GOE 571 AIRFOIL (goe571-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 571 AIRFOIL (goe571-il) Reynolds number: 500,000 Max Cl/Cd: 56.09 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe571-il-500000.txt Download as CSV file: xf-goe571-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 571 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
------ -------- --------- --------- -------- -------- --------
-13.250 0.4983 0.10906 0.10403 -0.1693 0.5888 0.0296
-13.000 0.4910 0.10620 0.10115 -0.1725 0.5863 0.0299
-12.750 0.4933 0.10339 0.09838 -0.1737 0.5841 0.0300
-12.500 0.5065 0.10132 0.09632 -0.1736 0.5803 0.0301
-12.250 0.5222 0.09933 0.09432 -0.1734 0.5758 0.0303
-12.000 0.5340 0.09744 0.09241 -0.1739 0.5723 0.0305
-11.750 0.5441 0.09547 0.09038 -0.1747 0.5680 0.0307
-11.500 0.5536 0.09356 0.08848 -0.1754 0.5645 0.0310
-11.250 0.5604 0.09159 0.08653 -0.1757 0.5610 0.0314
-11.000 0.5669 0.08955 0.08450 -0.1762 0.5571 0.0318
-10.750 0.5729 0.08748 0.08242 -0.1767 0.5533 0.0323
-10.500 0.5766 0.08512 0.08004 -0.1776 0.5500 0.0332
-10.250 0.5547 0.08013 0.07507 -0.1817 0.5485 0.0340
-10.000 0.5682 0.07873 0.07367 -0.1807 0.5442 0.0342
-9.750 0.5751 0.07695 0.07192 -0.1804 0.5403 0.0343
-9.500 0.5848 0.07555 0.07052 -0.1798 0.5359 0.0346
-9.250 0.5897 0.07381 0.06875 -0.1795 0.5317 0.0348
-9.000 0.5960 0.07220 0.06710 -0.1794 0.5274 0.0352
-8.750 0.5986 0.07036 0.06531 -0.1791 0.5240 0.0355
-8.500 0.5987 0.06837 0.06335 -0.1787 0.5200 0.0358
-8.250 0.5979 0.06645 0.06144 -0.1782 0.5161 0.0363
-8.000 0.5925 0.06424 0.05923 -0.1776 0.5125 0.0368
-7.750 0.4888 0.05305 0.04811 -0.1830 0.5148 0.0383
-7.500 0.4821 0.05202 0.04707 -0.1790 0.5109 0.0384
-7.250 0.4576 0.05084 0.04589 -0.1727 0.5075 0.0385
-7.000 0.4374 0.04985 0.04495 -0.1664 0.5046 0.0386
-6.750 0.4274 0.04883 0.04394 -0.1617 0.5008 0.0387
-6.500 0.4135 0.04761 0.04270 -0.1566 0.4971 0.0388
-6.250 0.4034 0.04653 0.04158 -0.1518 0.4937 0.0389
-6.000 0.3953 0.04548 0.04048 -0.1472 0.4901 0.0391
-5.750 0.3902 0.04445 0.03945 -0.1429 0.4870 0.0393
-5.500 0.3820 0.04313 0.03810 -0.1380 0.4840 0.0395
-5.250 0.3753 0.04179 0.03671 -0.1333 0.4809 0.0399
-5.000 0.3702 0.04051 0.03535 -0.1287 0.4777 0.0405
-4.750 0.3636 0.03903 0.03374 -0.1237 0.4748 0.0410
-4.500 0.3353 0.03540 0.02972 -0.1148 0.4727 0.0435
-4.250 0.3400 0.03442 0.02871 -0.1113 0.4701 0.0436
-4.000 0.3468 0.03373 0.02803 -0.1081 0.4672 0.0441
-3.750 0.3511 0.03281 0.02705 -0.1044 0.4645 0.0445
-3.500 0.3390 0.03080 0.02461 -0.0969 0.4623 0.0483
-3.250 0.3464 0.02998 0.02374 -0.0937 0.4593 0.0485
-3.000 0.3563 0.02940 0.02312 -0.0910 0.4563 0.0489
-2.750 0.3670 0.02881 0.02247 -0.0882 0.4536 0.0495
-2.500 0.3657 0.02779 0.02114 -0.0825 0.4517 0.0538
-2.250 0.3765 0.02719 0.02055 -0.0798 0.4492 0.0541
-2.000 0.3881 0.02677 0.02013 -0.0772 0.4468 0.0550
-1.750 0.3921 0.02627 0.01933 -0.0724 0.4447 0.0595
-1.500 0.4051 0.02572 0.01879 -0.0702 0.4423 0.0602
-1.250 0.4180 0.02534 0.01835 -0.0677 0.4399 0.0612
-1.000 0.4304 0.02502 0.01775 -0.0647 0.4373 0.0661
-0.750 0.4438 0.02454 0.01733 -0.0624 0.4356 0.0672
-0.500 0.4559 0.02433 0.01708 -0.0597 0.4339 0.0694
-0.250 0.4732 0.02054 0.01234 -0.0545 0.4324 0.0423
0.000 0.4928 0.02003 0.01178 -0.0531 0.4305 0.0415
0.250 0.5113 0.01974 0.01142 -0.0515 0.4282 0.0405
0.500 0.5309 0.01950 0.01109 -0.0501 0.4263 0.0400
0.750 0.5502 0.01939 0.01090 -0.0488 0.4241 0.0397
1.000 0.5726 0.01932 0.01076 -0.0481 0.4219 0.0397
