EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 58 AIRFOIL (e58-il) Reynolds number: 1,000,000 Max Cl/Cd: 235.69 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e58-il-1000000.txt Download as CSV file: xf-e58-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 58 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.0921 0.08863 0.08680 -0.1013 0.9675 0.0085
-8.250 -0.0828 0.08597 0.08414 -0.1028 0.9651 0.0088
-8.000 -0.0684 0.08293 0.08111 -0.1062 0.9632 0.0105
-7.750 -0.0531 0.07998 0.07816 -0.1102 0.9618 0.0107
-7.500 -0.0372 0.07670 0.07488 -0.1138 0.9607 0.0107
-7.250 -0.0202 0.07329 0.07147 -0.1178 0.9597 0.0108
-7.000 0.0010 0.06962 0.06779 -0.1229 0.9589 0.0108
-6.250 0.0310 0.05898 0.05717 -0.1317 0.9481 0.0113
-6.000 0.0602 0.05546 0.05364 -0.1384 0.9474 0.0115
-5.750 0.0931 0.05188 0.05004 -0.1460 0.9468 0.0119
-5.500 0.1288 0.04821 0.04634 -0.1543 0.9462 0.0130
-5.250 0.1734 0.04298 0.04105 -0.1664 0.9454 0.0140
-5.000 0.3062 0.01222 0.00830 -0.2157 0.9428 0.0117
-4.750 0.3291 0.01232 0.00842 -0.2148 0.9402 0.0126
-4.500 0.3604 0.01194 0.00794 -0.2157 0.9390 0.0139
-4.250 0.3956 0.01124 0.00710 -0.2174 0.9380 0.0153
-4.000 0.4338 0.00998 0.00564 -0.2200 0.9372 0.0171
-3.750 0.4675 0.00966 0.00531 -0.2213 0.9360 0.0189
-3.500 0.4999 0.00949 0.00511 -0.2222 0.9350 0.0214
-3.250 0.5334 0.00930 0.00488 -0.2234 0.9339 0.0230
-3.000 0.5703 0.00825 0.00371 -0.2256 0.9328 0.0264
-2.750 0.5983 0.00801 0.00346 -0.2257 0.9305 0.0290
-2.500 0.6250 0.00777 0.00319 -0.2255 0.9274 0.0308
-2.250 0.6550 0.00748 0.00287 -0.2259 0.9250 0.0326
-2.000 0.6860 0.00728 0.00265 -0.2266 0.9230 0.0341
-1.750 0.7186 0.00695 0.00226 -0.2276 0.9211 0.0359
-1.500 0.7507 0.00674 0.00200 -0.2285 0.9191 0.0375
-1.250 0.7757 0.00664 0.00189 -0.2278 0.9152 0.0390
-1.000 0.8040 0.00651 0.00175 -0.2278 0.9119 0.0427
-0.750 0.8353 0.00619 0.00165 -0.2287 0.9089 0.1249
-0.500 0.8665 0.00605 0.00156 -0.2293 0.9058 0.1535
-0.250 0.8911 0.00600 0.00155 -0.2286 0.9004 0.1666
0.000 0.9192 0.00590 0.00150 -0.2286 0.8959 0.1811
0.250 0.9492 0.00579 0.00145 -0.2291 0.8918 0.2078
0.500 0.9758 0.00565 0.00151 -0.2289 0.8864 0.2869
0.750 1.0063 0.00535 0.00157 -0.2297 0.8809 0.4641
1.000 1.0349 0.00496 0.00170 -0.2300 0.8697 0.7210
1.250 1.0555 0.00462 0.00174 -0.2281 0.8557 0.9172
1.500 1.0799 0.00459 0.00171 -0.2272 0.8409 1.0000
1.750 1.1054 0.00469 0.00172 -0.2265 0.8162 1.0000
2.000 1.1245 0.00509 0.00180 -0.2244 0.7529 1.0000
2.250 1.1322 0.00606 0.00221 -0.2199 0.6477 1.0000
2.500 1.1458 0.00687 0.00261 -0.2169 0.5701 1.0000
2.750 1.1581 0.00788 0.00308 -0.2139 0.4608 1.0000
3.000 1.1759 0.00864 0.00348 -0.2120 0.3911 1.0000
3.250 1.1955 0.00933 0.00382 -0.2105 0.3206 1.0000
3.500 1.2146 0.01011 0.00424 -0.2089 0.2553 1.0000
3.750 1.2350 0.01079 0.00462 -0.2076 0.1941 1.0000
4.000 1.2563 0.01140 0.00503 -0.2065 0.1564 1.0000
4.250 1.2791 0.01184 0.00536 -0.2056 0.1309 1.0000
4.500 1.2973 0.01285 0.00598 -0.2039 0.0598 1.0000
4.750 1.3169 0.01375 0.00660 -0.2024 0.0174 1.0000
5.000 1.3395 0.01428 0.00716 -0.2013 0.0128 1.0000
5.250 1.3624 0.01474 0.00769 -0.2003 0.0118 1.0000
5.500 1.3850 0.01521 0.00822 -0.1992 0.0109 1.0000
5.750 1.4072 0.01574 0.00879 -0.1981 0.0097 1.0000
6.000 1.4284 0.01638 0.00948 -0.1969 0.0087 1.0000
6.250 1.4464 0.01749 0.01071 -0.1949 0.0078 1.0000
6.500 1.4619 0.01882 0.01217 -0.1925 0.0074 1.0000
6.750 1.4816 0.01943 0.01286 -0.1909 0.0072 1.0000
7.000 1.5006 0.02009 0.01358 -0.1892 0.0068 1.0000
7.250 1.5179 0.02095 0.01451 -0.1872 0.0064 1.0000
7.500 1.5340 0.02198 0.01562 -0.1850 0.0060 1.0000
7.750 1.5500 0.02310 0.01683 -0.1828 0.0057 1.0000
8.000 1.5665 0.02410 0.01790 -0.1808 0.0054 1.0000
8.250 1.5826 0.02505 0.01891 -0.1789 0.0051 1.0000
8.500 1.5974 0.02637 0.02032 -0.1767 0.0049 1.0000
8.750 1.6121 0.02857 0.02264 -0.1746 0.0046 1.0000
9.000 1.6328 0.03181 0.02610 -0.1736 0.0045 1.0000
9.250 1.6579 0.03685 0.03151 -0.1736 0.0044 1.0000
9.500 1.6620 0.03751 0.03234 -0.1694 0.0043 1.0000
9.750 1.6684 0.03825 0.03321 -0.1658 0.0042 1.0000
10.000 1.6745 0.04048 0.03566 -0.1624 0.0041 1.0000
10.250 1.6783 0.04266 0.03806 -0.1586 0.0039 1.0000
10.500 1.6780 0.04557 0.04123 -0.1546 0.0038 1.0000
10.750 1.6736 0.04883 0.04475 -0.1501 0.0037 1.0000
11.000 1.6636 0.05271 0.04894 -0.1453 0.0036 1.0000
11.250 1.6502 0.05692 0.05344 -0.1406 0.0035 1.0000
11.500 1.6279 0.06240 0.05923 -0.1359 0.0036 1.0000
11.750 1.6066 0.06772 0.06483 -0.1320 0.0036 1.0000
12.000 1.5791 0.07425 0.07165 -0.1288 0.0036 1.0000
12.250 1.5541 0.08063 0.07826 -0.1271 0.0036 1.0000
12.500 1.5298 0.08730 0.08513 -0.1268 0.0037 1.0000
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Polar data table (+)
Polar graphs
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