EPPLER 379 AIRFOIL (e379-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 379 AIRFOIL (e379-il) Reynolds number: 200,000 Max Cl/Cd: 91.16 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e379-il-200000-n5.txt Download as CSV file: xf-e379-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 379 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.4399 0.15112 0.14770 0.0258 1.0000 0.0055
-12.000 -0.4312 0.14817 0.14476 0.0242 1.0000 0.0055
-11.750 -0.4225 0.14518 0.14179 0.0227 1.0000 0.0055
-11.500 -0.4135 0.14225 0.13887 0.0212 1.0000 0.0056
-11.250 -0.4047 0.13932 0.13596 0.0197 1.0000 0.0055
-11.000 -0.3959 0.13643 0.13309 0.0182 1.0000 0.0054
-10.750 -0.3868 0.13354 0.13022 0.0167 1.0000 0.0055
-10.500 -0.3778 0.13071 0.12741 0.0153 1.0000 0.0055
-10.250 -0.3688 0.12787 0.12459 0.0138 1.0000 0.0056
-10.000 -0.3596 0.12510 0.12184 0.0124 1.0000 0.0056
-9.750 -0.3503 0.12232 0.11909 0.0109 1.0000 0.0056
-9.500 -0.3411 0.11960 0.11640 0.0095 1.0000 0.0057
-9.250 -0.3317 0.11689 0.11368 0.0081 1.0000 0.0058
-9.000 -0.3223 0.11424 0.11105 0.0067 1.0000 0.0057
-8.750 -0.3127 0.11162 0.10847 0.0053 1.0000 0.0059
-8.500 -0.3031 0.10903 0.10593 0.0039 1.0000 0.0060
-8.250 -0.2938 0.10648 0.10342 0.0025 1.0000 0.0060
-8.000 -0.2836 0.10401 0.10098 0.0009 0.9608 0.0061
-7.750 -0.2612 0.10069 0.09753 -0.0038 0.8803 0.0063
-7.500 -0.2522 0.09840 0.09513 -0.0047 0.8377 0.0062
-7.000 -0.2311 0.09426 0.09080 -0.0071 0.7819 0.0065
-6.750 -0.2187 0.09189 0.08834 -0.0088 0.7606 0.0066
-6.500 -0.2034 0.08983 0.08619 -0.0111 0.7430 0.0067
-6.250 -0.1870 0.08766 0.08394 -0.0136 0.7268 0.0068
-6.000 -0.1697 0.08535 0.08156 -0.0162 0.7128 0.0069
-5.750 -0.1501 0.08319 0.07932 -0.0193 0.6997 0.0069
-5.500 -0.1307 0.08068 0.07669 -0.0222 0.6879 0.0070
-5.250 -0.1076 0.07847 0.07440 -0.0258 0.6773 0.0070
-5.000 -0.0838 0.07612 0.07197 -0.0294 0.6671 0.0071
-4.750 -0.0590 0.07368 0.06946 -0.0329 0.6576 0.0071
-4.500 -0.0328 0.07119 0.06689 -0.0366 0.6485 0.0071
-2.500 0.1827 0.04678 0.04190 -0.0606 0.5927 0.0086
-2.250 0.2139 0.04441 0.03941 -0.0635 0.5872 0.0102
-2.000 0.2482 0.04207 0.03699 -0.0667 0.5810 0.0110
-1.750 0.2824 0.03992 0.03470 -0.0696 0.5756 0.0121
-1.500 0.3172 0.03792 0.03259 -0.0722 0.5700 0.0132
-1.250 0.3526 0.03613 0.03062 -0.0747 0.5645 0.0140
-1.000 0.3881 0.03452 0.02884 -0.0769 0.5596 0.0142
-0.750 0.4235 0.03293 0.02713 -0.0789 0.5540 0.0144
-0.500 0.4581 0.03141 0.02544 -0.0805 0.5491 0.0145
-0.250 0.4919 0.02990 0.02378 -0.0819 0.5444 0.0145
0.000 0.5249 0.02837 0.02210 -0.0831 0.5391 0.0146
0.250 0.5570 0.02688 0.02043 -0.0839 0.5346 0.0146
0.500 0.5886 0.02539 0.01879 -0.0847 0.5300 0.0146
0.750 0.6194 0.02391 0.01717 -0.0853 0.5251 0.0146
1.000 0.6471 0.02218 0.01528 -0.0860 0.5209 0.0131
