GOE 155 (SSW D.1) AIRFOIL (goe155-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 155 (SSW D.1) AIRFOIL (goe155-il) Reynolds number: 500,000 Max Cl/Cd: 98.61 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe155-il-500000-n5.txt Download as CSV file: xf-goe155-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 155 (SSW D.1) 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3446 0.10616 0.10397 -0.0143 0.9368 0.0078
-8.750 -0.3398 0.10298 0.10077 -0.0153 0.9250 0.0078
-8.500 -0.3349 0.09975 0.09752 -0.0165 0.9140 0.0078
-8.250 -0.3300 0.09649 0.09424 -0.0177 0.9029 0.0078
-7.750 -0.3194 0.08977 0.08746 -0.0210 0.8797 0.0081
-7.500 -0.3139 0.08627 0.08394 -0.0233 0.8686 0.0082
-7.250 -0.3044 0.08107 0.07870 -0.0285 0.8585 0.0088
-7.000 -0.2895 0.07890 0.07650 -0.0307 0.8491 0.0091
-6.750 -0.2732 0.07608 0.07363 -0.0339 0.8401 0.0095
-6.500 -0.2557 0.07223 0.06972 -0.0384 0.8317 0.0097
-6.250 -0.2355 0.06811 0.06554 -0.0436 0.8229 0.0100
-6.000 -0.2129 0.06377 0.06112 -0.0491 0.8146 0.0104
-5.750 -0.1869 0.05871 0.05596 -0.0555 0.8058 0.0110
-5.500 -0.1566 0.05301 0.05014 -0.0629 0.7977 0.0121
-5.250 -0.1316 0.05098 0.04802 -0.0654 0.7893 0.0127
-5.000 -0.1020 0.04723 0.04415 -0.0699 0.7818 0.0133
-4.750 -0.0579 0.03761 0.03415 -0.0795 0.7756 0.0156
-4.500 -0.0325 0.03658 0.03305 -0.0806 0.7683 0.0161
-4.250 -0.0055 0.03500 0.03134 -0.0820 0.7605 0.0167
-4.000 0.0253 0.03187 0.02801 -0.0843 0.7534 0.0177
-3.750 0.0650 0.02281 0.01824 -0.0881 0.7478 0.0207
-3.500 0.0922 0.02182 0.01710 -0.0887 0.7419 0.0212
-3.250 0.1203 0.02053 0.01566 -0.0893 0.7362 0.0218
-3.000 0.1489 0.01904 0.01392 -0.0898 0.7307 0.0229
-2.750 0.1784 0.01695 0.01146 -0.0902 0.7260 0.0241
-2.500 0.2077 0.01538 0.00956 -0.0905 0.7207 0.0247
-2.250 0.2365 0.01434 0.00824 -0.0907 0.7146 0.0251
-2.000 0.2651 0.01369 0.00737 -0.0908 0.7072 0.0257
-1.750 0.2936 0.01333 0.00683 -0.0908 0.6990 0.0262
-1.500 0.3223 0.01271 0.00605 -0.0910 0.6914 0.0263
-1.250 0.3507 0.01215 0.00532 -0.0910 0.6837 0.0265
-1.000 0.3792 0.01122 0.00426 -0.0912 0.6754 0.0270
-0.750 0.4073 0.01070 0.00365 -0.0913 0.6669 0.0275
-0.500 0.4356 0.01035 0.00325 -0.0914 0.6560 0.0281
-0.250 0.4637 0.01009 0.00292 -0.0914 0.6423 0.0285
0.000 0.4918 0.00989 0.00266 -0.0915 0.6265 0.0289
0.250 0.5198 0.00973 0.00245 -0.0915 0.6101 0.0294
0.500 0.5479 0.00963 0.00229 -0.0916 0.5941 0.0303
0.750 0.5759 0.00955 0.00214 -0.0917 0.5758 0.0307
1.000 0.6038 0.00952 0.00202 -0.0917 0.5544 0.0308
1.250 0.6315 0.00955 0.00195 -0.0917 0.5316 0.0309
1.500 0.6591 0.00962 0.00192 -0.0918 0.5072 0.0312
1.750 0.6866 0.00973 0.00192 -0.0918 0.4839 0.0316
2.000 0.7140 0.00986 0.00195 -0.0918 0.4629 0.0322
2.250 0.7415 0.00999 0.00199 -0.0919 0.4458 0.0331
2.500 0.7690 0.01011 0.00205 -0.0919 0.4318 0.0341
2.750 0.7966 0.01022 0.00213 -0.0920 0.4194 0.0363
