GOE 365 AIRFOIL (goe365-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 365 AIRFOIL (goe365-il) Reynolds number: 100,000 Max Cl/Cd: 55.84 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe365-il-100000.txt Download as CSV file: xf-goe365-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 365 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3310 0.11154 0.10728 -0.0289 1.0000 0.0698
-8.500 -0.3528 0.11153 0.10737 -0.0263 1.0000 0.0701
-8.250 -0.3772 0.11155 0.10748 -0.0243 1.0000 0.0703
-8.000 -0.3879 0.11009 0.10608 -0.0292 0.9950 0.0707
-7.750 -0.3791 0.10488 0.10086 -0.0375 0.9876 0.0714
-7.500 -0.3450 0.09933 0.09531 -0.0299 0.9887 0.0745
-7.250 -0.3298 0.09585 0.09182 -0.0334 0.9837 0.0775
-7.000 -0.3235 0.09261 0.08859 -0.0376 0.9757 0.0805
-6.750 -0.3159 0.08902 0.08489 -0.0498 0.9645 0.0844
-6.500 -0.3004 0.08347 0.07929 -0.0557 0.9592 0.0866
-6.250 -0.2896 0.08080 0.07669 -0.0520 0.9540 0.0896
-6.000 -0.2728 0.07756 0.07341 -0.0549 0.9477 0.0953
-5.750 -0.2637 0.07410 0.06957 -0.0641 0.9376 0.1014
-5.500 -0.2455 0.07009 0.06574 -0.0624 0.9337 0.1044
-5.250 -0.2181 0.06687 0.06242 -0.0661 0.9293 0.1125
-5.000 -0.2121 0.06383 0.05916 -0.0678 0.9199 0.1189
-4.750 -0.1788 0.06153 0.05646 -0.0731 0.9149 0.1335
-4.500 -0.1754 0.05826 0.05344 -0.0697 0.9073 0.1365
-4.250 -0.1490 0.05542 0.05046 -0.0721 0.9017 0.1526
-4.000 -0.1307 0.05334 0.04824 -0.0727 0.8947 0.1682
-3.750 -0.1104 0.05101 0.04586 -0.0730 0.8878 0.1853
-3.500 -0.0760 0.04854 0.04325 -0.0758 0.8837 0.2173
-3.250 -0.0713 0.04710 0.04186 -0.0727 0.8741 0.2351
-1.750 0.1488 0.03390 0.02581 -0.0808 0.8379 0.1307
-1.250 0.2135 0.03117 0.02249 -0.0806 0.8239 0.1092
-1.000 0.2674 0.02976 0.02064 -0.0836 0.8210 0.1021
-0.750 0.2821 0.02960 0.02038 -0.0811 0.8099 0.1017
-0.500 0.3335 0.02800 0.01878 -0.0845 0.8066 0.1041
-0.250 0.3554 0.02767 0.01844 -0.0831 0.7969 0.1042
0.000 0.3996 0.02655 0.01736 -0.0850 0.7920 0.1057
0.250 0.4388 0.02567 0.01651 -0.0861 0.7867 0.1091
0.500 0.4633 0.02540 0.01621 -0.0850 0.7768 0.1144
0.750 0.5128 0.02429 0.01506 -0.0877 0.7734 0.1297
1.000 0.6015 0.02156 0.01457 -0.0990 0.7656 1.0000
1.250 0.6435 0.02105 0.01374 -0.1004 0.7590 1.0000
1.500 0.6667 0.02097 0.01351 -0.0988 0.7465 1.0000
1.750 0.6985 0.02054 0.01291 -0.0984 0.7349 1.0000
2.000 0.7418 0.01972 0.01189 -0.0997 0.7256 1.0000
2.250 0.7625 0.01969 0.01178 -0.0978 0.7118 1.0000
2.500 0.7844 0.01971 0.01172 -0.0962 0.6990 1.0000
2.750 0.8105 0.01961 0.01153 -0.0953 0.6875 1.0000
3.000 0.8440 0.01929 0.01109 -0.0955 0.6776 1.0000
3.250 0.8637 0.01935 0.01112 -0.0936 0.6638 1.0000
3.500 0.8851 0.01936 0.01109 -0.0920 0.6504 1.0000
3.750 0.9079 0.01932 0.01101 -0.0906 0.6370 1.0000
4.000 0.9317 0.01925 0.01090 -0.0893 0.6237 1.0000
4.250 0.9556 0.01915 0.01075 -0.0881 0.6097 1.0000
