EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Reynolds number: 1,000,000 Max Cl/Cd: 66.02 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ea81006-il-1000000-n5.txt Download as CSV file: xf-ea81006-il-1000000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8531 0.09011 0.08845 0.0029 1.0000 0.0028
-12.000 -0.8921 0.07777 0.07614 -0.0056 1.0000 0.0028
-11.500 -1.0491 0.04351 0.04134 -0.0219 1.0000 0.0025
-11.250 -1.0689 0.03662 0.03403 -0.0197 1.0000 0.0025
-11.000 -1.0719 0.03169 0.02873 -0.0178 1.0000 0.0025
-10.750 -1.0642 0.02844 0.02519 -0.0162 1.0000 0.0026
-10.500 -1.0512 0.02604 0.02256 -0.0150 1.0000 0.0026
-10.250 -1.0367 0.02391 0.02018 -0.0137 1.0000 0.0027
-10.000 -1.0193 0.02230 0.01837 -0.0126 1.0000 0.0027
-9.750 -1.0008 0.02084 0.01669 -0.0116 1.0000 0.0028
-9.500 -0.9807 0.01964 0.01533 -0.0107 1.0000 0.0029
-9.250 -0.9597 0.01861 0.01414 -0.0099 1.0000 0.0029
-9.000 -0.9402 0.01721 0.01252 -0.0089 1.0000 0.0031
-8.750 -0.9195 0.01602 0.01115 -0.0079 1.0000 0.0033
-8.500 -0.8971 0.01523 0.01025 -0.0072 1.0000 0.0036
-8.250 -0.8740 0.01463 0.00959 -0.0065 1.0000 0.0039
-8.000 -0.8509 0.01400 0.00889 -0.0058 1.0000 0.0042
-7.750 -0.8277 0.01340 0.00820 -0.0051 1.0000 0.0045
-7.500 -0.8043 0.01286 0.00757 -0.0043 1.0000 0.0048
-7.250 -0.7809 0.01236 0.00700 -0.0036 1.0000 0.0050
-7.000 -0.7573 0.01192 0.00650 -0.0028 1.0000 0.0052
-6.750 -0.7343 0.01137 0.00585 -0.0019 1.0000 0.0058
-6.500 -0.7113 0.01087 0.00529 -0.0010 1.0000 0.0067
-6.250 -0.6882 0.01043 0.00482 -0.0001 1.0000 0.0078
-6.000 -0.6649 0.01007 0.00439 0.0008 1.0000 0.0089
-5.750 -0.6414 0.00976 0.00405 0.0016 1.0000 0.0104
-5.500 -0.6186 0.00939 0.00368 0.0026 1.0000 0.0153
-5.250 -0.5956 0.00910 0.00346 0.0035 1.0000 0.0250
-5.000 -0.5660 0.00897 0.00333 0.0030 0.9988 0.0298
-4.750 -0.5343 0.00894 0.00333 0.0021 0.9969 0.0328
-4.500 -0.5024 0.00885 0.00325 0.0011 0.9947 0.0346
-4.250 -0.4683 0.00880 0.00320 -0.0003 0.9902 0.0360
-4.000 -0.4326 0.00890 0.00330 -0.0021 0.9856 0.0381
-3.750 -0.3983 0.00861 0.00300 -0.0036 0.9788 0.0390
-3.500 -0.3627 0.00837 0.00275 -0.0054 0.9698 0.0398
-3.250 -0.3288 0.00818 0.00254 -0.0067 0.9539 0.0404
-3.000 -0.2991 0.00806 0.00237 -0.0070 0.9263 0.0412
-2.750 -0.2743 0.00800 0.00219 -0.0061 0.8854 0.0420
-2.500 -0.2507 0.00800 0.00201 -0.0050 0.8335 0.0426
-2.250 -0.2277 0.00812 0.00183 -0.0039 0.7548 0.0432
-2.000 -0.2089 0.00894 0.00174 -0.0023 0.4770 0.0438
-1.750 -0.1867 0.00970 0.00168 -0.0016 0.1995 0.0443
-1.500 -0.1614 0.00997 0.00162 -0.0012 0.0632 0.0447
-1.250 -0.1345 0.00990 0.00152 -0.0010 0.0590 0.0454
-1.000 -0.1079 0.00980 0.00139 -0.0008 0.0563 0.0469
