EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Reynolds number: 200,000 Max Cl/Cd: 33.88 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ea61009-il-200000-n5.txt Download as CSV file: xf-ea61009-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.9218 0.06342 0.05978 -0.0143 1.0000 0.0197
-11.500 -0.9680 0.05453 0.05061 -0.0189 1.0000 0.0192
-11.250 -1.0043 0.04858 0.04436 -0.0187 1.0000 0.0191
-11.000 -1.0209 0.04453 0.04000 -0.0169 1.0000 0.0198
-10.750 -1.0314 0.03979 0.03481 -0.0154 1.0000 0.0210
-10.500 -1.0344 0.03505 0.02948 -0.0138 1.0000 0.0227
-10.250 -1.0252 0.03196 0.02596 -0.0127 1.0000 0.0242
-10.000 -1.0031 0.03206 0.02607 -0.0123 1.0000 0.0252
-9.750 -0.9813 0.03192 0.02586 -0.0120 1.0000 0.0265
-9.500 -0.9620 0.03079 0.02449 -0.0113 1.0000 0.0291
-9.250 -0.9440 0.02894 0.02220 -0.0106 1.0000 0.0317
-9.000 -0.9210 0.02881 0.02208 -0.0103 1.0000 0.0330
-8.750 -0.8984 0.02835 0.02151 -0.0100 1.0000 0.0347
-8.500 -0.8760 0.02759 0.02051 -0.0095 1.0000 0.0372
-8.250 -0.8529 0.02697 0.01956 -0.0090 1.0000 0.0391
-8.000 -0.8309 0.02540 0.01786 -0.0088 1.0000 0.0409
-7.750 -0.8072 0.02470 0.01711 -0.0086 1.0000 0.0427
-7.500 -0.7831 0.02397 0.01625 -0.0083 1.0000 0.0448
-7.250 -0.7586 0.02316 0.01528 -0.0080 1.0000 0.0472
-7.000 -0.7336 0.02224 0.01415 -0.0077 1.0000 0.0489
-6.750 -0.7082 0.02140 0.01313 -0.0073 1.0000 0.0500
-6.500 -0.6825 0.02073 0.01230 -0.0071 1.0000 0.0508
-6.250 -0.6572 0.01958 0.01110 -0.0069 1.0000 0.0518
-6.000 -0.6320 0.01866 0.01021 -0.0066 1.0000 0.0528
-5.750 -0.6066 0.01792 0.00947 -0.0064 1.0000 0.0538
-5.500 -0.5813 0.01726 0.00880 -0.0061 1.0000 0.0549
-5.250 -0.5560 0.01665 0.00819 -0.0059 1.0000 0.0560
-5.000 -0.5309 0.01612 0.00763 -0.0056 1.0000 0.0576
-4.750 -0.5061 0.01567 0.00715 -0.0052 1.0000 0.0593
-4.500 -0.4773 0.01522 0.00667 -0.0057 0.9885 0.0604
-4.250 -0.4446 0.01453 0.00602 -0.0071 0.9720 0.0615
-3.750 -0.3800 0.01356 0.00507 -0.0095 0.9487 0.0643
-3.500 -0.3492 0.01322 0.00470 -0.0102 0.9387 0.0659
-3.000 -0.2911 0.01267 0.00408 -0.0108 0.9206 0.0695
-2.750 -0.2633 0.01241 0.00380 -0.0107 0.9126 0.0721
-2.500 -0.2368 0.01220 0.00357 -0.0104 0.9031 0.0759
-2.250 -0.2108 0.01195 0.00335 -0.0100 0.8941 0.0831
-2.000 -0.1872 0.01132 0.00314 -0.0094 0.8851 0.1697
-1.750 -0.1623 0.01092 0.00297 -0.0089 0.8752 0.2299
-1.500 -0.1410 0.01009 0.00280 -0.0080 0.8631 0.3902
-1.250 -0.1195 0.00952 0.00270 -0.0066 0.8485 0.5019
-1.000 -0.0980 0.00915 0.00273 -0.0049 0.8316 0.6127
-0.750 -0.0742 0.00903 0.00276 -0.0035 0.8153 0.6649
