EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Reynolds number: 500,000 Max Cl/Cd: 43.26 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ea81006-il-500000.txt Download as CSV file: xf-ea81006-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.7341 0.06404 0.06172 -0.0154 1.0000 0.0150
-8.500 -0.7452 0.05724 0.05476 -0.0170 1.0000 0.0147
-8.250 -0.7535 0.05030 0.04759 -0.0174 1.0000 0.0143
-8.000 -0.7614 0.04243 0.03931 -0.0164 1.0000 0.0141
-7.750 -0.7670 0.03407 0.03033 -0.0141 1.0000 0.0144
-7.500 -0.7729 0.02357 0.01879 -0.0102 1.0000 0.0153
-7.250 -0.7611 0.01857 0.01306 -0.0078 1.0000 0.0162
-7.000 -0.7401 0.01718 0.01149 -0.0068 1.0000 0.0176
-6.750 -0.7156 0.01699 0.01124 -0.0062 1.0000 0.0190
-6.500 -0.6923 0.01620 0.01030 -0.0053 1.0000 0.0209
-6.250 -0.6694 0.01543 0.00941 -0.0044 1.0000 0.0233
-6.000 -0.6434 0.01622 0.01025 -0.0040 1.0000 0.0252
-5.750 -0.6190 0.01619 0.01012 -0.0032 1.0000 0.0283
-5.500 -0.5960 0.01557 0.00938 -0.0023 1.0000 0.0313
-5.250 -0.5722 0.01588 0.00970 -0.0017 1.0000 0.0336
-5.000 -0.5483 0.01600 0.00978 -0.0009 1.0000 0.0370
-4.500 -0.5030 0.01470 0.00831 0.0010 1.0000 0.0427
-4.250 -0.4803 0.01441 0.00804 0.0019 1.0000 0.0456
-4.000 -0.4576 0.01402 0.00759 0.0029 1.0000 0.0481
-3.750 -0.4349 0.01381 0.00732 0.0040 1.0000 0.0501
-3.500 -0.4124 0.01378 0.00721 0.0052 1.0000 0.0512
-3.250 -0.3912 0.01221 0.00564 0.0063 1.0000 0.0535
-3.000 -0.3699 0.01154 0.00501 0.0076 1.0000 0.0551
-2.750 -0.3486 0.01108 0.00455 0.0088 1.0000 0.0560
-2.500 -0.3274 0.01068 0.00417 0.0101 1.0000 0.0568
-2.250 -0.2886 0.01023 0.00373 0.0076 0.9964 0.0577
-2.000 -0.2440 0.00977 0.00329 0.0040 0.9901 0.0588
-1.750 -0.1966 0.00935 0.00290 -0.0002 0.9829 0.0609
-1.500 -0.1532 0.00897 0.00253 -0.0035 0.9708 0.0627
-1.250 -0.1127 0.00867 0.00224 -0.0059 0.9475 0.0638
-1.000 -0.0804 0.00837 0.00184 -0.0064 0.8951 0.0660
-0.750 -0.0611 0.00856 0.00156 -0.0039 0.7627 0.0677
-0.500 -0.0507 0.01039 0.00153 -0.0008 0.2401 0.0691
-0.250 -0.0267 0.01090 0.00151 -0.0003 0.0781 0.0710
0.000 0.0000 0.01089 0.00149 0.0000 0.0735 0.0735
0.250 0.0267 0.01089 0.00150 0.0003 0.0710 0.0782
0.500 0.0506 0.01038 0.00153 0.0009 0.0691 0.2434
0.750 0.0610 0.00856 0.00156 0.0039 0.0677 0.7632
1.000 0.0804 0.00837 0.00184 0.0064 0.0660 0.8944
1.250 0.1125 0.00867 0.00224 0.0060 0.0638 0.9471
1.500 0.1527 0.00896 0.00253 0.0036 0.0627 0.9706
1.750 0.1962 0.00935 0.00290 0.0003 0.0609 0.9828
2.000 0.2440 0.00977 0.00329 -0.0040 0.0588 0.9901
2.250 0.2891 0.01023 0.00373 -0.0077 0.0577 0.9966
2.500 0.3272 0.01068 0.00416 -0.0101 0.0568 1.0000
2.750 0.3485 0.01107 0.00455 -0.0088 0.0561 1.0000
3.000 0.3698 0.01154 0.00501 -0.0075 0.0551 1.0000
3.250 0.3910 0.01222 0.00565 -0.0063 0.0534 1.0000
3.500 0.4122 0.01377 0.00720 -0.0052 0.0512 1.0000
3.750 0.4347 0.01382 0.00733 -0.0040 0.0502 1.0000
4.000 0.4574 0.01403 0.00761 -0.0029 0.0481 1.0000
4.250 0.4802 0.01441 0.00803 -0.0018 0.0455 1.0000
4.500 0.5028 0.01470 0.00831 -0.0009 0.0427 1.0000
4.750 0.5240 0.01664 0.01030 0.0000 0.0394 1.0000
5.000 0.5482 0.01600 0.00977 0.0009 0.0370 1.0000
5.250 0.5720 0.01588 0.00970 0.0017 0.0336 1.0000
5.500 0.5959 0.01557 0.00938 0.0024 0.0313 1.0000
5.750 0.6188 0.01621 0.01014 0.0033 0.0283 1.0000
6.000 0.6432 0.01628 0.01032 0.0041 0.0252 1.0000
6.250 0.6692 0.01547 0.00946 0.0044 0.0234 1.0000
6.500 0.6923 0.01616 0.01026 0.0053 0.0209 1.0000
6.750 0.7154 0.01708 0.01134 0.0062 0.0190 1.0000
7.000 0.7398 0.01730 0.01162 0.0068 0.0177 1.0000
7.250 0.7615 0.01844 0.01291 0.0078 0.0162 1.0000
7.500 0.7730 0.02357 0.01879 0.0102 0.0153 1.0000
7.750 0.7681 0.03380 0.03004 0.0140 0.0144 1.0000
8.000 0.7625 0.04224 0.03911 0.0163 0.0141 1.0000
8.250 0.7536 0.05040 0.04769 0.0173 0.0143 1.0000
8.500 0.7452 0.05728 0.05479 0.0169 0.0148 1.0000
8.750 0.7349 0.06396 0.06164 0.0153 0.0151 1.0000
9.000 0.7198 0.07035 0.06811 0.0129 0.0153 1.0000
9.250 0.6718 0.08322 0.08097 0.0063 0.0219 1.0000
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Polar data table (+)
Polar graphs
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