EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Reynolds number: 1,000,000 Max Cl/Cd: 67.43 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea81006-il-1000000.txt Download as CSV file: xf-ea81006-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER EA 8(-1)-006 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.6729 0.12269 0.12098 0.0174 1.0000 0.0110
-11.250 -0.6728 0.11776 0.11606 0.0154 1.0000 0.0111
-9.750 -0.9719 0.02776 0.02459 -0.0153 1.0000 0.0069
-9.500 -0.9623 0.02423 0.02066 -0.0136 1.0000 0.0071
-9.250 -0.9453 0.02230 0.01847 -0.0125 1.0000 0.0073
-9.000 -0.9265 0.02075 0.01670 -0.0114 1.0000 0.0074
-8.750 -0.9091 0.01876 0.01441 -0.0101 1.0000 0.0076
-8.500 -0.8942 0.01615 0.01140 -0.0084 1.0000 0.0081
-8.250 -0.8730 0.01507 0.01015 -0.0074 1.0000 0.0086
-8.000 -0.8503 0.01433 0.00933 -0.0066 1.0000 0.0091
-7.750 -0.8275 0.01364 0.00854 -0.0058 1.0000 0.0097
-7.500 -0.8043 0.01303 0.00784 -0.0050 1.0000 0.0104
-7.250 -0.7807 0.01257 0.00730 -0.0043 1.0000 0.0109
-7.000 -0.7593 0.01166 0.00626 -0.0031 1.0000 0.0122
-6.750 -0.7364 0.01111 0.00567 -0.0022 1.0000 0.0137
-6.500 -0.7131 0.01069 0.00520 -0.0013 1.0000 0.0151
-6.250 -0.6900 0.01022 0.00471 -0.0003 1.0000 0.0182
-6.000 -0.6654 0.01008 0.00459 0.0003 1.0000 0.0218
-5.750 -0.6386 0.01032 0.00494 0.0006 1.0000 0.0265
-5.500 -0.6115 0.01072 0.00535 0.0009 1.0000 0.0289
-5.250 -0.5850 0.01113 0.00576 0.0013 1.0000 0.0297
-5.000 -0.5633 0.01059 0.00520 0.0024 1.0000 0.0322
-4.750 -0.5389 0.01071 0.00533 0.0031 1.0000 0.0336
-4.500 -0.5153 0.01077 0.00538 0.0040 1.0000 0.0352
-4.250 -0.4925 0.01076 0.00534 0.0050 1.0000 0.0368
-4.000 -0.4703 0.01063 0.00519 0.0061 1.0000 0.0380
-3.750 -0.4438 0.01081 0.00535 0.0064 0.9996 0.0391
-3.500 -0.4096 0.00990 0.00443 0.0047 0.9975 0.0415
-3.250 -0.3663 0.00944 0.00400 0.0012 0.9935 0.0435
-3.000 -0.3293 0.00906 0.00364 -0.0009 0.9884 0.0446
-2.750 -0.2905 0.00868 0.00326 -0.0033 0.9838 0.0455
-2.500 -0.2550 0.00832 0.00290 -0.0050 0.9746 0.0462
-2.250 -0.2194 0.00800 0.00257 -0.0067 0.9609 0.0470
-2.000 -0.1882 0.00776 0.00227 -0.0072 0.9284 0.0479
-1.750 -0.1646 0.00768 0.00204 -0.0059 0.8771 0.0490
-1.500 -0.1436 0.00787 0.00187 -0.0041 0.7776 0.0503
-1.250 -0.1304 0.00949 0.00185 -0.0017 0.2740 0.0509
-1.000 -0.1064 0.01006 0.00186 -0.0012 0.0677 0.0515
-0.750 -0.0803 0.00983 0.00159 -0.0008 0.0619 0.0533
-0.500 -0.0537 0.00975 0.00150 -0.0005 0.0597 0.0547
-0.250 -0.0269 0.00971 0.00145 -0.0003 0.0582 0.0559
0.000 0.0000 0.00970 0.00144 0.0000 0.0570 0.0570
0.250 0.0268 0.00971 0.00145 0.0003 0.0559 0.0582
0.500 0.0536 0.00975 0.00150 0.0005 0.0547 0.0598
0.750 0.0803 0.00983 0.00159 0.0008 0.0533 0.0619
1.000 0.1064 0.01006 0.00185 0.0012 0.0514 0.0673
1.250 0.1304 0.00950 0.00185 0.0017 0.0509 0.2720
1.500 0.1437 0.00788 0.00187 0.0041 0.0503 0.7752
1.750 0.1646 0.00768 0.00204 0.0059 0.0490 0.8772
2.000 0.1882 0.00776 0.00227 0.0072 0.0479 0.9280
2.250 0.2194 0.00800 0.00257 0.0067 0.0470 0.9609
2.500 0.2549 0.00833 0.00290 0.0050 0.0462 0.9745
2.750 0.2906 0.00868 0.00326 0.0033 0.0455 0.9839
3.000 0.3294 0.00907 0.00364 0.0008 0.0446 0.9885
3.250 0.3669 0.00943 0.00399 -0.0013 0.0434 0.9938
3.500 0.4087 0.00989 0.00443 -0.0045 0.0415 0.9974
3.750 0.4441 0.01082 0.00536 -0.0064 0.0391 0.9996
4.000 0.4700 0.01063 0.00520 -0.0061 0.0380 1.0000
4.250 0.4923 0.01074 0.00532 -0.0049 0.0368 1.0000
4.500 0.5152 0.01075 0.00536 -0.0039 0.0351 1.0000
4.750 0.5388 0.01069 0.00531 -0.0031 0.0336 1.0000
5.000 0.5631 0.01059 0.00520 -0.0023 0.0322 1.0000
5.250 0.5849 0.01113 0.00576 -0.0012 0.0297 1.0000
5.500 0.6114 0.01072 0.00535 -0.0009 0.0289 1.0000
5.750 0.6384 0.01034 0.00496 -0.0006 0.0266 1.0000
6.000 0.6651 0.01009 0.00460 -0.0003 0.0219 1.0000
6.250 0.6898 0.01023 0.00471 0.0004 0.0183 1.0000
6.500 0.7129 0.01069 0.00520 0.0013 0.0151 1.0000
6.750 0.7362 0.01112 0.00568 0.0022 0.0137 1.0000
7.000 0.7592 0.01164 0.00624 0.0031 0.0123 1.0000
7.250 0.7806 0.01256 0.00730 0.0043 0.0109 1.0000
7.500 0.8043 0.01301 0.00781 0.0050 0.0104 1.0000
7.750 0.8274 0.01363 0.00853 0.0058 0.0098 1.0000
8.000 0.8503 0.01433 0.00933 0.0066 0.0092 1.0000
8.250 0.8730 0.01507 0.01015 0.0074 0.0086 1.0000
8.500 0.8942 0.01619 0.01144 0.0084 0.0081 1.0000
8.750 0.9100 0.01859 0.01421 0.0100 0.0076 1.0000
9.000 0.9275 0.02057 0.01649 0.0113 0.0074 1.0000
9.250 0.9466 0.02207 0.01821 0.0123 0.0073 1.0000
9.500 0.9630 0.02413 0.02055 0.0135 0.0071 1.0000
9.750 0.9740 0.02735 0.02414 0.0151 0.0069 1.0000
10.250 0.7765 0.07559 0.07404 0.0059 0.0076 1.0000
10.500 0.7670 0.08183 0.08025 0.0013 0.0076 1.0000
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Polar data table (+)
Polar graphs
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