EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER E851 AIRFOIL (e851-il) Reynolds number: 200,000 Max Cl/Cd: 73.47 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e851-il-200000-n5.txt Download as CSV file: xf-e851-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER E851 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4702 0.08592 0.08244 -0.0391 1.0000 0.0061
-8.500 -0.4777 0.08229 0.07888 -0.0395 1.0000 0.0060
-8.250 -0.4886 0.07867 0.07533 -0.0395 1.0000 0.0059
-8.000 -0.5033 0.07541 0.07215 -0.0390 1.0000 0.0059
-7.750 -0.5129 0.07003 0.06682 -0.0433 0.9974 0.0058
-7.500 -0.5029 0.06063 0.05725 -0.0572 0.9898 0.0057
-7.250 -0.4915 0.05369 0.05006 -0.0647 0.9835 0.0055
-7.000 -0.4792 0.04753 0.04359 -0.0693 0.9775 0.0053
-6.750 -0.4633 0.04195 0.03763 -0.0723 0.9720 0.0051
-6.500 -0.4405 0.03657 0.03179 -0.0752 0.9688 0.0048
-6.250 -0.4235 0.03222 0.02696 -0.0753 0.9627 0.0046
-6.000 -0.3985 0.02799 0.02215 -0.0762 0.9593 0.0044
-5.750 -0.3701 0.02461 0.01822 -0.0770 0.9572 0.0043
-5.500 -0.3412 0.02198 0.01514 -0.0776 0.9553 0.0043
-5.250 -0.3185 0.02010 0.01297 -0.0767 0.9507 0.0043
-5.000 -0.2907 0.01845 0.01107 -0.0768 0.9480 0.0045
-4.750 -0.2613 0.01706 0.00949 -0.0774 0.9460 0.0048
-4.500 -0.2303 0.01588 0.00802 -0.0783 0.9443 0.0055
-4.250 -0.1995 0.01500 0.00691 -0.0792 0.9423 0.0069
-4.000 -0.1752 0.01435 0.00621 -0.0788 0.9378 0.0166
-3.750 -0.1465 0.01317 0.00549 -0.0798 0.9353 0.1021
-3.500 -0.1163 0.01227 0.00498 -0.0812 0.9334 0.2219
-3.250 -0.0863 0.01134 0.00481 -0.0825 0.9318 0.4303
-3.000 -0.0537 0.01116 0.00473 -0.0836 0.9305 0.4994
-2.750 -0.0295 0.01114 0.00468 -0.0830 0.9261 0.5280
-2.500 -0.0007 0.01109 0.00455 -0.0832 0.9231 0.5558
-2.250 0.0303 0.01104 0.00451 -0.0839 0.9210 0.5852
-2.000 0.0626 0.01097 0.00442 -0.0849 0.9193 0.6029
-1.750 0.0960 0.01091 0.00431 -0.0861 0.9179 0.6138
-1.500 0.1228 0.01090 0.00427 -0.0861 0.9142 0.6237
-1.250 0.1503 0.01088 0.00422 -0.0861 0.9106 0.6328
-1.000 0.1812 0.01083 0.00417 -0.0868 0.9081 0.6422
-0.750 0.2136 0.01077 0.00411 -0.0879 0.9062 0.6524
-0.500 0.2467 0.01071 0.00406 -0.0891 0.9045 0.6631
-0.250 0.2708 0.01074 0.00414 -0.0884 0.8996 0.6732
0.000 0.2998 0.01071 0.00417 -0.0888 0.8961 0.6843
0.250 0.3316 0.01065 0.00416 -0.0896 0.8935 0.6961
0.500 0.3649 0.01056 0.00415 -0.0908 0.8914 0.7088
0.750 0.3883 0.01059 0.00427 -0.0899 0.8853 0.7221
1.000 0.4188 0.01052 0.00430 -0.0904 0.8813 0.7367
1.250 0.4526 0.01039 0.00428 -0.0916 0.8781 0.7526
1.500 0.4760 0.01038 0.00446 -0.0906 0.8704 0.7700
1.750 0.5094 0.01019 0.00440 -0.0914 0.8652 0.7888
2.000 0.5340 0.01008 0.00443 -0.0904 0.8549 0.8104
2.250 0.5609 0.00989 0.00438 -0.0897 0.8428 0.8353
2.500 0.5895 0.00960 0.00421 -0.0891 0.8249 0.8644
2.750 0.6150 0.00936 0.00415 -0.0878 0.7992 0.9039
3.000 0.6488 0.00919 0.00411 -0.0886 0.7689 0.9902
3.250 0.6752 0.00919 0.00406 -0.0879 0.7217 1.0000
3.500 0.6932 0.00976 0.00390 -0.0852 0.5670 1.0000
3.750 0.6944 0.01130 0.00440 -0.0801 0.3729 1.0000
4.000 0.7028 0.01275 0.00503 -0.0770 0.2117 1.0000
4.250 0.7193 0.01377 0.00563 -0.0753 0.1208 1.0000
4.500 0.7391 0.01458 0.00624 -0.0741 0.0727 1.0000
4.750 0.7582 0.01551 0.00692 -0.0727 0.0318 1.0000
5.000 0.7773 0.01656 0.00788 -0.0711 0.0103 1.0000
5.250 0.7975 0.01751 0.00884 -0.0698 0.0051 1.0000
5.500 0.8177 0.01856 0.01019 -0.0683 0.0042 1.0000
5.750 0.8373 0.01985 0.01166 -0.0667 0.0037 1.0000
6.000 0.8579 0.02134 0.01332 -0.0653 0.0034 1.0000
6.250 0.8803 0.02314 0.01532 -0.0642 0.0032 1.0000
6.500 0.9043 0.02531 0.01775 -0.0632 0.0031 1.0000
6.750 0.9279 0.02788 0.02068 -0.0622 0.0030 1.0000
7.000 0.9487 0.03093 0.02416 -0.0606 0.0030 1.0000
7.250 0.9651 0.03447 0.02819 -0.0584 0.0031 1.0000
7.500 0.9767 0.03843 0.03266 -0.0555 0.0031 1.0000
7.750 0.9833 0.04277 0.03748 -0.0522 0.0032 1.0000
8.000 0.9859 0.04723 0.04235 -0.0487 0.0033 1.0000
8.250 0.9846 0.05174 0.04722 -0.0452 0.0034 1.0000
8.500 0.9798 0.05611 0.05190 -0.0418 0.0035 1.0000
8.750 0.9713 0.06026 0.05630 -0.0384 0.0036 1.0000
9.000 0.9574 0.06395 0.06017 -0.0347 0.0036 1.0000
9.250 0.9414 0.06769 0.06408 -0.0317 0.0037 1.0000
9.500 0.9237 0.07181 0.06836 -0.0299 0.0037 1.0000
9.750 0.9055 0.07644 0.07312 -0.0295 0.0038 1.0000
10.000 0.8867 0.08192 0.07872 -0.0309 0.0038 1.0000
10.250 0.8694 0.08845 0.08534 -0.0343 0.0038 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER E851 AIRFOIL (e851-il)