EPPLER 521 AIRFOIL (e521-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 521 AIRFOIL (e521-il) Reynolds number: 100,000 Max Cl/Cd: 34.71 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e521-il-100000-n5.txt Download as CSV file: xf-e521-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 521 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.500 -0.7603 0.12473 0.11976 -0.0025 1.0000 0.0222
-15.250 -0.7900 0.11242 0.10735 -0.0098 1.0000 0.0216
-15.000 -0.8193 0.10114 0.09590 -0.0168 1.0000 0.0213
-14.750 -0.8417 0.09221 0.08680 -0.0225 1.0000 0.0210
-14.500 -0.8617 0.08442 0.07881 -0.0272 1.0000 0.0210
-14.250 -0.8796 0.07761 0.07178 -0.0310 1.0000 0.0210
-14.000 -0.8951 0.07171 0.06564 -0.0339 1.0000 0.0213
-13.750 -0.9074 0.06661 0.06031 -0.0360 1.0000 0.0215
-13.500 -0.9168 0.06216 0.05562 -0.0373 1.0000 0.0219
-13.250 -0.9242 0.05820 0.05138 -0.0382 1.0000 0.0225
-13.000 -0.9299 0.05465 0.04770 -0.0386 1.0000 0.0235
-12.750 -0.9315 0.05183 0.04480 -0.0388 1.0000 0.0244
-12.500 -0.9321 0.04926 0.04213 -0.0387 1.0000 0.0255
-12.250 -0.9314 0.04681 0.03954 -0.0384 1.0000 0.0270
-12.000 -0.9285 0.04441 0.03693 -0.0377 1.0000 0.0287
-11.750 -0.9226 0.04224 0.03449 -0.0369 1.0000 0.0305
-11.500 -0.9192 0.04003 0.03226 -0.0361 1.0000 0.0324
-11.250 -0.9138 0.03825 0.03042 -0.0352 1.0000 0.0346
-11.000 -0.9060 0.03658 0.02861 -0.0342 1.0000 0.0375
-10.750 -0.8955 0.03501 0.02684 -0.0331 1.0000 0.0406
-10.500 -0.8901 0.03332 0.02519 -0.0319 1.0000 0.0440
-10.250 -0.8804 0.03194 0.02368 -0.0306 1.0000 0.0480
-10.000 -0.8700 0.03057 0.02219 -0.0293 1.0000 0.0524
-9.750 -0.8620 0.02927 0.02091 -0.0278 1.0000 0.0575
-9.500 -0.8506 0.02816 0.01970 -0.0263 1.0000 0.0640
-9.250 -0.8431 0.02700 0.01858 -0.0243 1.0000 0.0705
-9.000 -0.8334 0.02605 0.01761 -0.0224 1.0000 0.0787
-8.750 -0.8229 0.02521 0.01672 -0.0205 1.0000 0.0890
-8.500 -0.8166 0.02441 0.01600 -0.0179 1.0000 0.0989
-8.250 -0.8163 0.02385 0.01548 -0.0141 0.9995 0.1085
-8.000 -0.7907 0.02294 0.01452 -0.0151 0.9865 0.1277
-7.750 -0.7646 0.02205 0.01369 -0.0161 0.9774 0.1501
-7.500 -0.7375 0.02125 0.01292 -0.0172 0.9697 0.1749
-7.250 -0.7119 0.02052 0.01223 -0.0179 0.9621 0.2020
-7.000 -0.6856 0.01982 0.01160 -0.0186 0.9558 0.2318
-6.750 -0.6616 0.01921 0.01103 -0.0187 0.9491 0.2620
-6.500 -0.6372 0.01863 0.01054 -0.0189 0.9430 0.2948
-6.250 -0.6128 0.01812 0.01009 -0.0189 0.9374 0.3292
-6.000 -0.5912 0.01767 0.00974 -0.0182 0.9311 0.3638
-5.750 -0.5668 0.01725 0.00941 -0.0180 0.9263 0.4013
-5.500 -0.5456 0.01691 0.00917 -0.0171 0.9205 0.4380
-5.250 -0.5228 0.01664 0.00900 -0.0163 0.9153 0.4759
-5.000 -0.4974 0.01646 0.00893 -0.0158 0.9113 0.5136
-4.750 -0.4748 0.01641 0.00896 -0.0147 0.9059 0.5481
-4.500 -0.4493 0.01645 0.00905 -0.0140 0.9013 0.5799
-4.250 -0.4220 0.01657 0.00916 -0.0135 0.8975 0.6079
-4.000 -0.3966 0.01677 0.00935 -0.0127 0.8929 0.6304
-3.750 -0.3724 0.01693 0.00945 -0.0119 0.8878 0.6515
-3.500 -0.3452 0.01714 0.00959 -0.0114 0.8839 0.6687
-3.250 -0.3200 0.01734 0.00973 -0.0106 0.8794 0.6841
-3.000 -0.2964 0.01754 0.00986 -0.0097 0.8740 0.6983
-2.750 -0.2707 0.01771 0.00994 -0.0090 0.8697 0.7113
-2.500 -0.2467 0.01786 0.01002 -0.0082 0.8649 0.7238
-2.250 -0.2227 0.01806 0.01018 -0.0072 0.8593 0.7344
-2.000 -0.1956 0.01823 0.01031 -0.0067 0.8549 0.7436
-1.750 -0.1725 0.01834 0.01036 -0.0057 0.8494 0.7539
-1.500 -0.1481 0.01847 0.01046 -0.0049 0.8436 0.7627
-1.250 -0.1211 0.01855 0.01050 -0.0043 0.8393 0.7705
