GOE 624 AIRFOIL (goe624-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 624 AIRFOIL (goe624-il) Reynolds number: 50,000 Max Cl/Cd: 22.81 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe624-il-50000-n5.txt Download as CSV file: xf-goe624-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 624 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.2667 0.13335 0.12641 -0.0450 1.0000 0.1343
-10.750 -0.2770 0.13188 0.12504 -0.0436 1.0000 0.1344
-10.250 -0.2972 0.12254 0.11580 -0.0423 1.0000 0.0945
-10.000 -0.2882 0.11953 0.11278 -0.0427 0.9967 0.0921
-9.750 -0.2741 0.11515 0.10838 -0.0467 0.9899 0.0910
-9.500 -0.2612 0.11063 0.10384 -0.0510 0.9831 0.0896
-9.250 -0.2524 0.10588 0.09908 -0.0555 0.9751 0.0876
-9.000 -0.2677 0.09780 0.09098 -0.0647 0.9648 0.0830
-8.750 -0.2545 0.09347 0.08663 -0.0687 0.9574 0.0824
-8.500 -0.2484 0.08906 0.08221 -0.0727 0.9477 0.0819
-8.250 -0.2450 0.08351 0.07663 -0.0790 0.9387 0.0817
-8.000 -0.2482 0.07759 0.07065 -0.0848 0.9266 0.0817
-7.750 -0.2543 0.07182 0.06474 -0.0892 0.9148 0.0818
-7.500 -0.2544 0.06531 0.05794 -0.0945 0.9063 0.0822
-7.250 -0.2603 0.06070 0.05303 -0.0955 0.8939 0.0824
-7.000 -0.2536 0.05549 0.04731 -0.0979 0.8862 0.0828
-6.750 -0.2452 0.05242 0.04396 -0.0977 0.8758 0.0835
-6.500 -0.2189 0.04991 0.04126 -0.0994 0.8704 0.0848
-6.250 -0.2071 0.04829 0.03946 -0.0982 0.8595 0.0864
-6.000 -0.1818 0.04575 0.03655 -0.0994 0.8533 0.0893
-5.750 -0.1672 0.04346 0.03377 -0.0987 0.8432 0.0921
-5.500 -0.1409 0.04110 0.03096 -0.0993 0.8363 0.0948
-5.250 -0.1170 0.03981 0.02954 -0.0991 0.8279 0.0974
-5.000 -0.0922 0.03848 0.02796 -0.0989 0.8195 0.1018
-4.750 -0.0585 0.03676 0.02588 -0.0999 0.8144 0.1079
-4.500 -0.0415 0.03618 0.02524 -0.0984 0.8032 0.1125
-4.250 -0.0090 0.03485 0.02359 -0.0990 0.7975 0.1209
-4.000 0.0111 0.03439 0.02316 -0.0979 0.7876 0.1280
-3.750 0.0407 0.03354 0.02216 -0.0981 0.7808 0.1386
-3.500 0.0668 0.03300 0.02145 -0.0977 0.7731 0.1510
-3.000 0.1249 0.03182 0.02023 -0.0977 0.7596 0.1781
-2.750 0.1437 0.03169 0.02009 -0.0964 0.7487 0.1898
-2.500 0.1765 0.03103 0.01932 -0.0967 0.7429 0.2057
-2.250 0.1978 0.03091 0.01915 -0.0957 0.7332 0.2209
-2.000 0.2272 0.03036 0.01863 -0.0956 0.7266 0.2410
-1.750 0.2512 0.03005 0.01841 -0.0950 0.7183 0.2651
-1.500 0.2765 0.02964 0.01814 -0.0944 0.7104 0.3011
-1.250 0.3078 0.02879 0.01756 -0.0946 0.7056 0.3664
-1.000 0.3216 0.02868 0.01793 -0.0925 0.6946 0.4555
-0.750 0.3439 0.02759 0.01775 -0.0898 0.6895 0.6682
