GOE 802 A AIRFOIL (goe802a-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 802 A AIRFOIL (goe802a-il) Reynolds number: 200,000 Max Cl/Cd: 149.11 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe802a-il-200000-n5.txt Download as CSV file: xf-goe802a-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 802 A 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.1596 0.09181 0.08833 -0.0573 0.9613 0.0272
-8.250 -0.1449 0.08858 0.08509 -0.0615 0.9509 0.0280
-8.000 -0.1324 0.08572 0.08221 -0.0661 0.9387 0.0283
-7.750 -0.1132 0.08240 0.07884 -0.0729 0.9300 0.0285
-7.500 -0.1005 0.07896 0.07540 -0.0755 0.9203 0.0287
-7.250 -0.0896 0.07575 0.07218 -0.0752 0.9138 0.0291
-7.000 -0.0748 0.07294 0.06935 -0.0772 0.9064 0.0298
-6.750 -0.0583 0.07007 0.06644 -0.0801 0.8979 0.0305
-6.500 -0.0388 0.06704 0.06335 -0.0835 0.8913 0.0314
-6.250 -0.0180 0.06403 0.06029 -0.0873 0.8828 0.0325
-6.000 0.0133 0.06086 0.05697 -0.0938 0.8752 0.0336
-5.750 0.0450 0.05768 0.05360 -0.0996 0.8683 0.0339
-5.500 0.0565 0.05452 0.05048 -0.0991 0.8607 0.0344
-5.250 0.0760 0.05198 0.04787 -0.1002 0.8543 0.0353
-5.000 0.1000 0.04949 0.04529 -0.1024 0.8469 0.0365
-4.750 0.1265 0.04695 0.04263 -0.1048 0.8393 0.0379
-4.250 0.1914 0.04225 0.03745 -0.1103 0.8245 0.0399
-4.000 0.2117 0.03952 0.03468 -0.1110 0.8171 0.0404
-3.750 0.2359 0.03747 0.03254 -0.1119 0.8091 0.0414
-3.500 0.2629 0.03577 0.03069 -0.1129 0.8001 0.0442
-3.250 0.2979 0.03485 0.02934 -0.1140 0.7915 0.0458
-3.000 0.3205 0.03222 0.02669 -0.1146 0.7807 0.0464
-2.750 0.3460 0.03047 0.02481 -0.1152 0.7702 0.0473
-2.500 0.3730 0.02902 0.02319 -0.1156 0.7589 0.0488
-2.250 0.4031 0.02840 0.02226 -0.1158 0.7477 0.0515
-2.000 0.4328 0.02789 0.02137 -0.1158 0.7372 0.0518
-1.750 0.4580 0.02556 0.01902 -0.1164 0.7254 0.0523
-1.500 0.4842 0.02418 0.01753 -0.1167 0.7141 0.0539
-1.250 0.5131 0.02423 0.01727 -0.1165 0.7012 0.0577
-1.000 0.5407 0.02337 0.01619 -0.1165 0.6882 0.0577
-0.500 0.5971 0.02000 0.01217 -0.1165 0.6639 0.0395
-0.250 0.6241 0.01921 0.01122 -0.1165 0.6507 0.0392
0.000 0.6509 0.01855 0.01037 -0.1164 0.6374 0.0390
0.250 0.6777 0.01797 0.00958 -0.1162 0.6237 0.0390
0.500 0.7044 0.01744 0.00885 -0.1160 0.6104 0.0398
0.750 0.7307 0.01721 0.00857 -0.1160 0.5982 0.0414
1.000 0.7570 0.01691 0.00814 -0.1158 0.5860 0.0425
1.250 0.7834 0.01653 0.00763 -0.1156 0.5744 0.0425
1.500 0.8094 0.01623 0.00720 -0.1154 0.5629 0.0428
1.750 0.8354 0.01594 0.00683 -0.1151 0.5509 0.0434
2.000 0.8611 0.01568 0.00650 -0.1149 0.5400 0.0445
2.250 0.8867 0.01546 0.00621 -0.1146 0.5293 0.0460
2.500 0.9127 0.01557 0.00632 -0.1144 0.5194 0.0491
2.750 0.9382 0.01557 0.00629 -0.1142 0.5094 0.0529
3.000 0.9638 0.01555 0.00627 -0.1139 0.4999 0.0566
3.250 0.9891 0.01562 0.00630 -0.1136 0.4897 0.0628
