GOE 210 (DAIMLER) AIRFOIL (goe210-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 210 (DAIMLER) AIRFOIL (goe210-il) Reynolds number: 500,000 Max Cl/Cd: 86.84 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe210-il-500000-n5.txt Download as CSV file: xf-goe210-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 210 (DAIMLER) 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 -0.2508 0.09074 0.08866 -0.0284 1.0000 0.0086
-8.750 -0.2493 0.08766 0.08560 -0.0286 1.0000 0.0089
-8.500 -0.2487 0.08457 0.08254 -0.0286 1.0000 0.0090
-8.250 -0.2431 0.08074 0.07872 -0.0303 0.9978 0.0091
-8.000 -0.2350 0.07649 0.07447 -0.0329 0.9945 0.0091
-7.750 -0.3391 0.08952 0.08740 -0.0253 1.0000 0.0074
-7.500 -0.3356 0.08694 0.08484 -0.0249 0.9987 0.0071
-7.250 -0.3175 0.08269 0.08060 -0.0305 0.9928 0.0069
-7.000 -0.2959 0.07802 0.07592 -0.0373 0.9869 0.0068
-6.000 -0.1921 0.05726 0.05502 -0.0678 0.9485 0.0072
-5.750 -0.1646 0.05237 0.05002 -0.0740 0.9346 0.0071
-5.500 -0.1353 0.04711 0.04464 -0.0800 0.9174 0.0071
-5.250 -0.1038 0.04112 0.03843 -0.0858 0.8974 0.0072
-5.000 -0.0710 0.03410 0.03106 -0.0910 0.8765 0.0075
-4.750 -0.0425 0.03030 0.02696 -0.0932 0.8575 0.0079
-4.500 -0.0148 0.02765 0.02404 -0.0944 0.8420 0.0085
-4.250 0.0187 0.01395 0.00877 -0.0970 0.8314 0.0107
-4.000 0.0458 0.01267 0.00714 -0.0969 0.8200 0.0118
-3.750 0.0732 0.01261 0.00698 -0.0967 0.8100 0.0129
-3.500 0.1007 0.01209 0.00625 -0.0965 0.8009 0.0150
-3.250 0.1283 0.01140 0.00535 -0.0964 0.7929 0.0166
-3.000 0.1560 0.01139 0.00530 -0.0964 0.7850 0.0179
-2.750 0.1837 0.01131 0.00513 -0.0963 0.7767 0.0203
-2.500 0.2113 0.01105 0.00471 -0.0961 0.7678 0.0226
-2.250 0.2390 0.01095 0.00451 -0.0959 0.7582 0.0239
-2.000 0.2663 0.01025 0.00369 -0.0958 0.7498 0.0260
-1.750 0.2938 0.00993 0.00330 -0.0957 0.7413 0.0275
-1.500 0.3212 0.00971 0.00302 -0.0956 0.7305 0.0295
-1.250 0.3484 0.00950 0.00271 -0.0954 0.7159 0.0310
-1.000 0.3756 0.00931 0.00243 -0.0951 0.7007 0.0318
-0.750 0.4032 0.00915 0.00218 -0.0950 0.6891 0.0324
-0.500 0.4308 0.00901 0.00199 -0.0949 0.6799 0.0329
-0.250 0.4585 0.00891 0.00185 -0.0948 0.6689 0.0335
0.000 0.4860 0.00884 0.00171 -0.0947 0.6559 0.0339
0.250 0.5136 0.00877 0.00158 -0.0946 0.6418 0.0339
0.500 0.5410 0.00874 0.00148 -0.0944 0.6265 0.0340
0.750 0.5685 0.00872 0.00140 -0.0943 0.6118 0.0343
1.000 0.5956 0.00875 0.00134 -0.0941 0.5937 0.0351
1.250 0.6226 0.00881 0.00131 -0.0939 0.5692 0.0373
1.500 0.6479 0.00904 0.00134 -0.0934 0.5188 0.0439
1.750 0.6710 0.00951 0.00152 -0.0927 0.4299 0.1018
2.250 0.7160 0.00856 0.00206 -0.0914 0.3194 1.0000
2.500 0.7417 0.00890 0.00222 -0.0911 0.2870 1.0000
2.750 0.7675 0.00924 0.00238 -0.0909 0.2559 1.0000
3.000 0.7932 0.00958 0.00257 -0.0906 0.2273 1.0000
3.250 0.8190 0.00989 0.00275 -0.0903 0.2049 1.0000
