GOE 627 AIRFOIL (goe627-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 627 AIRFOIL (goe627-il) Reynolds number: 50,000 Max Cl/Cd: 29.14 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe627-il-50000-n5.txt Download as CSV file: xf-goe627-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 627 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3680 0.11091 0.10385 -0.0444 1.0000 0.0730
-10.500 -0.3706 0.10779 0.10079 -0.0437 1.0000 0.0724
-10.250 -0.3765 0.10442 0.09748 -0.0434 1.0000 0.0716
-10.000 -0.3867 0.10062 0.09376 -0.0435 1.0000 0.0710
-9.750 -0.4010 0.09646 0.08969 -0.0439 1.0000 0.0706
-9.500 -0.4191 0.09191 0.08522 -0.0446 1.0000 0.0700
-9.250 -0.4462 0.08647 0.07986 -0.0462 1.0000 0.0690
-9.000 -0.4816 0.08221 0.07564 -0.0453 1.0000 0.0684
-8.750 -0.5217 0.07870 0.07214 -0.0419 1.0000 0.0673
-8.500 -0.5545 0.07484 0.06820 -0.0389 1.0000 0.0668
-8.250 -0.5817 0.07102 0.06423 -0.0357 1.0000 0.0663
-8.000 -0.5984 0.06783 0.06089 -0.0327 1.0000 0.0664
-7.750 -0.6079 0.06469 0.05756 -0.0303 0.9993 0.0669
-7.500 -0.5963 0.06061 0.05312 -0.0318 0.9916 0.0679
-7.250 -0.5833 0.05669 0.04877 -0.0328 0.9840 0.0684
-7.000 -0.5689 0.05321 0.04483 -0.0331 0.9761 0.0688
-6.750 -0.5497 0.05019 0.04134 -0.0335 0.9690 0.0695
-6.500 -0.5305 0.04762 0.03827 -0.0334 0.9610 0.0712
-6.250 -0.5087 0.04530 0.03537 -0.0333 0.9536 0.0728
-6.000 -0.4835 0.04335 0.03311 -0.0336 0.9462 0.0741
-5.750 -0.4550 0.04170 0.03126 -0.0342 0.9395 0.0753
-5.500 -0.4282 0.04029 0.02966 -0.0344 0.9314 0.0769
-5.250 -0.3946 0.03903 0.02818 -0.0357 0.9251 0.0799
-5.000 -0.3691 0.03793 0.02681 -0.0353 0.9158 0.0825
-4.750 -0.3304 0.03672 0.02544 -0.0372 0.9101 0.0853
-4.500 -0.3061 0.03581 0.02452 -0.0366 0.8998 0.0876
-4.250 -0.2659 0.03488 0.02350 -0.0386 0.8938 0.0918
-4.000 -0.2416 0.03419 0.02270 -0.0378 0.8826 0.0964
-3.750 -0.2091 0.03341 0.02197 -0.0386 0.8740 0.1025
-3.500 -0.1771 0.03268 0.02115 -0.0392 0.8646 0.1096
-3.250 -0.1497 0.03201 0.02048 -0.0390 0.8539 0.1210
-3.000 -0.1121 0.03098 0.01964 -0.0407 0.8461 0.1486
-2.750 -0.0918 0.02997 0.01909 -0.0395 0.8340 0.2095
-2.500 -0.0688 0.02879 0.01863 -0.0387 0.8240 0.3555
-2.250 -0.0469 0.02754 0.01843 -0.0367 0.8142 0.5459
-2.000 0.0787 0.02788 0.01972 -0.0489 0.8129 0.9041
-1.750 0.1621 0.02808 0.01954 -0.0579 0.8026 0.9464
-1.500 0.2468 0.02757 0.01867 -0.0676 0.7957 0.9719
-1.250 0.3120 0.02710 0.01793 -0.0745 0.7824 0.9899
-1.000 0.3667 0.02654 0.01715 -0.0795 0.7687 1.0000
-0.750 0.3925 0.02610 0.01652 -0.0791 0.7534 1.0000
