GOE 623 AIRFOIL (goe623-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 623 AIRFOIL (goe623-il) Reynolds number: 500,000 Max Cl/Cd: 102.63 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe623-il-500000.txt Download as CSV file: xf-goe623-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 623 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3951 0.09445 0.09217 -0.0376 1.0000 0.0312
-9.750 -0.5584 0.05335 0.05112 -0.0643 1.0000 0.0240
-9.500 -0.5953 0.04006 0.03719 -0.0744 0.9916 0.0236
-9.250 -0.5870 0.03245 0.02874 -0.0794 0.9846 0.0241
-9.000 -0.5692 0.02800 0.02366 -0.0814 0.9786 0.0249
-8.750 -0.5415 0.02612 0.02173 -0.0830 0.9743 0.0258
-8.500 -0.5090 0.02542 0.02105 -0.0847 0.9711 0.0267
-8.250 -0.4839 0.02396 0.01937 -0.0851 0.9634 0.0277
-8.000 -0.4568 0.02225 0.01735 -0.0857 0.9578 0.0286
-7.750 -0.4319 0.02114 0.01596 -0.0855 0.9495 0.0293
-7.500 -0.4071 0.01946 0.01396 -0.0854 0.9426 0.0302
-7.250 -0.3841 0.01803 0.01247 -0.0849 0.9338 0.0313
-7.000 -0.3582 0.01728 0.01163 -0.0847 0.9270 0.0323
-6.750 -0.3333 0.01653 0.01077 -0.0841 0.9186 0.0332
-6.500 -0.3077 0.01580 0.00991 -0.0836 0.9114 0.0340
-6.250 -0.2820 0.01525 0.00925 -0.0832 0.9030 0.0350
-6.000 -0.2556 0.01486 0.00872 -0.0827 0.8957 0.0358
-5.750 -0.2307 0.01383 0.00759 -0.0821 0.8875 0.0366
-5.500 -0.2056 0.01301 0.00670 -0.0816 0.8803 0.0377
-5.250 -0.1795 0.01252 0.00617 -0.0812 0.8723 0.0389
-5.000 -0.1531 0.01213 0.00573 -0.0808 0.8647 0.0402
-4.750 -0.1264 0.01175 0.00529 -0.0805 0.8562 0.0414
-4.500 -0.0997 0.01140 0.00486 -0.0801 0.8482 0.0424
-4.250 -0.0726 0.01110 0.00450 -0.0798 0.8395 0.0433
-4.000 -0.0461 0.01065 0.00398 -0.0794 0.8315 0.0453
-3.750 -0.0189 0.01035 0.00368 -0.0792 0.8219 0.0480
-3.500 0.0085 0.01014 0.00340 -0.0789 0.8129 0.0506
-3.250 0.0359 0.00988 0.00312 -0.0787 0.8024 0.0553
-3.000 0.0632 0.00960 0.00288 -0.0785 0.7918 0.0696
-2.750 0.0899 0.00923 0.00273 -0.0783 0.7813 0.1293
-2.500 0.1176 0.00913 0.00264 -0.0781 0.7692 0.1527
-2.250 0.1453 0.00903 0.00252 -0.0780 0.7561 0.1664
-2.000 0.1730 0.00894 0.00241 -0.0779 0.7427 0.1791
-1.750 0.2007 0.00886 0.00230 -0.0778 0.7291 0.1922
-1.500 0.2282 0.00877 0.00221 -0.0776 0.7151 0.2088
-1.250 0.2554 0.00863 0.00213 -0.0775 0.7014 0.2432
-1.000 0.2821 0.00841 0.00208 -0.0774 0.6876 0.3169
-0.750 0.3080 0.00808 0.00204 -0.0771 0.6731 0.4275
-0.500 0.3323 0.00759 0.00205 -0.0766 0.6596 0.5980
-0.250 0.3528 0.00703 0.00214 -0.0748 0.6481 0.7987
0.000 0.3750 0.00689 0.00223 -0.0728 0.6371 0.9189
0.250 0.4125 0.00698 0.00225 -0.0745 0.6245 0.9674
0.500 0.4581 0.00710 0.00226 -0.0783 0.6123 0.9908
0.750 0.5028 0.00718 0.00227 -0.0819 0.6023 1.0000
1.000 0.5279 0.00729 0.00228 -0.0814 0.5944 1.0000
1.250 0.5535 0.00735 0.00230 -0.0809 0.5861 1.0000
1.500 0.5789 0.00747 0.00233 -0.0804 0.5790 1.0000
1.750 0.6047 0.00753 0.00238 -0.0799 0.5715 1.0000
2.250 0.6562 0.00773 0.00248 -0.0790 0.5560 1.0000
2.500 0.6819 0.00786 0.00254 -0.0786 0.5481 1.0000
