GOE 459 AIRFOIL (goe459-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 459 AIRFOIL (goe459-il) Reynolds number: 100,000 Max Cl/Cd: 40.46 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe459-il-100000.txt Download as CSV file: xf-goe459-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 459 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
------ -------- --------- --------- -------- -------- --------
-13.000 -0.7951 0.08291 0.07730 -0.0459 1.0000 0.0816
-12.750 -0.8261 0.07562 0.06993 -0.0490 1.0000 0.0809
-12.500 -0.8572 0.06958 0.06377 -0.0507 1.0000 0.0802
-12.250 -0.8891 0.06454 0.05855 -0.0505 1.0000 0.0798
-12.000 -0.9188 0.06047 0.05429 -0.0485 1.0000 0.0796
-11.750 -0.9476 0.05716 0.05077 -0.0446 1.0000 0.0798
-11.500 -0.9728 0.05432 0.04769 -0.0394 1.0000 0.0800
-11.250 -0.9956 0.05139 0.04441 -0.0342 1.0000 0.0809
-11.000 -1.0151 0.04893 0.04149 -0.0286 1.0000 0.0820
-10.750 -0.9951 0.04657 0.03937 -0.0285 1.0000 0.0870
-10.500 -0.9951 0.04464 0.03722 -0.0251 1.0000 0.0906
-10.250 -1.0024 0.04243 0.03453 -0.0204 1.0000 0.0935
-10.000 -0.9962 0.04005 0.03201 -0.0178 1.0000 0.0972
-9.750 -0.9812 0.03897 0.03094 -0.0161 1.0000 0.1021
-9.500 -0.9770 0.03746 0.02905 -0.0125 1.0000 0.1063
-9.250 -0.9653 0.03567 0.02714 -0.0103 1.0000 0.1105
-9.000 -0.9494 0.03465 0.02609 -0.0086 1.0000 0.1154
-8.750 -0.9382 0.03334 0.02447 -0.0058 1.0000 0.1194
-8.500 -0.9243 0.03171 0.02254 -0.0035 1.0000 0.1223
-8.250 -0.9033 0.03025 0.02115 -0.0026 1.0000 0.1261
-7.750 -0.8668 0.02793 0.01847 0.0009 1.0000 0.1337
-7.500 -0.8455 0.02659 0.01702 0.0020 1.0000 0.1372
-7.250 -0.8241 0.02550 0.01602 0.0031 1.0000 0.1419
-7.000 -0.8054 0.02463 0.01504 0.0048 1.0000 0.1480
-6.750 -0.7852 0.02354 0.01400 0.0061 1.0000 0.1544
-6.500 -0.7668 0.02267 0.01313 0.0079 1.0000 0.1630
-6.250 -0.7500 0.02168 0.01229 0.0099 1.0000 0.1734
-6.000 -0.7359 0.02074 0.01149 0.0123 1.0000 0.1902
-5.750 -0.7254 0.01978 0.01079 0.0154 1.0000 0.2187
-5.500 -0.7165 0.01894 0.01021 0.0187 1.0000 0.2597
-5.250 -0.7066 0.01824 0.00973 0.0219 1.0000 0.3008
-5.000 -0.6957 0.01761 0.00932 0.0249 1.0000 0.3393
-4.750 -0.6844 0.01707 0.00902 0.0278 1.0000 0.3789
-4.500 -0.6728 0.01666 0.00884 0.0308 1.0000 0.4220
-4.250 -0.6608 0.01634 0.00872 0.0338 1.0000 0.4683
-4.000 -0.6485 0.01607 0.00863 0.0368 1.0000 0.5143
-3.750 -0.6361 0.01587 0.00861 0.0398 1.0000 0.5604
-3.500 -0.6235 0.01576 0.00868 0.0429 1.0000 0.6056
-3.250 -0.6099 0.01572 0.00881 0.0459 1.0000 0.6472
-3.000 -0.5964 0.01572 0.00891 0.0489 1.0000 0.6865
-2.750 -0.5806 0.01578 0.00909 0.0515 1.0000 0.7228
-2.500 -0.5639 0.01590 0.00931 0.0540 1.0000 0.7585
-2.250 -0.5451 0.01607 0.00956 0.0561 1.0000 0.7921
-2.000 -0.5222 0.01631 0.00983 0.0573 1.0000 0.8231
-1.750 -0.4934 0.01664 0.01017 0.0572 1.0000 0.8522
-1.500 -0.4373 0.01739 0.01086 0.0517 0.9936 0.8789
-1.250 -0.3702 0.01830 0.01170 0.0445 0.9881 0.8991
-1.000 -0.2950 0.01938 0.01270 0.0358 0.9840 0.9155