1.250 0.6785 0.02087 0.01373 -0.0863 0.5162 0.0102
1.500 0.7102 0.01964 0.01230 -0.0864 0.5115 0.0090
1.750 0.7414 0.01842 0.01081 -0.0862 0.5076 0.0083
2.000 0.7721 0.01717 0.00931 -0.0860 0.5032 0.0079
2.250 0.8019 0.01600 0.00791 -0.0857 0.4985 0.0079
2.500 0.8306 0.01502 0.00665 -0.0854 0.4945 0.0084
3.000 0.8857 0.01391 0.00519 -0.0848 0.4853 0.0235
3.250 0.9126 0.01354 0.00470 -0.0846 0.4810 0.0331
3.500 0.9395 0.01332 0.00441 -0.0843 0.4766 0.0324
3.750 0.9666 0.01328 0.00435 -0.0841 0.4713 0.0324
4.000 0.9935 0.01332 0.00429 -0.0839 0.4669 0.0333
4.250 1.0204 0.01340 0.00434 -0.0838 0.4620 0.0361
4.500 1.0472 0.01346 0.00449 -0.0836 0.4565 0.0813
4.750 1.0737 0.01357 0.00464 -0.0834 0.4519 0.1672
5.000 1.1091 0.01280 0.00491 -0.0855 0.4447 1.0000
5.250 1.1350 0.01298 0.00512 -0.0852 0.4387 1.0000
5.500 1.1612 0.01314 0.00533 -0.0850 0.4309 1.0000
5.750 1.1870 0.01333 0.00549 -0.0847 0.4243 1.0000
6.000 1.2129 0.01351 0.00574 -0.0845 0.4161 1.0000
6.250 1.2386 0.01371 0.00597 -0.0842 0.4084 1.0000
6.500 1.2643 0.01390 0.00622 -0.0840 0.3998 1.0000
6.750 1.2872 0.01412 0.00614 -0.0835 0.3380 1.0000
7.000 1.2837 0.01959 0.00919 -0.0827 0.0176 1.0000
7.250 1.3053 0.02036 0.01020 -0.0821 0.0055 1.0000
7.500 1.3271 0.02108 0.01103 -0.0815 0.0040 1.0000
7.750 1.3485 0.02185 0.01202 -0.0808 0.0032 1.0000
8.000 1.3687 0.02274 0.01315 -0.0800 0.0029 1.0000
8.250 1.3876 0.02377 0.01442 -0.0791 0.0027 1.0000
8.500 1.4044 0.02500 0.01593 -0.0780 0.0026 1.0000
8.750 1.4189 0.02640 0.01759 -0.0768 0.0026 1.0000
9.000 1.4305 0.02799 0.01944 -0.0753 0.0025 1.0000
9.250 1.4381 0.02981 0.02150 -0.0737 0.0025 1.0000
9.500 1.4409 0.03178 0.02368 -0.0717 0.0025 1.0000
9.750 1.4375 0.03404 0.02612 -0.0696 0.0026 1.0000
10.000 1.4339 0.03675 0.02900 -0.0683 0.0026 1.0000
10.250 1.4301 0.03982 0.03223 -0.0676 0.0026 1.0000
10.500 1.4278 0.04291 0.03548 -0.0671 0.0026 1.0000
10.750 1.4245 0.04624 0.03896 -0.0667 0.0026 1.0000
11.000 1.4227 0.04946 0.04232 -0.0662 0.0027 1.0000
11.250 1.4211 0.05272 0.04574 -0.0656 0.0027 1.0000
11.500 1.4201 0.05598 0.04915 -0.0648 0.0028 1.0000
11.750 1.4205 0.05922 0.05275 -0.0641 0.0029 1.0000
12.000 1.4205 0.06259 0.05632 -0.0633 0.0029 1.0000
12.250 1.4200 0.06616 0.06011 -0.0627 0.0030 1.0000
12.500 1.4181 0.07006 0.06425 -0.0624 0.0031 1.0000
12.750 1.4141 0.07438 0.06882 -0.0624 0.0032 1.0000
13.000 1.4082 0.07906 0.07374 -0.0630 0.0032 1.0000
13.250 1.4002 0.08416 0.07908 -0.0639 0.0033 1.0000
13.500 1.3904 0.08963 0.08478 -0.0654 0.0034 1.0000
13.750 1.3793 0.09555 0.09094 -0.0675 0.0034 1.0000
14.000 1.3676 0.10181 0.09742 -0.0700 0.0035 1.0000
14.250 1.3552 0.10848 0.10431 -0.0730 0.0035 1.0000
14.500 1.3427 0.11547 0.11150 -0.0765 0.0036 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 379 AIRFOIL (e379-il)