2.750 0.7083 0.00793 0.00261 -0.0783 0.5397 1.0000
3.000 0.7343 0.00806 0.00268 -0.0779 0.5319 1.0000
3.250 0.7609 0.00815 0.00277 -0.0776 0.5230 1.0000
3.500 0.7872 0.00828 0.00285 -0.0773 0.5143 1.0000
3.750 0.8136 0.00839 0.00294 -0.0770 0.5040 1.0000
4.000 0.8400 0.00851 0.00305 -0.0768 0.4926 1.0000
4.250 0.8661 0.00865 0.00316 -0.0764 0.4800 1.0000
4.500 0.8921 0.00880 0.00327 -0.0761 0.4658 1.0000
4.750 0.9179 0.00897 0.00341 -0.0758 0.4491 1.0000
5.000 0.9432 0.00919 0.00355 -0.0753 0.4284 1.0000
5.250 0.9676 0.00946 0.00373 -0.0748 0.4027 1.0000
5.500 0.9910 0.00983 0.00397 -0.0741 0.3730 1.0000
5.750 1.0138 0.01026 0.00425 -0.0733 0.3443 1.0000
6.000 1.0364 0.01072 0.00457 -0.0726 0.3190 1.0000
6.250 1.0595 0.01114 0.00490 -0.0719 0.2977 1.0000
6.500 1.0825 0.01156 0.00525 -0.0712 0.2796 1.0000
6.750 1.1057 0.01196 0.00559 -0.0705 0.2647 1.0000
7.000 1.1287 0.01236 0.00594 -0.0699 0.2518 1.0000
7.250 1.1513 0.01277 0.00630 -0.0691 0.2399 1.0000
7.500 1.1738 0.01319 0.00666 -0.0684 0.2284 1.0000
7.750 1.1972 0.01352 0.00699 -0.0678 0.2187 1.0000
8.000 1.2194 0.01392 0.00737 -0.0671 0.2104 1.0000
8.250 1.2423 0.01427 0.00773 -0.0664 0.2024 1.0000
8.500 1.2642 0.01466 0.00813 -0.0656 0.1957 1.0000
8.750 1.2868 0.01500 0.00849 -0.0649 0.1890 1.0000
9.000 1.3076 0.01543 0.00891 -0.0640 0.1817 1.0000
9.250 1.3296 0.01576 0.00928 -0.0632 0.1744 1.0000
9.500 1.3495 0.01621 0.00973 -0.0621 0.1674 1.0000
9.750 1.3701 0.01660 0.01014 -0.0612 0.1589 1.0000
10.000 1.3888 0.01704 0.01060 -0.0599 0.1486 1.0000
10.250 1.4040 0.01761 0.01112 -0.0581 0.1329 1.0000
10.500 1.4093 0.01880 0.01203 -0.0551 0.0924 1.0000
10.750 1.4007 0.02094 0.01385 -0.0505 0.0514 1.0000
11.000 1.4043 0.02243 0.01530 -0.0477 0.0379 1.0000
11.250 1.4103 0.02382 0.01669 -0.0454 0.0311 1.0000
11.500 1.4164 0.02526 0.01816 -0.0434 0.0274 1.0000
11.750 1.4253 0.02654 0.01951 -0.0418 0.0251 1.0000
12.000 1.4295 0.02825 0.02126 -0.0400 0.0232 1.0000
12.250 1.4357 0.02986 0.02296 -0.0385 0.0221 1.0000
12.500 1.4420 0.03152 0.02471 -0.0372 0.0211 1.0000
12.750 1.4467 0.03338 0.02666 -0.0360 0.0202 1.0000
13.000 1.4483 0.03560 0.02896 -0.0348 0.0195 1.0000
13.250 1.4442 0.03847 0.03193 -0.0337 0.0188 1.0000
13.500 1.4405 0.04144 0.03502 -0.0328 0.0183 1.0000
13.750 1.4423 0.04394 0.03763 -0.0323 0.0179 1.0000
14.000 1.4420 0.04677 0.04057 -0.0320 0.0174 1.0000
14.250 1.4406 0.04979 0.04370 -0.0318 0.0170 1.0000
14.500 1.4378 0.05308 0.04709 -0.0318 0.0165 1.0000
14.750 1.4337 0.05665 0.05075 -0.0320 0.0162 1.0000
15.000 1.4280 0.06050 0.05469 -0.0324 0.0159 1.0000
15.250 1.4202 0.06469 0.05898 -0.0329 0.0156 1.0000
15.500 1.4100 0.06931 0.06370 -0.0336 0.0154 1.0000
15.750 1.3983 0.07417 0.06866 -0.0344 0.0151 1.0000
16.000 1.3846 0.07928 0.07387 -0.0353 0.0149 1.0000
16.250 1.3801 0.08338 0.07807 -0.0362 0.0148 1.0000
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Polar data table (+)
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