GOE 451 AIRFOIL (modified line 6) (goe451-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 451 AIRFOIL (modified line 6) (goe451-il) Reynolds number: 100,000 Max Cl/Cd: 29.16 at α=1.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe451-il-100000.txt Download as CSV file: xf-goe451-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 451 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4165 0.09864 0.09371 -0.0220 1.0000 0.0354
-7.500 -0.4156 0.09623 0.09138 -0.0221 1.0000 0.0372
-7.250 -0.4140 0.09399 0.08921 -0.0225 1.0000 0.0388
-7.000 -0.4088 0.09162 0.08690 -0.0244 1.0000 0.0401
-6.750 -0.4006 0.08974 0.08508 -0.0277 1.0000 0.0412
-6.500 -0.3877 0.08803 0.08340 -0.0329 1.0000 0.0418
-6.250 -0.3703 0.08595 0.08133 -0.0385 1.0000 0.0422
-6.000 -0.3517 0.08313 0.07849 -0.0432 1.0000 0.0424
-5.750 -0.3322 0.08005 0.07536 -0.0472 1.0000 0.0425
-5.500 -0.3327 0.07377 0.06922 -0.0446 1.0000 0.0432
-5.250 -0.3291 0.06923 0.06465 -0.0425 1.0000 0.0442
-5.000 -0.3183 0.06554 0.06099 -0.0428 1.0000 0.0455
-4.750 -0.3025 0.06201 0.05746 -0.0445 1.0000 0.0471
-4.500 -0.2826 0.05856 0.05399 -0.0473 1.0000 0.0491
-4.250 -0.2580 0.05510 0.05047 -0.0509 1.0000 0.0517
-3.750 -0.1914 0.04732 0.04241 -0.0612 0.9989 0.0580
-3.500 -0.1728 0.04414 0.03921 -0.0620 0.9989 0.0621
-3.250 -0.1338 0.04073 0.03550 -0.0672 0.9985 0.0711
-3.000 -0.1119 0.03808 0.03277 -0.0681 0.9987 0.0807
-2.750 -0.0735 0.03446 0.02896 -0.0728 0.9956 0.0997
-2.500 -0.0389 0.03191 0.02627 -0.0763 0.9923 0.1393
-2.000 0.0685 0.02451 0.01762 -0.0871 0.9803 0.1214
-1.750 0.1201 0.02058 0.01286 -0.0905 0.9722 0.0923
-1.500 0.1722 0.01750 0.00913 -0.0945 0.9587 0.1030
-1.250 0.2228 0.01520 0.00626 -0.0978 0.9360 0.0948
-1.000 0.2646 0.01372 0.00464 -0.1001 0.9034 0.0940
-0.750 0.3140 0.01269 0.00313 -0.1035 0.8118 0.0985
-0.500 0.3417 0.01281 0.00243 -0.1027 0.6606 0.1128
-0.250 0.3542 0.01510 0.00291 -0.1009 0.0572 0.2799
0.000 0.3737 0.01440 0.00391 -0.0994 0.0377 1.0000
0.250 0.3941 0.01542 0.00493 -0.0986 0.0311 1.0000
0.500 0.4206 0.01516 0.00457 -0.0988 0.0264 1.0000
0.750 0.4352 0.01698 0.00618 -0.0973 0.0231 1.0000
1.000 0.4649 0.01620 0.00547 -0.0981 0.0210 1.0000
1.250 0.4867 0.01669 0.00588 -0.0977 0.0191 1.0000
1.500 0.5064 0.01754 0.00659 -0.0970 0.0176 1.0000
1.750 0.5257 0.01946 0.00799 -0.0964 0.0165 1.0000
2.000 0.5510 0.02151 0.00960 -0.0969 0.0159 1.0000
2.250 0.5729 0.02067 0.00903 -0.0964 0.0154 1.0000
2.500 0.5961 0.02076 0.00920 -0.0961 0.0144 1.0000
2.750 0.6201 0.02163 0.01018 -0.0960 0.0132 1.0000
3.000 0.6449 0.02314 0.01155 -0.0962 0.0117 1.0000
3.250 0.6720 0.02588 0.01404 -0.0970 0.0110 1.0000
3.750 0.7229 0.03129 0.01930 -0.0977 0.0105 1.0000
4.000 0.7435 0.02902 0.01750 -0.0962 0.0102 1.0000
4.250 0.7666 0.02930 0.01800 -0.0954 0.0096 1.0000
4.500 0.7902 0.03061 0.01948 -0.0950 0.0089 1.0000
4.750 0.8135 0.03250 0.02156 -0.0946 0.0083 1.0000
5.000 0.8364 0.03459 0.02404 -0.0942 0.0081 1.0000
5.250 0.8582 0.03715 0.02685 -0.0937 0.0079 1.0000
5.500 0.8792 0.04058 0.03051 -0.0934 0.0077 1.0000
5.750 0.8994 0.04559 0.03569 -0.0935 0.0075 1.0000
6.000 0.9207 0.05426 0.04424 -0.0945 0.0074 1.0000
8.250 1.0546 0.06865 0.06193 -0.0844 0.0070 1.0000
8.500 1.0670 0.06974 0.06338 -0.0826 0.0069 1.0000
8.750 1.0780 0.07178 0.06574 -0.0810 0.0067 1.0000
9.000 1.0864 0.07457 0.06881 -0.0795 0.0064 1.0000
9.250 1.0922 0.07780 0.07230 -0.0782 0.0062 1.0000
9.500 1.0954 0.08128 0.07602 -0.0768 0.0061 1.0000
9.750 1.0962 0.08489 0.07986 -0.0755 0.0059 1.0000
10.000 1.0941 0.08859 0.08377 -0.0742 0.0058 1.0000
10.250 1.0893 0.09231 0.08767 -0.0730 0.0058 1.0000
10.500 1.0814 0.09604 0.09155 -0.0719 0.0058 1.0000
10.750 1.0689 0.09949 0.09514 -0.0705 0.0058 1.0000
11.000 1.0550 0.10324 0.09898 -0.0699 0.0059 1.0000
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Polar data table (+)
Polar graphs
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