EPPLER 554 AIRFOIL (e554-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: EPPLER 554 AIRFOIL (e554-il) Reynolds number: 50,000 Max Cl/Cd: 4.75 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e554-il-50000.txt Download as CSV file: xf-e554-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.4675 0.09661 0.08984 -0.0573 1.0000 0.1150
-11.500 -0.4631 0.09300 0.08627 -0.0572 1.0000 0.1168
-11.250 -0.6067 0.07541 0.06848 -0.0712 1.0000 0.1009
-11.000 -0.6236 0.07118 0.06422 -0.0714 1.0000 0.1002
-10.750 -0.6459 0.06754 0.06055 -0.0706 1.0000 0.0997
-10.500 -0.6679 0.06446 0.05742 -0.0689 1.0000 0.0992
-10.250 -0.6859 0.06129 0.05415 -0.0672 1.0000 0.0988
-10.000 -0.7012 0.05830 0.05101 -0.0651 1.0000 0.0986
-9.750 -0.7132 0.05552 0.04805 -0.0627 1.0000 0.0987
-9.500 -0.7234 0.05302 0.04533 -0.0600 1.0000 0.0993
-9.250 -0.7326 0.05082 0.04285 -0.0571 1.0000 0.1002
-9.000 -0.7338 0.04847 0.04042 -0.0545 1.0000 0.1025
-8.750 -0.7276 0.04673 0.03878 -0.0521 1.0000 0.1062
-8.500 -0.7273 0.04511 0.03703 -0.0494 1.0000 0.1096
-8.250 -0.7273 0.04351 0.03514 -0.0468 1.0000 0.1130
-8.000 -0.7187 0.04194 0.03369 -0.0445 1.0000 0.1188
-7.750 -0.7132 0.04068 0.03231 -0.0423 1.0000 0.1265
-7.500 -0.7043 0.03950 0.03134 -0.0398 1.0000 0.1353
-7.250 -0.6975 0.03832 0.03024 -0.0373 1.0000 0.1476
-7.000 -0.6922 0.03716 0.02923 -0.0349 1.0000 0.1627
-6.750 -0.6761 0.03530 0.02774 -0.0350 0.9957 0.1938
-6.500 -0.6450 0.03417 0.02882 -0.0365 0.9825 0.3504
-6.250 -0.5768 0.04863 0.04320 -0.0252 0.9641 0.4572
-6.000 -0.4698 0.06084 0.05491 -0.0163 0.9531 0.5071
-5.750 -0.3628 0.06499 0.05849 -0.0174 0.9451 0.5633
-5.500 -0.2198 0.06631 0.05922 -0.0242 0.9408 0.6537
-2.750 0.1105 0.05231 0.04343 -0.0536 0.8250 0.7779
-2.500 0.0900 0.05277 0.04391 -0.0484 0.8134 0.7713
-2.250 0.0674 0.05263 0.04373 -0.0436 0.8062 0.7635
-2.000 -0.0861 0.05482 0.04623 -0.0196 0.7944 0.7491
-1.750 -0.0620 0.05406 0.04535 -0.0212 0.7890 0.7522
-1.500 -0.1303 0.05439 0.04577 -0.0108 0.7834 0.7498
-1.250 -0.1468 0.05418 0.04554 -0.0074 0.7786 0.7512
-1.000 -0.1352 0.05379 0.04506 -0.0077 0.7735 0.7544
-0.750 -0.1229 0.05353 0.04470 -0.0084 0.7691 0.7580
-0.500 -0.1326 0.05357 0.04468 -0.0071 0.7664 0.7606
-0.250 -0.1300 0.05370 0.04478 -0.0060 0.7637 0.7633
0.000 -0.1209 0.05389 0.04492 -0.0058 0.7607 0.7671
0.250 -0.1065 0.05415 0.04511 -0.0066 0.7581 0.7715
0.500 -0.0899 0.05453 0.04540 -0.0084 0.7567 0.7756
0.750 -0.0716 0.05500 0.04581 -0.0100 0.7552 0.7795