3.500 0.8449 0.01017 0.00294 -0.0901 0.1901 1.0000
3.750 0.8711 0.01041 0.00313 -0.0898 0.1796 1.0000
4.000 0.8972 0.01066 0.00334 -0.0896 0.1710 1.0000
4.250 0.9234 0.01089 0.00354 -0.0894 0.1639 1.0000
4.500 0.9493 0.01114 0.00376 -0.0891 0.1577 1.0000
4.750 0.9756 0.01135 0.00399 -0.0889 0.1518 1.0000
5.000 1.0012 0.01163 0.00423 -0.0886 0.1442 1.0000
5.250 1.0273 0.01183 0.00446 -0.0884 0.1380 1.0000
5.500 1.0524 0.01215 0.00473 -0.0880 0.1222 1.0000
5.750 1.0727 0.01309 0.00530 -0.0871 0.0651 1.0000
6.000 1.0969 0.01351 0.00570 -0.0866 0.0585 1.0000
6.250 1.1214 0.01389 0.00608 -0.0861 0.0530 1.0000
6.500 1.1464 0.01417 0.00641 -0.0858 0.0500 1.0000
6.750 1.1702 0.01459 0.00684 -0.0853 0.0437 1.0000
7.000 1.1938 0.01504 0.00724 -0.0847 0.0322 1.0000
7.250 1.2144 0.01587 0.00795 -0.0837 0.0126 1.0000
7.500 1.2376 0.01636 0.00852 -0.0830 0.0104 1.0000
7.750 1.2599 0.01693 0.00920 -0.0822 0.0090 1.0000
8.000 1.2805 0.01770 0.01010 -0.0812 0.0075 1.0000
8.250 1.3019 0.01832 0.01082 -0.0803 0.0068 1.0000
8.500 1.3229 0.01898 0.01158 -0.0793 0.0062 1.0000
8.750 1.3431 0.01970 0.01240 -0.0782 0.0057 1.0000
9.000 1.3621 0.02052 0.01331 -0.0771 0.0053 1.0000
9.250 1.3779 0.02165 0.01455 -0.0754 0.0049 1.0000
9.500 1.3946 0.02259 0.01560 -0.0739 0.0046 1.0000
9.750 1.4116 0.02343 0.01657 -0.0725 0.0042 1.0000
10.000 1.4265 0.02441 0.01766 -0.0708 0.0039 1.0000
10.250 1.4391 0.02551 0.01888 -0.0689 0.0037 1.0000
10.500 1.4484 0.02663 0.02011 -0.0664 0.0036 1.0000
10.750 1.4559 0.02784 0.02141 -0.0638 0.0034 1.0000
11.000 1.4611 0.02923 0.02291 -0.0611 0.0033 1.0000
11.250 1.4634 0.03090 0.02471 -0.0584 0.0032 1.0000
11.500 1.4612 0.03304 0.02699 -0.0555 0.0031 1.0000
11.750 1.4597 0.03528 0.02938 -0.0532 0.0030 1.0000
12.000 1.4602 0.03748 0.03174 -0.0513 0.0030 1.0000
12.250 1.4596 0.03994 0.03436 -0.0498 0.0029 1.0000
12.500 1.4570 0.04275 0.03733 -0.0486 0.0029 1.0000
12.750 1.4534 0.04584 0.04059 -0.0477 0.0028 1.0000
13.000 1.4480 0.04927 0.04423 -0.0471 0.0028 1.0000
13.250 1.4413 0.05303 0.04816 -0.0468 0.0027 1.0000
13.500 1.4331 0.05714 0.05244 -0.0469 0.0027 1.0000
13.750 1.4237 0.06160 0.05707 -0.0474 0.0026 1.0000
14.000 1.4131 0.06645 0.06209 -0.0484 0.0026 1.0000
14.250 1.4013 0.07171 0.06753 -0.0498 0.0026 1.0000
14.500 1.3888 0.07744 0.07343 -0.0517 0.0026 1.0000
14.750 1.3757 0.08359 0.07975 -0.0541 0.0025 1.0000
15.000 1.3616 0.09021 0.08652 -0.0570 0.0025 1.0000
15.250 1.3475 0.09717 0.09365 -0.0602 0.0025 1.0000
15.500 1.3327 0.10449 0.10111 -0.0637 0.0025 1.0000
15.750 1.3173 0.11207 0.10885 -0.0674 0.0025 1.0000
16.000 1.3026 0.11968 0.11659 -0.0713 0.0025 1.0000
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Polar data table (+)
Polar graphs
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