EPPLER 862 STRUT AIRFOIL (e862-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 862 STRUT AIRFOIL (e862-il) Reynolds number: 200,000 Max Cl/Cd: 20.73 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e862-il-200000.txt Download as CSV file: xf-e862-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 862 STRUT 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.3652 0.11662 0.10999 -0.0071 0.6118 0.2738
-11.750 -0.4169 0.10650 0.09978 -0.0097 0.6118 0.2862
-11.500 -0.3763 0.10847 0.10167 -0.0094 0.6038 0.2854
-11.250 -0.3297 0.11103 0.10423 -0.0090 0.5962 0.2838
-11.000 -0.3276 0.10789 0.10107 -0.0098 0.5920 0.2885
-10.750 -0.2997 0.10804 0.10119 -0.0099 0.5859 0.2896
-10.500 -0.2763 0.10763 0.10072 -0.0102 0.5806 0.2914
-10.250 -0.2637 0.10592 0.09892 -0.0107 0.5760 0.2946
-10.000 -0.2902 0.09946 0.09241 -0.0120 0.5746 0.3036
-9.750 -0.2619 0.09954 0.09252 -0.0122 0.5697 0.3048
-9.500 -0.2385 0.09905 0.09203 -0.0125 0.5651 0.3066
-9.250 -0.2319 0.09660 0.08955 -0.0131 0.5618 0.3109
-9.000 -0.2523 0.09111 0.08400 -0.0141 0.5600 0.3191
-8.750 -0.2243 0.09122 0.08407 -0.0142 0.5559 0.3203
-8.500 -0.7223 0.03927 0.03157 -0.0120 0.5755 0.3810
-8.250 -0.6473 0.04209 0.03455 -0.0146 0.5690 0.3806
-8.000 -0.5882 0.04437 0.03689 -0.0159 0.5630 0.3806
-7.750 -0.5364 0.04638 0.03891 -0.0166 0.5579 0.3806
-7.500 -0.4888 0.04828 0.04077 -0.0170 0.5533 0.3807
-7.250 -0.4440 0.05004 0.04254 -0.0174 0.5488 0.3807
-7.000 -0.6884 0.03372 0.02566 -0.0010 0.5538 0.3986
-6.750 -0.6796 0.03280 0.02462 0.0006 0.5504 0.4015
-6.500 -0.6434 0.03292 0.02486 -0.0007 0.5464 0.4030
-6.250 -0.6104 0.03304 0.02504 -0.0015 0.5420 0.4047
-6.000 -0.5799 0.03313 0.02515 -0.0020 0.5380 0.4066
-5.750 -0.5550 0.03301 0.02501 -0.0019 0.5346 0.4090
-5.500 -0.5372 0.03257 0.02448 -0.0011 0.5317 0.4120
-5.250 -0.5333 0.03151 0.02322 0.0012 0.5292 0.4160
-5.000 -0.5285 0.03058 0.02205 0.0033 0.5264 0.4194
-4.750 -0.4962 0.03036 0.02197 0.0024 0.5232 0.4213
-4.500 -0.4650 0.03041 0.02210 0.0017 0.5195 0.4232
-4.250 -0.4356 0.03043 0.02216 0.0013 0.5159 0.4252
-4.000 -0.4087 0.03032 0.02205 0.0011 0.5128 0.4274
-3.750 -0.3840 0.03009 0.02176 0.0012 0.5101 0.4300
-3.500 -0.3615 0.02978 0.02136 0.0015 0.5075 0.4331
-3.250 -0.3429 0.02943 0.02083 0.0022 0.5049 0.4366
-3.000 -0.3237 0.02906 0.02035 0.0028 0.5022 0.4394
-2.750 -0.2949 0.02876 0.02017 0.0023 0.4991 0.4414
-2.500 -0.2659 0.02873 0.02022 0.0018 0.4958 0.4433
-2.250 -0.2376 0.02875 0.02028 0.0014 0.4928 0.4454
-2.000 -0.2098 0.02875 0.02029 0.0012 0.4900 0.4479
-1.750 -0.1830 0.02866 0.02016 0.0010 0.4875 0.4507
-1.500 -0.1569 0.02848 0.01990 0.0009 0.4851 0.4537
