GOE 511 AIRFOIL (goe511-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 511 AIRFOIL (goe511-il) Reynolds number: 200,000 Max Cl/Cd: 57.64 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe511-il-200000.txt Download as CSV file: xf-goe511-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 511 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 0.0500 0.09612 0.09267 -0.0964 0.9265 0.0640
-8.750 0.0445 0.09278 0.08934 -0.1009 0.9176 0.0673
-8.500 0.0455 0.08804 0.08462 -0.1061 0.9107 0.0684
-8.250 0.0852 0.08472 0.08128 -0.1075 0.9091 0.0699
-8.000 0.1132 0.08095 0.07750 -0.1123 0.9071 0.0725
-7.250 0.1962 0.05711 0.05380 -0.1199 0.8580 0.0835
-7.000 0.0990 0.06870 0.06525 -0.1132 0.8617 0.0789
-6.750 0.1224 0.06531 0.06181 -0.1161 0.8567 0.0816
-6.250 0.1028 0.05892 0.05523 -0.1133 0.8233 0.0873
-6.000 0.1183 0.05690 0.05320 -0.1118 0.8082 0.0885
-5.750 0.1415 0.05443 0.05065 -0.1128 0.7930 0.0915
-5.500 0.1578 0.05008 0.04601 -0.1159 0.7768 0.0987
-5.250 0.2039 0.04755 0.04334 -0.1206 0.7593 0.1019
-5.000 0.2295 0.04371 0.03907 -0.1242 0.7397 0.1110
-4.750 0.2638 0.04194 0.03711 -0.1264 0.7183 0.1134
-4.500 0.2837 0.04050 0.03543 -0.1260 0.6984 0.1186
-4.250 0.2341 0.02715 0.02043 -0.1140 0.6926 0.0756
-4.000 0.2526 0.02538 0.01835 -0.1124 0.6768 0.0743
-3.750 0.2642 0.02382 0.01645 -0.1090 0.6616 0.0725
-3.500 0.2757 0.02203 0.01403 -0.1052 0.6484 0.0695
-3.250 0.2961 0.02114 0.01275 -0.1033 0.6358 0.0689
-3.000 0.3150 0.02048 0.01184 -0.1012 0.6234 0.0688
-2.750 0.3338 0.01986 0.01106 -0.0991 0.6118 0.0691
-2.250 0.3766 0.01901 0.00996 -0.0961 0.5908 0.0715
-2.000 0.4053 0.01871 0.00946 -0.0961 0.5817 0.0733
-1.750 0.4237 0.01840 0.00909 -0.0939 0.5724 0.0745
-1.500 0.4557 0.01817 0.00866 -0.0946 0.5639 0.0759
-1.250 0.4729 0.01782 0.00835 -0.0923 0.5554 0.0776
-1.000 0.4982 0.01764 0.00814 -0.0917 0.5473 0.0802
-0.750 0.5197 0.01756 0.00803 -0.0903 0.5394 0.0837
-0.500 0.5382 0.01742 0.00789 -0.0882 0.5316 0.0882
-0.250 0.5680 0.01742 0.00779 -0.0886 0.5241 0.0958
0.000 0.5801 0.01727 0.00772 -0.0852 0.5167 0.1045
0.250 0.5994 0.01696 0.00774 -0.0835 0.5100 0.1686
0.500 0.6262 0.01741 0.00817 -0.0832 0.5033 0.2520
0.750 0.6401 0.01758 0.00835 -0.0803 0.4964 0.2697
1.000 0.6622 0.01778 0.00847 -0.0791 0.4902 0.2860
1.250 0.6875 0.01801 0.00863 -0.0787 0.4841 0.3023
1.750 0.7215 0.01824 0.00885 -0.0744 0.4721 0.3271
2.000 0.7502 0.01842 0.00895 -0.0747 0.4663 0.3429
2.250 0.7605 0.01849 0.00912 -0.0713 0.4607 0.3559
2.500 0.7789 0.01860 0.00921 -0.0695 0.4553 0.3679
2.750 0.8022 0.01864 0.00924 -0.0686 0.4507 0.3787
3.000 0.8236 0.01880 0.00939 -0.0675 0.4462 0.3899
3.250 0.8363 0.01890 0.00957 -0.0646 0.4416 0.4022
3.500 1.1381 0.01990 0.01154 -0.1229 0.4281 1.0000
3.750 1.1496 0.02014 0.01179 -0.1198 0.4245 1.0000
4.000 1.1629 0.02036 0.01199 -0.1171 0.4207 1.0000
