GOE 591 AIRFOIL (goe591-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 591 AIRFOIL (goe591-il) Reynolds number: 50,000 Max Cl/Cd: 34.41 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe591-il-50000-n5.txt Download as CSV file: xf-goe591-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 591 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3274 0.10943 0.10337 -0.0329 1.0000 0.1165
-8.000 -0.3485 0.10890 0.10299 -0.0305 1.0000 0.1167
-7.750 -0.3698 0.10825 0.10248 -0.0282 1.0000 0.1169
-7.500 -0.3863 0.10687 0.10120 -0.0279 0.9997 0.1171
-7.250 -0.3648 0.10137 0.09571 -0.0287 0.9951 0.1184
-6.750 -0.3259 0.09302 0.08730 -0.0344 0.9808 0.1261
-6.250 -0.2936 0.07612 0.06999 -0.0585 0.9573 0.0722
-6.000 -0.2722 0.07174 0.06552 -0.0619 0.9500 0.0710
-5.750 -0.2550 0.06727 0.06092 -0.0655 0.9402 0.0705
-5.500 -0.2288 0.06153 0.05483 -0.0722 0.9325 0.0717
-5.250 -0.2088 0.05685 0.04979 -0.0756 0.9226 0.0721
-5.000 -0.1834 0.05245 0.04500 -0.0787 0.9150 0.0720
-4.750 -0.1581 0.04839 0.04047 -0.0812 0.9066 0.0722
-4.500 -0.1324 0.04572 0.03759 -0.0827 0.8988 0.0746
-4.250 -0.1037 0.04359 0.03524 -0.0842 0.8911 0.0777
-4.000 -0.0761 0.04084 0.03203 -0.0854 0.8829 0.0789
-3.750 -0.0430 0.03816 0.02881 -0.0872 0.8760 0.0801
-3.500 -0.0145 0.03610 0.02627 -0.0877 0.8677 0.0818
-3.250 0.0217 0.03422 0.02374 -0.0893 0.8612 0.0865
-3.000 0.0494 0.03293 0.02228 -0.0895 0.8525 0.0902
-2.750 0.0863 0.03154 0.02062 -0.0909 0.8462 0.0939
-2.500 0.1152 0.03052 0.01931 -0.0909 0.8374 0.0982
-2.250 0.1535 0.02934 0.01788 -0.0923 0.8312 0.1041
-2.000 0.1812 0.02878 0.01719 -0.0921 0.8218 0.1148
-1.750 0.2193 0.02795 0.01626 -0.0935 0.8156 0.1326
-1.500 0.2454 0.02747 0.01581 -0.0931 0.8056 0.1523
-1.250 0.2822 0.02679 0.01520 -0.0944 0.7995 0.1989
-1.000 0.3053 0.02633 0.01494 -0.0937 0.7888 0.2516
-0.750 0.3400 0.02510 0.01430 -0.0948 0.7831 0.3708
-0.500 0.3892 0.02351 0.01409 -0.0979 0.7740 1.0000
-0.250 0.4256 0.02342 0.01358 -0.0989 0.7674 1.0000
0.000 0.4458 0.02373 0.01364 -0.0975 0.7555 1.0000
0.250 0.4707 0.02392 0.01359 -0.0968 0.7454 1.0000
0.500 0.5024 0.02393 0.01335 -0.0971 0.7376 1.0000
0.750 0.5232 0.02426 0.01351 -0.0958 0.7262 1.0000
1.000 0.5512 0.02438 0.01344 -0.0956 0.7174 1.0000
1.250 0.5776 0.02454 0.01346 -0.0951 0.7081 1.0000
1.500 0.5995 0.02487 0.01367 -0.0940 0.6974 1.0000
1.750 0.6328 0.02481 0.01345 -0.0945 0.6905 1.0000
2.000 0.6516 0.02526 0.01383 -0.0930 0.6791 1.0000
2.250 0.6754 0.02556 0.01403 -0.0922 0.6694 1.0000
2.500 0.7051 0.02562 0.01399 -0.0922 0.6614 1.0000
2.750 0.7247 0.02609 0.01442 -0.0909 0.6506 1.0000
