GOE 624 AIRFOIL (goe624-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 624 AIRFOIL (goe624-il) Reynolds number: 500,000 Max Cl/Cd: 99.52 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe624-il-500000.txt Download as CSV file: xf-goe624-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 624 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6274 0.04102 0.03765 -0.1083 0.9487 0.0340
-11.750 -0.6437 0.03767 0.03398 -0.1068 0.9306 0.0341
-11.500 -0.6459 0.03491 0.03095 -0.1054 0.9182 0.0342
-11.250 -0.6395 0.03273 0.02856 -0.1043 0.9075 0.0344
-11.000 -0.6279 0.03095 0.02660 -0.1034 0.8986 0.0346
-10.750 -0.6139 0.02935 0.02481 -0.1025 0.8899 0.0349
-10.500 -0.5974 0.02793 0.02321 -0.1017 0.8820 0.0352
-10.250 -0.5793 0.02660 0.02172 -0.1009 0.8737 0.0355
-10.000 -0.5605 0.02533 0.02024 -0.1001 0.8665 0.0358
-9.750 -0.5398 0.02415 0.01890 -0.0996 0.8587 0.0362
-9.500 -0.5186 0.02310 0.01764 -0.0989 0.8518 0.0367
-9.250 -0.4962 0.02207 0.01642 -0.0984 0.8447 0.0372
-9.000 -0.4732 0.02108 0.01523 -0.0978 0.8370 0.0375
-8.750 -0.4495 0.02024 0.01419 -0.0973 0.8297 0.0379
-8.500 -0.4247 0.01951 0.01329 -0.0968 0.8211 0.0382
-8.000 -0.3763 0.01748 0.01105 -0.0959 0.8042 0.0391
-7.750 -0.3508 0.01687 0.01036 -0.0955 0.7963 0.0396
-7.500 -0.3244 0.01632 0.00977 -0.0953 0.7886 0.0402
-7.250 -0.2978 0.01585 0.00922 -0.0951 0.7810 0.0409
-7.000 -0.2710 0.01541 0.00869 -0.0948 0.7740 0.0417
-6.750 -0.2438 0.01497 0.00818 -0.0947 0.7662 0.0425
-6.500 -0.2167 0.01459 0.00767 -0.0944 0.7594 0.0431
-6.250 -0.1901 0.01398 0.00704 -0.0942 0.7519 0.0439
-6.000 -0.1632 0.01353 0.00657 -0.0940 0.7444 0.0449
-5.750 -0.1357 0.01321 0.00620 -0.0939 0.7372 0.0461
-5.500 -0.1079 0.01290 0.00585 -0.0938 0.7295 0.0474
-5.250 -0.0800 0.01266 0.00550 -0.0937 0.7230 0.0487
-5.000 -0.0524 0.01226 0.00514 -0.0937 0.7156 0.0508
-4.750 -0.0243 0.01205 0.00488 -0.0936 0.7086 0.0534
-4.500 0.0038 0.01180 0.00461 -0.0936 0.7017 0.0567
-4.250 0.0323 0.01161 0.00440 -0.0936 0.6941 0.0617
-4.000 0.0605 0.01145 0.00422 -0.0935 0.6875 0.0698
-3.750 0.0891 0.01127 0.00409 -0.0936 0.6806 0.0801
-3.500 0.1179 0.01122 0.00401 -0.0937 0.6738 0.0881
-3.250 0.1463 0.01109 0.00389 -0.0938 0.6675 0.0955
-3.000 0.1751 0.01101 0.00377 -0.0938 0.6604 0.1015
-2.750 0.2033 0.01092 0.00365 -0.0939 0.6536 0.1078
-2.500 0.2322 0.01085 0.00356 -0.0940 0.6465 0.1127
-2.250 0.2604 0.01070 0.00342 -0.0940 0.6394 0.1196
-2.000 0.2890 0.01066 0.00334 -0.0941 0.6330 0.1258
-1.750 0.3175 0.01051 0.00324 -0.0942 0.6258 0.1360
-1.500 0.3457 0.01043 0.00316 -0.0942 0.6189 0.1496
-1.250 0.3741 0.01029 0.00311 -0.0943 0.6116 0.1730
-1.000 0.4019 0.01010 0.00306 -0.0944 0.6038 0.2186
-0.750 0.4295 0.00990 0.00304 -0.0944 0.5963 0.2813
-0.500 0.4570 0.00966 0.00302 -0.0945 0.5881 0.3545
-0.250 0.4838 0.00940 0.00302 -0.0944 0.5806 0.4541
0.000 0.5094 0.00897 0.00306 -0.0941 0.5720 0.5990
0.250 0.5320 0.00857 0.00314 -0.0929 0.5638 0.7611
0.500 0.5532 0.00835 0.00326 -0.0909 0.5546 0.8829
0.750 0.5837 0.00841 0.00332 -0.0910 0.5456 0.9495
1.000 0.6246 0.00851 0.00337 -0.0936 0.5354 0.9793
1.250 0.6698 0.00866 0.00341 -0.0973 0.5254 0.9940
1.500 0.7100 0.00880 0.00345 -0.1001 0.5153 1.0000
1.750 0.7343 0.00891 0.00350 -0.0994 0.5070 1.0000
2.000 0.7583 0.00905 0.00355 -0.0987 0.4991 1.0000
2.250 0.7827 0.00918 0.00362 -0.0980 0.4919 1.0000
