EPPLER 521 AIRFOIL (e521-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 521 AIRFOIL (e521-il) Reynolds number: 50,000 Max Cl/Cd: 29.23 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e521-il-50000.txt Download as CSV file: xf-e521-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7380 0.09183 0.08526 -0.0306 1.0000 0.1298
-11.750 -0.7569 0.08485 0.07827 -0.0335 1.0000 0.1289
-11.500 -0.7808 0.07831 0.07170 -0.0359 1.0000 0.1274
-11.250 -0.8093 0.07242 0.06572 -0.0373 1.0000 0.1253
-11.000 -0.8409 0.06728 0.06044 -0.0371 1.0000 0.1231
-10.750 -0.8695 0.06314 0.05611 -0.0351 1.0000 0.1221
-10.500 -0.8868 0.05947 0.05219 -0.0328 1.0000 0.1234
-10.250 -0.9005 0.05593 0.04833 -0.0304 1.0000 0.1259
-10.000 -0.9136 0.05244 0.04436 -0.0274 1.0000 0.1286
-9.750 -0.9016 0.04936 0.04125 -0.0262 1.0000 0.1358
-9.500 -0.9008 0.04614 0.03759 -0.0238 1.0000 0.1414
-9.250 -0.8890 0.04370 0.03509 -0.0221 1.0000 0.1525
-9.000 -0.8789 0.04122 0.03246 -0.0202 1.0000 0.1639
-8.750 -0.8677 0.03900 0.03013 -0.0181 1.0000 0.1779
-8.500 -0.8576 0.03718 0.02826 -0.0157 1.0000 0.1948
-8.250 -0.8496 0.03539 0.02637 -0.0128 1.0000 0.2125
-8.000 -0.8407 0.03395 0.02500 -0.0100 1.0000 0.2334
-7.750 -0.8343 0.03257 0.02361 -0.0066 1.0000 0.2561
-7.500 -0.8258 0.03137 0.02258 -0.0035 1.0000 0.2824
-7.250 -0.8176 0.03020 0.02154 -0.0003 1.0000 0.3111
-7.000 -0.8103 0.02912 0.02056 0.0031 1.0000 0.3433
-6.750 -0.7998 0.02815 0.01982 0.0062 1.0000 0.3791
-6.500 -0.7911 0.02725 0.01905 0.0096 1.0000 0.4182
-6.250 -0.7786 0.02664 0.01871 0.0129 1.0000 0.4611
-6.000 -0.7649 0.02635 0.01862 0.0165 1.0000 0.5062
-5.750 -0.7468 0.02677 0.01925 0.0202 1.0000 0.5507
-5.500 -0.7243 0.02799 0.02055 0.0240 1.0000 0.5908
-5.250 -0.6980 0.02973 0.02225 0.0278 1.0000 0.6250
-5.000 -0.6688 0.03164 0.02404 0.0313 1.0000 0.6551
-4.750 -0.6411 0.03324 0.02544 0.0345 1.0000 0.6833
-4.500 -0.6148 0.03446 0.02648 0.0373 1.0000 0.7107
-4.250 -0.5903 0.03530 0.02714 0.0399 1.0000 0.7380
-4.000 -0.5497 0.03635 0.02793 0.0402 1.0000 0.7639
-3.750 -0.4934 0.03722 0.02853 0.0376 1.0000 0.7894
-3.500 -0.4331 0.03751 0.02856 0.0332 1.0000 0.8155
-3.250 -0.3888 0.03733 0.02815 0.0305 1.0000 0.8409
-3.000 -0.3397 0.03702 0.02765 0.0266 1.0000 0.8653
-2.750 -0.2837 0.03652 0.02694 0.0211 1.0000 0.8883
-2.500 -0.2486 0.03598 0.02627 0.0186 1.0000 0.9112
-2.250 -0.1897 0.03521 0.02534 0.0117 1.0000 0.9323
-2.000 -0.1468 0.03456 0.02458 0.0070 1.0000 0.9531
-1.750 -0.0898 0.03366 0.02357 -0.0005 1.0000 0.9720
-1.500 -0.0324 0.03278 0.02261 -0.0084 1.0000 0.9907
-1.250 0.0013 0.03223 0.02201 -0.0123 1.0000 1.0000
-1.000 0.0006 0.03211 0.02188 -0.0098 1.0000 1.0000
-0.750 0.0002 0.03202 0.02177 -0.0074 1.0000 1.0000
