E62 (5.62%) (e62-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: E62 (5.62%) (e62-il) Reynolds number: 500,000 Max Cl/Cd: 156.78 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e62-il-500000.txt Download as CSV file: xf-e62-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: E62 (5.62%)
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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3660 0.12498 0.12273 -0.0234 1.0000 0.0089
-9.500 -0.3648 0.12260 0.12037 -0.0232 1.0000 0.0089
-9.250 -0.3643 0.12029 0.11809 -0.0229 1.0000 0.0089
-7.500 -0.3000 0.09671 0.09456 -0.0387 0.9892 0.0091
-7.250 -0.2842 0.09274 0.09058 -0.0412 0.9872 0.0093
-7.000 -0.2666 0.08919 0.08702 -0.0445 0.9855 0.0096
-6.750 -0.2554 0.08621 0.08406 -0.0468 0.9802 0.0098
-6.500 -0.2348 0.08243 0.08028 -0.0517 0.9762 0.0101
-6.250 -0.2079 0.07821 0.07605 -0.0585 0.9738 0.0105
-6.000 -0.1872 0.07453 0.07236 -0.0636 0.9690 0.0108
-5.750 -0.1629 0.07052 0.06834 -0.0697 0.9643 0.0113
-5.500 -0.1300 0.06602 0.06380 -0.0781 0.9619 0.0121
-5.250 -0.0758 0.06053 0.05826 -0.0930 0.9602 0.0139
-5.000 -0.0482 0.05658 0.05427 -0.0989 0.9526 0.0141
-4.750 -0.0063 0.04879 0.04641 -0.1113 0.9502 0.0148
-4.500 0.0228 0.04627 0.04386 -0.1150 0.9488 0.0161
-4.250 0.0683 0.04196 0.03946 -0.1239 0.9479 0.0174
-4.000 0.1030 0.03792 0.03531 -0.1294 0.9413 0.0187
-3.750 0.1580 0.03283 0.02997 -0.1385 0.9397 0.0211
-3.500 0.2084 0.02930 0.02611 -0.1446 0.9385 0.0219
-3.250 0.2526 0.02357 0.02013 -0.1525 0.9377 0.0245
-3.000 0.2936 0.02143 0.01779 -0.1564 0.9368 0.0282
-2.750 0.3402 0.01831 0.01414 -0.1610 0.9362 0.0344
-2.500 0.3784 0.01683 0.01258 -0.1639 0.9354 0.0386
-2.000 0.4568 0.01134 0.00607 -0.1671 0.9302 0.0185
-1.750 0.4928 0.01023 0.00481 -0.1686 0.9277 0.0187
-1.500 0.5293 0.00970 0.00422 -0.1703 0.9254 0.0212
-1.250 0.5686 0.00893 0.00337 -0.1727 0.9236 0.0210
-1.000 0.6042 0.00845 0.00281 -0.1742 0.9197 0.0215
-0.750 0.6369 0.00808 0.00236 -0.1750 0.9131 0.0269
-0.500 0.6770 0.00731 0.00216 -0.1780 0.9092 0.2376
-0.250 0.7096 0.00657 0.00224 -0.1795 0.9006 0.5634
0.000 0.7338 0.00568 0.00216 -0.1783 0.8924 1.0000
0.250 0.7664 0.00565 0.00203 -0.1792 0.8825 1.0000
0.500 0.7963 0.00568 0.00197 -0.1795 0.8710 1.0000
0.750 0.8258 0.00572 0.00193 -0.1797 0.8592 1.0000
1.000 0.8546 0.00579 0.00192 -0.1798 0.8466 1.0000
1.250 0.8827 0.00588 0.00193 -0.1798 0.8337 1.0000
1.500 0.9102 0.00598 0.00196 -0.1796 0.8207 1.0000
1.750 0.9372 0.00609 0.00203 -0.1793 0.8073 1.0000
2.000 0.9637 0.00621 0.00209 -0.1789 0.7930 1.0000
2.250 0.9898 0.00634 0.00218 -0.1784 0.7780 1.0000
2.500 1.0155 0.00649 0.00227 -0.1779 0.7624 1.0000
2.750 1.0410 0.00664 0.00238 -0.1773 0.7461 1.0000
3.000 1.0660 0.00682 0.00253 -0.1766 0.7278 1.0000
3.250 1.0903 0.00700 0.00266 -0.1758 0.7061 1.0000
3.500 1.1109 0.00733 0.00280 -0.1741 0.6618 1.0000
3.750 1.1271 0.00790 0.00301 -0.1715 0.5849 1.0000
4.000 1.1417 0.00875 0.00337 -0.1688 0.4902 1.0000
4.250 1.1567 0.00977 0.00389 -0.1665 0.3912 1.0000
4.500 1.1727 0.01085 0.00446 -0.1645 0.2923 1.0000
4.750 1.1885 0.01205 0.00510 -0.1625 0.1909 1.0000
5.000 1.2038 0.01336 0.00585 -0.1605 0.0945 1.0000
5.250 1.2188 0.01478 0.00679 -0.1583 0.0194 1.0000
5.500 1.2409 0.01537 0.00742 -0.1572 0.0140 1.0000
5.750 1.2631 0.01593 0.00808 -0.1561 0.0127 1.0000
6.000 1.2843 0.01661 0.00885 -0.1548 0.0116 1.0000
6.250 1.3040 0.01747 0.00980 -0.1532 0.0108 1.0000
6.500 1.3216 0.01855 0.01100 -0.1512 0.0103 1.0000
6.750 1.3376 0.01979 0.01237 -0.1489 0.0101 1.0000
7.000 1.3534 0.02107 0.01379 -0.1467 0.0101 1.0000
7.250 1.3697 0.02240 0.01523 -0.1446 0.0101 1.0000
7.500 1.3872 0.02374 0.01668 -0.1427 0.0103 1.0000
7.750 1.4061 0.02522 0.01829 -0.1410 0.0107 1.0000
8.000 1.4269 0.02720 0.02046 -0.1395 0.0110 1.0000
8.250 1.4476 0.02917 0.02257 -0.1384 0.0108 1.0000
8.500 1.4755 0.03388 0.02777 -0.1375 0.0125 1.0000
9.750 1.4250 0.05396 0.05047 -0.1081 0.0192 1.0000
10.000 1.4046 0.05726 0.05396 -0.1023 0.0189 1.0000
10.250 1.3818 0.06117 0.05808 -0.0972 0.0186 1.0000
10.500 1.3567 0.06569 0.06280 -0.0929 0.0184 1.0000
10.750 1.3291 0.07084 0.06815 -0.0896 0.0183 1.0000
11.000 1.2969 0.07709 0.07462 -0.0874 0.0190 1.0000
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Polar data table (+)
Polar graphs
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