EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: EPPLER 59 AIRFOIL (e59-il) Reynolds number: 500,000 Max Cl/Cd: 155.56 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e59-il-500000.txt Download as CSV file: xf-e59-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 59 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4177 0.11875 0.11650 -0.0158 1.0000 0.0132
-9.000 -0.4166 0.11636 0.11414 -0.0154 1.0000 0.0134
-8.750 -0.4160 0.11399 0.11178 -0.0149 1.0000 0.0137
-8.500 -0.4054 0.11060 0.10840 -0.0175 0.9989 0.0142
-8.250 -0.3926 0.10707 0.10487 -0.0207 0.9972 0.0144
-8.000 -0.3795 0.10341 0.10121 -0.0243 0.9955 0.0152
-7.750 -0.3669 0.09984 0.09764 -0.0277 0.9932 0.0157
-7.500 -0.3554 0.09643 0.09423 -0.0320 0.9896 0.0165
-7.250 -0.3434 0.09313 0.09094 -0.0383 0.9858 0.0169
-7.000 -0.3284 0.08939 0.08720 -0.0427 0.9819 0.0170
-6.750 -0.3113 0.08507 0.08289 -0.0478 0.9772 0.0170
-6.500 -0.2853 0.07994 0.07774 -0.0556 0.9743 0.0171
-6.250 -0.2701 0.07439 0.07219 -0.0597 0.9729 0.0179
-6.000 -0.2591 0.07192 0.06973 -0.0606 0.9681 0.0185
-5.750 -0.2370 0.06835 0.06615 -0.0657 0.9643 0.0195
-5.500 -0.2018 0.06338 0.06113 -0.0753 0.9620 0.0215
-5.250 -0.1376 0.05627 0.05390 -0.0941 0.9604 0.0236
-5.000 -0.0887 0.04933 0.04683 -0.1066 0.9596 0.0237
-4.750 -0.0605 0.04019 0.03751 -0.1162 0.9540 0.0247
-4.500 -0.0255 0.03682 0.03403 -0.1213 0.9522 0.0256
-4.250 0.0171 0.03282 0.02982 -0.1278 0.9511 0.0272
-4.000 0.1163 0.01845 0.01381 -0.1478 0.9596 0.0229
-3.750 0.1575 0.01571 0.01050 -0.1513 0.9597 0.0247
-3.500 0.1843 0.01498 0.00969 -0.1514 0.9564 0.0270
-3.250 0.2142 0.01453 0.00915 -0.1518 0.9537 0.0298
-3.000 0.2460 0.01443 0.00895 -0.1526 0.9516 0.0327
-2.750 0.2822 0.01313 0.00752 -0.1546 0.9512 0.0372
-2.500 0.3165 0.01287 0.00723 -0.1560 0.9498 0.0418
-2.250 0.3517 0.01263 0.00694 -0.1575 0.9487 0.0445
-2.000 0.3881 0.01193 0.00623 -0.1595 0.9480 0.0505
-1.750 0.4238 0.01165 0.00594 -0.1612 0.9472 0.0548
-1.500 0.4607 0.01131 0.00558 -0.1631 0.9465 0.0609
-1.250 0.4837 0.01110 0.00538 -0.1620 0.9405 0.0699
-1.000 0.5217 0.01057 0.00499 -0.1641 0.9381 0.1076
-0.750 0.5606 0.01001 0.00479 -0.1666 0.9366 0.2283
-0.500 0.5998 0.00939 0.00469 -0.1693 0.9355 0.4219
-0.250 0.6394 0.00881 0.00457 -0.1719 0.9345 0.5986
0.000 0.6786 0.00789 0.00431 -0.1739 0.9334 0.8373
0.250 0.6979 0.00733 0.00390 -0.1712 0.9243 1.0000
0.500 0.7441 0.00680 0.00333 -0.1747 0.9208 1.0000