4.250 1.1792 0.02059 0.01215 -0.1151 0.4173 1.0000
4.500 1.2019 0.02088 0.01233 -0.1143 0.4140 1.0000
4.750 1.2265 0.02128 0.01265 -0.1140 0.4107 1.0000
5.000 1.2334 0.02153 0.01297 -0.1101 0.4078 1.0000
5.250 1.2456 0.02181 0.01327 -0.1073 0.4049 1.0000
5.500 1.2611 0.02211 0.01355 -0.1051 0.4020 1.0000
5.750 1.2799 0.02240 0.01380 -0.1036 0.3994 1.0000
6.000 1.3039 0.02274 0.01409 -0.1033 0.3969 1.0000
6.250 1.3433 0.02333 0.01455 -0.1061 0.3941 1.0000
6.500 1.3470 0.02366 0.01497 -0.1017 0.3921 1.0000
6.750 1.3534 0.02401 0.01541 -0.0978 0.3896 1.0000
7.000 1.3638 0.02438 0.01583 -0.0948 0.3871 1.0000
7.250 1.3762 0.02469 0.01615 -0.0923 0.3841 1.0000
7.500 1.3941 0.02499 0.01642 -0.0907 0.3814 1.0000
7.750 1.4201 0.02529 0.01663 -0.0908 0.3784 1.0000
8.000 1.4431 0.02583 0.01715 -0.0904 0.3756 1.0000
8.250 1.4416 0.02625 0.01769 -0.0854 0.3732 1.0000
8.500 1.4461 0.02670 0.01822 -0.0815 0.3705 1.0000
8.750 1.4547 0.02707 0.01863 -0.0785 0.3673 1.0000
9.000 1.4704 0.02736 0.01891 -0.0767 0.3643 1.0000
9.250 1.4947 0.02767 0.01917 -0.0766 0.3617 1.0000
9.500 1.5302 0.02823 0.01964 -0.0786 0.3590 1.0000
9.750 1.5252 0.02885 0.02043 -0.0733 0.3570 1.0000
10.000 1.5240 0.02953 0.02124 -0.0690 0.3546 1.0000
10.250 1.5281 0.03017 0.02198 -0.0656 0.3519 1.0000
10.500 1.5384 0.03076 0.02261 -0.0633 0.3495 1.0000
10.750 1.5548 0.03115 0.02301 -0.0620 0.3468 1.0000
11.000 1.5806 0.03146 0.02328 -0.0622 0.3444 1.0000
11.250 1.6156 0.03197 0.02372 -0.0640 0.3418 1.0000
11.500 1.6019 0.03299 0.02495 -0.0581 0.3396 1.0000
11.750 1.5933 0.03411 0.02622 -0.0534 0.3372 1.0000
12.000 1.5906 0.03513 0.02738 -0.0497 0.3343 1.0000
12.250 1.5982 0.03585 0.02815 -0.0475 0.3315 1.0000
12.500 1.6190 0.03612 0.02839 -0.0470 0.3286 1.0000
12.750 1.6590 0.03603 0.02820 -0.0492 0.3256 1.0000
13.000 1.6463 0.03754 0.02989 -0.0445 0.3230 1.0000
13.250 1.6222 0.03954 0.03210 -0.0388 0.3200 1.0000
13.500 1.6122 0.04121 0.03391 -0.0352 0.3168 1.0000
13.750 1.6201 0.04207 0.03480 -0.0336 0.3138 1.0000
14.000 1.6493 0.04188 0.03456 -0.0341 0.3108 1.0000
14.250 1.6791 0.04202 0.03466 -0.0348 0.3076 1.0000
14.500 1.6289 0.04590 0.03885 -0.0281 0.3042 1.0000
14.750 1.5998 0.04927 0.04241 -0.0241 0.3002 1.0000
15.000 1.6093 0.05014 0.04330 -0.0232 0.2966 1.0000
15.250 1.6500 0.04893 0.04199 -0.0242 0.2935 1.0000
15.500 1.6274 0.05225 0.04546 -0.0214 0.2895 1.0000
15.750 1.5601 0.05970 0.05321 -0.0173 0.2837 1.0000
16.000 1.5803 0.05974 0.05323 -0.0171 0.2799 1.0000
16.250 1.6326 0.05696 0.05031 -0.0180 0.2767 1.0000
16.500 1.4926 0.07324 0.06707 -0.0145 0.2677 1.0000
16.750 1.5223 0.07219 0.06599 -0.0144 0.2645 1.0000
17.000 1.5791 0.06815 0.06181 -0.0146 0.2617 1.0000
17.250 1.2055 0.12049 0.11475 -0.0203 0.2293 1.0000
17.500 1.2394 0.11817 0.11244 -0.0196 0.2294 1.0000
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