3.000 0.7528 0.02622 0.01448 -0.0906 0.6424 1.0000
3.250 0.7763 0.02654 0.01477 -0.0898 0.6326 1.0000
3.500 0.7975 0.02697 0.01517 -0.0887 0.6226 1.0000
3.750 0.8291 0.02695 0.01509 -0.0889 0.6150 1.0000
4.000 0.8465 0.02754 0.01569 -0.0873 0.6039 1.0000
4.250 0.8705 0.02785 0.01600 -0.0865 0.5946 1.0000
4.500 0.8976 0.02802 0.01616 -0.0861 0.5858 1.0000
4.750 0.9157 0.02859 0.01676 -0.0847 0.5751 1.0000
5.000 0.9433 0.02875 0.01693 -0.0843 0.5665 1.0000
5.250 0.9644 0.02920 0.01742 -0.0832 0.5565 1.0000
5.500 0.9834 0.02976 0.01804 -0.0819 0.5464 1.0000
5.750 1.0147 0.02979 0.01808 -0.0820 0.5385 1.0000
6.000 1.0281 0.03061 0.01901 -0.0800 0.5276 1.0000
6.250 1.0498 0.03110 0.01956 -0.0790 0.5184 1.0000
6.500 1.0751 0.03141 0.01994 -0.0785 0.5094 1.0000
6.750 1.0890 0.03224 0.02090 -0.0766 0.4990 1.0000
7.000 1.1152 0.03254 0.02125 -0.0761 0.4904 1.0000
7.250 1.1320 0.03324 0.02206 -0.0746 0.4802 1.0000
7.500 1.1467 0.03405 0.02302 -0.0728 0.4700 1.0000
7.750 1.1762 0.03418 0.02321 -0.0727 0.4613 1.0000
8.000 1.1854 0.03518 0.02436 -0.0703 0.4501 1.0000
8.250 1.1979 0.03598 0.02528 -0.0682 0.4389 1.0000
8.500 1.2174 0.03636 0.02575 -0.0668 0.4278 1.0000
8.750 1.2385 0.03665 0.02614 -0.0655 0.4165 1.0000
9.000 1.2396 0.03786 0.02751 -0.0622 0.4050 1.0000
9.250 1.2482 0.03881 0.02858 -0.0598 0.3944 1.0000
9.500 1.2700 0.03916 0.02901 -0.0587 0.3846 1.0000
9.750 1.2699 0.04065 0.03068 -0.0557 0.3742 1.0000
10.000 1.2752 0.04198 0.03220 -0.0533 0.3645 1.0000
10.250 1.2978 0.04229 0.03261 -0.0524 0.3552 1.0000
10.500 1.2894 0.04453 0.03505 -0.0493 0.3453 1.0000
10.750 1.2956 0.04595 0.03663 -0.0474 0.3362 1.0000
11.000 1.3110 0.04675 0.03756 -0.0461 0.3271 1.0000
11.250 1.2995 0.04960 0.04060 -0.0436 0.3178 1.0000
11.500 1.3065 0.05088 0.04200 -0.0420 0.3077 1.0000
11.750 1.3163 0.05161 0.04283 -0.0404 0.2961 1.0000
12.000 1.3001 0.05526 0.04667 -0.0387 0.2864 1.0000
12.250 1.2947 0.05777 0.04929 -0.0374 0.2752 1.0000
12.500 1.2952 0.05963 0.05123 -0.0362 0.2635 1.0000
12.750 1.2954 0.06150 0.05313 -0.0351 0.2510 1.0000
13.000 1.2805 0.06583 0.05765 -0.0349 0.2413 1.0000
13.250 1.2728 0.06934 0.06129 -0.0347 0.2313 1.0000
13.500 1.2717 0.07185 0.06388 -0.0343 0.2200 1.0000
13.750 1.2654 0.07511 0.06720 -0.0343 0.2075 1.0000
14.000 1.2444 0.08134 0.07364 -0.0358 0.1983 1.0000
14.250 1.2312 0.08626 0.07866 -0.0370 0.1859 1.0000
14.500 1.2188 0.09130 0.08378 -0.0384 0.1731 1.0000
14.750 1.2064 0.09656 0.08913 -0.0401 0.1596 1.0000
15.000 1.1958 0.10167 0.09432 -0.0418 0.1455 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 591 AIRFOIL (goe591-il)