GOE 269 AIRFOIL (goe269-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 269 AIRFOIL (goe269-il) Reynolds number: 200,000 Max Cl/Cd: 79.19 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe269-il-200000-n5.txt Download as CSV file: xf-goe269-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 269 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3466 0.10072 0.09722 -0.0166 1.0000 0.0320
-8.000 -0.3433 0.09788 0.09442 -0.0188 1.0000 0.0321
-7.750 -0.3388 0.09485 0.09143 -0.0213 1.0000 0.0322
-7.500 -0.3305 0.09142 0.08803 -0.0247 1.0000 0.0322
-7.000 -0.3144 0.08410 0.08077 -0.0277 1.0000 0.0329
-6.750 -0.3066 0.08199 0.07870 -0.0254 1.0000 0.0340
-6.500 -0.2973 0.07955 0.07629 -0.0266 0.9987 0.0352
-6.250 -0.2692 0.07550 0.07222 -0.0334 0.9885 0.0373
-5.750 -0.1783 0.06444 0.06085 -0.0614 0.9634 0.0403
-5.500 -0.1469 0.05977 0.05608 -0.0670 0.9526 0.0403
-5.250 -0.1259 0.05624 0.05253 -0.0682 0.9423 0.0398
-4.750 -0.0569 0.04769 0.04367 -0.0794 0.9202 0.0398
-4.500 -0.0296 0.04456 0.04043 -0.0819 0.9067 0.0393
-4.250 0.0015 0.04114 0.03681 -0.0853 0.8921 0.0390
-3.750 0.0705 0.03350 0.02854 -0.0921 0.8617 0.0400
-3.500 0.1003 0.03056 0.02531 -0.0937 0.8452 0.0400
-3.250 0.1314 0.02749 0.02189 -0.0952 0.8291 0.0402
-3.000 0.1618 0.02471 0.01876 -0.0965 0.8136 0.0406
-2.750 0.1885 0.02384 0.01776 -0.0967 0.7982 0.0416
-2.500 0.2158 0.02311 0.01689 -0.0969 0.7835 0.0428
-2.250 0.2461 0.02125 0.01469 -0.0975 0.7697 0.0434
-2.000 0.2767 0.01934 0.01239 -0.0980 0.7568 0.0439
-1.750 0.3067 0.01779 0.01047 -0.0984 0.7445 0.0448
-1.500 0.3362 0.01647 0.00879 -0.0985 0.7328 0.0462
-1.250 0.3654 0.01536 0.00732 -0.0985 0.7214 0.0477
-1.000 0.3939 0.01471 0.00644 -0.0985 0.7102 0.0501
-0.750 0.4218 0.01455 0.00624 -0.0984 0.6993 0.0519
-0.250 0.4781 0.01379 0.00522 -0.0981 0.6782 0.0566
0.000 0.5063 0.01343 0.00472 -0.0980 0.6680 0.0596
0.500 0.5621 0.01315 0.00438 -0.0977 0.6466 0.0675
0.750 0.5899 0.01302 0.00423 -0.0976 0.6360 0.0727
1.000 0.6177 0.01292 0.00412 -0.0975 0.6253 0.0776
1.250 0.6455 0.01284 0.00400 -0.0974 0.6134 0.0832
1.500 0.6733 0.01270 0.00390 -0.0973 0.6010 0.0881
1.750 0.7011 0.01265 0.00385 -0.0972 0.5886 0.0930
2.000 0.7288 0.01260 0.00379 -0.0971 0.5752 0.0977
2.250 0.7564 0.01257 0.00377 -0.0970 0.5603 0.1019
2.500 0.7840 0.01260 0.00378 -0.0968 0.5445 0.1063
2.750 0.8113 0.01266 0.00380 -0.0966 0.5259 0.1125
3.250 0.8646 0.01287 0.00394 -0.0960 0.4797 0.1409
