EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Reynolds number: 100,000 Max Cl/Cd: 27.23 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ea61009-il-100000-n5.txt Download as CSV file: xf-ea61009-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.7793 0.08372 0.07875 -0.0031 1.0000 0.0379
-10.750 -0.8071 0.07416 0.06911 -0.0108 1.0000 0.0373
-10.500 -0.8331 0.06742 0.06225 -0.0152 1.0000 0.0370
-10.250 -0.8589 0.06220 0.05686 -0.0168 1.0000 0.0369
-10.000 -0.8807 0.05753 0.05190 -0.0165 1.0000 0.0374
-9.750 -0.8974 0.05249 0.04634 -0.0158 1.0000 0.0385
-9.500 -0.8996 0.04863 0.04209 -0.0151 1.0000 0.0400
-9.250 -0.8842 0.04714 0.04062 -0.0149 1.0000 0.0420
-9.000 -0.8723 0.04488 0.03816 -0.0143 1.0000 0.0441
-8.750 -0.8626 0.04175 0.03459 -0.0136 1.0000 0.0462
-8.500 -0.8514 0.03897 0.03115 -0.0126 1.0000 0.0494
-8.250 -0.8360 0.03654 0.02817 -0.0119 1.0000 0.0520
-8.000 -0.8152 0.03465 0.02619 -0.0117 1.0000 0.0539
-7.750 -0.7944 0.03299 0.02426 -0.0113 1.0000 0.0569
-7.500 -0.7737 0.03130 0.02204 -0.0106 1.0000 0.0598
-7.250 -0.7505 0.02944 0.01989 -0.0101 1.0000 0.0611
-7.000 -0.7259 0.02778 0.01816 -0.0099 1.0000 0.0624
-6.750 -0.7010 0.02641 0.01668 -0.0096 1.0000 0.0640
-6.500 -0.6759 0.02518 0.01531 -0.0093 1.0000 0.0656
-6.250 -0.6505 0.02410 0.01403 -0.0089 1.0000 0.0671
-6.000 -0.6250 0.02316 0.01292 -0.0086 1.0000 0.0685
-5.750 -0.6005 0.02205 0.01191 -0.0083 1.0000 0.0708
-5.500 -0.5759 0.02119 0.01106 -0.0079 1.0000 0.0737
-5.250 -0.5512 0.02042 0.01022 -0.0075 1.0000 0.0758
-5.000 -0.5267 0.01970 0.00945 -0.0070 1.0000 0.0769
-4.750 -0.5036 0.01888 0.00870 -0.0064 1.0000 0.0782
-4.500 -0.4807 0.01819 0.00804 -0.0057 1.0000 0.0797
-4.250 -0.4580 0.01760 0.00746 -0.0051 1.0000 0.0815
-4.000 -0.4359 0.01709 0.00691 -0.0042 1.0000 0.0836
-3.750 -0.4147 0.01666 0.00644 -0.0032 1.0000 0.0858
-3.500 -0.3953 0.01620 0.00600 -0.0020 1.0000 0.0887
-3.250 -0.3765 0.01580 0.00563 -0.0007 1.0000 0.0939
-3.000 -0.3550 0.01537 0.00528 0.0000 0.9979 0.1034
-2.750 -0.3223 0.01471 0.00499 -0.0016 0.9883 0.1508
-2.500 -0.2904 0.01383 0.00468 -0.0032 0.9800 0.2669
-2.250 -0.2594 0.01285 0.00450 -0.0046 0.9727 0.4401
-2.000 -0.2311 0.01235 0.00463 -0.0046 0.9641 0.5750
-1.750 -0.2005 0.01226 0.00488 -0.0045 0.9569 0.6684
-1.500 -0.1706 0.01234 0.00511 -0.0043 0.9488 0.7268
-1.250 -0.1397 0.01245 0.00526 -0.0044 0.9403 0.7643
-1.000 -0.1079 0.01255 0.00537 -0.0046 0.9312 0.7907
-0.750 -0.0789 0.01264 0.00547 -0.0042 0.9203 0.8119
