EPPLER 336 AIRFOIL (e336-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 336 AIRFOIL (e336-il) Reynolds number: 50,000 Max Cl/Cd: 16.12 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e336-il-50000.txt Download as CSV file: xf-e336-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 336 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4121 0.10220 0.09616 0.0007 1.0000 0.2798
-8.500 -0.4159 0.09944 0.09351 0.0009 1.0000 0.2981
-8.250 -0.4298 0.09717 0.09138 0.0010 1.0000 0.3160
-8.000 -0.3959 0.09287 0.08709 0.0039 1.0000 0.3445
-7.750 -0.3840 0.08976 0.08406 0.0061 1.0000 0.3757
-7.500 -0.3824 0.08762 0.08202 0.0087 1.0000 0.4116
-7.250 -0.3438 0.08414 0.07854 0.0128 1.0000 0.4712
-7.000 -0.3132 0.08121 0.07565 0.0163 1.0000 0.5338
-6.750 -0.2830 0.07827 0.07276 0.0189 1.0000 0.5952
-4.500 -0.4816 0.04773 0.04041 0.0028 0.9646 0.1357
-4.250 -0.4292 0.04344 0.03539 -0.0011 0.9468 0.1179
-4.000 -0.3762 0.03973 0.03095 -0.0048 0.9299 0.1115
-3.750 -0.3163 0.03695 0.02731 -0.0088 0.9139 0.1067
-3.500 0.0697 0.02484 0.01801 -0.0541 0.9425 1.0000
-3.250 0.1346 0.02412 0.01650 -0.0617 0.9104 1.0000
-3.000 0.1752 0.02387 0.01572 -0.0645 0.8791 1.0000
-2.750 0.2003 0.02396 0.01543 -0.0644 0.8520 1.0000
-2.500 0.2212 0.02417 0.01534 -0.0634 0.8287 1.0000
-2.250 0.2407 0.02444 0.01534 -0.0622 0.8075 1.0000
-2.000 0.2599 0.02473 0.01539 -0.0609 0.7887 1.0000
-1.750 0.2792 0.02505 0.01549 -0.0596 0.7715 1.0000
-1.500 0.2987 0.02541 0.01565 -0.0584 0.7553 1.0000
-1.250 0.3186 0.02583 0.01590 -0.0574 0.7403 1.0000
-1.000 0.3396 0.02630 0.01622 -0.0568 0.7267 1.0000
-0.750 0.3599 0.02680 0.01658 -0.0559 0.7139 1.0000
-0.500 0.3796 0.02725 0.01687 -0.0546 0.7027 1.0000
-0.250 0.3999 0.02785 0.01737 -0.0540 0.6908 1.0000
0.000 0.4201 0.02856 0.01802 -0.0535 0.6800 1.0000
0.250 0.4403 0.02903 0.01834 -0.0521 0.6715 1.0000
0.500 0.4599 0.02996 0.01927 -0.0521 0.6607 1.0000
0.750 0.4796 0.03064 0.01987 -0.0511 0.6524 1.0000
1.000 0.4986 0.03159 0.02080 -0.0507 0.6434 1.0000
1.250 0.5176 0.03244 0.02161 -0.0499 0.6358 1.0000
1.500 0.5353 0.03354 0.02272 -0.0494 0.6275 1.0000
1.750 0.5542 0.03437 0.02350 -0.0483 0.6209 1.0000
2.000 0.5683 0.03596 0.02516 -0.0482 0.6127 1.0000
2.250 0.5883 0.03661 0.02575 -0.0467 0.6069 1.0000
2.500 0.5964 0.03885 0.02810 -0.0466 0.5992 1.0000
2.750 0.6126 0.03995 0.02920 -0.0454 0.5927 1.0000
3.000 0.6220 0.04184 0.03113 -0.0444 0.5865 1.0000
3.250 0.6240 0.04433 0.03368 -0.0434 0.5802 1.0000
3.500 0.6449 0.04502 0.03435 -0.0419 0.5745 1.0000
3.750 0.6296 0.04879 0.03821 -0.0403 0.5687 1.0000
4.000 0.6193 0.05187 0.04132 -0.0382 0.5640 1.0000
4.250 0.6469 0.05233 0.04180 -0.0370 0.5581 1.0000
4.500 0.6152 0.05668 0.04617 -0.0340 0.5550 1.0000
4.750 0.5807 0.06052 0.04999 -0.0300 0.5541 1.0000
5.000 0.5536 0.06358 0.05298 -0.0258 0.5548 1.0000
5.250 0.5375 0.06639 0.05577 -0.0228 0.5556 1.0000
5.500 0.5296 0.06915 0.05856 -0.0209 0.5581 1.0000
5.750 0.3548 0.07685 0.06619 -0.0150 0.7086 1.0000
6.000 0.3635 0.07857 0.06790 -0.0137 0.6948 1.0000
6.250 0.3755 0.08061 0.06993 -0.0129 0.6828 1.0000
6.500 0.4076 0.08402 0.07337 -0.0145 0.6714 1.0000
6.750 0.4038 0.08475 0.07408 -0.0117 0.6573 1.0000
7.000 0.4060 0.08621 0.07555 -0.0098 0.6438 1.0000
7.250 0.4124 0.08812 0.07748 -0.0085 0.6323 1.0000
7.500 0.4456 0.09196 0.08136 -0.0101 0.6224 1.0000
7.750 0.4398 0.09266 0.08206 -0.0074 0.6085 1.0000
8.000 0.4390 0.09414 0.08354 -0.0056 0.5961 1.0000
8.250 0.4515 0.09679 0.08621 -0.0052 0.5870 1.0000
8.500 0.4695 0.09942 0.08889 -0.0050 0.5749 1.0000
8.750 0.4643 0.10060 0.09008 -0.0031 0.5623 1.0000
9.000 0.4696 0.10287 0.09236 -0.0022 0.5520 1.0000
9.250 0.4969 0.10655 0.09618 -0.0028 0.5419 1.0000
9.500 0.4860 0.10735 0.09696 -0.0008 0.5298 1.0000
9.750 0.4909 0.10976 0.09939 0.0000 0.5202 1.0000
10.000 0.5179 0.11361 0.10337 -0.0006 0.5097 1.0000
10.250 0.5039 0.11435 0.10410 0.0012 0.4989 1.0000
10.500 0.5159 0.11752 0.10733 0.0013 0.4906 1.0000
10.750 0.5318 0.12042 0.11031 0.0015 0.4784 1.0000
11.000 0.5213 0.12178 0.11167 0.0025 0.4691 1.0000
11.250 0.5472 0.12634 0.11634 0.0019 0.4607 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 336 AIRFOIL (e336-il)