EPPLER 545 AIRFOIL (e545-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 545 AIRFOIL (e545-il) Reynolds number: 50,000 Max Cl/Cd: 19.16 at α=11.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e545-il-50000-n5.txt Download as CSV file: xf-e545-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 545 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
------ -------- --------- --------- -------- -------- --------
-13.000 -0.4304 0.11398 0.10669 -0.0572 1.0000 0.0366
-12.750 -0.4396 0.10731 0.10006 -0.0602 1.0000 0.0360
-12.500 -0.4569 0.09964 0.09242 -0.0640 1.0000 0.0352
-12.250 -0.5104 0.08760 0.08020 -0.0708 1.0000 0.0329
-12.000 -0.5285 0.08277 0.07532 -0.0721 1.0000 0.0327
-11.750 -0.5501 0.07813 0.07059 -0.0729 1.0000 0.0326
-11.500 -0.5682 0.07431 0.06668 -0.0728 1.0000 0.0324
-11.250 -0.5882 0.07068 0.06294 -0.0721 1.0000 0.0323
-11.000 -0.6074 0.06756 0.05970 -0.0706 1.0000 0.0323
-10.750 -0.6258 0.06482 0.05684 -0.0683 1.0000 0.0323
-10.500 -0.6454 0.06242 0.05431 -0.0653 1.0000 0.0323
-10.250 -0.6640 0.06060 0.05239 -0.0614 1.0000 0.0324
-10.000 -0.6867 0.05916 0.05085 -0.0564 1.0000 0.0325
-9.750 -0.7059 0.05780 0.04936 -0.0514 1.0000 0.0326
-9.500 -0.7233 0.05627 0.04765 -0.0466 1.0000 0.0328
-9.250 -0.7328 0.05492 0.04624 -0.0426 1.0000 0.0334
-9.000 -0.7397 0.05371 0.04498 -0.0389 1.0000 0.0343
-8.750 -0.7161 0.05082 0.04167 -0.0407 0.9925 0.0359
-8.500 -0.6886 0.04792 0.03830 -0.0420 0.9851 0.0378
-8.250 -0.6516 0.04568 0.03588 -0.0440 0.9798 0.0398
-8.000 -0.6088 0.04363 0.03356 -0.0455 0.9752 0.0436
-7.750 -0.5562 0.04221 0.03203 -0.0477 0.9726 0.0504
-7.500 -0.5054 0.04119 0.03085 -0.0492 0.9697 0.0600
-7.250 -0.4765 0.04007 0.02980 -0.0497 0.9615 0.0701
-7.000 -0.4500 0.03870 0.02848 -0.0507 0.9541 0.0865
-6.750 -0.4369 0.03727 0.02719 -0.0502 0.9429 0.1074
-6.500 -0.4313 0.03556 0.02578 -0.0491 0.9314 0.1370
-6.250 -0.4307 0.03332 0.02407 -0.0482 0.9205 0.1938
-6.000 -0.4021 0.03605 0.02899 -0.0407 0.9137 0.4933
-5.750 -0.4055 0.03502 0.02769 -0.0390 0.9011 0.5531
-5.500 -0.3872 0.03626 0.02862 -0.0364 0.8912 0.5983
-5.250 -0.3382 0.03879 0.03074 -0.0359 0.8876 0.6359
-5.000 -0.3192 0.03995 0.03166 -0.0327 0.8776 0.6599
-4.750 -0.2676 0.04134 0.03270 -0.0336 0.8738 0.6829
-4.500 -0.2252 0.04145 0.03247 -0.0355 0.8696 0.6978
-4.250 -0.1955 0.04145 0.03224 -0.0354 0.8612 0.7049
-4.000 -0.1710 0.04100 0.03156 -0.0356 0.8541 0.7137
-3.750 -0.1487 0.04057 0.03092 -0.0355 0.8470 0.7206
-3.500 -0.1247 0.04033 0.03050 -0.0352 0.8387 0.7261
-3.250 -0.1048 0.03971 0.02968 -0.0351 0.8328 0.7334
-3.000 -0.0984 0.03956 0.02942 -0.0323 0.8226 0.7386
-2.750 -0.0669 0.03917 0.02886 -0.0334 0.8170 0.7426
-2.500 -0.0481 0.03881 0.02835 -0.0328 0.8105 0.7478
-2.250 -0.0582 0.03852 0.02800 -0.0279 0.8001 0.7543
-2.000 -0.0209 0.03814 0.02744 -0.0299 0.7961 0.7569
-1.750 -0.0153 0.03815 0.02739 -0.0269 0.7870 0.7607
-1.500 0.0020 0.03787 0.02699 -0.0260 0.7806 0.7647
-1.250 0.0257 0.03737 0.02636 -0.0263 0.7763 0.7687
-1.000 0.0107 0.03747 0.02644 -0.0204 0.7653 0.7732
-0.750 0.0390 0.03723 0.02610 -0.0211 0.7606 0.7756
-0.500 0.0626 0.03704 0.02582 -0.0210 0.7557 0.7784
-0.250 0.0575 0.03719 0.02595 -0.0168 0.7462 0.7820
0.000 0.0806 0.03690 0.02557 -0.0169 0.7416 0.7852