1.250 0.6017 0.01923 0.01062 -0.0485 0.4195 0.0402
1.500 0.6170 0.01926 0.01067 -0.0466 0.4183 0.0405
1.750 0.6331 0.01933 0.01077 -0.0450 0.4169 0.0409
2.000 0.6506 0.01941 0.01086 -0.0436 0.4154 0.0412
2.250 0.6682 0.01951 0.01097 -0.0422 0.4135 0.0417
2.500 0.6865 0.01964 0.01110 -0.0411 0.4116 0.0420
2.750 0.7058 0.01979 0.01124 -0.0402 0.4099 0.0425
3.000 0.7269 0.01992 0.01137 -0.0397 0.4082 0.0435
3.250 0.7501 0.02007 0.01151 -0.0395 0.4063 0.0446
3.500 0.7777 0.02021 0.01162 -0.0401 0.4045 0.0455
3.750 0.8110 0.02034 0.01171 -0.0416 0.4024 0.0470
4.000 0.8282 0.02066 0.01207 -0.0407 0.4013 0.0494
4.250 0.8434 0.02106 0.01252 -0.0396 0.3998 0.0504
4.500 0.8637 0.02142 0.01292 -0.0393 0.3983 0.0557
4.750 1.1886 0.02119 0.01499 -0.1021 0.3943 1.0000
5.000 1.2040 0.02162 0.01539 -0.1008 0.3928 1.0000
5.250 1.2189 0.02208 0.01582 -0.0995 0.3913 1.0000
5.500 1.2360 0.02251 0.01621 -0.0985 0.3898 1.0000
5.750 1.2564 0.02288 0.01651 -0.0980 0.3878 1.0000
6.000 1.2838 0.02316 0.01670 -0.0986 0.3858 1.0000
6.250 1.2934 0.02381 0.01739 -0.0967 0.3849 1.0000
6.500 1.3026 0.02452 0.01814 -0.0948 0.3838 1.0000
6.750 1.3130 0.02523 0.01888 -0.0932 0.3825 1.0000
7.000 1.3231 0.02599 0.01966 -0.0917 0.3809 1.0000
7.250 1.3341 0.02674 0.02043 -0.0903 0.3792 1.0000
7.500 1.3463 0.02747 0.02115 -0.0891 0.3775 1.0000
7.750 1.3605 0.02813 0.02181 -0.0882 0.3760 1.0000
8.000 1.3740 0.02883 0.02249 -0.0873 0.3744 1.0000
8.250 1.3900 0.02944 0.02307 -0.0866 0.3729 1.0000
8.500 1.4121 0.02978 0.02335 -0.0867 0.3708 1.0000
8.750 1.4317 0.03030 0.02385 -0.0864 0.3692 1.0000
9.000 1.4327 0.03169 0.02533 -0.0845 0.3680 1.0000
9.250 1.4349 0.03311 0.02682 -0.0827 0.3663 1.0000
9.500 1.4413 0.03436 0.02812 -0.0815 0.3647 1.0000
9.750 1.4484 0.03559 0.02938 -0.0803 0.3629 1.0000
10.000 1.4583 0.03667 0.03049 -0.0795 0.3612 1.0000
10.250 1.4700 0.03765 0.03147 -0.0789 0.3596 1.0000
10.500 1.4817 0.03864 0.03245 -0.0783 0.3579 1.0000
10.750 1.4978 0.03933 0.03311 -0.0780 0.3562 1.0000
11.000 1.5231 0.03938 0.03310 -0.0784 0.3542 1.0000
11.250 1.5203 0.04144 0.03526 -0.0768 0.3526 1.0000
11.500 1.5131 0.04391 0.03784 -0.0751 0.3506 1.0000
11.750 1.5097 0.04620 0.04020 -0.0738 0.3483 1.0000
12.000 1.5131 0.04797 0.04202 -0.0730 0.3462 1.0000
12.250 1.5208 0.04942 0.04347 -0.0725 0.3441 1.0000
12.500 1.5319 0.05057 0.04463 -0.0722 0.3424 1.0000
12.750 1.5481 0.05126 0.04529 -0.0721 0.3406 1.0000
13.000 1.5723 0.05124 0.04522 -0.0725 0.3388 1.0000
13.250 1.5494 0.05541 0.04955 -0.0705 0.3363 1.0000
13.500 1.5316 0.05929 0.05356 -0.0689 0.3333 1.0000
13.750 1.5284 0.06188 0.05621 -0.0682 0.3307 1.0000
14.000 1.5354 0.06350 0.05784 -0.0679 0.3284 1.0000
14.250 1.5481 0.06454 0.05887 -0.0679 0.3265 1.0000
14.500 1.5697 0.06468 0.05896 -0.0682 0.3247 1.0000
14.750 1.5476 0.06916 0.06357 -0.0667 0.3215 1.0000
15.000 1.5132 0.07510 0.06967 -0.0650 0.3172 1.0000
15.250 1.5149 0.07736 0.07195 -0.0648 0.3145 1.0000
15.500 1.5261 0.07861 0.07318 -0.0649 0.3121 1.0000
15.750 1.5553 0.07789 0.07238 -0.0656 0.3101 1.0000
16.000 1.4868 0.08773 0.08248 -0.0631 0.3038 1.0000
16.250 1.4787 0.09118 0.08598 -0.0627 0.3001 1.0000
16.500 1.4939 0.09200 0.08676 -0.0630 0.2979 1.0000
16.750 1.5208 0.09144 0.08614 -0.0637 0.2960 1.0000
17.000 1.4385 0.10336 0.09833 -0.0614 0.2883 1.0000
17.250 1.4501 0.10461 0.09955 -0.0618 0.2850 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 571 AIRFOIL (goe571-il)