GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 113 AIRFOIL (goe113-il) Reynolds number: 500,000 Max Cl/Cd: 86.79 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe113-il-500000.txt Download as CSV file: xf-goe113-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 113 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4586 0.09905 0.09684 -0.0077 1.0000 0.0119
-8.000 -0.4540 0.09550 0.09332 -0.0099 1.0000 0.0119
-7.750 -0.4502 0.09195 0.08979 -0.0121 1.0000 0.0119
-7.500 -0.4433 0.08798 0.08584 -0.0158 1.0000 0.0120
-7.250 -0.4334 0.08367 0.08153 -0.0201 1.0000 0.0120
-7.000 -0.4287 0.07759 0.07547 -0.0245 1.0000 0.0122
-6.750 -0.4251 0.07355 0.07144 -0.0247 1.0000 0.0125
-6.500 -0.4161 0.07036 0.06825 -0.0257 1.0000 0.0129
-6.250 -0.4053 0.06697 0.06484 -0.0277 1.0000 0.0134
-6.000 -0.3927 0.06337 0.06122 -0.0300 1.0000 0.0141
-5.750 -0.3783 0.05963 0.05743 -0.0322 1.0000 0.0149
-5.500 -0.3604 0.05583 0.05356 -0.0346 1.0000 0.0164
-5.250 -0.3283 0.05290 0.05043 -0.0374 1.0000 0.0178
-5.000 -0.3101 0.04946 0.04686 -0.0379 1.0000 0.0179
-4.750 -0.2923 0.04599 0.04324 -0.0380 1.0000 0.0180
-4.500 -0.2742 0.04257 0.03967 -0.0379 1.0000 0.0180
-4.250 -0.2558 0.03921 0.03611 -0.0375 1.0000 0.0180
-4.000 -0.2216 0.03495 0.03157 -0.0402 0.9981 0.0181
-2.750 -0.0166 0.00509 0.00020 -0.0555 0.9723 0.0281
-1.750 0.1065 0.01216 0.00640 -0.0594 0.9496 0.0392
-1.500 0.1389 0.01110 0.00522 -0.0600 0.9302 0.0391
-1.250 0.1677 0.00990 0.00392 -0.0598 0.9027 0.0399
-1.000 0.1937 0.00930 0.00318 -0.0590 0.8624 0.0404
-0.750 0.2184 0.00903 0.00269 -0.0580 0.8150 0.0408
-0.500 0.2437 0.00884 0.00230 -0.0573 0.7751 0.0423
-0.250 0.2700 0.00873 0.00201 -0.0568 0.7426 0.0426
0.000 0.2969 0.00865 0.00179 -0.0565 0.7155 0.0425
0.250 0.3240 0.00858 0.00160 -0.0563 0.6930 0.0434
0.500 0.3512 0.00855 0.00145 -0.0561 0.6702 0.0458
0.750 0.3785 0.00855 0.00134 -0.0559 0.6480 0.0498
1.000 0.4058 0.00857 0.00129 -0.0557 0.6280 0.0599
1.250 0.4284 0.00626 0.00138 -0.0551 0.6115 1.0000
1.500 0.4557 0.00638 0.00138 -0.0549 0.5937 1.0000
1.750 0.4830 0.00650 0.00140 -0.0548 0.5774 1.0000
2.000 0.5104 0.00663 0.00145 -0.0547 0.5629 1.0000
2.250 0.5377 0.00675 0.00150 -0.0546 0.5482 1.0000
2.500 0.5651 0.00688 0.00157 -0.0546 0.5333 1.0000
2.750 0.5923 0.00702 0.00165 -0.0545 0.5177 1.0000
3.000 0.6195 0.00717 0.00176 -0.0544 0.4980 1.0000
3.250 0.6457 0.00744 0.00184 -0.0541 0.4525 1.0000
3.500 0.6713 0.00780 0.00195 -0.0539 0.3927 1.0000
3.750 0.6960 0.00838 0.00217 -0.0536 0.3109 1.0000
4.000 0.7204 0.00905 0.00249 -0.0533 0.2438 1.0000
4.250 0.7460 0.00950 0.00279 -0.0531 0.2079 1.0000
4.500 0.7716 0.00993 0.00304 -0.0529 0.1686 1.0000
4.750 0.7960 0.01060 0.00337 -0.0526 0.1121 1.0000
5.000 0.8176 0.01180 0.00408 -0.0519 0.0225 1.0000
5.250 0.8437 0.01220 0.00457 -0.0515 0.0179 1.0000
5.500 0.8695 0.01267 0.00523 -0.0510 0.0162 1.0000
5.750 0.8946 0.01325 0.00594 -0.0505 0.0156 1.0000
6.000 0.9190 0.01394 0.00675 -0.0499 0.0148 1.0000
6.250 0.9427 0.01471 0.00760 -0.0492 0.0134 1.0000
6.500 0.9653 0.01564 0.00865 -0.0483 0.0128 1.0000
6.750 0.9867 0.01674 0.00989 -0.0473 0.0122 1.0000
7.000 1.0070 0.01805 0.01130 -0.0460 0.0118 1.0000
7.250 1.0271 0.01956 0.01294 -0.0445 0.0120 1.0000
7.500 1.0474 0.02169 0.01525 -0.0427 0.0137 1.0000
9.000 1.1590 0.03633 0.03130 -0.0341 0.0144 1.0000
9.250 1.1716 0.03854 0.03377 -0.0324 0.0132 1.0000
9.500 1.1804 0.04111 0.03659 -0.0306 0.0123 1.0000
9.750 1.1862 0.04389 0.03959 -0.0287 0.0117 1.0000
10.000 1.1895 0.04683 0.04272 -0.0268 0.0113 1.0000
10.250 1.1912 0.05004 0.04607 -0.0249 0.0109 1.0000
10.500 1.1853 0.05352 0.04975 -0.0227 0.0106 1.0000
10.750 1.1723 0.05685 0.05326 -0.0197 0.0105 1.0000
11.000 1.1558 0.06034 0.05691 -0.0174 0.0104 1.0000
11.250 1.1389 0.06417 0.06092 -0.0167 0.0103 1.0000
11.500 1.1213 0.06871 0.06562 -0.0176 0.0103 1.0000
11.750 1.1042 0.07387 0.07094 -0.0200 0.0103 1.0000
12.000 1.0885 0.07947 0.07670 -0.0236 0.0104 1.0000
12.250 1.0713 0.08619 0.08357 -0.0284 0.0104 1.0000
12.500 1.0537 0.09386 0.09141 -0.0343 0.0106 1.0000
12.750 1.0338 0.10333 0.10103 -0.0415 0.0108 1.0000
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Polar data table (+)
Polar graphs
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