-0.500 -0.0499 0.00897 0.00282 -0.0022 0.8005 0.7097
-0.250 -0.0254 0.00895 0.00286 -0.0010 0.7831 0.7454
0.000 0.0000 0.00895 0.00287 0.0000 0.7660 0.7660
0.250 0.0254 0.00895 0.00286 0.0010 0.7455 0.7832
0.500 0.0499 0.00897 0.00282 0.0022 0.7101 0.8004
0.750 0.0742 0.00903 0.00276 0.0035 0.6646 0.8152
1.000 0.0979 0.00916 0.00273 0.0049 0.6124 0.8316
1.250 0.1195 0.00952 0.00270 0.0066 0.5017 0.8485
1.500 0.1410 0.01009 0.00280 0.0080 0.3907 0.8630
1.750 0.1623 0.01092 0.00297 0.0089 0.2304 0.8753
2.000 0.1871 0.01131 0.00314 0.0094 0.1706 0.8851
2.250 0.2108 0.01195 0.00335 0.0100 0.0831 0.8941
2.500 0.2367 0.01220 0.00357 0.0104 0.0759 0.9032
2.750 0.2633 0.01241 0.00380 0.0108 0.0721 0.9126
3.000 0.2910 0.01267 0.00408 0.0108 0.0695 0.9206
3.500 0.3491 0.01322 0.00470 0.0102 0.0659 0.9387
3.750 0.3800 0.01356 0.00507 0.0095 0.0643 0.9487
4.250 0.4446 0.01453 0.00602 0.0071 0.0615 0.9718
4.500 0.4772 0.01522 0.00667 0.0057 0.0604 0.9883
4.750 0.5061 0.01567 0.00715 0.0052 0.0593 1.0000
5.000 0.5309 0.01613 0.00764 0.0056 0.0576 1.0000
5.250 0.5560 0.01664 0.00818 0.0059 0.0560 1.0000
5.500 0.5812 0.01726 0.00880 0.0062 0.0549 1.0000
5.750 0.6065 0.01792 0.00947 0.0064 0.0538 1.0000
6.000 0.6319 0.01865 0.01020 0.0067 0.0529 1.0000
6.250 0.6571 0.01957 0.01109 0.0069 0.0518 1.0000
6.500 0.6823 0.02072 0.01229 0.0071 0.0508 1.0000
6.750 0.7080 0.02140 0.01313 0.0074 0.0500 1.0000
7.000 0.7334 0.02223 0.01415 0.0077 0.0489 1.0000
7.250 0.7584 0.02315 0.01527 0.0081 0.0472 1.0000
7.500 0.7828 0.02396 0.01625 0.0084 0.0448 1.0000
7.750 0.8069 0.02469 0.01711 0.0086 0.0427 1.0000
8.000 0.8305 0.02539 0.01785 0.0088 0.0409 1.0000
8.250 0.8525 0.02696 0.01955 0.0091 0.0391 1.0000
8.500 0.8756 0.02757 0.02049 0.0096 0.0372 1.0000
8.750 0.8980 0.02833 0.02148 0.0101 0.0347 1.0000
9.000 0.9206 0.02879 0.02206 0.0104 0.0330 1.0000
9.250 0.9435 0.02893 0.02220 0.0107 0.0317 1.0000
9.500 0.9616 0.03076 0.02445 0.0114 0.0291 1.0000
9.750 0.9806 0.03194 0.02589 0.0121 0.0266 1.0000
10.000 1.0027 0.03198 0.02598 0.0125 0.0252 1.0000
10.250 1.0246 0.03194 0.02594 0.0128 0.0242 1.0000
10.500 1.0341 0.03494 0.02936 0.0139 0.0228 1.0000
10.750 1.0307 0.03978 0.03480 0.0155 0.0210 1.0000
11.000 1.0189 0.04472 0.04020 0.0171 0.0200 1.0000
11.250 1.0074 0.04816 0.04391 0.0187 0.0190 1.0000
11.500 0.9673 0.05456 0.05064 0.0190 0.0193 1.0000
11.750 0.9193 0.06378 0.06014 0.0141 0.0198 1.0000
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Polar data table (+)
Polar graphs
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