-1.000 -0.1006 0.01860 0.01051 -0.0031 0.8322 0.7804
-0.750 -0.0728 0.01871 0.01062 -0.0026 0.8269 0.7861
-0.500 -0.0508 0.01870 0.01057 -0.0015 0.8207 0.7953
-0.250 -0.0243 0.01879 0.01067 -0.0009 0.8141 0.8006
0.000 0.0000 0.01872 0.01057 0.0000 0.8087 0.8088
0.500 0.0508 0.01870 0.01057 0.0015 0.7953 0.8207
0.750 0.0728 0.01871 0.01061 0.0025 0.7861 0.8269
1.000 0.1007 0.01860 0.01051 0.0031 0.7804 0.8322
1.250 0.1211 0.01855 0.01050 0.0043 0.7705 0.8393
1.500 0.1481 0.01847 0.01046 0.0049 0.7627 0.8436
1.750 0.1726 0.01834 0.01036 0.0057 0.7539 0.8494
2.000 0.1957 0.01823 0.01031 0.0067 0.7436 0.8549
2.250 0.2227 0.01806 0.01018 0.0072 0.7344 0.8593
2.500 0.2468 0.01786 0.01003 0.0082 0.7238 0.8649
2.750 0.2707 0.01771 0.00994 0.0090 0.7113 0.8697
3.000 0.2964 0.01754 0.00986 0.0097 0.6982 0.8740
3.250 0.3201 0.01735 0.00973 0.0106 0.6842 0.8794
3.500 0.3452 0.01714 0.00959 0.0114 0.6687 0.8839
3.750 0.3724 0.01693 0.00945 0.0119 0.6515 0.8878
4.000 0.3966 0.01677 0.00936 0.0127 0.6304 0.8929
4.250 0.4219 0.01657 0.00916 0.0135 0.6079 0.8975
4.500 0.4492 0.01646 0.00905 0.0140 0.5799 0.9013
4.750 0.4748 0.01641 0.00897 0.0147 0.5483 0.9059
5.000 0.4973 0.01646 0.00893 0.0158 0.5136 0.9113
5.250 0.5228 0.01664 0.00900 0.0163 0.4759 0.9153
5.500 0.5455 0.01691 0.00917 0.0171 0.4379 0.9205
5.750 0.5668 0.01725 0.00941 0.0180 0.4014 0.9263
6.000 0.5911 0.01767 0.00974 0.0183 0.3636 0.9311
6.250 0.6128 0.01812 0.01009 0.0189 0.3294 0.9374
6.500 0.6373 0.01863 0.01053 0.0189 0.2952 0.9429
6.750 0.6616 0.01921 0.01103 0.0187 0.2620 0.9491
7.000 0.6855 0.01982 0.01160 0.0186 0.2317 0.9557
7.250 0.7119 0.02052 0.01223 0.0179 0.2022 0.9621
7.500 0.7375 0.02125 0.01292 0.0172 0.1751 0.9697
7.750 0.7645 0.02205 0.01368 0.0162 0.1499 0.9774
8.000 0.7906 0.02294 0.01452 0.0151 0.1278 0.9865
8.250 0.8162 0.02385 0.01548 0.0141 0.1087 0.9994
8.500 0.8166 0.02441 0.01600 0.0179 0.0988 1.0000
8.750 0.8228 0.02521 0.01672 0.0205 0.0889 1.0000
9.000 0.8333 0.02604 0.01759 0.0224 0.0786 1.0000
9.250 0.8429 0.02700 0.01857 0.0244 0.0704 1.0000
9.500 0.8505 0.02815 0.01969 0.0263 0.0640 1.0000
9.750 0.8619 0.02925 0.02089 0.0278 0.0576 1.0000
10.000 0.8701 0.03054 0.02217 0.0293 0.0524 1.0000
10.250 0.8803 0.03193 0.02367 0.0307 0.0481 1.0000
10.500 0.8899 0.03331 0.02517 0.0319 0.0439 1.0000
10.750 0.8955 0.03500 0.02684 0.0331 0.0407 1.0000
11.000 0.9059 0.03656 0.02860 0.0342 0.0375 1.0000
11.250 0.9138 0.03824 0.03041 0.0352 0.0346 1.0000
11.500 0.9190 0.04001 0.03223 0.0361 0.0323 1.0000
11.750 0.9225 0.04219 0.03446 0.0369 0.0303 1.0000
12.000 0.9284 0.04439 0.03692 0.0378 0.0286 1.0000
12.250 0.9318 0.04679 0.03952 0.0384 0.0271 1.0000
12.500 0.9327 0.04920 0.04207 0.0387 0.0256 1.0000
12.750 0.9320 0.05179 0.04476 0.0388 0.0244 1.0000
13.000 0.9302 0.05464 0.04768 0.0387 0.0235 1.0000
13.250 0.9248 0.05817 0.05134 0.0383 0.0226 1.0000
13.500 0.9176 0.06210 0.05555 0.0374 0.0219 1.0000
13.750 0.9072 0.06669 0.06040 0.0359 0.0214 1.0000
14.000 0.8949 0.07178 0.06572 0.0338 0.0211 1.0000
14.250 0.8805 0.07751 0.07168 0.0310 0.0210 1.0000
14.500 0.8621 0.08444 0.07882 0.0271 0.0209 1.0000
14.750 0.8421 0.09222 0.08681 0.0224 0.0210 1.0000
15.000 0.8192 0.10130 0.09607 0.0166 0.0213 1.0000
15.250 0.7888 0.11289 0.10783 0.0094 0.0216 1.0000
15.500 0.7612 0.12468 0.11971 0.0024 0.0221 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 521 AIRFOIL (e521-il)