EPPLER 559 AIRFOIL (e559-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 559 AIRFOIL (e559-il) Reynolds number: 50,000 Max Cl/Cd: 30.49 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e559-il-50000-n5.txt Download as CSV file: xf-e559-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 559 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.3834 0.10686 0.09942 -0.0512 1.0000 0.0749
-11.250 -0.4110 0.09766 0.09033 -0.0549 1.0000 0.0763
-11.000 -0.4601 0.08494 0.07771 -0.0607 1.0000 0.0768
-10.750 -0.4210 0.09026 0.08310 -0.0559 1.0000 0.0806
-10.500 -0.4515 0.08206 0.07501 -0.0588 1.0000 0.0822
-10.250 -0.5114 0.07029 0.06328 -0.0640 1.0000 0.0817
-10.000 -0.5584 0.06255 0.05549 -0.0663 1.0000 0.0811
-9.750 -0.5930 0.05791 0.05082 -0.0661 1.0000 0.0810
-9.500 -0.6193 0.05412 0.04699 -0.0659 1.0000 0.0816
-9.250 -0.6362 0.05069 0.04348 -0.0661 1.0000 0.0833
-9.000 -0.6481 0.04757 0.04020 -0.0658 1.0000 0.0860
-8.750 -0.6528 0.04497 0.03749 -0.0653 1.0000 0.0898
-8.500 -0.6503 0.04341 0.03599 -0.0644 1.0000 0.0947
-8.250 -0.6474 0.04101 0.03339 -0.0645 1.0000 0.1019
-8.000 -0.6311 0.03935 0.03170 -0.0659 0.9973 0.1122
-7.750 -0.5958 0.03763 0.02990 -0.0708 0.9884 0.1296
-7.500 -0.5593 0.03694 0.02925 -0.0746 0.9792 0.1491
-7.250 -0.5223 0.03594 0.02801 -0.0789 0.9698 0.1722
-7.000 -0.4863 0.03593 0.02800 -0.0812 0.9600 0.1898
-6.500 -0.4187 0.03506 0.02684 -0.0852 0.9392 0.2208
-6.250 -0.3811 0.03456 0.02614 -0.0877 0.9303 0.2360
-6.000 -0.3475 0.03394 0.02529 -0.0897 0.9195 0.2505
-5.750 -0.3155 0.03397 0.02531 -0.0902 0.9088 0.2612
-5.500 -0.2756 0.03352 0.02470 -0.0927 0.9009 0.2740
-5.250 -0.2452 0.03292 0.02390 -0.0938 0.8892 0.2863
-5.000 -0.2095 0.03217 0.02287 -0.0959 0.8795 0.2996
-4.750 -0.1735 0.03199 0.02267 -0.0970 0.8706 0.3097
-4.500 -0.1413 0.03146 0.02198 -0.0981 0.8597 0.3205
-4.250 -0.0993 0.03075 0.02101 -0.1009 0.8526 0.3340
-4.000 -0.0698 0.03050 0.02072 -0.1011 0.8409 0.3430
-3.750 -0.0288 0.02994 0.01999 -0.1034 0.8334 0.3549
-3.500 0.0044 0.02938 0.01920 -0.1049 0.8221 0.3670
-3.250 0.0359 0.02916 0.01896 -0.1051 0.8117 0.3755
-3.000 0.0744 0.02860 0.01822 -0.1070 0.8030 0.3871
-2.750 0.1039 0.02834 0.01785 -0.1073 0.7911 0.3972
-2.500 0.1407 0.02793 0.01733 -0.1087 0.7824 0.4076
-2.250 0.1724 0.02759 0.01681 -0.1096 0.7710 0.4192
-2.000 0.2011 0.02744 0.01664 -0.1094 0.7598 0.4281
-1.750 0.2385 0.02705 0.01609 -0.1109 0.7512 0.4396
-1.500 0.2649 0.02696 0.01593 -0.1107 0.7388 0.4500
-1.250 0.2952 0.02680 0.01570 -0.1110 0.7285 0.4602
-1.000 0.3284 0.02659 0.01535 -0.1118 0.7188 0.4722
-0.750 0.3539 0.02659 0.01534 -0.1113 0.7073 0.4824
-0.500 0.3872 0.02641 0.01506 -0.1120 0.6984 0.4942
