EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Reynolds number: 50,000 Max Cl/Cd: 21.43 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ea61009-il-50000-n5.txt Download as CSV file: xf-ea61009-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER EA 6(-1)-009 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.7727 0.08345 0.07632 -0.0107 1.0000 0.0687
-10.000 -0.7923 0.07785 0.07066 -0.0136 1.0000 0.0685
-9.750 -0.8124 0.07286 0.06553 -0.0149 1.0000 0.0685
-9.500 -0.8297 0.06768 0.06017 -0.0159 1.0000 0.0687
-9.250 -0.8258 0.06396 0.05627 -0.0159 1.0000 0.0700
-9.000 -0.8262 0.05974 0.05177 -0.0160 1.0000 0.0708
-8.750 -0.8234 0.05566 0.04735 -0.0158 1.0000 0.0719
-8.500 -0.8182 0.05202 0.04327 -0.0155 1.0000 0.0749
-8.250 -0.8099 0.04857 0.03937 -0.0151 1.0000 0.0784
-8.000 -0.7935 0.04585 0.03644 -0.0147 1.0000 0.0809
-7.750 -0.7776 0.04265 0.03277 -0.0142 1.0000 0.0821
-7.500 -0.7590 0.03977 0.02941 -0.0136 1.0000 0.0835
-7.250 -0.7382 0.03723 0.02637 -0.0131 1.0000 0.0851
-7.000 -0.7154 0.03500 0.02379 -0.0126 1.0000 0.0868
-6.750 -0.6912 0.03315 0.02184 -0.0123 1.0000 0.0890
-6.500 -0.6674 0.03171 0.02020 -0.0119 1.0000 0.0937
-6.250 -0.6428 0.03025 0.01851 -0.0115 1.0000 0.0983
-6.000 -0.6173 0.02883 0.01708 -0.0111 1.0000 0.1010
-5.750 -0.5914 0.02760 0.01576 -0.0106 1.0000 0.1032
-5.500 -0.5658 0.02650 0.01457 -0.0100 1.0000 0.1057
-5.250 -0.5407 0.02553 0.01344 -0.0093 1.0000 0.1085
-5.000 -0.5164 0.02464 0.01244 -0.0085 1.0000 0.1116
-4.750 -0.4937 0.02374 0.01159 -0.0078 1.0000 0.1164
-4.500 -0.4708 0.02297 0.01080 -0.0071 1.0000 0.1245
-4.250 -0.4487 0.02218 0.01004 -0.0064 1.0000 0.1341
-4.000 -0.4268 0.02138 0.00932 -0.0057 1.0000 0.1467
-3.750 -0.4056 0.02050 0.00863 -0.0049 1.0000 0.1719
-3.500 -0.3865 0.01945 0.00792 -0.0040 1.0000 0.2300
-3.250 -0.3725 0.01815 0.00742 -0.0023 1.0000 0.3594
-3.000 -0.3616 0.01724 0.00755 0.0013 1.0000 0.5283
-2.750 -0.3482 0.01714 0.00790 0.0054 1.0000 0.6510
-2.500 -0.3334 0.01721 0.00806 0.0089 1.0000 0.7145
-2.250 -0.3187 0.01729 0.00812 0.0121 1.0000 0.7609
-2.000 -0.3034 0.01736 0.00816 0.0152 1.0000 0.7978
-1.750 -0.2871 0.01741 0.00819 0.0182 1.0000 0.8327
-1.500 -0.2682 0.01745 0.00819 0.0207 1.0000 0.8682
-1.250 -0.2291 0.01765 0.00835 0.0202 1.0000 0.9136
-1.000 -0.1490 0.01790 0.00846 0.0117 1.0000 0.9494
-0.750 -0.1003 0.01780 0.00826 0.0071 1.0000 0.9641
-0.500 -0.0565 0.01767 0.00808 0.0031 1.0000 0.9782
-0.250 -0.0109 0.01755 0.00794 -0.0013 1.0000 0.9946
0.000 0.0000 0.01752 0.00790 0.0000 1.0000 1.0000
0.250 0.0109 0.01755 0.00794 0.0013 0.9946 1.0000
0.500 0.0566 0.01767 0.00808 -0.0031 0.9782 1.0000
0.750 0.0999 0.01779 0.00825 -0.0070 0.9642 1.0000
1.000 0.1486 0.01790 0.00846 -0.0116 0.9495 1.0000
1.250 0.2291 0.01765 0.00835 -0.0202 0.9136 1.0000
1.500 0.2683 0.01745 0.00820 -0.0207 0.8684 1.0000
1.750 0.2872 0.01741 0.00819 -0.0182 0.8329 1.0000
2.000 0.3035 0.01736 0.00817 -0.0152 0.7979 1.0000
2.500 0.3335 0.01721 0.00806 -0.0089 0.7147 1.0000
2.750 0.3483 0.01714 0.00790 -0.0054 0.6513 1.0000
3.000 0.3616 0.01724 0.00754 -0.0013 0.5276 1.0000
3.250 0.3726 0.01815 0.00742 0.0023 0.3599 1.0000
3.500 0.3865 0.01945 0.00792 0.0040 0.2296 1.0000
3.750 0.4056 0.02050 0.00863 0.0049 0.1719 1.0000
4.000 0.4268 0.02137 0.00932 0.0057 0.1468 1.0000
4.250 0.4487 0.02217 0.01004 0.0064 0.1342 1.0000
4.500 0.4708 0.02297 0.01079 0.0071 0.1244 1.0000
4.750 0.4936 0.02374 0.01159 0.0078 0.1164 1.0000
5.000 0.5163 0.02464 0.01244 0.0085 0.1116 1.0000
5.250 0.5406 0.02553 0.01343 0.0093 0.1086 1.0000
5.500 0.5657 0.02650 0.01456 0.0100 0.1057 1.0000
5.750 0.5913 0.02759 0.01576 0.0106 0.1032 1.0000
6.000 0.6172 0.02883 0.01708 0.0111 0.1010 1.0000
6.250 0.6427 0.03024 0.01851 0.0115 0.0983 1.0000
6.500 0.6673 0.03171 0.02020 0.0120 0.0937 1.0000
6.750 0.6911 0.03314 0.02184 0.0123 0.0890 1.0000
7.000 0.7152 0.03499 0.02378 0.0126 0.0869 1.0000
7.250 0.7381 0.03722 0.02636 0.0131 0.0851 1.0000
7.500 0.7588 0.03976 0.02940 0.0137 0.0835 1.0000
7.750 0.7774 0.04264 0.03276 0.0142 0.0821 1.0000
8.000 0.7933 0.04584 0.03642 0.0147 0.0809 1.0000
8.250 0.8096 0.04855 0.03936 0.0151 0.0784 1.0000
8.500 0.8180 0.05200 0.04325 0.0155 0.0749 1.0000
8.750 0.8232 0.05563 0.04732 0.0159 0.0719 1.0000
9.000 0.8261 0.05970 0.05173 0.0160 0.0708 1.0000
9.250 0.8254 0.06395 0.05626 0.0160 0.0700 1.0000
9.500 0.8294 0.06765 0.06004 0.0160 0.0687 1.0000
9.750 0.8120 0.07284 0.06552 0.0150 0.0685 1.0000
10.000 0.7914 0.07788 0.07069 0.0136 0.0685 1.0000
10.250 0.7724 0.08346 0.07633 0.0107 0.0687 1.0000
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Polar data table (+)
Polar graphs
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