3.500 0.8851 0.01156 0.00405 -0.0945 0.4596 1.0000
3.750 0.9115 0.01186 0.00419 -0.0942 0.4405 1.0000
4.000 0.9378 0.01216 0.00437 -0.0939 0.4235 1.0000
4.250 0.9642 0.01245 0.00458 -0.0936 0.4097 1.0000
4.500 0.9906 0.01275 0.00481 -0.0933 0.3974 1.0000
4.750 1.0172 0.01301 0.00507 -0.0931 0.3858 1.0000
5.000 1.0437 0.01329 0.00533 -0.0928 0.3753 1.0000
5.250 1.0698 0.01360 0.00561 -0.0925 0.3646 1.0000
5.500 1.0961 0.01388 0.00592 -0.0923 0.3530 1.0000
5.750 1.1221 0.01417 0.00623 -0.0920 0.3403 1.0000
6.000 1.1477 0.01450 0.00655 -0.0916 0.3257 1.0000
6.250 1.1732 0.01484 0.00690 -0.0913 0.3104 1.0000
6.500 1.1980 0.01523 0.00728 -0.0908 0.2915 1.0000
6.750 1.2223 0.01567 0.00767 -0.0904 0.2668 1.0000
7.000 1.2457 0.01621 0.00813 -0.0898 0.2388 1.0000
7.250 1.2681 0.01686 0.00867 -0.0891 0.2091 1.0000
7.500 1.2894 0.01764 0.00932 -0.0884 0.1814 1.0000
7.750 1.3104 0.01843 0.01001 -0.0876 0.1609 1.0000
8.000 1.3317 0.01917 0.01071 -0.0868 0.1462 1.0000
8.250 1.3527 0.01989 0.01141 -0.0859 0.1328 1.0000
8.500 1.3738 0.02058 0.01213 -0.0851 0.1209 1.0000
8.750 1.3944 0.02127 0.01286 -0.0842 0.1109 1.0000
9.000 1.4143 0.02202 0.01361 -0.0832 0.0995 1.0000
9.250 1.4339 0.02276 0.01436 -0.0822 0.0842 1.0000
9.750 1.4644 0.02498 0.01641 -0.0793 0.0452 1.0000
10.000 1.4749 0.02646 0.01781 -0.0772 0.0297 1.0000
10.250 1.4847 0.02789 0.01923 -0.0750 0.0235 1.0000
10.500 1.4933 0.02927 0.02068 -0.0727 0.0206 1.0000
10.750 1.4999 0.03056 0.02211 -0.0700 0.0191 1.0000
11.000 1.5043 0.03204 0.02372 -0.0673 0.0180 1.0000
11.250 1.5068 0.03380 0.02560 -0.0649 0.0171 1.0000
11.500 1.5063 0.03594 0.02789 -0.0626 0.0163 1.0000
11.750 1.5068 0.03816 0.03029 -0.0608 0.0157 1.0000
12.000 1.5080 0.04044 0.03274 -0.0594 0.0152 1.0000
12.250 1.5079 0.04300 0.03547 -0.0584 0.0147 1.0000
12.500 1.5067 0.04585 0.03847 -0.0576 0.0141 1.0000
12.750 1.5040 0.04902 0.04179 -0.0572 0.0137 1.0000
13.000 1.5001 0.05247 0.04540 -0.0571 0.0133 1.0000
13.250 1.4945 0.05627 0.04934 -0.0573 0.0130 1.0000
13.500 1.4876 0.06036 0.05358 -0.0576 0.0128 1.0000
13.750 1.4792 0.06474 0.05810 -0.0582 0.0126 1.0000
14.000 1.4697 0.06940 0.06290 -0.0591 0.0124 1.0000
14.250 1.4593 0.07426 0.06788 -0.0600 0.0122 1.0000
14.500 1.4486 0.07928 0.07303 -0.0611 0.0120 1.0000
14.750 1.4381 0.08441 0.07828 -0.0624 0.0119 1.0000
15.000 1.4296 0.08934 0.08332 -0.0636 0.0117 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 269 AIRFOIL (goe269-il)