-0.500 0.4190 0.02568 0.01592 -0.0786 0.7377 1.0000
-0.250 0.4452 0.02531 0.01537 -0.0781 0.7210 1.0000
0.000 0.4704 0.02502 0.01489 -0.0773 0.7036 1.0000
0.250 0.4945 0.02479 0.01448 -0.0763 0.6853 1.0000
0.500 0.5175 0.02463 0.01414 -0.0752 0.6662 1.0000
0.750 0.5399 0.02453 0.01386 -0.0739 0.6470 1.0000
1.000 0.5613 0.02450 0.01364 -0.0725 0.6278 1.0000
1.250 0.5817 0.02453 0.01349 -0.0710 0.6087 1.0000
1.500 0.6009 0.02462 0.01341 -0.0692 0.5902 1.0000
1.750 0.6192 0.02477 0.01338 -0.0674 0.5720 1.0000
2.000 0.6365 0.02495 0.01340 -0.0654 0.5543 1.0000
2.250 0.6536 0.02518 0.01347 -0.0633 0.5375 1.0000
2.500 0.6704 0.02544 0.01358 -0.0613 0.5213 1.0000
2.750 0.6879 0.02573 0.01371 -0.0594 0.5062 1.0000
3.000 0.7063 0.02602 0.01383 -0.0576 0.4922 1.0000
3.250 0.7237 0.02637 0.01405 -0.0558 0.4787 1.0000
3.500 0.7396 0.02679 0.01438 -0.0537 0.4659 1.0000
3.750 0.7582 0.02719 0.01465 -0.0522 0.4542 1.0000
4.000 0.7784 0.02758 0.01491 -0.0508 0.4433 1.0000
4.250 0.7939 0.02809 0.01538 -0.0488 0.4324 1.0000
4.500 0.8170 0.02851 0.01564 -0.0480 0.4233 1.0000
4.750 0.8320 0.02909 0.01623 -0.0460 0.4137 1.0000
5.000 0.8569 0.02955 0.01654 -0.0456 0.4061 1.0000
5.250 0.8699 0.03022 0.01726 -0.0433 0.3976 1.0000
5.500 0.8934 0.03072 0.01765 -0.0427 0.3905 1.0000
5.750 0.9078 0.03142 0.01839 -0.0407 0.3832 1.0000
6.000 0.9256 0.03205 0.01901 -0.0392 0.3763 1.0000
6.250 0.9508 0.03263 0.01948 -0.0389 0.3707 1.0000
6.500 0.9604 0.03351 0.02049 -0.0363 0.3646 1.0000
6.750 0.9782 0.03423 0.02122 -0.0349 0.3590 1.0000
7.000 1.0065 0.03480 0.02167 -0.0351 0.3542 1.0000
7.250 1.0095 0.03587 0.02292 -0.0316 0.3486 1.0000
7.500 1.0221 0.03677 0.02390 -0.0296 0.3435 1.0000
7.750 1.0445 0.03747 0.02457 -0.0290 0.3391 1.0000
8.000 1.0601 0.03842 0.02556 -0.0275 0.3350 1.0000
8.250 1.0582 0.03972 0.02703 -0.0235 0.3306 1.0000
8.500 1.0660 0.04081 0.02822 -0.0210 0.3263 1.0000
8.750 1.0867 0.04160 0.02900 -0.0203 0.3223 1.0000
9.000 1.1024 0.04260 0.03003 -0.0189 0.3186 1.0000
9.250 1.0863 0.04457 0.03222 -0.0137 0.3148 1.0000
9.500 1.0798 0.04643 0.03423 -0.0101 0.3109 1.0000
9.750 1.0881 0.04776 0.03564 -0.0083 0.3072 1.0000
10.000 1.1188 0.04821 0.03607 -0.0087 0.3037 1.0000
10.250 1.0927 0.05118 0.03924 -0.0039 0.3003 1.0000
10.500 1.0301 0.05700 0.04534 0.0019 0.2959 1.0000
10.750 0.9909 0.06246 0.05095 0.0044 0.2910 1.0000
11.000 1.0169 0.06267 0.05117 0.0049 0.2880 1.0000
11.500 0.8511 0.08853 0.07731 0.0020 0.2706 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 627 AIRFOIL (goe627-il)