EPPLER 554 AIRFOIL (e554-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 554 AIRFOIL (e554-il) Reynolds number: 200,000 Max Cl/Cd: 68.98 at α=10.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e554-il-200000.txt Download as CSV file: xf-e554-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 554 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.750 -0.5175 0.11353 0.10966 -0.0477 1.0000 0.0282
-14.500 -0.5271 0.10718 0.10329 -0.0506 1.0000 0.0280
-14.250 -0.6387 0.08352 0.07903 -0.0661 1.0000 0.0241
-14.000 -0.6425 0.07944 0.07492 -0.0673 1.0000 0.0239
-13.750 -0.6616 0.07407 0.06941 -0.0700 1.0000 0.0238
-13.500 -0.6771 0.06925 0.06446 -0.0719 1.0000 0.0237
-13.250 -0.6935 0.06479 0.05984 -0.0736 1.0000 0.0237
-13.000 -0.7047 0.06097 0.05588 -0.0747 1.0000 0.0236
-12.750 -0.7182 0.05708 0.05183 -0.0755 1.0000 0.0235
-12.500 -0.7313 0.05356 0.04812 -0.0759 1.0000 0.0235
-12.250 -0.7401 0.05057 0.04495 -0.0758 1.0000 0.0235
-12.000 -0.7488 0.04768 0.04188 -0.0752 1.0000 0.0235
-11.750 -0.7588 0.04512 0.03911 -0.0741 1.0000 0.0237
-11.500 -0.7677 0.04297 0.03678 -0.0724 1.0000 0.0238
-11.250 -0.7692 0.04105 0.03471 -0.0704 1.0000 0.0238
-11.000 -0.7751 0.03939 0.03291 -0.0676 1.0000 0.0239
-10.750 -0.7800 0.03784 0.03119 -0.0645 1.0000 0.0241
-10.500 -0.7720 0.03611 0.02936 -0.0623 1.0000 0.0242
-10.250 -0.7560 0.03460 0.02789 -0.0604 1.0000 0.0247
-10.000 -0.7537 0.03358 0.02690 -0.0574 1.0000 0.0252
-9.750 -0.7472 0.03260 0.02592 -0.0553 0.9989 0.0257
-9.500 -0.7060 0.03104 0.02428 -0.0582 0.9940 0.0270
-9.250 -0.6678 0.02951 0.02267 -0.0603 0.9872 0.0286
-9.000 -0.6314 0.02799 0.02124 -0.0635 0.9766 0.0312
-8.750 -0.5982 0.02631 0.01958 -0.0665 0.9629 0.0343
-8.500 -0.5652 0.02453 0.01778 -0.0697 0.9484 0.0386
-8.250 -0.5321 0.02249 0.01576 -0.0736 0.9337 0.0448
-8.000 -0.4931 0.02049 0.01379 -0.0788 0.9191 0.0571
-7.750 -0.4552 0.01829 0.01170 -0.0840 0.9013 0.0817
-7.500 -0.4224 0.01595 0.00969 -0.0885 0.8803 0.1460
-7.250 -0.4006 0.01359 0.00825 -0.0910 0.8575 0.3212
-7.000 -0.3700 0.01375 0.00842 -0.0915 0.8363 0.3884
-6.750 -0.3403 0.01413 0.00858 -0.0915 0.8164 0.4149
-6.500 -0.3118 0.01457 0.00884 -0.0912 0.7980 0.4317
-6.250 -0.2841 0.01496 0.00899 -0.0909 0.7811 0.4456
-6.000 -0.2560 0.01553 0.00952 -0.0901 0.7659 0.4555
-5.750 -0.2284 0.01593 0.00975 -0.0896 0.7521 0.4663