3.000 0.8240 0.01034 0.00223 -0.0920 0.4079 0.0404
3.250 0.8514 0.01044 0.00237 -0.0920 0.3966 0.0724
3.750 0.9056 0.01066 0.00273 -0.0922 0.3673 0.1774
4.250 0.9555 0.00969 0.00320 -0.0916 0.3270 1.0000
4.500 0.9819 0.00996 0.00340 -0.0916 0.3092 1.0000
4.750 1.0080 0.01028 0.00362 -0.0914 0.2887 1.0000
5.000 1.0337 0.01063 0.00388 -0.0913 0.2690 1.0000
5.250 1.0594 0.01097 0.00416 -0.0911 0.2535 1.0000
5.500 1.0848 0.01133 0.00445 -0.0909 0.2379 1.0000
5.750 1.1101 0.01170 0.00475 -0.0907 0.2233 1.0000
6.000 1.1349 0.01211 0.00509 -0.0905 0.2064 1.0000
6.250 1.1599 0.01248 0.00543 -0.0902 0.1897 1.0000
6.500 1.1798 0.01351 0.00604 -0.0895 0.1155 1.0000
6.750 1.1954 0.01508 0.00719 -0.0881 0.0371 1.0000
7.000 1.2184 0.01563 0.00773 -0.0876 0.0294 1.0000
7.250 1.2420 0.01607 0.00824 -0.0871 0.0247 1.0000
7.500 1.2645 0.01662 0.00881 -0.0864 0.0199 1.0000
7.750 1.2869 0.01716 0.00940 -0.0858 0.0169 1.0000
8.000 1.3088 0.01771 0.00997 -0.0851 0.0144 1.0000
8.250 1.3288 0.01846 0.01076 -0.0841 0.0121 1.0000
8.500 1.3492 0.01910 0.01149 -0.0832 0.0113 1.0000
8.750 1.3689 0.01977 0.01224 -0.0822 0.0103 1.0000
9.000 1.3877 0.02047 0.01300 -0.0812 0.0095 1.0000
9.250 1.4045 0.02129 0.01386 -0.0798 0.0086 1.0000
9.500 1.4183 0.02231 0.01498 -0.0781 0.0079 1.0000
9.750 1.4323 0.02319 0.01596 -0.0764 0.0075 1.0000
10.000 1.4421 0.02416 0.01704 -0.0740 0.0072 1.0000
10.250 1.4505 0.02525 0.01824 -0.0716 0.0069 1.0000
10.500 1.4586 0.02645 0.01953 -0.0695 0.0066 1.0000
10.750 1.4660 0.02777 0.02094 -0.0675 0.0063 1.0000
11.000 1.4725 0.02924 0.02250 -0.0657 0.0061 1.0000
11.250 1.4769 0.03096 0.02431 -0.0639 0.0059 1.0000
11.500 1.4774 0.03313 0.02658 -0.0621 0.0056 1.0000
11.750 1.4756 0.03564 0.02922 -0.0604 0.0054 1.0000
12.000 1.4797 0.03769 0.03141 -0.0592 0.0053 1.0000
12.250 1.4827 0.03993 0.03378 -0.0581 0.0051 1.0000
12.500 1.4844 0.04235 0.03633 -0.0572 0.0049 1.0000
12.750 1.4850 0.04498 0.03909 -0.0564 0.0048 1.0000
13.000 1.4846 0.04778 0.04201 -0.0557 0.0046 1.0000
13.250 1.4827 0.05082 0.04518 -0.0552 0.0045 1.0000
13.500 1.4804 0.05400 0.04848 -0.0548 0.0044 1.0000
13.750 1.4773 0.05734 0.05195 -0.0545 0.0043 1.0000
14.000 1.4737 0.06087 0.05560 -0.0545 0.0042 1.0000
14.250 1.4700 0.06455 0.05941 -0.0547 0.0042 1.0000
14.500 1.4659 0.06841 0.06339 -0.0551 0.0041 1.0000
14.750 1.4612 0.07246 0.06758 -0.0557 0.0040 1.0000
15.000 1.4559 0.07674 0.07198 -0.0565 0.0040 1.0000
15.250 1.4504 0.08118 0.07654 -0.0576 0.0039 1.0000
15.500 1.4437 0.08586 0.08135 -0.0587 0.0038 1.0000
15.750 1.4361 0.09077 0.08638 -0.0600 0.0038 1.0000
16.000 1.4279 0.09587 0.09161 -0.0615 0.0037 1.0000
16.250 1.4184 0.10126 0.09714 -0.0632 0.0037 1.0000
16.500 1.4079 0.10687 0.10289 -0.0651 0.0037 1.0000
16.750 1.3964 0.11278 0.10894 -0.0671 0.0036 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 155 (SSW D.1) AIRFOIL (goe155-il)