EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
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Airfoil: EPPLER 59 AIRFOIL (e59-il) Reynolds number: 1,000,000 Max Cl/Cd: 108.78 at α=-1° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e59-il-1000000-n5.txt Download as CSV file: xf-e59-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.1651 0.08378 0.08197 -0.0858 0.9733 0.0036
-8.000 -0.1555 0.07996 0.07816 -0.0886 0.9712 0.0040
-7.750 -0.1449 0.07499 0.07320 -0.0930 0.9693 0.0043
-7.250 -0.1278 0.06705 0.06528 -0.0995 0.9623 0.0049
-7.000 -0.1091 0.06534 0.06357 -0.1023 0.9603 0.0052
-6.750 -0.0872 0.06264 0.06087 -0.1069 0.9587 0.0055
-6.500 -0.0611 0.05888 0.05710 -0.1137 0.9574 0.0058
-6.250 -0.0277 0.05392 0.05211 -0.1235 0.9561 0.0061
-6.000 -0.0127 0.05030 0.04848 -0.1276 0.9498 0.0065
-5.750 0.0248 0.04458 0.04271 -0.1386 0.9470 0.0073
-5.500 0.0785 0.03613 0.03411 -0.1553 0.9454 0.0076
-5.250 0.1606 0.01667 0.01363 -0.1814 0.9443 0.0082
-5.000 0.1974 0.01443 0.01102 -0.1847 0.9435 0.0086
-4.750 0.2307 0.01381 0.01025 -0.1863 0.9427 0.0089
-4.500 0.2633 0.01104 0.00703 -0.1887 0.9406 0.0099
-4.250 0.2899 0.01042 0.00631 -0.1888 0.9376 0.0104
-4.000 0.3196 0.00981 0.00556 -0.1894 0.9351 0.0108
-3.750 0.3508 0.00919 0.00482 -0.1903 0.9329 0.0111
-3.500 0.3834 0.00858 0.00411 -0.1915 0.9312 0.0113
-3.250 0.4159 0.00815 0.00361 -0.1927 0.9298 0.0118
-3.000 0.4459 0.00773 0.00313 -0.1933 0.9277 0.0120
-2.750 0.4732 0.00735 0.00268 -0.1932 0.9242 0.0119
-2.500 0.5031 0.00700 0.00224 -0.1936 0.9206 0.0118
-2.250 0.5361 0.00667 0.00185 -0.1948 0.9167 0.0117
-2.000 0.5652 0.00645 0.00158 -0.1950 0.9110 0.0118
-1.750 0.5954 0.00627 0.00135 -0.1955 0.9049 0.0119
-1.500 0.6276 0.00610 0.00103 -0.1962 0.8777 0.0125
-1.250 0.6557 0.00610 0.00084 -0.1961 0.8419 0.0143
-1.000 0.6799 0.00625 0.00079 -0.1952 0.8031 0.0154
-0.750 0.7007 0.00653 0.00080 -0.1935 0.7531 0.0174
-0.500 0.7192 0.00695 0.00090 -0.1914 0.6867 0.0196
-0.250 0.7301 0.00796 0.00125 -0.1879 0.5350 0.0279
0.000 0.7545 0.00819 0.00138 -0.1873 0.4977 0.0602
0.250 0.7797 0.00834 0.00151 -0.1869 0.4556 0.1238
0.500 0.7951 0.00948 0.00194 -0.1848 0.2757 0.1647
0.750 0.8196 0.00982 0.00216 -0.1844 0.2138 0.2301
1.000 0.8417 0.01051 0.00264 -0.1838 0.0783 0.4019
1.250 0.8674 0.01077 0.00288 -0.1836 0.0352 0.4770
1.500 0.8947 0.01082 0.00307 -0.1836 0.0310 0.5543
1.750 0.9220 0.01087 0.00325 -0.1836 0.0295 0.6230
2.000 0.9496 0.01076 0.00353 -0.1837 0.0283 0.8016
2.250 0.9665 0.01057 0.00365 -0.1811 0.0275 1.0000
