GOE 413 AIRFOIL (goe413-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 413 AIRFOIL (goe413-il) Reynolds number: 200,000 Max Cl/Cd: 74.53 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe413-il-200000-n5.txt Download as CSV file: xf-goe413-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 413 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5854 0.04867 0.04342 -0.0960 0.8919 0.0477
-11.000 -0.5980 0.04361 0.03804 -0.1029 0.8783 0.0478
-10.750 -0.5938 0.04075 0.03492 -0.1054 0.8690 0.0481
-10.500 -0.5845 0.03838 0.03230 -0.1070 0.8603 0.0484
-10.250 -0.5718 0.03631 0.02996 -0.1079 0.8534 0.0487
-10.000 -0.5555 0.03444 0.02786 -0.1087 0.8455 0.0490
-9.750 -0.5377 0.03280 0.02597 -0.1092 0.8381 0.0494
-9.500 -0.5179 0.03132 0.02424 -0.1095 0.8311 0.0497
-9.250 -0.4966 0.02997 0.02267 -0.1099 0.8232 0.0501
-9.000 -0.4745 0.02878 0.02123 -0.1100 0.8164 0.0505
-8.750 -0.4510 0.02772 0.01993 -0.1102 0.8093 0.0510
-8.500 -0.4269 0.02673 0.01874 -0.1103 0.8015 0.0515
-8.250 -0.4024 0.02570 0.01766 -0.1102 0.7947 0.0519
-8.000 -0.3769 0.02481 0.01674 -0.1102 0.7861 0.0524
-7.750 -0.3514 0.02400 0.01586 -0.1101 0.7780 0.0529
-7.500 -0.3254 0.02324 0.01502 -0.1101 0.7700 0.0533
-7.250 -0.2991 0.02251 0.01422 -0.1101 0.7613 0.0538
-7.000 -0.2727 0.02184 0.01345 -0.1100 0.7534 0.0543
-6.750 -0.2459 0.02119 0.01275 -0.1100 0.7435 0.0548
-6.500 -0.2191 0.02059 0.01205 -0.1099 0.7347 0.0553
-6.250 -0.1919 0.02005 0.01143 -0.1099 0.7244 0.0559
-6.000 -0.1645 0.01957 0.01084 -0.1098 0.7157 0.0566
-5.750 -0.1374 0.01899 0.01024 -0.1098 0.7060 0.0573
-5.500 -0.1105 0.01842 0.00967 -0.1097 0.6976 0.0581
-5.250 -0.0830 0.01794 0.00918 -0.1098 0.6881 0.0588
-4.750 -0.0275 0.01714 0.00828 -0.1098 0.6716 0.0602
-4.500 0.0005 0.01680 0.00786 -0.1099 0.6644 0.0610
-4.250 0.0288 0.01650 0.00750 -0.1100 0.6578 0.0618
-4.000 0.0574 0.01623 0.00717 -0.1101 0.6507 0.0627
-3.750 0.0857 0.01591 0.00679 -0.1102 0.6442 0.0638
-3.500 0.1144 0.01564 0.00647 -0.1104 0.6383 0.0653
-3.250 0.1433 0.01541 0.00620 -0.1106 0.6323 0.0670
-3.000 0.1721 0.01524 0.00595 -0.1107 0.6266 0.0686
-2.500 0.2301 0.01487 0.00550 -0.1110 0.6159 0.0729
-2.250 0.2591 0.01471 0.00533 -0.1112 0.6105 0.0765
-2.000 0.2880 0.01446 0.00513 -0.1114 0.6053 0.0878
-1.750 0.3168 0.01416 0.00506 -0.1117 0.6005 0.1493
-1.500 0.3459 0.01406 0.00506 -0.1119 0.5950 0.1729
-1.250 0.3748 0.01401 0.00503 -0.1120 0.5893 0.1902
-1.000 0.4037 0.01399 0.00500 -0.1121 0.5841 0.2035
-0.750 0.4326 0.01398 0.00499 -0.1122 0.5792 0.2175
-0.500 0.4616 0.01393 0.00502 -0.1123 0.5737 0.2354
-0.250 0.4904 0.01390 0.00504 -0.1124 0.5685 0.2531
0.000 0.5193 0.01390 0.00505 -0.1125 0.5639 0.2727
0.250 0.5481 0.01390 0.00510 -0.1126 0.5594 0.2950
0.500 0.5769 0.01387 0.00518 -0.1127 0.5541 0.3187
0.750 0.6056 0.01387 0.00525 -0.1128 0.5489 0.3438
1.000 0.6343 0.01389 0.00530 -0.1129 0.5443 0.3706
1.250 0.6628 0.01389 0.00538 -0.1129 0.5398 0.4001
1.500 0.6913 0.01387 0.00551 -0.1130 0.5342 0.4333
