GOE 359 AIRFOIL (goe359-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 359 AIRFOIL (goe359-il) Reynolds number: 50,000 Max Cl/Cd: 41.45 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe359-il-50000-n5.txt Download as CSV file: xf-goe359-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 359 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
------ -------- --------- --------- -------- -------- --------
-7.250 -0.2609 0.10120 0.09522 -0.0274 1.0000 0.0692
-7.000 -0.2655 0.09965 0.09382 -0.0267 1.0000 0.0714
-6.750 -0.2729 0.09878 0.09310 -0.0268 1.0000 0.0735
-6.500 -0.2823 0.09849 0.09296 -0.0277 1.0000 0.0745
-6.250 -0.2893 0.09809 0.09266 -0.0293 1.0000 0.0749
-6.000 -0.2904 0.09353 0.08821 -0.0228 1.0000 0.0778
-5.750 -0.2671 0.09055 0.08521 -0.0300 0.9918 0.0863
-5.500 -0.2404 0.08569 0.08031 -0.0355 0.9818 0.0907
-5.250 -0.1937 0.08270 0.07714 -0.0508 0.9677 0.1013
-5.000 -0.1785 0.07724 0.07175 -0.0494 0.9588 0.1048
-4.500 -0.1128 0.06923 0.06357 -0.0617 0.9370 0.1272
-4.250 -0.0748 0.06530 0.05952 -0.0690 0.9268 0.1436
-4.000 -0.0428 0.06116 0.05528 -0.0735 0.9161 0.1520
-3.750 0.0005 0.05713 0.05105 -0.0810 0.9073 0.1643
-3.250 0.1165 0.04626 0.03903 -0.0999 0.8889 0.0871
-3.000 0.1562 0.04240 0.03495 -0.1038 0.8814 0.0817
-2.750 0.1948 0.03951 0.03163 -0.1072 0.8702 0.0837
-2.500 0.2338 0.03674 0.02842 -0.1102 0.8600 0.0832
-2.250 0.2760 0.03397 0.02513 -0.1133 0.8519 0.0806
-2.000 0.3107 0.03191 0.02255 -0.1147 0.8402 0.0791
-1.750 0.3447 0.03020 0.02035 -0.1158 0.8292 0.0786
-1.500 0.3783 0.02887 0.01864 -0.1167 0.8188 0.0806
-1.250 0.4120 0.02785 0.01717 -0.1173 0.8085 0.0863
-1.000 0.4413 0.02697 0.01589 -0.1173 0.7965 0.0874
-0.750 0.4703 0.02618 0.01482 -0.1170 0.7852 0.0878
-0.500 0.5014 0.02545 0.01382 -0.1170 0.7749 0.0889
-0.250 0.5314 0.02489 0.01303 -0.1169 0.7643 0.0905
0.000 0.5587 0.02454 0.01253 -0.1165 0.7527 0.0928
0.250 0.5865 0.02425 0.01208 -0.1161 0.7418 0.0960
0.500 0.6158 0.02396 0.01159 -0.1159 0.7324 0.1009
0.750 0.6423 0.02382 0.01143 -0.1155 0.7213 0.1088
1.000 0.6690 0.02371 0.01132 -0.1152 0.7106 0.1258
1.250 0.6948 0.02238 0.01136 -0.1152 0.7019 0.4901
1.500 0.7213 0.02201 0.01121 -0.1142 0.6915 1.0000
1.750 0.7469 0.02238 0.01135 -0.1137 0.6812 1.0000
2.000 0.7745 0.02264 0.01139 -0.1134 0.6725 1.0000
2.250 0.8000 0.02302 0.01164 -0.1129 0.6625 1.0000
2.500 0.8251 0.02343 0.01193 -0.1124 0.6526 1.0000
2.750 0.8521 0.02366 0.01203 -0.1119 0.6432 1.0000
3.000 0.8771 0.02392 0.01221 -0.1111 0.6314 1.0000
3.250 0.9011 0.02420 0.01241 -0.1101 0.6183 1.0000
3.500 0.9251 0.02447 0.01262 -0.1092 0.6054 1.0000
3.750 0.9494 0.02476 0.01289 -0.1084 0.5933 1.0000
4.000 0.9748 0.02503 0.01313 -0.1077 0.5828 1.0000
4.250 0.9993 0.02539 0.01350 -0.1070 0.5725 1.0000
4.500 1.0221 0.02588 0.01405 -0.1062 0.5616 1.0000
4.750 1.0465 0.02628 0.01451 -0.1055 0.5516 1.0000
5.000 1.0716 0.02660 0.01486 -0.1048 0.5418 1.0000
5.250 1.0934 0.02715 0.01552 -0.1039 0.5304 1.0000
5.500 1.1162 0.02763 0.01614 -0.1030 0.5196 1.0000
5.750 1.1403 0.02800 0.01658 -0.1022 0.5091 1.0000
6.000 1.1636 0.02842 0.01712 -0.1014 0.4981 1.0000
6.250 1.1846 0.02902 0.01787 -0.1003 0.4861 1.0000
6.500 1.2062 0.02957 0.01862 -0.0993 0.4744 1.0000
6.750 1.2285 0.03007 0.01925 -0.0983 0.4628 1.0000
7.000 1.2503 0.03050 0.01980 -0.0971 0.4500 1.0000
7.250 1.2662 0.03077 0.02015 -0.0950 0.4289 1.0000
7.500 1.2809 0.03090 0.02027 -0.0926 0.4051 1.0000
7.750 1.2943 0.03147 0.02102 -0.0904 0.3843 1.0000
8.000 1.3087 0.03206 0.02170 -0.0884 0.3658 1.0000
8.250 1.3202 0.03281 0.02255 -0.0862 0.3451 1.0000
8.500 1.3313 0.03365 0.02349 -0.0841 0.3257 1.0000
8.750 1.3382 0.03461 0.02444 -0.0815 0.3016 1.0000
9.000 1.3415 0.03579 0.02550 -0.0785 0.2776 1.0000
9.250 1.3447 0.03720 0.02691 -0.0759 0.2579 1.0000
9.500 1.3477 0.03882 0.02855 -0.0735 0.2408 1.0000
9.750 1.3511 0.04058 0.03040 -0.0715 0.2267 1.0000
10.000 1.3530 0.04260 0.03262 -0.0697 0.2119 1.0000
10.250 1.3529 0.04496 0.03511 -0.0682 0.1940 1.0000
10.500 1.3513 0.04766 0.03795 -0.0671 0.1695 1.0000
10.750 1.3432 0.05113 0.04133 -0.0662 0.1335 1.0000
11.250 1.3118 0.06057 0.05025 -0.0657 0.0649 1.0000
11.500 1.2986 0.06543 0.05506 -0.0659 0.0562 1.0000
11.750 1.2885 0.07005 0.05977 -0.0661 0.0479 1.0000
12.000 1.2804 0.07457 0.06446 -0.0665 0.0392 1.0000
12.250 1.2720 0.07923 0.06929 -0.0671 0.0349 1.0000
12.500 1.2632 0.08405 0.07429 -0.0679 0.0328 1.0000
12.750 1.2536 0.08913 0.07954 -0.0690 0.0314 1.0000
13.000 1.2439 0.09437 0.08493 -0.0702 0.0305 1.0000
13.250 1.2347 0.09968 0.09041 -0.0716 0.0299 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 359 AIRFOIL (goe359-il)