-0.500 0.4092 0.02739 0.01785 -0.0955 0.6805 0.9955
-0.250 0.4367 0.02743 0.01757 -0.0952 0.6736 1.0000
0.000 0.4564 0.02778 0.01769 -0.0940 0.6652 1.0000
0.250 0.4791 0.02802 0.01771 -0.0932 0.6573 1.0000
0.500 0.5109 0.02788 0.01730 -0.0933 0.6525 1.0000
0.750 0.5222 0.02871 0.01803 -0.0912 0.6415 1.0000
1.000 0.5528 0.02868 0.01778 -0.0913 0.6364 1.0000
1.250 0.5666 0.02948 0.01848 -0.0896 0.6267 1.0000
1.500 0.5936 0.02965 0.01848 -0.0893 0.6205 1.0000
1.750 0.6181 0.02998 0.01866 -0.0887 0.6141 1.0000
2.000 0.6337 0.03077 0.01938 -0.0873 0.6052 1.0000
2.250 0.6647 0.03079 0.01924 -0.0874 0.6006 1.0000
2.500 0.6755 0.03190 0.02032 -0.0855 0.5914 1.0000
2.750 0.7002 0.03226 0.02058 -0.0851 0.5853 1.0000
3.000 0.7338 0.03217 0.02034 -0.0855 0.5815 1.0000
3.250 0.7350 0.03388 0.02208 -0.0828 0.5711 1.0000
3.500 0.7637 0.03405 0.02216 -0.0827 0.5663 1.0000
3.750 0.7806 0.03493 0.02299 -0.0816 0.5598 1.0000
4.000 0.7914 0.03615 0.02421 -0.0798 0.5518 1.0000
4.250 0.8230 0.03616 0.02415 -0.0801 0.5479 1.0000
4.500 0.8214 0.03820 0.02622 -0.0773 0.5388 1.0000
4.750 0.8424 0.03882 0.02681 -0.0766 0.5332 1.0000
5.000 0.8777 0.03860 0.02653 -0.0771 0.5299 1.0000
5.250 0.8564 0.04184 0.02985 -0.0728 0.5188 1.0000
5.500 0.8862 0.04190 0.02987 -0.0727 0.5147 1.0000
5.750 0.9112 0.04229 0.03024 -0.0723 0.5102 1.0000
6.000 0.8919 0.04579 0.03382 -0.0689 0.4994 1.0000
6.250 0.9284 0.04523 0.03323 -0.0691 0.4964 1.0000
6.750 0.9315 0.04964 0.03771 -0.0657 0.4804 1.0000
7.000 0.9650 0.04917 0.03723 -0.0656 0.4772 1.0000
7.500 0.9543 0.05518 0.04335 -0.0624 0.4585 1.0000
8.000 0.9326 0.06313 0.05143 -0.0602 0.4366 1.0000
8.250 0.9509 0.06386 0.05219 -0.0594 0.4307 1.0000
8.500 0.9848 0.06279 0.05113 -0.0586 0.4280 1.0000
9.000 0.9884 0.06814 0.05659 -0.0571 0.4109 1.0000
9.250 1.0255 0.06661 0.05509 -0.0562 0.4090 1.0000
9.750 1.0237 0.07267 0.06127 -0.0549 0.3914 1.0000
10.250 1.0229 0.07886 0.06758 -0.0539 0.3743 1.0000
10.500 0.9924 0.08615 0.07498 -0.0544 0.3599 1.0000
10.750 1.0209 0.08548 0.07435 -0.0534 0.3575 1.0000
11.000 0.9913 0.09281 0.08177 -0.0541 0.3437 1.0000
11.250 1.0196 0.09209 0.08110 -0.0531 0.3412 1.0000
11.500 0.9889 0.09982 0.08891 -0.0542 0.3279 1.0000
11.750 1.0137 0.09967 0.08882 -0.0533 0.3254 1.0000
12.000 0.9827 0.10773 0.09694 -0.0549 0.3130 1.0000
12.250 1.0065 0.10766 0.09695 -0.0541 0.3101 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 624 AIRFOIL (goe624-il)