0.750 0.7723 0.00654 0.00305 -0.1743 0.9107 1.0000
1.000 0.8050 0.00626 0.00274 -0.1749 0.8988 1.0000
1.250 0.8410 0.00600 0.00244 -0.1763 0.8816 1.0000
1.500 0.8758 0.00587 0.00226 -0.1774 0.8582 1.0000
1.750 0.9116 0.00586 0.00208 -0.1788 0.8170 1.0000
2.000 0.9373 0.00625 0.00208 -0.1779 0.7415 1.0000
2.250 0.9455 0.00737 0.00241 -0.1734 0.5898 1.0000
2.500 0.9612 0.00824 0.00282 -0.1710 0.5109 1.0000
2.750 0.9817 0.00885 0.00313 -0.1695 0.4408 1.0000
3.000 0.9954 0.01023 0.00368 -0.1671 0.2800 1.0000
3.250 1.0156 0.01113 0.00412 -0.1658 0.1887 1.0000
3.500 1.0328 0.01249 0.00480 -0.1640 0.0614 1.0000
3.750 1.0569 0.01297 0.00521 -0.1632 0.0474 1.0000
4.000 1.0815 0.01337 0.00562 -0.1626 0.0442 1.0000
4.250 1.1058 0.01382 0.00609 -0.1618 0.0419 1.0000
4.500 1.1293 0.01437 0.00671 -0.1609 0.0396 1.0000
4.750 1.1526 0.01492 0.00733 -0.1599 0.0383 1.0000
5.000 1.1759 0.01544 0.00791 -0.1590 0.0371 1.0000
5.250 1.1987 0.01602 0.00855 -0.1580 0.0350 1.0000
5.500 1.2207 0.01672 0.00930 -0.1569 0.0328 1.0000
5.750 1.2412 0.01764 0.01026 -0.1555 0.0304 1.0000
6.000 1.2582 0.01943 0.01218 -0.1535 0.0274 1.0000
6.250 1.2822 0.01979 0.01260 -0.1528 0.0259 1.0000
6.500 1.3046 0.02054 0.01342 -0.1517 0.0238 1.0000
6.750 1.3267 0.02124 0.01416 -0.1507 0.0217 1.0000
7.000 1.3465 0.02388 0.01683 -0.1496 0.0190 1.0000
7.250 1.3691 0.02400 0.01708 -0.1485 0.0181 1.0000
7.500 1.3915 0.02487 0.01807 -0.1475 0.0167 1.0000
7.750 1.4132 0.02579 0.01908 -0.1464 0.0155 1.0000
8.000 1.4338 0.02654 0.01990 -0.1452 0.0146 1.0000
8.250 1.4534 0.02809 0.02151 -0.1441 0.0136 1.0000
8.500 1.4733 0.03034 0.02403 -0.1427 0.0127 1.0000
8.750 1.4926 0.03192 0.02585 -0.1412 0.0119 1.0000
9.000 1.5099 0.03431 0.02852 -0.1394 0.0113 1.0000
9.250 1.5249 0.03649 0.03095 -0.1374 0.0108 1.0000
9.500 1.5374 0.03869 0.03341 -0.1350 0.0104 1.0000
9.750 1.5479 0.04055 0.03547 -0.1325 0.0100 1.0000
10.000 1.5558 0.04241 0.03752 -0.1298 0.0097 1.0000
10.250 1.5557 0.04574 0.04117 -0.1260 0.0095 1.0000
10.500 1.5437 0.04968 0.04545 -0.1205 0.0094 1.0000
10.750 1.5279 0.05374 0.04984 -0.1149 0.0094 1.0000
11.000 1.5074 0.05816 0.05457 -0.1096 0.0093 1.0000
11.250 1.4849 0.06307 0.05979 -0.1050 0.0094 1.0000
11.500 1.4625 0.06821 0.06522 -0.1015 0.0095 1.0000
12.000 1.3091 0.10082 0.09893 -0.1046 0.0115 1.0000
12.250 1.2837 0.11052 0.10878 -0.1110 0.0115 1.0000
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