0.250 0.0787 0.03694 0.02557 -0.0135 0.7337 0.7895
0.500 0.0933 0.03701 0.02561 -0.0121 0.7270 0.7917
0.750 0.1228 0.03684 0.02538 -0.0130 0.7231 0.7935
1.000 0.1282 0.03709 0.02562 -0.0104 0.7158 0.7966
1.250 0.1379 0.03727 0.02577 -0.0085 0.7086 0.7999
1.500 0.1670 0.03713 0.02559 -0.0094 0.7048 0.8027
2.000 0.1842 0.03776 0.02621 -0.0061 0.6903 0.8084
2.250 0.2152 0.03769 0.02612 -0.0070 0.6866 0.8104
2.750 0.2348 0.03870 0.02718 -0.0039 0.6719 0.8170
3.000 0.2665 0.03861 0.02708 -0.0050 0.6682 0.8198
3.500 0.2911 0.03985 0.02839 -0.0033 0.6532 0.8258
3.750 0.3244 0.03971 0.02828 -0.0043 0.6496 0.8280
4.000 0.3246 0.04091 0.02954 -0.0021 0.6395 0.8317
4.250 0.3508 0.04105 0.02972 -0.0025 0.6339 0.8350
4.500 0.3891 0.04066 0.02941 -0.0039 0.6305 0.8382
4.750 0.3868 0.04210 0.03090 -0.0019 0.6181 0.8426
5.000 0.4230 0.04163 0.03052 -0.0026 0.6137 0.8454
5.500 0.4633 0.04222 0.03129 -0.0018 0.5964 0.8527
6.000 0.5090 0.04240 0.03166 -0.0013 0.5786 0.8608
6.500 0.5533 0.04245 0.03195 -0.0002 0.5606 0.8706
7.000 0.5934 0.04298 0.03277 0.0008 0.5427 0.8817
7.250 0.5972 0.04445 0.03438 0.0020 0.5296 0.8886
7.500 0.6333 0.04367 0.03376 0.0016 0.5248 0.8943
7.750 0.6361 0.04531 0.03556 0.0027 0.5111 0.9024
8.000 0.6742 0.04438 0.03485 0.0022 0.5067 0.9102
8.250 0.6781 0.04609 0.03674 0.0029 0.4923 0.9215
8.750 0.7347 0.04631 0.03743 0.0012 0.4731 0.9513
9.000 0.7475 0.04742 0.03871 0.0007 0.4581 1.0000
9.500 0.8108 0.04628 0.03790 -0.0001 0.4377 1.0000
9.750 0.8230 0.04747 0.03920 -0.0002 0.4225 1.0000
10.000 0.8375 0.04848 0.04034 -0.0002 0.4078 1.0000
10.250 0.8560 0.04907 0.04107 -0.0002 0.3936 1.0000
10.500 0.8793 0.04910 0.04122 -0.0001 0.3796 1.0000
10.750 0.9053 0.04881 0.04105 0.0002 0.3652 1.0000
11.000 0.9277 0.04883 0.04116 0.0005 0.3494 1.0000
11.250 0.9453 0.04934 0.04172 0.0010 0.3322 1.0000
11.500 0.9599 0.05013 0.04255 0.0015 0.3141 1.0000
11.750 0.9719 0.05119 0.04364 0.0021 0.2956 1.0000
12.000 0.9817 0.05250 0.04494 0.0027 0.2767 1.0000
12.250 0.9891 0.05407 0.04649 0.0032 0.2578 1.0000
12.500 0.9947 0.05583 0.04820 0.0038 0.2391 1.0000
12.750 0.9985 0.05788 0.05019 0.0042 0.2213 1.0000
13.000 1.0002 0.06028 0.05255 0.0045 0.2044 1.0000
13.250 1.0008 0.06292 0.05518 0.0046 0.1884 1.0000
13.500 0.9989 0.06604 0.05837 0.0044 0.1730 1.0000
13.750 0.9963 0.06938 0.06182 0.0039 0.1589 1.0000
14.000 0.9948 0.07259 0.06507 0.0034 0.1465 1.0000
14.250 0.9948 0.07556 0.06797 0.0029 0.1359 1.0000
14.500 0.9965 0.07837 0.07065 0.0025 0.1262 1.0000
14.750 0.9973 0.08160 0.07390 0.0020 0.1167 1.0000
15.000 0.9979 0.08496 0.07731 0.0014 0.1081 1.0000
15.250 1.0003 0.08809 0.08044 0.0009 0.1003 1.0000
15.500 1.0002 0.09169 0.08417 -0.0001 0.0935 1.0000
15.750 0.9978 0.09583 0.08849 -0.0014 0.0876 1.0000
16.000 1.0018 0.09874 0.09136 -0.0021 0.0820 1.0000
16.250 0.9937 0.10410 0.09702 -0.0042 0.0779 1.0000
16.500 0.9988 0.10689 0.09977 -0.0051 0.0733 1.0000
16.750 0.9857 0.11340 0.10661 -0.0081 0.0706 1.0000
17.000 0.9819 0.11810 0.11143 -0.0103 0.0672 1.0000
17.250 0.9767 0.12318 0.11662 -0.0128 0.0645 1.0000
17.500 0.9558 0.13196 0.12567 -0.0177 0.0635 1.0000
17.750 0.9267 0.14328 0.13719 -0.0242 0.0633 1.0000
18.000 0.8864 0.15894 0.15291 -0.0333 0.0636 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 545 AIRFOIL (e545-il)