4.500 0.9789 0.01907 0.01062 -0.0868 0.5947 1.0000
4.750 1.0015 0.01903 0.01054 -0.0854 0.5787 1.0000
5.000 1.0239 0.01903 0.01047 -0.0840 0.5616 1.0000
5.250 1.0464 0.01908 0.01042 -0.0826 0.5436 1.0000
5.500 1.0695 0.01921 0.01042 -0.0814 0.5252 1.0000
5.750 1.0895 0.01951 0.01063 -0.0797 0.5053 1.0000
6.000 1.1082 0.01990 0.01093 -0.0779 0.4843 1.0000
6.250 1.1285 0.02033 0.01120 -0.0764 0.4638 1.0000
6.500 1.1466 0.02087 0.01164 -0.0746 0.4431 1.0000
6.750 1.1634 0.02147 0.01214 -0.0727 0.4229 1.0000
7.000 1.1813 0.02209 0.01265 -0.0710 0.4046 1.0000
7.250 1.1991 0.02272 0.01319 -0.0694 0.3880 1.0000
7.500 1.2172 0.02340 0.01382 -0.0678 0.3735 1.0000
7.750 1.2359 0.02410 0.01449 -0.0665 0.3607 1.0000
8.000 1.2572 0.02483 0.01516 -0.0656 0.3500 1.0000
8.250 1.2779 0.02554 0.01584 -0.0646 0.3400 1.0000
8.500 1.2954 0.02631 0.01670 -0.0631 0.3306 1.0000
8.750 1.3196 0.02702 0.01731 -0.0628 0.3221 1.0000
9.000 1.3329 0.02776 0.01821 -0.0607 0.3135 1.0000
9.250 1.3581 0.02856 0.01894 -0.0606 0.3064 1.0000
9.500 1.3719 0.02940 0.02000 -0.0587 0.2995 1.0000
9.750 1.3994 0.03021 0.02072 -0.0590 0.2931 1.0000
10.000 1.4117 0.03119 0.02194 -0.0569 0.2873 1.0000
10.250 1.4312 0.03207 0.02293 -0.0560 0.2813 1.0000
10.500 1.4551 0.03308 0.02397 -0.0558 0.2756 1.0000
10.750 1.4627 0.03405 0.02518 -0.0531 0.2695 1.0000
11.000 1.4864 0.03474 0.02576 -0.0528 0.2615 1.0000
11.250 1.4845 0.03561 0.02690 -0.0487 0.2547 1.0000
11.500 1.5012 0.03614 0.02736 -0.0473 0.2460 1.0000
11.750 1.4947 0.03702 0.02847 -0.0426 0.2397 1.0000
12.000 1.5127 0.03760 0.02895 -0.0416 0.2317 1.0000
12.250 1.5042 0.03875 0.03038 -0.0370 0.2263 1.0000
12.500 1.5196 0.03937 0.03095 -0.0358 0.2193 1.0000
12.750 1.5148 0.04068 0.03251 -0.0323 0.2139 1.0000
13.000 1.5132 0.04174 0.03371 -0.0294 0.2077 1.0000
13.250 1.5210 0.04267 0.03466 -0.0276 0.2015 1.0000
13.500 1.5112 0.04439 0.03666 -0.0244 0.1965 1.0000
13.750 1.5175 0.04525 0.03753 -0.0228 0.1903 1.0000
14.000 1.5069 0.04720 0.03971 -0.0202 0.1850 1.0000
14.250 1.4978 0.04909 0.04177 -0.0181 0.1790 1.0000
14.500 1.4954 0.05066 0.04339 -0.0166 0.1730 1.0000
14.750 1.4812 0.05353 0.04657 -0.0151 0.1678 1.0000
15.000 1.4749 0.05561 0.04870 -0.0141 0.1613 1.0000
15.250 1.4574 0.05929 0.05266 -0.0134 0.1551 1.0000
15.500 1.4455 0.06258 0.05609 -0.0131 0.1483 1.0000
15.750 1.4281 0.06707 0.06082 -0.0135 0.1410 1.0000
16.000 1.4120 0.07165 0.06550 -0.0143 0.1329 1.0000
16.250 1.3922 0.07742 0.07152 -0.0159 0.1239 1.0000
16.500 1.3723 0.08354 0.07780 -0.0178 0.1150 1.0000
16.750 1.3524 0.08994 0.08425 -0.0201 0.1073 1.0000
17.000 1.3290 0.09719 0.09153 -0.0230 0.0995 1.0000
17.250 1.3077 0.10430 0.09868 -0.0259 0.0937 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 365 AIRFOIL (goe365-il)