-1.250 -0.1322 0.02834 0.01962 0.0009 0.4826 0.4566
-1.000 -0.1090 0.02836 0.01954 0.0011 0.4800 0.4592
-0.750 -0.0833 0.02805 0.01933 0.0009 0.4772 0.4616
-0.500 -0.0554 0.02810 0.01949 0.0005 0.4742 0.4638
-0.250 -0.0275 0.02822 0.01969 0.0002 0.4714 0.4661
0.000 0.0000 0.02828 0.01977 0.0000 0.4687 0.4687
0.250 0.0275 0.02822 0.01969 -0.0002 0.4661 0.4714
0.500 0.0553 0.02810 0.01949 -0.0005 0.4638 0.4742
0.750 0.0833 0.02805 0.01933 -0.0009 0.4616 0.4772
1.000 0.1091 0.02836 0.01954 -0.0011 0.4592 0.4799
1.250 0.1322 0.02834 0.01962 -0.0009 0.4566 0.4826
1.500 0.1569 0.02848 0.01990 -0.0009 0.4537 0.4851
1.750 0.1830 0.02865 0.02016 -0.0010 0.4507 0.4875
2.000 0.2098 0.02874 0.02029 -0.0012 0.4479 0.4900
2.250 0.2376 0.02875 0.02028 -0.0015 0.4454 0.4928
2.500 0.2659 0.02873 0.02022 -0.0018 0.4433 0.4958
2.750 0.2949 0.02876 0.02017 -0.0023 0.4414 0.4991
3.000 0.3238 0.02906 0.02035 -0.0028 0.4394 0.5022
3.250 0.3429 0.02943 0.02083 -0.0022 0.4365 0.5049
3.500 0.3616 0.02978 0.02136 -0.0015 0.4331 0.5075
3.750 0.3842 0.03008 0.02176 -0.0012 0.4300 0.5101
4.000 0.4088 0.03032 0.02205 -0.0011 0.4274 0.5128
4.250 0.4357 0.03043 0.02216 -0.0013 0.4252 0.5159
4.500 0.4651 0.03041 0.02210 -0.0018 0.4232 0.5195
4.750 0.4963 0.03037 0.02197 -0.0024 0.4213 0.5232
5.000 0.5286 0.03058 0.02205 -0.0034 0.4194 0.5264
5.250 0.5335 0.03151 0.02321 -0.0012 0.4160 0.5292
5.500 0.5375 0.03256 0.02447 0.0011 0.4120 0.5317
5.750 0.5554 0.03301 0.02501 0.0019 0.4090 0.5346
6.000 0.5803 0.03312 0.02514 0.0020 0.4066 0.5380
6.250 0.6109 0.03303 0.02503 0.0014 0.4047 0.5420
6.500 0.6438 0.03292 0.02485 0.0006 0.4030 0.5464
6.750 0.6800 0.03280 0.02462 -0.0007 0.4015 0.5504
7.000 0.6888 0.03371 0.02565 0.0010 0.3986 0.5538
7.250 0.4448 0.05004 0.04254 0.0172 0.3807 0.5488
7.500 0.4897 0.04828 0.04076 0.0168 0.3807 0.5533
7.750 0.5371 0.04639 0.03891 0.0164 0.3806 0.5579
8.000 0.5891 0.04437 0.03688 0.0157 0.3805 0.5630
8.250 0.6481 0.04209 0.03455 0.0145 0.3806 0.5690
8.500 0.7230 0.03927 0.03158 0.0119 0.3809 0.5755
8.750 0.2259 0.09121 0.08405 0.0139 0.3204 0.5559
9.000 0.2540 0.09110 0.08399 0.0138 0.3191 0.5601
9.250 0.2840 0.09073 0.08364 0.0137 0.3182 0.5647
9.500 0.3158 0.09019 0.08310 0.0137 0.3175 0.5697
9.750 0.3486 0.08958 0.08245 0.0136 0.3169 0.5748
10.000 0.2921 0.09945 0.09240 0.0118 0.3036 0.5747
10.250 0.2879 0.10318 0.09618 0.0110 0.2979 0.5778
10.500 0.3374 0.10034 0.09342 0.0114 0.3000 0.5849
10.750 0.3794 0.09843 0.09153 0.0117 0.3009 0.5919
11.000 0.3297 0.10788 0.10106 0.0095 0.2885 0.5922
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Polar data table (+)
Polar graphs
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