1.000 -0.0648 0.05559 0.04638 -0.0097 0.7559 0.7837
1.250 -0.0647 0.05637 0.04715 -0.0094 0.7611 0.7881
1.500 -0.0485 0.05737 0.04810 -0.0115 0.7665 0.7930
1.750 -0.0233 0.05855 0.04923 -0.0139 0.7703 0.7977
2.000 -0.1629 0.05759 0.04860 -0.0003 0.9312 0.7969
2.250 -0.1399 0.05852 0.04948 -0.0020 0.9228 0.8022
2.500 -0.1027 0.06043 0.05130 -0.0064 0.9120 0.8093
2.750 -0.0839 0.06082 0.05166 -0.0075 0.8999 0.8150
3.000 -0.0545 0.06263 0.05342 -0.0102 0.8927 0.8219
3.250 -0.0262 0.06372 0.05447 -0.0132 0.8795 0.8286
3.500 -0.0097 0.06441 0.05515 -0.0134 0.8678 0.8351
3.750 0.0311 0.06734 0.05803 -0.0181 0.8593 0.8450
4.000 0.0403 0.06715 0.05787 -0.0172 0.8461 0.8522
4.250 0.0622 0.06853 0.05923 -0.0188 0.8360 0.8614
4.500 0.0940 0.07062 0.06135 -0.0214 0.8250 0.8723
4.750 0.1051 0.07102 0.06178 -0.0211 0.8125 0.8824
5.000 0.1366 0.07359 0.06439 -0.0238 0.8051 0.8974
5.250 0.1538 0.07425 0.06512 -0.0244 0.7908 0.9125
5.500 0.1695 0.07523 0.06619 -0.0253 0.7788 0.9302
5.750 0.2176 0.07876 0.06983 -0.0318 0.7697 0.9652
6.000 0.2388 0.07976 0.07085 -0.0345 0.7547 1.0000
6.250 0.2573 0.08132 0.07241 -0.0372 0.7419 1.0000
6.500 0.3037 0.08554 0.07660 -0.0439 0.7339 1.0000
6.750 0.3251 0.08710 0.07817 -0.0467 0.7194 1.0000
7.000 0.3416 0.08896 0.08003 -0.0490 0.7068 1.0000
7.250 0.3898 0.09374 0.08477 -0.0553 0.6988 1.0000
7.500 0.4022 0.09495 0.08598 -0.0567 0.6841 1.0000
7.750 0.4137 0.09691 0.08794 -0.0581 0.6720 1.0000
8.000 0.4622 0.10208 0.09308 -0.0633 0.6638 1.0000
8.250 0.4634 0.10275 0.09377 -0.0631 0.6492 1.0000
8.500 0.4718 0.10488 0.09590 -0.0639 0.6378 1.0000
8.750 0.5166 0.10991 0.10094 -0.0677 0.6289 1.0000
9.000 0.5122 0.11054 0.10158 -0.0670 0.6148 1.0000
9.250 0.5206 0.11304 0.10410 -0.0677 0.6046 1.0000
9.500 0.5585 0.11750 0.10861 -0.0703 0.5943 1.0000
9.750 0.5529 0.11851 0.10965 -0.0698 0.5812 1.0000
10.000 0.5662 0.12168 0.11285 -0.0709 0.5724 1.0000
10.250 0.5927 0.12517 0.11639 -0.0724 0.5603 1.0000
10.500 0.5893 0.12684 0.11811 -0.0724 0.5484 1.0000
10.750 0.6184 0.13152 0.12285 -0.0742 0.5407 1.0000
11.000 0.6225 0.13316 0.12455 -0.0744 0.5271 1.0000
11.250 0.6236 0.13563 0.12706 -0.0750 0.5173 1.0000
11.500 0.6635 0.14112 0.13266 -0.0769 0.5074 1.0000
11.750 0.6492 0.14158 0.13315 -0.0768 0.4953 1.0000
12.000 0.6646 0.14551 0.13715 -0.0780 0.4879 1.0000
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Polar data table (+)
Polar graphs
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