-0.500 0.0000 0.03194 0.02169 -0.0049 1.0000 1.0000
-0.250 0.0000 0.03190 0.02164 -0.0025 1.0000 1.0000
0.000 0.0000 0.03189 0.02163 0.0000 1.0000 1.0000
0.250 0.0000 0.03190 0.02164 0.0024 1.0000 1.0000
0.500 0.0000 0.03194 0.02168 0.0049 1.0000 1.0000
0.750 -0.0002 0.03201 0.02177 0.0074 1.0000 1.0000
1.000 -0.0006 0.03210 0.02187 0.0098 1.0000 1.0000
1.250 -0.0013 0.03222 0.02200 0.0123 1.0000 1.0000
1.500 0.0326 0.03278 0.02260 0.0084 0.9906 1.0000
1.750 0.0896 0.03365 0.02357 0.0005 0.9721 1.0000
2.000 0.1468 0.03455 0.02457 -0.0070 0.9531 1.0000
2.250 0.1901 0.03519 0.02534 -0.0117 0.9322 1.0000
2.500 0.2482 0.03596 0.02625 -0.0185 0.9113 1.0000
2.750 0.2838 0.03650 0.02692 -0.0212 0.8883 1.0000
3.000 0.3399 0.03701 0.02763 -0.0267 0.8653 1.0000
3.250 0.3896 0.03731 0.02814 -0.0306 0.8408 1.0000
3.500 0.4335 0.03750 0.02854 -0.0333 0.8154 1.0000
3.750 0.4934 0.03721 0.02853 -0.0376 0.7894 1.0000
4.000 0.5498 0.03633 0.02792 -0.0402 0.7639 1.0000
4.250 0.5903 0.03530 0.02714 -0.0399 0.7380 1.0000
4.500 0.6144 0.03448 0.02649 -0.0373 0.7108 1.0000
4.750 0.6410 0.03325 0.02545 -0.0345 0.6834 1.0000
5.000 0.6688 0.03165 0.02405 -0.0313 0.6552 1.0000
5.250 0.6981 0.02973 0.02225 -0.0279 0.6250 1.0000
5.500 0.7244 0.02799 0.02056 -0.0241 0.5907 1.0000
5.750 0.7471 0.02676 0.01924 -0.0202 0.5506 1.0000
6.000 0.7651 0.02634 0.01861 -0.0165 0.5061 1.0000
6.250 0.7788 0.02664 0.01870 -0.0130 0.4611 1.0000
6.500 0.7914 0.02726 0.01905 -0.0096 0.4182 1.0000
6.750 0.8000 0.02815 0.01982 -0.0062 0.3790 1.0000
7.000 0.8104 0.02911 0.02055 -0.0031 0.3432 1.0000
7.250 0.8178 0.03020 0.02153 0.0003 0.3111 1.0000
7.500 0.8261 0.03138 0.02259 0.0035 0.2824 1.0000
7.750 0.8344 0.03257 0.02361 0.0066 0.2560 1.0000
8.000 0.8408 0.03394 0.02499 0.0099 0.2334 1.0000
8.250 0.8496 0.03539 0.02637 0.0128 0.2124 1.0000
8.500 0.8576 0.03718 0.02826 0.0157 0.1946 1.0000
8.750 0.8677 0.03901 0.03013 0.0181 0.1779 1.0000
9.000 0.8789 0.04121 0.03245 0.0202 0.1640 1.0000
9.250 0.8889 0.04368 0.03507 0.0222 0.1524 1.0000
9.500 0.9008 0.04613 0.03758 0.0238 0.1414 1.0000
9.750 0.9015 0.04934 0.04123 0.0263 0.1359 1.0000
10.000 0.9138 0.05241 0.04432 0.0274 0.1286 1.0000
10.250 0.9005 0.05589 0.04828 0.0304 0.1258 1.0000
10.500 0.8869 0.05941 0.05212 0.0329 0.1231 1.0000
10.750 0.8697 0.06312 0.05609 0.0351 0.1220 1.0000
11.000 0.8412 0.06724 0.06040 0.0371 0.1231 1.0000
11.250 0.8094 0.07240 0.06571 0.0373 0.1253 1.0000
11.500 0.7804 0.07833 0.07172 0.0358 0.1274 1.0000
11.750 0.7567 0.08489 0.07832 0.0334 0.1289 1.0000
12.000 0.7377 0.09189 0.08532 0.0305 0.1298 1.0000
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Polar data table (+)
Polar graphs
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