3.500 1.0147 0.01580 0.00645 -0.1133 0.4798 0.0729
4.000 1.0650 0.01616 0.00668 -0.1126 0.4575 0.0921
4.250 1.0889 0.01624 0.00672 -0.1121 0.4456 0.0939
4.500 1.1131 0.01641 0.00688 -0.1117 0.4331 0.0944
4.750 1.1371 0.01651 0.00701 -0.1113 0.4216 0.0949
5.250 1.1843 0.01682 0.00735 -0.1104 0.3988 0.0964
5.500 1.2073 0.01707 0.00759 -0.1098 0.3891 0.0967
5.750 1.2313 0.01722 0.00782 -0.1094 0.3803 0.0971
6.000 1.2542 0.01742 0.00805 -0.1089 0.3725 0.0978
6.250 1.2777 0.01763 0.00832 -0.1084 0.3641 0.0986
6.500 1.2996 0.01795 0.00865 -0.1077 0.3545 0.0993
6.750 1.3222 0.01818 0.00893 -0.1071 0.3447 0.1010
7.000 1.3429 0.01854 0.00925 -0.1062 0.3340 0.1030
7.250 1.3647 0.01883 0.00959 -0.1055 0.3231 0.1049
7.750 1.4056 0.01953 0.01030 -0.1036 0.3025 0.1104
8.000 1.3703 0.00919 0.00032 -0.0938 0.2903 0.1145
8.250 1.3856 0.00955 0.00068 -0.0921 0.2794 0.1165
8.500 1.4631 0.02079 0.01157 -0.1002 0.2702 0.1175
8.750 1.4792 0.02133 0.01210 -0.0986 0.2566 0.1200
9.000 1.4921 0.02193 0.01267 -0.0965 0.2408 0.1244
9.500 1.5108 0.02254 0.01430 -0.0915 0.2023 1.0000
9.750 1.5190 0.02358 0.01527 -0.0891 0.1855 1.0000
10.250 1.5370 0.02573 0.01738 -0.0849 0.1623 1.0000
10.500 1.5451 0.02692 0.01856 -0.0829 0.1539 1.0000
10.750 1.5551 0.02803 0.01971 -0.0812 0.1454 1.0000
11.000 1.5616 0.02941 0.02109 -0.0793 0.1372 1.0000
11.250 1.5709 0.03064 0.02239 -0.0777 0.1287 1.0000
11.500 1.5763 0.03221 0.02397 -0.0760 0.1216 1.0000
11.750 1.5853 0.03353 0.02537 -0.0746 0.1146 1.0000
12.000 1.5899 0.03527 0.02712 -0.0730 0.1069 1.0000
12.250 1.5961 0.03692 0.02883 -0.0717 0.0993 1.0000
12.500 1.5987 0.03895 0.03088 -0.0703 0.0911 1.0000
12.750 1.6023 0.04094 0.03292 -0.0691 0.0831 1.0000
13.000 1.6032 0.04326 0.03527 -0.0679 0.0762 1.0000
13.250 1.6038 0.04567 0.03773 -0.0668 0.0705 1.0000
13.500 1.6019 0.04845 0.04054 -0.0659 0.0654 1.0000
13.750 1.6008 0.05120 0.04336 -0.0651 0.0603 1.0000
14.000 1.5961 0.05449 0.04669 -0.0645 0.0551 1.0000
14.250 1.5945 0.05754 0.04983 -0.0641 0.0505 1.0000
14.500 1.5900 0.06104 0.05341 -0.0638 0.0440 1.0000
14.750 1.5811 0.06521 0.05759 -0.0638 0.0272 1.0000
15.000 1.5657 0.07039 0.06275 -0.0641 0.0222 1.0000
15.250 1.5533 0.07534 0.06775 -0.0646 0.0200 1.0000
15.500 1.5442 0.07996 0.07249 -0.0652 0.0189 1.0000
15.750 1.5355 0.08461 0.07728 -0.0660 0.0182 1.0000
16.000 1.5267 0.08937 0.08219 -0.0669 0.0176 1.0000
16.250 1.5170 0.09433 0.08729 -0.0680 0.0171 1.0000
16.500 1.5069 0.09946 0.09257 -0.0692 0.0167 1.0000
16.750 1.4963 0.10477 0.09802 -0.0707 0.0164 1.0000
17.000 1.4855 0.11021 0.10361 -0.0724 0.0161 1.0000
17.250 1.4742 0.11581 0.10937 -0.0743 0.0158 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 802 A AIRFOIL (goe802a-il)