EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Reynolds number: 200,000 Max Cl/Cd: 31.75 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea61009-il-200000.txt Download as CSV file: xf-ea61009-il-200000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.7263 0.08246 0.07899 -0.0030 1.0000 0.0689
-9.500 -0.7519 0.07415 0.07065 -0.0100 1.0000 0.0679
-9.250 -0.7787 0.06846 0.06489 -0.0124 1.0000 0.0674
-9.000 -0.7970 0.06312 0.05937 -0.0138 1.0000 0.0685
-8.750 -0.8251 0.06163 0.05702 -0.0132 1.0000 0.0724
-8.500 -0.8100 0.05388 0.04968 -0.0144 1.0000 0.0763
-8.250 -0.7973 0.05159 0.04731 -0.0141 1.0000 0.0810
-8.000 -0.7950 0.04776 0.04318 -0.0139 1.0000 0.0897
-7.750 -0.7956 0.03687 0.03040 -0.0106 1.0000 0.0596
-7.500 -0.7809 0.03145 0.02461 -0.0102 1.0000 0.0611
-7.250 -0.7586 0.02907 0.02220 -0.0101 1.0000 0.0635
-7.000 -0.7354 0.02752 0.02051 -0.0097 1.0000 0.0668
-6.750 -0.7117 0.02577 0.01847 -0.0092 1.0000 0.0687
-6.500 -0.6869 0.02460 0.01699 -0.0086 1.0000 0.0709
-6.250 -0.6614 0.02380 0.01593 -0.0081 1.0000 0.0723
-6.000 -0.6363 0.02149 0.01344 -0.0079 1.0000 0.0735
-5.750 -0.6109 0.01970 0.01164 -0.0077 1.0000 0.0758
-5.500 -0.5849 0.01859 0.01049 -0.0074 1.0000 0.0773
-5.250 -0.5589 0.01764 0.00950 -0.0071 1.0000 0.0787
-5.000 -0.5334 0.01680 0.00864 -0.0067 1.0000 0.0803
-4.750 -0.5084 0.01608 0.00792 -0.0062 1.0000 0.0826
-4.500 -0.4845 0.01542 0.00724 -0.0056 1.0000 0.0842
-4.250 -0.4628 0.01484 0.00666 -0.0046 1.0000 0.0853
-4.000 -0.4444 0.01432 0.00614 -0.0031 1.0000 0.0864
-3.750 -0.4288 0.01367 0.00553 -0.0012 1.0000 0.0884
-3.500 -0.4124 0.01319 0.00508 0.0005 1.0000 0.0911
-3.250 -0.3946 0.01285 0.00474 0.0019 1.0000 0.0947
-3.000 -0.3693 0.01255 0.00442 0.0020 0.9981 0.0989
-2.750 -0.3315 0.01214 0.00408 -0.0004 0.9923 0.1072
-2.500 -0.2946 0.01152 0.00382 -0.0028 0.9862 0.1626
-2.250 -0.2605 0.01015 0.00353 -0.0052 0.9798 0.3986
-2.000 -0.2287 0.00922 0.00367 -0.0062 0.9727 0.6308
-1.750 -0.1917 0.00917 0.00389 -0.0075 0.9665 0.7127
-1.500 -0.1552 0.00925 0.00406 -0.0087 0.9587 0.7553
-1.250 -0.1197 0.00934 0.00419 -0.0096 0.9494 0.7831
-1.000 -0.0862 0.00943 0.00428 -0.0098 0.9372 0.8068
-0.750 -0.0613 0.00954 0.00442 -0.0079 0.9209 0.8270
-0.500 -0.0406 0.00962 0.00450 -0.0051 0.9032 0.8442
-0.250 -0.0208 0.00966 0.00453 -0.0024 0.8870 0.8593
0.000 0.0000 0.00967 0.00454 0.0000 0.8728 0.8727
0.250 0.0208 0.00966 0.00453 0.0024 0.8594 0.8871
0.500 0.0406 0.00962 0.00449 0.0051 0.8440 0.9031
0.750 0.0614 0.00953 0.00441 0.0078 0.8268 0.9204
1.000 0.0862 0.00943 0.00428 0.0098 0.8068 0.9371
1.250 0.1197 0.00934 0.00419 0.0096 0.7830 0.9495
1.500 0.1553 0.00925 0.00406 0.0087 0.7552 0.9588
1.750 0.1918 0.00917 0.00389 0.0075 0.7124 0.9666
2.000 0.2287 0.00922 0.00367 0.0062 0.6309 0.9728
2.250 0.2604 0.01017 0.00353 0.0053 0.3970 0.9799
2.500 0.2945 0.01153 0.00382 0.0028 0.1613 0.9863
2.750 0.3315 0.01214 0.00408 0.0004 0.1072 0.9924
3.000 0.3693 0.01255 0.00442 -0.0020 0.0988 0.9982
3.250 0.3944 0.01284 0.00474 -0.0019 0.0948 1.0000
3.500 0.4122 0.01319 0.00509 -0.0004 0.0911 1.0000
3.750 0.4286 0.01367 0.00554 0.0012 0.0884 1.0000
4.000 0.4442 0.01433 0.00614 0.0031 0.0864 1.0000
4.250 0.4627 0.01484 0.00665 0.0046 0.0853 1.0000
4.500 0.4844 0.01542 0.00724 0.0056 0.0842 1.0000
4.750 0.5084 0.01608 0.00792 0.0062 0.0826 1.0000
5.000 0.5334 0.01680 0.00864 0.0067 0.0803 1.0000
5.250 0.5589 0.01764 0.00949 0.0071 0.0787 1.0000
5.500 0.5849 0.01859 0.01048 0.0074 0.0773 1.0000
5.750 0.6109 0.01970 0.01164 0.0077 0.0758 1.0000
6.000 0.6364 0.02149 0.01343 0.0079 0.0735 1.0000
6.250 0.6614 0.02378 0.01591 0.0081 0.0722 1.0000
6.500 0.6870 0.02462 0.01701 0.0086 0.0709 1.0000
6.750 0.7118 0.02578 0.01847 0.0091 0.0687 1.0000
7.000 0.7356 0.02753 0.02052 0.0097 0.0668 1.0000
7.250 0.7588 0.02912 0.02226 0.0100 0.0637 1.0000
7.500 0.7810 0.03147 0.02463 0.0102 0.0611 1.0000
7.750 0.7950 0.03725 0.03080 0.0105 0.0595 1.0000
8.000 0.7954 0.04774 0.04315 0.0139 0.0896 1.0000
8.250 0.7975 0.05161 0.04733 0.0140 0.0809 1.0000
8.500 0.8107 0.05386 0.04966 0.0144 0.0762 1.0000
8.750 0.8252 0.06163 0.05703 0.0131 0.0724 1.0000
9.000 0.7979 0.06317 0.05941 0.0137 0.0689 1.0000
9.250 0.7790 0.06851 0.06495 0.0123 0.0674 1.0000
9.500 0.7519 0.07422 0.07073 0.0098 0.0677 1.0000
9.750 0.7263 0.08264 0.07917 0.0028 0.0690 1.0000
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Polar data table (+)
Polar graphs
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