1.750 0.7195 0.01385 0.00563 -0.1130 0.5289 0.4719
2.000 0.7476 0.01387 0.00574 -0.1129 0.5242 0.5123
2.250 0.7757 0.01391 0.00587 -0.1128 0.5192 0.5475
2.500 0.8036 0.01394 0.00601 -0.1127 0.5133 0.5800
2.750 0.8310 0.01397 0.00611 -0.1125 0.5078 0.6126
3.000 0.8577 0.01398 0.00622 -0.1121 0.5030 0.6509
3.250 0.8828 0.01392 0.00643 -0.1113 0.4970 0.7112
3.500 0.9051 0.01384 0.00657 -0.1098 0.4906 0.8012
3.750 0.9259 0.01364 0.00652 -0.1077 0.4846 0.9781
4.000 0.9542 0.01379 0.00666 -0.1078 0.4763 1.0000
4.250 0.9814 0.01397 0.00675 -0.1077 0.4677 1.0000
4.500 1.0086 0.01415 0.00691 -0.1076 0.4589 1.0000
4.750 1.0354 0.01436 0.00705 -0.1074 0.4503 1.0000
5.000 1.0620 0.01457 0.00724 -0.1072 0.4412 1.0000
5.250 1.0878 0.01481 0.00741 -0.1069 0.4311 1.0000
5.500 1.1137 0.01506 0.00763 -0.1066 0.4210 1.0000
6.000 1.1641 0.01562 0.00812 -0.1058 0.4017 1.0000
6.250 1.1882 0.01595 0.00839 -0.1053 0.3914 1.0000
6.500 1.2119 0.01630 0.00869 -0.1047 0.3801 1.0000
6.750 1.2352 0.01667 0.00903 -0.1041 0.3700 1.0000
7.000 1.2571 0.01710 0.00939 -0.1033 0.3595 1.0000
7.250 1.2794 0.01751 0.00979 -0.1025 0.3489 1.0000
7.500 1.2997 0.01800 0.01022 -0.1015 0.3384 1.0000
7.750 1.3196 0.01851 0.01069 -0.1005 0.3273 1.0000
8.000 1.3390 0.01902 0.01119 -0.0994 0.3179 1.0000
8.250 1.3563 0.01961 0.01174 -0.0980 0.3082 1.0000
8.500 1.3727 0.02017 0.01230 -0.0965 0.2992 1.0000
8.750 1.3865 0.02089 0.01299 -0.0948 0.2897 1.0000
9.000 1.4014 0.02160 0.01372 -0.0933 0.2788 1.0000
9.250 1.4141 0.02247 0.01457 -0.0917 0.2690 1.0000
9.500 1.4261 0.02343 0.01551 -0.0902 0.2577 1.0000
9.750 1.4371 0.02450 0.01657 -0.0887 0.2453 1.0000
10.000 1.4448 0.02584 0.01788 -0.0871 0.2303 1.0000
10.250 1.4493 0.02748 0.01945 -0.0855 0.2125 1.0000
10.500 1.4515 0.02939 0.02128 -0.0838 0.1941 1.0000
10.750 1.4502 0.03166 0.02346 -0.0822 0.1723 1.0000
11.000 1.4415 0.03468 0.02633 -0.0804 0.1440 1.0000
11.250 1.4240 0.03866 0.03009 -0.0785 0.1099 1.0000
11.500 1.4155 0.04201 0.03336 -0.0773 0.0956 1.0000
11.750 1.4131 0.04488 0.03623 -0.0765 0.0895 1.0000
12.000 1.4117 0.04771 0.03910 -0.0757 0.0857 1.0000
12.250 1.4119 0.05047 0.04191 -0.0752 0.0833 1.0000
12.500 1.4129 0.05324 0.04474 -0.0748 0.0814 1.0000
12.750 1.4131 0.05615 0.04773 -0.0744 0.0798 1.0000
13.000 1.4124 0.05922 0.05088 -0.0742 0.0784 1.0000
13.250 1.4107 0.06247 0.05419 -0.0740 0.0773 1.0000
13.500 1.4079 0.06590 0.05769 -0.0739 0.0764 1.0000
13.750 1.4066 0.06920 0.06107 -0.0739 0.0756 1.0000
14.000 1.4065 0.07236 0.06431 -0.0739 0.0749 1.0000
14.250 1.4059 0.07563 0.06768 -0.0740 0.0742 1.0000
14.500 1.4053 0.07894 0.07108 -0.0741 0.0736 1.0000
14.750 1.4046 0.08228 0.07449 -0.0743 0.0730 1.0000
15.000 1.4039 0.08563 0.07793 -0.0745 0.0724 1.0000
15.250 1.4038 0.08892 0.08128 -0.0748 0.0718 1.0000
15.500 1.4038 0.09222 0.08464 -0.0751 0.0713 1.0000
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Polar data table (+)
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