GOE 414 AIRFOIL (goe414-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 414 AIRFOIL (goe414-il) Reynolds number: 50,000 Max Cl/Cd: 28.74 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe414-il-50000-n5.txt Download as CSV file: xf-goe414-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 414 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3435 0.11377 0.10764 -0.0398 1.0000 0.0634
-9.000 -0.3529 0.11147 0.10543 -0.0383 1.0000 0.0630
-8.750 -0.3636 0.10920 0.10325 -0.0368 1.0000 0.0625
-8.500 -0.3620 0.10516 0.09924 -0.0395 0.9958 0.0620
-8.250 -0.3573 0.10046 0.09453 -0.0439 0.9889 0.0619
-8.000 -0.3562 0.09579 0.08988 -0.0484 0.9808 0.0623
-7.750 -0.3530 0.09019 0.08429 -0.0549 0.9721 0.0631
-7.500 -0.3465 0.08335 0.07741 -0.0630 0.9630 0.0638
-7.250 -0.3430 0.07637 0.07038 -0.0698 0.9527 0.0642
-7.000 -0.3379 0.06855 0.06235 -0.0767 0.9436 0.0644
-6.750 -0.3283 0.06047 0.05384 -0.0834 0.9356 0.0650
-6.500 -0.3102 0.05897 0.05231 -0.0838 0.9277 0.0677
-6.250 -0.2877 0.05546 0.04851 -0.0869 0.9210 0.0717
-6.000 -0.2722 0.05026 0.04275 -0.0896 0.9127 0.0742
-5.750 -0.2474 0.04475 0.03627 -0.0930 0.9066 0.0776
-5.500 -0.2229 0.04350 0.03497 -0.0937 0.8996 0.0824
-5.250 -0.1952 0.04119 0.03219 -0.0950 0.8929 0.0878
-5.000 -0.1610 0.03823 0.02846 -0.0972 0.8882 0.0931
-4.500 -0.1017 0.03575 0.02551 -0.0989 0.8730 0.1103
-4.250 -0.0707 0.03435 0.02370 -0.0996 0.8638 0.1217
-4.000 -0.0400 0.03331 0.02254 -0.1002 0.8545 0.1344
-3.750 -0.0016 0.03240 0.02145 -0.1021 0.8474 0.1541
-3.500 0.0237 0.03203 0.02108 -0.1019 0.8366 0.1745
-3.250 0.0643 0.03126 0.02010 -0.1041 0.8316 0.1965
-3.000 0.0864 0.03087 0.01963 -0.1031 0.8211 0.2069
-2.750 0.1261 0.03011 0.01871 -0.1049 0.8167 0.2229
-2.500 0.1483 0.02991 0.01843 -0.1040 0.8060 0.2389
-2.250 0.1818 0.02938 0.01781 -0.1048 0.7996 0.2604
-2.000 0.2084 0.02904 0.01748 -0.1045 0.7906 0.2843
-1.750 0.2377 0.02864 0.01709 -0.1047 0.7830 0.3213
-1.500 0.2661 0.02820 0.01675 -0.1048 0.7749 0.3689
-1.250 0.2913 0.02780 0.01653 -0.1043 0.7662 0.4249
-1.000 0.3195 0.02704 0.01617 -0.1040 0.7592 0.5142
-0.750 0.3623 0.02580 0.01602 -0.1055 0.7512 1.0000
-0.500 0.3939 0.02573 0.01560 -0.1060 0.7437 1.0000
-0.250 0.4163 0.02597 0.01557 -0.1051 0.7333 1.0000
0.000 0.4513 0.02579 0.01509 -0.1060 0.7273 1.0000
0.250 0.4715 0.02611 0.01523 -0.1049 0.7162 1.0000
0.500 0.4986 0.02621 0.01513 -0.1046 0.7076 1.0000
