GOE 123 AIRFOIL (goe123-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 123 AIRFOIL (goe123-il) Reynolds number: 100,000 Max Cl/Cd: 65.76 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe123-il-100000.txt Download as CSV file: xf-goe123-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 123 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3083 0.10234 0.09762 -0.0264 1.0000 0.0495
-7.500 -0.3197 0.10239 0.09778 -0.0251 1.0000 0.0499
-7.250 -0.3291 0.10256 0.09803 -0.0252 1.0000 0.0502
-7.000 -0.3347 0.10263 0.09817 -0.0266 1.0000 0.0505
-6.750 -0.3344 0.10234 0.09788 -0.0293 1.0000 0.0506
-6.500 -0.3356 0.09762 0.09325 -0.0266 1.0000 0.0511
-6.250 -0.3377 0.09247 0.08816 -0.0209 1.0000 0.0520
-6.000 -0.3380 0.08957 0.08530 -0.0186 1.0000 0.0530
-5.750 -0.3368 0.08720 0.08295 -0.0177 1.0000 0.0542
-5.500 -0.3335 0.08492 0.08070 -0.0177 1.0000 0.0557
-5.250 -0.3274 0.08265 0.07844 -0.0187 1.0000 0.0579
-5.000 -0.3107 0.08106 0.07680 -0.0233 1.0000 0.0615
-4.750 -0.2714 0.08093 0.07644 -0.0354 1.0000 0.0632
-4.500 -0.2754 0.07555 0.07118 -0.0312 1.0000 0.0640
-4.250 -0.2720 0.07200 0.06766 -0.0291 1.0000 0.0653
-4.000 -0.2609 0.06917 0.06484 -0.0293 1.0000 0.0674
-3.750 -0.2384 0.06637 0.06198 -0.0324 0.9991 0.0709
-3.500 -0.1669 0.06341 0.05867 -0.0473 0.9935 0.0774
-3.250 -0.1472 0.05892 0.05423 -0.0483 0.9886 0.0807
-3.000 -0.0676 0.05861 0.05335 -0.0622 0.9828 0.0909
-2.750 -0.0519 0.05283 0.04775 -0.0629 0.9776 0.0928
-2.500 -0.0153 0.04971 0.04458 -0.0670 0.9724 0.0987
-2.250 0.0341 0.04723 0.04183 -0.0737 0.9668 0.1076
-2.000 0.0844 0.04623 0.04047 -0.0796 0.9598 0.1198
-1.750 0.1207 0.04222 0.03654 -0.0835 0.9559 0.1252
-1.500 0.1604 0.04055 0.03464 -0.0872 0.9480 0.1354
-1.250 0.2078 0.03851 0.03239 -0.0922 0.9426 0.1484
-1.000 0.2459 0.03661 0.03037 -0.0954 0.9354 0.1625
-0.750 0.2902 0.03478 0.02840 -0.0996 0.9287 0.1893
-0.500 0.3308 0.03283 0.02637 -0.1029 0.9206 0.2205
-0.250 0.3769 0.03068 0.02419 -0.1071 0.9131 0.2791
0.000 0.4172 0.02889 0.02235 -0.1098 0.9036 0.3230
0.500 0.5306 0.02774 0.01981 -0.1153 0.8877 0.1334
0.750 0.5802 0.02581 0.01767 -0.1184 0.8824 0.1167
1.000 0.6140 0.02472 0.01640 -0.1187 0.8720 0.1088
1.250 0.6586 0.02357 0.01498 -0.1203 0.8662 0.1049
1.500 0.6877 0.02274 0.01415 -0.1199 0.8543 0.1058
1.750 0.7183 0.02195 0.01341 -0.1196 0.8433 0.1098
2.000 0.7539 0.02100 0.01251 -0.1196 0.8359 0.1228
2.250 0.7809 0.02037 0.01191 -0.1184 0.8229 0.1387
2.500 0.8098 0.01919 0.01099 -0.1177 0.8102 0.2388
2.750 0.8341 0.01761 0.01051 -0.1157 0.7973 1.0000
3.000 0.8614 0.01735 0.01014 -0.1144 0.7831 1.0000