-0.500 -0.0502 0.01273 0.00557 -0.0035 0.9089 0.8325
-0.250 -0.0245 0.01281 0.00567 -0.0020 0.8951 0.8560
0.000 0.0000 0.01283 0.00571 0.0000 0.8767 0.8765
0.250 0.0245 0.01281 0.00567 0.0020 0.8562 0.8949
0.500 0.0503 0.01273 0.00557 0.0035 0.8327 0.9088
0.750 0.0789 0.01264 0.00547 0.0042 0.8118 0.9203
1.000 0.1079 0.01256 0.00537 0.0046 0.7909 0.9312
1.250 0.1397 0.01245 0.00526 0.0044 0.7643 0.9403
1.500 0.1707 0.01234 0.00511 0.0043 0.7273 0.9488
2.000 0.2312 0.01235 0.00463 0.0045 0.5768 0.9641
2.250 0.2594 0.01285 0.00450 0.0046 0.4401 0.9727
2.500 0.2904 0.01382 0.00468 0.0032 0.2677 0.9799
2.750 0.3223 0.01471 0.00499 0.0016 0.1512 0.9882
3.000 0.3550 0.01537 0.00528 0.0000 0.1034 0.9978
3.250 0.3766 0.01579 0.00563 0.0007 0.0941 1.0000
3.500 0.3954 0.01620 0.00601 0.0020 0.0887 1.0000
3.750 0.4148 0.01666 0.00644 0.0032 0.0858 1.0000
4.000 0.4360 0.01709 0.00691 0.0042 0.0836 1.0000
4.250 0.4581 0.01760 0.00745 0.0050 0.0815 1.0000
4.500 0.4807 0.01819 0.00804 0.0057 0.0797 1.0000
4.750 0.5036 0.01888 0.00870 0.0064 0.0782 1.0000
5.000 0.5267 0.01970 0.00945 0.0070 0.0769 1.0000
5.250 0.5511 0.02042 0.01022 0.0075 0.0759 1.0000
5.500 0.5759 0.02119 0.01106 0.0080 0.0738 1.0000
5.750 0.6004 0.02205 0.01191 0.0083 0.0708 1.0000
6.000 0.6249 0.02316 0.01292 0.0086 0.0685 1.0000
6.250 0.6504 0.02409 0.01402 0.0090 0.0671 1.0000
6.500 0.6758 0.02518 0.01531 0.0093 0.0656 1.0000
6.750 0.7009 0.02640 0.01667 0.0096 0.0640 1.0000
7.000 0.7258 0.02778 0.01815 0.0099 0.0624 1.0000
7.250 0.7503 0.02943 0.01988 0.0102 0.0611 1.0000
7.500 0.7735 0.03129 0.02203 0.0106 0.0598 1.0000
7.750 0.7941 0.03296 0.02424 0.0113 0.0568 1.0000
8.000 0.8149 0.03464 0.02618 0.0117 0.0539 1.0000
8.250 0.8357 0.03652 0.02816 0.0120 0.0520 1.0000
8.500 0.8511 0.03895 0.03113 0.0127 0.0495 1.0000
8.750 0.8623 0.04174 0.03457 0.0137 0.0462 1.0000
9.000 0.8720 0.04485 0.03812 0.0144 0.0441 1.0000
9.250 0.8836 0.04717 0.04066 0.0150 0.0420 1.0000
9.500 0.8987 0.04871 0.04219 0.0152 0.0401 1.0000
9.750 0.8970 0.05245 0.04630 0.0159 0.0385 1.0000
10.000 0.8798 0.05751 0.05190 0.0166 0.0373 1.0000
10.250 0.8586 0.06216 0.05682 0.0169 0.0369 1.0000
10.500 0.8336 0.06731 0.06214 0.0153 0.0370 1.0000
10.750 0.8061 0.07429 0.06925 0.0107 0.0373 1.0000
11.000 0.7795 0.08365 0.07868 0.0032 0.0379 1.0000
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Polar data table (+)
Polar graphs
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