0.750 0.5276 0.02624 0.01496 -0.1047 0.6995 1.0000
1.000 0.5506 0.02651 0.01508 -0.1039 0.6895 1.0000
1.250 0.5842 0.02638 0.01477 -0.1046 0.6832 1.0000
1.500 0.6047 0.02680 0.01508 -0.1035 0.6727 1.0000
1.750 0.6400 0.02665 0.01477 -0.1044 0.6673 1.0000
2.000 0.6584 0.02721 0.01526 -0.1031 0.6566 1.0000
2.250 0.6868 0.02738 0.01532 -0.1032 0.6493 1.0000
2.500 0.7111 0.02773 0.01559 -0.1027 0.6407 1.0000
2.750 0.7366 0.02807 0.01585 -0.1024 0.6330 1.0000
3.000 0.7615 0.02845 0.01619 -0.1020 0.6254 1.0000
3.250 0.7875 0.02884 0.01652 -0.1018 0.6188 1.0000
3.500 0.8084 0.02945 0.01713 -0.1010 0.6109 1.0000
3.750 0.8428 0.02947 0.01709 -0.1019 0.6063 1.0000
4.000 0.8527 0.03063 0.01829 -0.0998 0.5967 1.0000
4.250 0.8845 0.03078 0.01841 -0.1003 0.5920 1.0000
4.500 0.8940 0.03205 0.01977 -0.0983 0.5837 1.0000
4.750 0.9204 0.03249 0.02022 -0.0982 0.5784 1.0000
5.000 0.9379 0.03339 0.02116 -0.0971 0.5720 1.0000
5.250 0.9530 0.03441 0.02225 -0.0957 0.5652 1.0000
5.500 0.9854 0.03455 0.02244 -0.0964 0.5613 1.0000
5.750 0.9825 0.03655 0.02455 -0.0930 0.5525 1.0000
6.000 1.0106 0.03685 0.02490 -0.0931 0.5477 1.0000
6.250 1.0122 0.03847 0.02662 -0.0901 0.5392 1.0000
6.500 1.0399 0.03860 0.02683 -0.0899 0.5331 1.0000
6.750 1.0398 0.04025 0.02857 -0.0868 0.5242 1.0000
7.000 1.0673 0.04035 0.02876 -0.0865 0.5183 1.0000
7.250 1.0641 0.04242 0.03093 -0.0835 0.5098 1.0000
7.500 1.0856 0.04297 0.03162 -0.0827 0.5041 1.0000
7.750 1.0879 0.04490 0.03367 -0.0805 0.4967 1.0000
8.000 1.0973 0.04639 0.03528 -0.0789 0.4899 1.0000
8.250 1.1371 0.04568 0.03472 -0.0795 0.4868 1.0000
8.500 1.1057 0.05037 0.03952 -0.0754 0.4754 1.0000
8.750 1.1437 0.04963 0.03897 -0.0757 0.4723 1.0000
9.250 1.1460 0.05437 0.04401 -0.0722 0.4575 1.0000
9.750 1.1253 0.06185 0.05173 -0.0688 0.4385 1.0000
10.250 1.1527 0.06458 0.05482 -0.0670 0.4262 1.0000
10.500 1.1174 0.07178 0.06208 -0.0659 0.4123 1.0000
11.000 1.1145 0.07870 0.06930 -0.0647 0.3960 1.0000
11.500 1.0683 0.09168 0.08240 -0.0650 0.3731 1.0000
13.000 1.1388 0.09184 0.08368 -0.0563 0.2871 1.0000
13.250 1.1534 0.09123 0.08324 -0.0550 0.2637 1.0000
13.500 1.1568 0.09292 0.08500 -0.0543 0.2321 1.0000
13.750 1.1526 0.09606 0.08801 -0.0542 0.1858 1.0000
14.000 1.1402 0.09955 0.09032 -0.0535 0.0990 1.0000
14.250 1.1265 0.10498 0.09560 -0.0544 0.0753 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 414 AIRFOIL (goe414-il)