-0.250 0.4139 0.02644 0.01502 -0.1119 0.6872 0.5063
0.000 0.4413 0.02646 0.01501 -0.1117 0.6772 0.5179
0.250 0.4719 0.02639 0.01489 -0.1119 0.6680 0.5304
0.500 0.4966 0.02653 0.01501 -0.1114 0.6574 0.5431
0.750 0.5296 0.02644 0.01485 -0.1120 0.6496 0.5581
1.000 0.5513 0.02667 0.01512 -0.1111 0.6387 0.5716
1.250 0.5818 0.02665 0.01507 -0.1113 0.6309 0.5872
1.500 0.6040 0.02689 0.01538 -0.1103 0.6209 0.6028
1.750 0.6326 0.02694 0.01544 -0.1103 0.6132 0.6208
2.000 0.6550 0.02718 0.01576 -0.1093 0.6038 0.6398
2.250 0.6821 0.02725 0.01587 -0.1089 0.5964 0.6616
2.500 0.7019 0.02753 0.01629 -0.1075 0.5875 0.6845
2.750 0.7294 0.02752 0.01633 -0.1070 0.5810 0.7138
3.000 0.7439 0.02790 0.01690 -0.1047 0.5718 0.7462
3.250 0.7677 0.02779 0.01690 -0.1033 0.5659 0.7926
3.500 0.7784 0.02810 0.01745 -0.1002 0.5574 0.8672
3.750 0.8137 0.02819 0.01748 -0.1017 0.5504 1.0000
4.000 0.8371 0.02891 0.01814 -0.1017 0.5422 1.0000
4.250 0.8652 0.02941 0.01855 -0.1022 0.5352 1.0000
4.500 0.8921 0.02996 0.01903 -0.1024 0.5287 1.0000
4.750 0.9120 0.03076 0.01983 -0.1016 0.5208 1.0000
5.000 0.9455 0.03101 0.01999 -0.1025 0.5155 1.0000
5.250 0.9571 0.03218 0.02124 -0.1007 0.5072 1.0000
5.500 0.9842 0.03268 0.02170 -0.1007 0.5012 1.0000
5.750 1.0050 0.03348 0.02251 -0.1000 0.4947 1.0000
6.000 1.0207 0.03449 0.02360 -0.0988 0.4875 1.0000
6.250 1.0531 0.03476 0.02383 -0.0993 0.4826 1.0000
6.500 1.0575 0.03632 0.02551 -0.0968 0.4747 1.0000
6.750 1.0803 0.03700 0.02624 -0.0963 0.4687 1.0000
7.000 1.1076 0.03753 0.02676 -0.0962 0.4637 1.0000
7.250 1.1039 0.03950 0.02890 -0.0930 0.4556 1.0000
7.500 1.1343 0.03979 0.02920 -0.0932 0.4505 1.0000
7.750 1.1315 0.04176 0.03131 -0.0901 0.4433 1.0000
8.000 1.1392 0.04313 0.03276 -0.0880 0.4368 1.0000
8.250 1.1828 0.04281 0.03244 -0.0895 0.4326 1.0000
8.500 1.1357 0.04720 0.03702 -0.0828 0.4230 1.0000
8.750 1.1667 0.04732 0.03718 -0.0828 0.4182 1.0000
9.000 1.1215 0.05269 0.04269 -0.0784 0.4082 1.0000
9.250 1.1413 0.05348 0.04356 -0.0776 0.4027 1.0000
9.500 1.1949 0.05173 0.04185 -0.0784 0.3997 1.0000
9.750 1.1128 0.06135 0.05160 -0.0749 0.3859 1.0000
10.000 1.1623 0.05915 0.04948 -0.0744 0.3836 1.0000
10.250 1.0940 0.06951 0.05993 -0.0742 0.3688 1.0000
10.750 1.0550 0.08142 0.07199 -0.0755 0.3476 1.0000
11.250 1.0806 0.08456 0.07531 -0.0746 0.3356 1.0000
11.750 1.0313 0.09886 0.08975 -0.0779 0.3121 1.0000
12.000 1.0474 0.09998 0.09097 -0.0774 0.3064 1.0000
12.250 1.0786 0.09876 0.08987 -0.0760 0.3035 1.0000
12.750 1.0365 0.11258 0.10383 -0.0801 0.2809 1.0000
13.000 1.0521 0.11374 0.10509 -0.0797 0.2752 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 559 AIRFOIL (e559-il)