 0.5024 1.0000
-0.500 0.2832 0.01598 0.00699 -0.0288 0.4947 1.0000
-0.250 0.3060 0.01601 0.00681 -0.0280 0.4886 1.0000
0.000 0.3293 0.01605 0.00674 -0.0273 0.4813 1.0000
0.250 0.3523 0.01610 0.00663 -0.0266 0.4753 1.0000
0.500 0.3756 0.01617 0.00659 -0.0259 0.4690 1.0000
0.750 0.3989 0.01625 0.00656 -0.0252 0.4628 1.0000
1.000 0.4222 0.01635 0.00650 -0.0244 0.4579 1.0000
1.250 0.4458 0.01646 0.00658 -0.0238 0.4517 1.0000
1.500 0.4692 0.01657 0.00659 -0.0231 0.4462 1.0000
1.750 0.4927 0.01670 0.00663 -0.0224 0.4412 1.0000
2.000 0.5163 0.01684 0.00676 -0.0217 0.4355 1.0000
2.250 0.5398 0.01699 0.00683 -0.0210 0.4308 1.0000
2.500 0.5633 0.01716 0.00694 -0.0204 0.4262 1.0000
2.750 0.5867 0.01734 0.00714 -0.0197 0.4207 1.0000
3.000 0.6101 0.01752 0.00726 -0.0190 0.4161 1.0000
3.250 0.6335 0.01773 0.00742 -0.0183 0.4121 1.0000
3.500 0.6567 0.01797 0.00773 -0.0176 0.4070 1.0000
3.750 0.6799 0.01820 0.00794 -0.0169 0.4025 1.0000
4.000 0.7032 0.01841 0.00808 -0.0162 0.3987 1.0000
4.250 0.7259 0.01871 0.00848 -0.0155 0.3937 1.0000
4.500 0.7488 0.01900 0.00882 -0.0147 0.3893 1.0000
4.750 0.7719 0.01925 0.00905 -0.0140 0.3855 1.0000
5.000 0.7945 0.01956 0.00940 -0.0132 0.3813 1.0000
5.250 0.8167 0.01991 0.00985 -0.0124 0.3764 1.0000
5.500 0.8393 0.02022 0.01020 -0.0116 0.3725 1.0000
5.750 0.8623 0.02050 0.01046 -0.0109 0.3692 1.0000
6.000 0.8835 0.02094 0.01106 -0.0100 0.3642 1.0000
6.250 0.9052 0.02132 0.01153 -0.0092 0.3597 1.0000
6.500 0.9276 0.02163 0.01188 -0.0083 0.3560 1.0000
6.750 0.9492 0.02205 0.01238 -0.0074 0.3520 1.0000
7.000 0.9694 0.02254 0.01304 -0.0065 0.3470 1.0000
7.250 0.9910 0.02289 0.01346 -0.0055 0.3428 1.0000
7.500 1.0137 0.02318 0.01376 -0.0047 0.3394 1.0000
7.750 1.0315 0.02381 0.01465 -0.0035 0.3335 1.0000
8.000 1.0520 0.02420 0.01515 -0.0025 0.3288 1.0000
8.250 1.0746 0.02442 0.01535 -0.0016 0.3249 1.0000
8.500 1.0906 0.02509 0.01632 -0.0002 0.3184 1.0000
8.750 1.1107 0.02539 0.01670 0.0009 0.3130 1.0000
9.000 1.1294 0.02581 0.01724 0.0022 0.3076 1.0000
9.250 1.1458 0.02635 0.01797 0.0036 0.3013 1.0000
9.500 1.1668 0.02653 0.01820 0.0047 0.2963 1.0000
9.750 1.1795 0.02727 0.01921 0.0065 0.2893 1.0000
10.000 1.1977 0.02751 0.01953 0.0079 0.2831 1.0000
10.250 1.2099 0.02818 0.02042 0.0097 0.2759 1.0000
10.500 1.2254 0.02850 0.02083 0.0114 0.2690 1.0000
10.750 1.2346 0.02925 0.02179 0.0135 0.2610 1.0000
11.000 1.2479 0.02957 0.02219 0.0154 0.2538 1.0000
11.250 1.2514 0.03057 0.02342 0.0180 0.2451 1.0000
11.500 1.2573 0.03126 0.02419 0.0205 0.2373 1.0000
11.750 1.2544 0.03236 0.02544 0.0237 0.2293 1.0000
12.000 1.2483 0.03358 0.02673 0.0270 0.2221 1.0000
12.250 1.2400 0.03530 0.02856 0.0293 0.2145 1.0000
12.500 1.2302 0.03768 0.03105 0.0306 0.2070 1.0000
12.750 1.2216 0.04024 0.03365 0.0312 0.1990 1.0000
13.000 1.2059 0.04405 0.03760 0.0309 0.1911 1.0000
13.250 1.1953 0.04744 0.04101 0.0306 0.1834 1.0000
13.500 1.1776 0.05200 0.04568 0.0296 0.1758 1.0000
13.750 1.1643 0.05611 0.04980 0.0288 0.1682 1.0000
14.000 1.1502 0.06044 0.05417 0.0277 0.1604 1.0000
14.250 1.1359 0.06495 0.05873 0.0266 0.1530 1.0000
14.500 1.1276 0.06857 0.06231 0.0258 0.1455 1.0000
14.750 1.1123 0.07355 0.06740 0.0243 0.1383 1.0000
15.000 1.1065 0.07709 0.07089 0.0234 0.1310 1.0000
15.250 1.0940 0.08198 0.07587 0.0218 0.1240 1.0000
15.500 1.0883 0.08573 0.07959 0.0207 0.1173 1.0000
15.750 1.0771 0.09061 0.08453 0.0190 0.1103 1.0000
16.000 1.0697 0.09487 0.08878 0.0175 0.1036 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 636 AIRFOIL (e636-il)