GOE 180 (MVA H.26) AIRFOIL (goe180-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 180 (MVA H.26) AIRFOIL (goe180-il) Reynolds number: 50,000 Max Cl/Cd: 40.55 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe180-il-50000-n5.txt Download as CSV file: xf-goe180-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 180 (MVA H.26) 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3412 0.10035 0.09367 -0.0258 1.0000 0.1127
-7.500 -0.3527 0.09919 0.09267 -0.0277 1.0000 0.1153
-7.250 -0.3596 0.09763 0.09123 -0.0324 1.0000 0.1163
-7.000 -0.3531 0.09328 0.08697 -0.0307 1.0000 0.1175
-6.750 -0.3453 0.08956 0.08329 -0.0281 1.0000 0.1196
-6.500 -0.3412 0.08666 0.08043 -0.0274 1.0000 0.1224
-6.250 -0.3394 0.08414 0.07797 -0.0285 1.0000 0.1276
-6.000 -0.3389 0.08166 0.07552 -0.0350 1.0000 0.1323
-5.750 -0.3333 0.07836 0.07232 -0.0296 1.0000 0.1368
-5.500 -0.3275 0.07581 0.06977 -0.0316 1.0000 0.1444
-5.250 -0.3019 0.06731 0.06089 -0.0422 1.0000 0.0847
-4.750 -0.2699 0.05791 0.05120 -0.0474 1.0000 0.0728
-4.500 -0.2558 0.05519 0.04842 -0.0476 1.0000 0.0742
-4.250 -0.2372 0.05204 0.04512 -0.0491 1.0000 0.0752
-4.000 -0.1960 0.04631 0.03854 -0.0558 0.9976 0.0712
-3.750 -0.1530 0.04227 0.03411 -0.0611 0.9900 0.0708
-3.500 -0.1105 0.03889 0.03034 -0.0659 0.9817 0.0728
-3.250 -0.0695 0.03652 0.02771 -0.0698 0.9723 0.0757
-3.000 -0.0251 0.03368 0.02431 -0.0738 0.9634 0.0759
-2.750 0.0178 0.03137 0.02148 -0.0772 0.9535 0.0766
-2.500 0.0584 0.02950 0.01919 -0.0798 0.9423 0.0780
-2.250 0.0993 0.02808 0.01736 -0.0823 0.9309 0.0826
-2.000 0.1418 0.02673 0.01558 -0.0848 0.9203 0.0864
-1.750 0.1834 0.02545 0.01418 -0.0871 0.9100 0.0897
-1.500 0.2223 0.02449 0.01304 -0.0887 0.8975 0.0943
-1.250 0.2607 0.02365 0.01193 -0.0901 0.8842 0.1009
-1.000 0.2975 0.02289 0.01116 -0.0915 0.8700 0.1140
-0.750 0.3330 0.02213 0.01048 -0.0927 0.8549 0.1416
-0.500 0.3674 0.02115 0.00990 -0.0938 0.8393 0.2248
-0.250 0.3953 0.01874 0.00952 -0.0928 0.8235 1.0000
0.000 0.4277 0.01879 0.00914 -0.0932 0.8066 1.0000
0.250 0.4593 0.01886 0.00888 -0.0934 0.7900 1.0000
0.500 0.4900 0.01895 0.00869 -0.0935 0.7739 1.0000
0.750 0.5192 0.01910 0.00860 -0.0934 0.7576 1.0000
1.000 0.5484 0.01926 0.00855 -0.0933 0.7425 1.0000
1.250 0.5773 0.01944 0.00853 -0.0931 0.7283 1.0000
1.500 0.6054 0.01968 0.00861 -0.0929 0.7144 1.0000
1.750 0.6333 0.01997 0.00876 -0.0928 0.7011 1.0000
2.000 0.6609 0.02028 0.00896 -0.0926 0.6885 1.0000
2.250 0.6885 0.02059 0.00916 -0.0924 0.6762 1.0000