3.250 0.8884 0.01706 0.00977 -0.1131 0.7679 1.0000
3.500 0.9153 0.01675 0.00938 -0.1117 0.7513 1.0000
3.750 0.9424 0.01642 0.00897 -0.1103 0.7335 1.0000
4.000 0.9672 0.01629 0.00882 -0.1088 0.7115 1.0000
4.250 0.9933 0.01612 0.00858 -0.1075 0.6896 1.0000
4.500 1.0180 0.01613 0.00854 -0.1062 0.6638 1.0000
4.750 1.0429 0.01618 0.00851 -0.1049 0.6371 1.0000
5.000 1.0675 0.01633 0.00858 -0.1037 0.6086 1.0000
5.250 1.0916 0.01660 0.00872 -0.1024 0.5792 1.0000
5.500 1.1154 0.01700 0.00897 -0.1013 0.5507 1.0000
5.750 1.1391 0.01750 0.00932 -0.1002 0.5240 1.0000
6.000 1.1623 0.01806 0.00984 -0.0992 0.4985 1.0000
6.250 1.1855 0.01867 0.01041 -0.0983 0.4756 1.0000
6.500 1.2079 0.01922 0.01088 -0.0971 0.4518 1.0000
6.750 1.2286 0.01972 0.01137 -0.0958 0.4254 1.0000
7.000 1.2492 0.02024 0.01192 -0.0945 0.4009 1.0000
7.250 1.2691 0.02078 0.01239 -0.0932 0.3766 1.0000
7.500 1.2878 0.02130 0.01299 -0.0917 0.3502 1.0000
7.750 1.3051 0.02186 0.01358 -0.0900 0.3217 1.0000
8.000 1.3206 0.02248 0.01422 -0.0881 0.2884 1.0000
8.250 1.3308 0.02336 0.01503 -0.0855 0.2184 1.0000
8.500 1.3250 0.02653 0.01708 -0.0812 0.1027 1.0000
8.750 1.3271 0.02900 0.01934 -0.0778 0.0837 1.0000
9.000 1.3321 0.03102 0.02141 -0.0746 0.0744 1.0000
9.250 1.3342 0.03316 0.02348 -0.0714 0.0679 1.0000
9.500 1.3447 0.03479 0.02525 -0.0688 0.0633 1.0000
9.750 1.3566 0.03663 0.02713 -0.0667 0.0596 1.0000
10.000 1.3780 0.03937 0.02971 -0.0659 0.0558 1.0000
10.250 1.3951 0.04111 0.03172 -0.0643 0.0529 1.0000
10.500 1.4164 0.04338 0.03418 -0.0634 0.0503 1.0000
10.750 1.4411 0.04625 0.03722 -0.0630 0.0487 1.0000
11.000 1.4625 0.04950 0.04070 -0.0622 0.0478 1.0000
11.250 1.4787 0.05306 0.04448 -0.0611 0.0468 1.0000
11.500 1.4900 0.05828 0.05000 -0.0602 0.0456 1.0000
11.750 1.4852 0.06140 0.05351 -0.0569 0.0452 1.0000
12.000 1.4766 0.06448 0.05695 -0.0534 0.0450 1.0000
12.250 1.4671 0.06820 0.06098 -0.0503 0.0451 1.0000
12.500 1.4573 0.07256 0.06562 -0.0479 0.0454 1.0000
12.750 1.4458 0.07658 0.06993 -0.0456 0.0457 1.0000
13.000 1.4282 0.07958 0.07321 -0.0434 0.0460 1.0000
13.250 1.4057 0.08328 0.07723 -0.0420 0.0464 1.0000
13.500 1.3774 0.08816 0.08244 -0.0419 0.0470 1.0000
13.750 1.3408 0.09465 0.08927 -0.0439 0.0476 1.0000
14.000 1.3023 0.10284 0.09778 -0.0481 0.0483 1.0000
14.250 1.2578 0.11393 0.10913 -0.0555 0.0496 1.0000
14.500 1.2127 0.12815 0.12355 -0.0655 0.0523 1.0000
14.750 1.1877 0.13952 0.13496 -0.0726 0.0541 1.0000
15.000 1.1746 0.14838 0.14384 -0.0772 0.0553 1.0000
15.250 0.8867 0.15398 0.14959 -0.0648 0.0658 1.0000
15.500 0.8821 0.15871 0.15433 -0.0662 0.0667 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 123 AIRFOIL (goe123-il)