GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: GOE 113 AIRFOIL (goe113-il) Reynolds number: 50,000 Max Cl/Cd: 37.7 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe113-il-50000.txt Download as CSV file: xf-goe113-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4797 0.10744 0.10089 -0.0058 1.0000 0.1469
-7.750 -0.4667 0.10254 0.09601 -0.0043 1.0000 0.1572
-7.500 -0.4766 0.10134 0.09494 -0.0099 1.0000 0.1613
-7.250 -0.4648 0.09666 0.09029 -0.0076 1.0000 0.1736
-7.000 -0.4566 0.09268 0.08637 -0.0073 1.0000 0.1837
-6.750 -0.4522 0.08926 0.08301 -0.0088 1.0000 0.1945
-6.500 -0.4467 0.08587 0.07968 -0.0100 1.0000 0.2079
-6.250 -0.4395 0.08248 0.07635 -0.0102 1.0000 0.2246
-6.000 -0.4371 0.08000 0.07390 -0.0144 1.0000 0.2447
-5.750 -0.4262 0.07591 0.06990 -0.0099 1.0000 0.2673
-5.500 -0.4195 0.07274 0.06679 -0.0085 1.0000 0.2953
-5.250 -0.4156 0.06987 0.06395 -0.0076 1.0000 0.3302
-5.000 -0.4097 0.06687 0.06104 -0.0026 1.0000 0.3727
-4.750 -0.0924 0.04378 0.03696 -0.0067 1.0000 1.0000
-4.500 -0.0800 0.04143 0.03465 -0.0082 1.0000 1.0000
-4.250 -0.0702 0.03933 0.03259 -0.0089 1.0000 0.9986
-4.000 -0.1255 0.04156 0.03506 0.0055 1.0000 0.9579
-3.750 -0.1743 0.04273 0.03647 0.0162 1.0000 0.9038
-3.500 -0.2211 0.04324 0.03723 0.0242 1.0000 0.8477
-3.250 -0.2686 0.04331 0.03761 0.0315 1.0000 0.8030
-3.000 -0.1786 0.03529 0.02717 -0.0423 1.0000 0.2151
-2.750 -0.1440 0.03230 0.02356 -0.0433 1.0000 0.1839
-2.500 -0.1124 0.02989 0.02052 -0.0435 1.0000 0.1666
-2.250 -0.0822 0.02787 0.01794 -0.0432 1.0000 0.1524
-2.000 -0.0537 0.02628 0.01587 -0.0425 1.0000 0.1450
-1.750 -0.0270 0.02479 0.01409 -0.0419 1.0000 0.1438
-1.500 -0.0013 0.02362 0.01266 -0.0413 1.0000 0.1498
-1.250 0.0237 0.02255 0.01145 -0.0403 1.0000 0.1513
-1.000 0.0499 0.02172 0.01043 -0.0396 1.0000 0.1539
-0.750 0.0747 0.02110 0.00969 -0.0389 1.0000 0.1588
-0.500 0.0990 0.02057 0.00914 -0.0386 1.0000 0.1672
-0.250 0.1237 0.02017 0.00878 -0.0387 1.0000 0.1875
0.000 0.1466 0.01724 0.00863 -0.0380 1.0000 0.9111
0.250 0.1630 0.01778 0.00859 -0.0365 1.0000 1.0000
0.500 0.1827 0.01842 0.00875 -0.0361 1.0000 1.0000
0.750 0.2017 0.01913 0.00917 -0.0361 1.0000 1.0000
1.000 0.2203 0.01990 0.00973 -0.0363 1.0000 1.0000
1.250 0.2388 0.02073 0.01039 -0.0366 1.0000 1.0000
1.500 0.2709 0.02165 0.01116 -0.0395 0.9945 1.0000
1.750 0.3235 0.02256 0.01192 -0.0460 0.9773 1.0000
2.000 0.3765 0.02336 0.01266 -0.0522 0.9593 1.0000
2.250 0.4330 0.02388 0.01318 -0.0583 0.9360 1.0000
2.500 0.4946 0.02393 0.01333 -0.0643 0.9079 1.0000
2.750 0.5521 0.02377 0.01329 -0.0690 0.8830 1.0000
3.000 0.5999 0.02366 0.01332 -0.0717 0.8618 1.0000
3.250 0.6380 0.02364 0.01351 -0.0725 0.8384 1.0000
3.500 0.6756 0.02343 0.01346 -0.0726 0.8138 1.0000
3.750 0.7086 0.02317 0.01333 -0.0712 0.7857 1.0000
4.000 0.7374 0.02291 0.01319 -0.0689 0.7545 1.0000
4.250 0.7639 0.02268 0.01312 -0.0659 0.7209 1.0000
4.500 0.7870 0.02267 0.01316 -0.0627 0.6824 1.0000
4.750 0.8095 0.02275 0.01325 -0.0595 0.6398 1.0000
5.000 0.8305 0.02310 0.01359 -0.0565 0.5904 1.0000
5.250 0.8531 0.02359 0.01398 -0.0539 0.5437 1.0000
5.500 0.8761 0.02436 0.01481 -0.0520 0.5026 1.0000
5.750 0.8974 0.02478 0.01522 -0.0499 0.4628 1.0000
6.000 0.9145 0.02453 0.01483 -0.0469 0.4154 1.0000
6.250 0.9365 0.02511 0.01553 -0.0454 0.3850 1.0000
6.500 0.9539 0.02530 0.01576 -0.0432 0.3445 1.0000
6.750 0.9645 0.02561 0.01586 -0.0404 0.2821 1.0000
7.000 0.9737 0.02719 0.01715 -0.0380 0.1878 1.0000
7.250 0.9846 0.02990 0.01935 -0.0356 0.1159 1.0000
7.500 1.0006 0.03218 0.02151 -0.0335 0.0962 1.0000
7.750 1.0204 0.03434 0.02376 -0.0317 0.0864 1.0000
8.000 1.0410 0.03657 0.02603 -0.0302 0.0787 1.0000
8.250 1.0623 0.03886 0.02843 -0.0289 0.0723 1.0000
8.500 1.0871 0.04220 0.03186 -0.0279 0.0699 1.0000
8.750 1.1089 0.04602 0.03617 -0.0267 0.0693 1.0000
9.000 1.1252 0.05029 0.04095 -0.0253 0.0694 1.0000
9.250 1.1366 0.05492 0.04604 -0.0238 0.0699 1.0000
9.500 1.1432 0.05965 0.05136 -0.0222 0.0706 1.0000
9.750 1.1310 0.06484 0.05735 -0.0201 0.0725 1.0000
10.000 1.1110 0.07050 0.06347 -0.0187 0.0741 1.0000
10.250 1.0883 0.07575 0.06904 -0.0179 0.0753 1.0000
10.500 1.0629 0.08098 0.07444 -0.0181 0.0762 1.0000
10.750 1.0361 0.08724 0.08084 -0.0209 0.0771 1.0000
11.000 1.0098 0.09495 0.08861 -0.0261 0.0782 1.0000
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Polar data table (+)
Polar graphs
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