-5.500 -0.2014 0.01639 0.01013 -0.0888 0.7379 0.4754
-5.250 -0.1746 0.01660 0.01021 -0.0883 0.7248 0.4832
-5.000 -0.1471 0.01705 0.01059 -0.0876 0.7131 0.4903
-4.750 -0.1204 0.01732 0.01071 -0.0871 0.7016 0.4998
-4.500 -0.0935 0.01783 0.01123 -0.0861 0.6904 0.5062
-4.250 -0.0658 0.01779 0.01096 -0.0862 0.6813 0.5128
-4.000 -0.0392 0.01754 0.01060 -0.0863 0.6711 0.5169
-3.750 -0.0115 0.01754 0.01049 -0.0861 0.6631 0.5197
-3.500 0.0154 0.01746 0.01035 -0.0860 0.6537 0.5229
-3.250 0.0436 0.01737 0.01008 -0.0862 0.6464 0.5266
-3.000 0.0711 0.01715 0.00973 -0.0866 0.6377 0.5311
-2.750 0.0993 0.01695 0.00938 -0.0869 0.6308 0.5341
-2.500 0.1267 0.01689 0.00929 -0.0869 0.6239 0.5366
-2.250 0.1543 0.01686 0.00921 -0.0868 0.6172 0.5397
-2.000 0.1831 0.01684 0.00905 -0.0871 0.6117 0.5430
-1.750 0.2108 0.01673 0.00889 -0.0874 0.6050 0.5465
-1.500 0.2398 0.01663 0.00863 -0.0880 0.5989 0.5505
-1.250 0.2682 0.01660 0.00853 -0.0881 0.5941 0.5532
-1.000 0.2953 0.01659 0.00856 -0.0881 0.5885 0.5560
-0.750 0.3233 0.01658 0.00853 -0.0882 0.5833 0.5592
-0.500 0.3523 0.01659 0.00843 -0.0886 0.5788 0.5627
-0.250 0.3812 0.01662 0.00837 -0.0891 0.5742 0.5666
0.000 0.4088 0.01658 0.00834 -0.0892 0.5689 0.5700
0.250 0.4366 0.01661 0.00838 -0.0892 0.5645 0.5731
0.500 0.4655 0.01669 0.00841 -0.0895 0.5608 0.5766
0.750 0.4938 0.01679 0.00849 -0.0898 0.5569 0.5803
1.000 0.5216 0.01684 0.00853 -0.0900 0.5522 0.5844
1.250 0.5498 0.01686 0.00854 -0.0903 0.5479 0.5879
1.500 0.5781 0.01694 0.00863 -0.0904 0.5444 0.5913
1.750 0.6071 0.01713 0.00879 -0.0907 0.5412 0.5955
2.000 0.6336 0.01723 0.00896 -0.0907 0.5369 0.6002
2.250 0.6616 0.01733 0.00906 -0.0910 0.5327 0.6046
2.500 0.6893 0.01740 0.00918 -0.0910 0.5291 0.6080
2.750 0.7183 0.01756 0.00931 -0.0913 0.5260 0.6122
3.000 0.7459 0.01777 0.00956 -0.0915 0.5225 0.6171
3.250 0.7724 0.01792 0.00978 -0.0915 0.5183 0.6224
3.500 0.7989 0.01805 0.01001 -0.0914 0.5145 0.6267
3.750 0.8271 0.01819 0.01018 -0.0915 0.5111 0.6319
4.000 0.8572 0.01841 0.01032 -0.0921 0.5080 0.6377
4.250 0.8824 0.01862 0.01068 -0.0919 0.5043 0.6425
4.500 0.9075 0.01882 0.01101 -0.0915 0.5001 0.6478
4.750 0.9348 0.01899 0.01123 -0.0916 0.4962 0.6546
5.000 0.9632 0.01913 0.01140 -0.0918 0.4929 0.6606
5.250 0.9918 0.01941 0.01169 -0.0921 0.4896 0.6674
5.500 1.0151 0.01966 0.01210 -0.0915 0.4850 0.6748
5.750 1.0400 0.01981 0.01240 -0.0911 0.4805 0.6811