2.500 0.9925 0.01077 0.00382 -0.1807 0.0270 1.0000
2.750 1.0185 0.01098 0.00402 -0.1804 0.0266 1.0000
3.000 1.0443 0.01120 0.00423 -0.1800 0.0263 1.0000
3.250 1.0700 0.01143 0.00448 -0.1796 0.0260 1.0000
3.500 1.0954 0.01167 0.00472 -0.1791 0.0254 1.0000
3.750 1.1202 0.01201 0.00506 -0.1786 0.0226 1.0000
4.000 1.1456 0.01224 0.00527 -0.1781 0.0219 1.0000
4.250 1.1714 0.01238 0.00539 -0.1778 0.0214 1.0000
4.500 1.1972 0.01252 0.00552 -0.1775 0.0200 1.0000
4.750 1.2212 0.01288 0.00570 -0.1769 0.0082 1.0000
5.000 1.2463 0.01312 0.00593 -0.1764 0.0039 1.0000
5.250 1.2707 0.01344 0.00628 -0.1758 0.0035 1.0000
5.500 1.2946 0.01381 0.00670 -0.1750 0.0030 1.0000
5.750 1.3184 0.01420 0.00713 -0.1743 0.0027 1.0000
6.000 1.3419 0.01461 0.00759 -0.1735 0.0025 1.0000
6.250 1.3650 0.01507 0.00810 -0.1726 0.0023 1.0000
6.500 1.3874 0.01560 0.00873 -0.1716 0.0020 1.0000
6.750 1.4088 0.01630 0.00955 -0.1703 0.0017 1.0000
7.000 1.4291 0.01714 0.01051 -0.1688 0.0015 1.0000
7.250 1.4506 0.01775 0.01119 -0.1677 0.0015 1.0000
7.500 1.4714 0.01845 0.01197 -0.1664 0.0014 1.0000
7.750 1.4914 0.01925 0.01286 -0.1650 0.0014 1.0000
8.000 1.5106 0.02014 0.01386 -0.1634 0.0014 1.0000
8.250 1.5289 0.02116 0.01502 -0.1617 0.0013 1.0000
8.500 1.5465 0.02226 0.01627 -0.1599 0.0013 1.0000
8.750 1.5632 0.02351 0.01767 -0.1579 0.0013 1.0000
9.000 1.5791 0.02486 0.01919 -0.1558 0.0013 1.0000
9.250 1.5937 0.02638 0.02089 -0.1535 0.0013 1.0000
9.500 1.6063 0.02803 0.02274 -0.1509 0.0012 1.0000
9.750 1.6170 0.02974 0.02466 -0.1481 0.0012 1.0000
10.000 1.6249 0.03172 0.02687 -0.1448 0.0012 1.0000
10.250 1.6308 0.03379 0.02916 -0.1413 0.0012 1.0000
10.500 1.6336 0.03613 0.03176 -0.1375 0.0012 1.0000
10.750 1.6333 0.03875 0.03465 -0.1334 0.0012 1.0000
11.000 1.6300 0.04159 0.03776 -0.1292 0.0012 1.0000
11.250 1.6237 0.04472 0.04115 -0.1250 0.0012 1.0000
11.500 1.6116 0.04857 0.04529 -0.1205 0.0013 1.0000
11.750 1.5984 0.05251 0.04949 -0.1166 0.0013 1.0000
12.000 1.5828 0.05685 0.05407 -0.1131 0.0013 1.0000
12.250 1.5666 0.06146 0.05890 -0.1105 0.0013 1.0000
12.500 1.5446 0.06717 0.06486 -0.1086 0.0013 1.0000
12.750 1.5192 0.07389 0.07181 -0.1081 0.0013 1.0000
13.000 1.4979 0.08060 0.07870 -0.1090 0.0013 1.0000
13.250 1.4769 0.08804 0.08632 -0.1114 0.0013 1.0000
13.500 1.4523 0.09750 0.09596 -0.1162 0.0013 1.0000
13.750 1.4262 0.10913 0.10776 -0.1235 0.0013 1.0000
14.000 1.4018 0.12270 0.12149 -0.1327 0.0013 1.0000
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Polar data table (+)
Polar graphs
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