GOE 359 AIRFOIL (goe359-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 359 AIRFOIL (goe359-il) Reynolds number: 500,000 Max Cl/Cd: 110.81 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe359-il-500000.txt Download as CSV file: xf-goe359-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2845 0.10996 0.10773 -0.0248 1.0000 0.0152
-9.000 -0.2815 0.10769 0.10549 -0.0267 1.0000 0.0154
-8.750 -0.2802 0.10540 0.10324 -0.0281 1.0000 0.0155
-8.500 -0.2000 0.08920 0.08729 -0.0308 1.0000 0.0159
-8.250 -0.1950 0.08609 0.08420 -0.0305 0.9995 0.0161
-8.000 -0.1799 0.08192 0.08003 -0.0336 0.9965 0.0164
-7.750 -0.1653 0.07780 0.07591 -0.0371 0.9917 0.0168
-7.500 -0.2775 0.09154 0.08958 -0.0257 1.0000 0.0159
-7.250 -0.2618 0.08801 0.08606 -0.0283 0.9971 0.0162
-7.000 -0.2369 0.08392 0.08196 -0.0341 0.9907 0.0166
-6.750 -0.2119 0.07988 0.07792 -0.0403 0.9799 0.0172
-6.500 -0.1863 0.07574 0.07377 -0.0468 0.9624 0.0180
-6.250 -0.1399 0.07029 0.06825 -0.0614 0.9426 0.0201
-6.000 -0.0901 0.06437 0.06221 -0.0757 0.9231 0.0204
-5.750 -0.0535 0.05949 0.05716 -0.0839 0.9013 0.0204
-5.500 -0.0419 0.05478 0.05236 -0.0852 0.8807 0.0211
-5.250 -0.0224 0.05230 0.04976 -0.0866 0.8629 0.0216
-5.000 0.0014 0.04952 0.04687 -0.0894 0.8472 0.0224
-4.750 0.0289 0.04643 0.04365 -0.0930 0.8330 0.0238
-4.500 0.0769 0.04253 0.03947 -0.1004 0.8204 0.0268
-4.250 0.1092 0.03868 0.03540 -0.1037 0.8084 0.0269
-4.000 0.1325 0.03268 0.02918 -0.1071 0.7972 0.0280
-3.500 0.1816 0.02930 0.02554 -0.1085 0.7721 0.0307
-3.250 0.2182 0.02743 0.02330 -0.1095 0.7598 0.0358
-3.000 0.2467 0.02208 0.01740 -0.1113 0.7488 0.0372
-2.750 0.2715 0.02066 0.01593 -0.1116 0.7364 0.0385
-2.500 0.2976 0.01957 0.01468 -0.1117 0.7240 0.0403
-2.250 0.3253 0.01840 0.01324 -0.1117 0.7114 0.0442
-2.000 0.3538 0.01657 0.01097 -0.1117 0.6984 0.0495
-1.750 0.3800 0.01573 0.01004 -0.1117 0.6842 0.0520
-1.500 0.4074 0.01512 0.00920 -0.1114 0.6702 0.0567
-1.250 0.4343 0.01417 0.00805 -0.1114 0.6575 0.0641
-1.000 0.4617 0.01391 0.00764 -0.1112 0.6453 0.0705
-0.500 0.5191 0.01193 0.00500 -0.1099 0.6217 0.0481
0.000 0.5734 0.01088 0.00377 -0.1094 0.5980 0.0460
0.250 0.6006 0.01058 0.00339 -0.1091 0.5863 0.0460
0.500 0.6276 0.01041 0.00316 -0.1088 0.5752 0.0466
0.750 0.6546 0.01032 0.00301 -0.1086 0.5646 0.0472
1.000 0.6816 0.01011 0.00276 -0.1083 0.5533 0.0472
1.250 0.7087 0.00994 0.00256 -0.1081 0.5420 0.0477
1.500 0.7357 0.00978 0.00234 -0.1079 0.5299 0.0497
1.750 0.7626 0.00975 0.00225 -0.1076 0.5173 0.0526
2.000 0.7894 0.00978 0.00221 -0.1074 0.5050 0.0559
2.500 0.8432 0.00952 0.00239 -0.1072 0.4836 0.2878
2.750 0.8675 0.00827 0.00255 -0.1067 0.4745 1.0000
3.000 0.8940 0.00843 0.00262 -0.1064 0.4652 1.0000
3.250 0.9209 0.00856 0.00272 -0.1062 0.4567 1.0000
3.500 0.9474 0.00873 0.00283 -0.1059 0.4481 1.0000
3.750 0.9738 0.00889 0.00294 -0.1057 0.4372 1.0000
4.000 1.0002 0.00905 0.00306 -0.1054 0.4257 1.0000
4.250 1.0261 0.00926 0.00320 -0.1051 0.4098 1.0000
4.500 1.0514 0.00950 0.00334 -0.1047 0.3883 1.0000
4.750 1.0768 0.00975 0.00350 -0.1043 0.3662 1.0000
5.000 1.1015 0.01005 0.00369 -0.1039 0.3382 1.0000
5.250 1.1234 0.01066 0.00401 -0.1030 0.2849 1.0000
5.500 1.1439 0.01143 0.00448 -0.1020 0.2380 1.0000
5.750 1.1655 0.01209 0.00492 -0.1012 0.2056 1.0000
6.000 1.1867 0.01277 0.00535 -0.1003 0.1630 1.0000
6.250 1.1981 0.01449 0.00642 -0.0981 0.0633 1.0000
6.500 1.2190 0.01517 0.00702 -0.0971 0.0452 1.0000
6.750 1.2419 0.01562 0.00744 -0.0964 0.0341 1.0000
7.000 1.2617 0.01638 0.00813 -0.0952 0.0177 1.0000
7.250 1.2843 0.01683 0.00866 -0.0944 0.0162 1.0000
7.500 1.3062 0.01732 0.00927 -0.0935 0.0149 1.0000
7.750 1.3270 0.01791 0.00995 -0.0924 0.0143 1.0000
8.000 1.3468 0.01856 0.01071 -0.0912 0.0137 1.0000
8.250 1.3652 0.01931 0.01156 -0.0898 0.0132 1.0000
8.500 1.3816 0.02018 0.01253 -0.0881 0.0128 1.0000
8.750 1.3957 0.02119 0.01366 -0.0861 0.0124 1.0000
9.000 1.4071 0.02231 0.01487 -0.0838 0.0122 1.0000
9.250 1.4131 0.02355 0.01620 -0.0806 0.0119 1.0000
9.500 1.4139 0.02508 0.01782 -0.0770 0.0116 1.0000
9.750 1.4109 0.02696 0.01979 -0.0733 0.0113 1.0000
10.000 1.4034 0.02939 0.02232 -0.0698 0.0110 1.0000
10.250 1.4004 0.03175 0.02476 -0.0673 0.0109 1.0000
10.500 1.4001 0.03406 0.02714 -0.0653 0.0109 1.0000
10.750 1.4021 0.03631 0.02945 -0.0635 0.0109 1.0000
11.000 1.4066 0.03838 0.03156 -0.0618 0.0109 1.0000
11.250 1.4134 0.04031 0.03355 -0.0603 0.0109 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 359 AIRFOIL (goe359-il)