6.000 1.0683 0.01992 0.01254 -0.0913 0.4766 0.6897
6.250 1.0978 0.02013 0.01274 -0.0917 0.4730 0.6977
6.500 1.1185 0.02037 0.01321 -0.0906 0.4676 0.7073
6.750 1.1431 0.02046 0.01343 -0.0901 0.4626 0.7163
7.000 1.1723 0.02052 0.01349 -0.0904 0.4584 0.7275
7.250 1.1961 0.02074 0.01386 -0.0898 0.4537 0.7383
7.500 1.2172 0.02089 0.01421 -0.0887 0.4479 0.7506
7.750 1.2437 0.02087 0.01425 -0.0885 0.4430 0.7651
8.000 1.2696 0.02097 0.01442 -0.0881 0.4380 0.7813
8.250 1.2870 0.02107 0.01477 -0.0864 0.4314 0.7993
8.500 1.3112 0.02094 0.01473 -0.0855 0.4259 0.8211
8.750 1.3309 0.02098 0.01492 -0.0840 0.4202 0.8495
9.000 1.3415 0.02087 0.01507 -0.0806 0.4137 0.9000
9.250 1.3714 0.02063 0.01481 -0.0810 0.4074 1.0000
9.500 1.3902 0.02088 0.01521 -0.0800 0.3992 1.0000
9.750 1.4157 0.02084 0.01514 -0.0797 0.3918 1.0000
10.000 1.4322 0.02107 0.01550 -0.0781 0.3831 1.0000
10.250 1.4540 0.02108 0.01547 -0.0772 0.3749 1.0000
10.500 1.4655 0.02135 0.01592 -0.0748 0.3649 1.0000
10.750 1.4801 0.02153 0.01612 -0.0728 0.3552 1.0000
11.000 1.4883 0.02179 0.01642 -0.0698 0.3444 1.0000
11.250 1.4932 0.02223 0.01696 -0.0664 0.3325 1.0000
11.500 1.4973 0.02281 0.01757 -0.0632 0.3199 1.0000
11.750 1.4991 0.02359 0.01836 -0.0599 0.3064 1.0000
12.000 1.4983 0.02464 0.01941 -0.0568 0.2921 1.0000
12.250 1.4949 0.02603 0.02078 -0.0538 0.2772 1.0000
12.500 1.4890 0.02778 0.02252 -0.0511 0.2621 1.0000
12.750 1.4812 0.02992 0.02463 -0.0488 0.2473 1.0000
13.000 1.4717 0.03246 0.02715 -0.0468 0.2328 1.0000
13.250 1.4610 0.03535 0.03000 -0.0453 0.2192 1.0000
13.500 1.4491 0.03858 0.03320 -0.0442 0.2059 1.0000
13.750 1.4382 0.04196 0.03659 -0.0435 0.1932 1.0000
14.000 1.4271 0.04553 0.04017 -0.0431 0.1813 1.0000
14.250 1.4158 0.04928 0.04392 -0.0429 0.1702 1.0000
14.500 1.4036 0.05327 0.04787 -0.0429 0.1602 1.0000
14.750 1.3936 0.05724 0.05185 -0.0432 0.1498 1.0000
15.000 1.3851 0.06119 0.05583 -0.0436 0.1401 1.0000
15.250 1.3755 0.06536 0.05997 -0.0442 0.1313 1.0000
15.500 1.3671 0.06955 0.06416 -0.0450 0.1227 1.0000
15.750 1.3608 0.07361 0.06826 -0.0458 0.1143 1.0000
16.000 1.3524 0.07795 0.07254 -0.0468 0.1072 1.0000
16.250 1.3473 0.08206 0.07672 -0.0479 0.0995 1.0000
16.500 1.3416 0.08624 0.08089 -0.0490 0.0929 1.0000
16.750 1.3363 0.09050 0.08517 -0.0502 0.0867 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 554 AIRFOIL (e554-il)