-0.750 -0.2258 0.02011 0.01335 0.0278 0.9781 0.9279
-0.500 -0.1510 0.02070 0.01388 0.0185 0.9726 0.9354
-0.250 -0.0824 0.02101 0.01415 0.0103 0.9638 0.9452
0.000 0.0001 0.02116 0.01428 0.0000 0.9548 0.9548
0.250 0.0824 0.02102 0.01415 -0.0103 0.9453 0.9640
0.500 0.1510 0.02070 0.01388 -0.0185 0.9354 0.9726
0.750 0.2261 0.02010 0.01334 -0.0279 0.9279 0.9781
1.000 0.2952 0.01937 0.01269 -0.0358 0.9154 0.9839
1.250 0.3702 0.01829 0.01169 -0.0445 0.8990 0.9880
1.500 0.4375 0.01739 0.01086 -0.0518 0.8792 0.9935
1.750 0.4935 0.01663 0.01016 -0.0572 0.8521 1.0000
2.000 0.5221 0.01631 0.00984 -0.0573 0.8232 1.0000
2.250 0.5450 0.01607 0.00955 -0.0561 0.7919 1.0000
2.500 0.5637 0.01589 0.00929 -0.0540 0.7579 1.0000
2.750 0.5805 0.01578 0.00908 -0.0515 0.7225 1.0000
3.000 0.5963 0.01572 0.00889 -0.0488 0.6862 1.0000
3.250 0.6097 0.01572 0.00880 -0.0459 0.6468 1.0000
3.500 0.6234 0.01576 0.00867 -0.0429 0.6055 1.0000
3.750 0.6360 0.01587 0.00862 -0.0398 0.5604 1.0000
4.000 0.6485 0.01607 0.00863 -0.0368 0.5147 1.0000
4.250 0.6607 0.01633 0.00871 -0.0338 0.4681 1.0000
4.500 0.6727 0.01665 0.00883 -0.0308 0.4219 1.0000
4.750 0.6843 0.01707 0.00902 -0.0278 0.3789 1.0000
5.000 0.6956 0.01760 0.00931 -0.0249 0.3394 1.0000
5.250 0.7066 0.01824 0.00972 -0.0219 0.3012 1.0000
5.500 0.7165 0.01894 0.01021 -0.0187 0.2599 1.0000
5.750 0.7253 0.01979 0.01079 -0.0154 0.2180 1.0000
6.000 0.7359 0.02074 0.01149 -0.0123 0.1906 1.0000
6.250 0.7500 0.02168 0.01228 -0.0099 0.1736 1.0000
6.500 0.7668 0.02267 0.01313 -0.0079 0.1629 1.0000
6.750 0.7853 0.02354 0.01401 -0.0062 0.1547 1.0000
7.000 0.8054 0.02463 0.01504 -0.0048 0.1481 1.0000
7.250 0.8243 0.02551 0.01603 -0.0031 0.1421 1.0000
7.500 0.8455 0.02658 0.01702 -0.0020 0.1372 1.0000
7.750 0.8669 0.02794 0.01847 -0.0009 0.1337 1.0000
8.000 0.8854 0.02911 0.01987 0.0009 0.1302 1.0000
8.250 0.9033 0.03025 0.02114 0.0026 0.1261 1.0000
8.500 0.9243 0.03171 0.02254 0.0035 0.1223 1.0000
8.750 0.9383 0.03335 0.02447 0.0058 0.1195 1.0000
9.000 0.9494 0.03463 0.02607 0.0086 0.1154 1.0000
9.250 0.9653 0.03567 0.02713 0.0103 0.1105 1.0000
9.500 0.9767 0.03743 0.02902 0.0125 0.1062 1.0000
9.750 0.9812 0.03894 0.03091 0.0161 0.1020 1.0000
10.000 0.9965 0.04002 0.03196 0.0177 0.0970 1.0000
10.250 1.0016 0.04236 0.03448 0.0205 0.0933 1.0000
10.500 0.9951 0.04458 0.03716 0.0251 0.0904 1.0000
10.750 0.9959 0.04652 0.03930 0.0284 0.0868 1.0000
11.000 1.0154 0.04899 0.04154 0.0285 0.0821 1.0000
11.250 0.9953 0.05139 0.04441 0.0342 0.0809 1.0000
11.500 0.9742 0.05437 0.04773 0.0392 0.0801 1.0000
11.750 0.9467 0.05718 0.05079 0.0446 0.0796 1.0000
12.000 0.9178 0.06054 0.05438 0.0485 0.0795 1.0000
12.250 0.8884 0.06461 0.05863 0.0505 0.0797 1.0000
12.500 0.8577 0.06963 0.06382 0.0506 0.0803 1.0000
12.750 0.8250 0.07584 0.07015 0.0488 0.0809 1.0000
13.000 0.7969 0.08286 0.07725 0.0459 0.0816 1.0000
13.250 0.4916 0.13544 0.13017 0.0179 0.1652 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 459 AIRFOIL (goe459-il)