-0.750 -0.0810 0.00974 0.00132 -0.0006 0.0550 0.0481
-0.500 -0.0540 0.00970 0.00128 -0.0004 0.0541 0.0492
-0.250 -0.0270 0.00968 0.00126 -0.0002 0.0529 0.0503
0.000 0.0000 0.00967 0.00125 0.0000 0.0515 0.0515
0.250 0.0270 0.00968 0.00126 0.0002 0.0503 0.0529
0.500 0.0540 0.00970 0.00128 0.0004 0.0492 0.0541
0.750 0.0810 0.00974 0.00132 0.0006 0.0481 0.0550
1.000 0.1079 0.00980 0.00139 0.0008 0.0469 0.0563
1.250 0.1345 0.00991 0.00152 0.0010 0.0454 0.0589
1.750 0.1867 0.00970 0.00168 0.0016 0.0443 0.1979
2.000 0.2088 0.00893 0.00174 0.0023 0.0438 0.4779
2.250 0.2277 0.00812 0.00183 0.0039 0.0432 0.7540
2.500 0.2507 0.00801 0.00201 0.0050 0.0426 0.8334
2.750 0.2744 0.00800 0.00219 0.0061 0.0419 0.8842
3.000 0.2991 0.00806 0.00237 0.0070 0.0412 0.9263
3.250 0.3286 0.00818 0.00254 0.0068 0.0404 0.9539
3.500 0.3627 0.00836 0.00275 0.0054 0.0397 0.9701
3.750 0.3977 0.00861 0.00301 0.0037 0.0391 0.9783
4.000 0.4321 0.00893 0.00333 0.0022 0.0382 0.9855
4.250 0.4682 0.00879 0.00318 0.0004 0.0359 0.9901
4.500 0.5019 0.00886 0.00325 -0.0010 0.0346 0.9945
4.750 0.5340 0.00893 0.00332 -0.0020 0.0326 0.9968
5.000 0.5656 0.00897 0.00334 -0.0029 0.0299 0.9987
5.250 0.5956 0.00910 0.00347 -0.0035 0.0252 1.0000
5.500 0.6185 0.00938 0.00367 -0.0026 0.0155 1.0000
5.750 0.6413 0.00976 0.00405 -0.0016 0.0105 1.0000
6.000 0.6648 0.01007 0.00439 -0.0007 0.0089 1.0000
6.250 0.6881 0.01043 0.00482 0.0001 0.0077 1.0000
6.500 0.7112 0.01086 0.00529 0.0010 0.0067 1.0000
6.750 0.7342 0.01137 0.00585 0.0019 0.0058 1.0000
7.000 0.7571 0.01192 0.00650 0.0029 0.0052 1.0000
7.250 0.7807 0.01236 0.00701 0.0036 0.0050 1.0000
7.500 0.8042 0.01285 0.00756 0.0044 0.0048 1.0000
7.750 0.8275 0.01340 0.00820 0.0051 0.0045 1.0000
8.000 0.8507 0.01400 0.00889 0.0058 0.0042 1.0000
8.250 0.8739 0.01463 0.00959 0.0065 0.0039 1.0000
8.500 0.8970 0.01524 0.01026 0.0072 0.0036 1.0000
8.750 0.9194 0.01605 0.01119 0.0079 0.0033 1.0000
9.000 0.9400 0.01724 0.01256 0.0089 0.0030 1.0000
9.250 0.9599 0.01856 0.01409 0.0099 0.0029 1.0000
9.500 0.9808 0.01962 0.01530 0.0107 0.0029 1.0000
9.750 1.0008 0.02083 0.01669 0.0116 0.0028 1.0000
10.000 1.0192 0.02231 0.01838 0.0126 0.0027 1.0000
10.250 1.0365 0.02398 0.02026 0.0137 0.0027 1.0000
10.500 1.0527 0.02576 0.02225 0.0148 0.0026 1.0000
10.750 1.0639 0.02850 0.02526 0.0162 0.0026 1.0000
11.000 1.0703 0.03202 0.02909 0.0179 0.0025 1.0000
11.250 1.0696 0.03652 0.03392 0.0197 0.0025 1.0000
11.500 1.0450 0.04417 0.04204 0.0220 0.0025 1.0000
12.000 0.9120 0.07345 0.07183 0.0091 0.0027 1.0000
12.250 0.8491 0.09119 0.08953 -0.0038 0.0029 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il)