2.500 0.7160 0.02089 0.00935 -0.0921 0.6636 1.0000
2.750 0.7430 0.02117 0.00954 -0.0916 0.6499 1.0000
3.000 0.7691 0.02146 0.00973 -0.0910 0.6348 1.0000
3.250 0.7948 0.02175 0.00994 -0.0903 0.6194 1.0000
3.500 0.8207 0.02207 0.01021 -0.0896 0.6051 1.0000
3.750 0.8468 0.02244 0.01056 -0.0891 0.5928 1.0000
4.000 0.8719 0.02289 0.01104 -0.0886 0.5809 1.0000
4.250 0.8973 0.02334 0.01154 -0.0881 0.5695 1.0000
4.500 0.9232 0.02376 0.01197 -0.0876 0.5584 1.0000
4.750 0.9488 0.02419 0.01243 -0.0871 0.5468 1.0000
5.000 0.9727 0.02469 0.01303 -0.0864 0.5342 1.0000
5.250 0.9967 0.02519 0.01364 -0.0857 0.5214 1.0000
5.500 1.0207 0.02569 0.01422 -0.0849 0.5085 1.0000
5.750 1.0443 0.02617 0.01479 -0.0840 0.4945 1.0000
6.000 1.0667 0.02662 0.01535 -0.0829 0.4782 1.0000
6.250 1.0881 0.02704 0.01586 -0.0816 0.4604 1.0000
6.500 1.1092 0.02745 0.01633 -0.0803 0.4421 1.0000
6.750 1.1302 0.02787 0.01682 -0.0790 0.4242 1.0000
7.000 1.1506 0.02839 0.01749 -0.0777 0.4070 1.0000
7.250 1.1704 0.02903 0.01829 -0.0764 0.3900 1.0000
7.500 1.1902 0.02970 0.01912 -0.0751 0.3738 1.0000
7.750 1.2097 0.03042 0.01999 -0.0739 0.3581 1.0000
8.000 1.2288 0.03116 0.02091 -0.0725 0.3429 1.0000
8.250 1.2474 0.03194 0.02183 -0.0712 0.3283 1.0000
8.500 1.2657 0.03277 0.02279 -0.0698 0.3147 1.0000
8.750 1.2840 0.03364 0.02379 -0.0684 0.3024 1.0000
9.000 1.3031 0.03456 0.02482 -0.0671 0.2916 1.0000
9.250 1.3204 0.03575 0.02628 -0.0657 0.2808 1.0000
9.500 1.3367 0.03691 0.02762 -0.0642 0.2695 1.0000
9.750 1.3452 0.03789 0.02865 -0.0618 0.2521 1.0000
10.000 1.3466 0.03894 0.02960 -0.0587 0.2315 1.0000
10.250 1.3420 0.04044 0.03119 -0.0553 0.2119 1.0000
10.500 1.3363 0.04214 0.03302 -0.0521 0.1949 1.0000
10.750 1.3312 0.04415 0.03522 -0.0495 0.1783 1.0000
11.000 1.3255 0.04650 0.03773 -0.0476 0.1612 1.0000
11.250 1.3182 0.04931 0.04062 -0.0462 0.1445 1.0000
11.500 1.3088 0.05275 0.04415 -0.0454 0.1270 1.0000
11.750 1.2973 0.05681 0.04821 -0.0452 0.1108 1.0000
12.000 1.2854 0.06130 0.05275 -0.0455 0.0953 1.0000
12.250 1.2740 0.06603 0.05754 -0.0460 0.0823 1.0000
12.500 1.2627 0.07090 0.06243 -0.0468 0.0740 1.0000
12.750 1.2511 0.07594 0.06747 -0.0479 0.0684 1.0000
13.000 1.2405 0.08093 0.07248 -0.0489 0.0641 1.0000
13.250 1.2306 0.08591 0.07753 -0.0500 0.0603 1.0000
13.500 1.2202 0.09099 0.08262 -0.0514 0.0575 1.0000
13.750 1.2115 0.09586 0.08751 -0.0525 0.0549 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 180 (MVA H.26) AIRFOIL (goe180-il)