2.500 0.8072 0.00933 0.00370 -0.0974 0.4846 1.0000
2.750 0.8317 0.00951 0.00380 -0.0968 0.4783 1.0000
3.000 0.8573 0.00964 0.00390 -0.0964 0.4722 1.0000
3.250 0.8824 0.00982 0.00401 -0.0960 0.4663 1.0000
3.500 0.9078 0.01001 0.00415 -0.0956 0.4605 1.0000
3.750 0.9338 0.01015 0.00427 -0.0953 0.4544 1.0000
4.000 0.9588 0.01036 0.00441 -0.0948 0.4478 1.0000
4.250 0.9846 0.01053 0.00455 -0.0945 0.4418 1.0000
4.500 1.0106 0.01069 0.00469 -0.0943 0.4360 1.0000
4.750 1.0359 0.01092 0.00486 -0.0939 0.4310 1.0000
5.000 1.0616 0.01113 0.00505 -0.0937 0.4266 1.0000
5.250 1.0879 0.01128 0.00521 -0.0935 0.4225 1.0000
5.500 1.1136 0.01147 0.00540 -0.0933 0.4182 1.0000
5.750 1.1388 0.01171 0.00559 -0.0929 0.4140 1.0000
6.000 1.1642 0.01195 0.00582 -0.0927 0.4101 1.0000
6.250 1.1899 0.01210 0.00601 -0.0924 0.4061 1.0000
6.500 1.2151 0.01230 0.00621 -0.0921 0.4018 1.0000
6.750 1.2394 0.01255 0.00643 -0.0917 0.3973 1.0000
7.000 1.2638 0.01281 0.00669 -0.0913 0.3930 1.0000
7.250 1.2886 0.01297 0.00691 -0.0909 0.3887 1.0000
7.500 1.3127 0.01319 0.00713 -0.0905 0.3843 1.0000
7.750 1.3355 0.01348 0.00740 -0.0898 0.3797 1.0000
8.000 1.3586 0.01371 0.00766 -0.0892 0.3747 1.0000
8.250 1.3812 0.01391 0.00790 -0.0885 0.3690 1.0000
8.500 1.4019 0.01421 0.00819 -0.0875 0.3636 1.0000
8.750 1.4226 0.01449 0.00849 -0.0865 0.3587 1.0000
9.000 1.4424 0.01472 0.00878 -0.0853 0.3534 1.0000
9.250 1.4598 0.01507 0.00913 -0.0838 0.3477 1.0000
9.500 1.4779 0.01543 0.00951 -0.0824 0.3414 1.0000
9.750 1.4957 0.01580 0.00991 -0.0811 0.3339 1.0000
10.000 1.5118 0.01629 0.01039 -0.0796 0.3272 1.0000
10.250 1.5300 0.01671 0.01087 -0.0785 0.3198 1.0000
10.500 1.5442 0.01734 0.01148 -0.0769 0.3126 1.0000
10.750 1.5618 0.01785 0.01204 -0.0759 0.3045 1.0000
11.000 1.5746 0.01861 0.01279 -0.0743 0.2962 1.0000
11.250 1.5887 0.01936 0.01356 -0.0730 0.2864 1.0000
11.500 1.6005 0.02027 0.01446 -0.0715 0.2763 1.0000
11.750 1.6091 0.02142 0.01557 -0.0697 0.2659 1.0000
12.000 1.6194 0.02252 0.01667 -0.0683 0.2555 1.0000
12.250 1.6284 0.02374 0.01790 -0.0668 0.2460 1.0000
12.750 1.6436 0.02651 0.02067 -0.0639 0.2288 1.0000
13.000 1.6477 0.02820 0.02234 -0.0624 0.2214 1.0000
13.250 1.6565 0.02960 0.02379 -0.0613 0.2145 1.0000
13.500 1.6587 0.03154 0.02572 -0.0598 0.2079 1.0000
13.750 1.6669 0.03304 0.02728 -0.0588 0.2020 1.0000
14.000 1.6699 0.03500 0.02926 -0.0576 0.1962 1.0000
14.250 1.6737 0.03697 0.03126 -0.0566 0.1910 1.0000
14.500 1.6793 0.03885 0.03319 -0.0557 0.1855 1.0000
14.750 1.6786 0.04133 0.03568 -0.0547 0.1802 1.0000
15.000 1.6834 0.04337 0.03780 -0.0540 0.1751 1.0000
15.250 1.6849 0.04579 0.04026 -0.0534 0.1697 1.0000
15.500 1.6817 0.04872 0.04320 -0.0526 0.1644 1.0000
15.750 1.6857 0.05098 0.04554 -0.0522 0.1588 1.0000
16.000 1.6818 0.05410 0.04868 -0.0517 0.1534 1.0000
16.250 1.6823 0.05681 0.05145 -0.0513 0.1482 1.0000
16.500 1.6800 0.05987 0.05457 -0.0510 0.1424 1.0000
16.750 1.6741 0.06338 0.05811 -0.0507 0.1372 1.0000
17.000 1.6713 0.06659 0.06138 -0.0506 0.1305 1.0000
17.250 1.6628 0.07055 0.06537 -0.0506 0.1246 1.0000
17.500 1.6561 0.07434 0.06920 -0.0507 0.1175 1.0000
17.750 1.6461 0.07857 0.07346 -0.0508 0.1115 1.0000
18.000 1.6353 0.08298 0.07790 -0.0512 0.1053 1.0000
18.250 1.6251 0.08738 0.08234 -0.0516 0.1004 1.0000
18.500 1.6144 0.09189 0.08688 -0.0521 0.0958 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 624 AIRFOIL (goe624-il)