GOE 445 AIRFOIL (goe445-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 445 AIRFOIL (goe445-il) Reynolds number: 1,000,000 Max Cl/Cd: 42.15 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe445-il-1000000-n5.txt Download as CSV file: xf-goe445-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 445 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5284 0.07952 0.07804 -0.0047 1.0000 0.0056
-8.750 -0.5327 0.07427 0.07281 -0.0068 1.0000 0.0056
-8.500 -0.5381 0.06890 0.06745 -0.0093 1.0000 0.0056
-8.250 -0.5456 0.06316 0.06172 -0.0128 1.0000 0.0056
-8.000 -0.5607 0.05688 0.05543 -0.0192 1.0000 0.0056
-7.750 -0.5768 0.05283 0.05132 -0.0198 1.0000 0.0056
-7.500 -0.5851 0.04832 0.04672 -0.0199 1.0000 0.0056
-7.250 -0.5887 0.04364 0.04194 -0.0197 1.0000 0.0056
-7.000 -0.5890 0.03901 0.03717 -0.0191 1.0000 0.0056
-6.750 -0.5865 0.03449 0.03250 -0.0182 1.0000 0.0057
-6.500 -0.5814 0.03021 0.02804 -0.0169 1.0000 0.0057
-6.250 -0.5743 0.02614 0.02378 -0.0153 1.0000 0.0057
-6.000 -0.5656 0.02244 0.01986 -0.0134 1.0000 0.0057
-5.750 -0.5555 0.01909 0.01627 -0.0114 1.0000 0.0057
-5.500 -0.5441 0.01610 0.01303 -0.0092 1.0000 0.0057
-5.250 -0.5316 0.01345 0.01007 -0.0069 1.0000 0.0057
-4.500 -0.5044 0.01773 0.01322 0.0005 0.9981 0.0051
-4.250 -0.4750 0.01496 0.01012 0.0005 0.9959 0.0043
-4.000 -0.4454 0.01260 0.00750 0.0008 0.9938 0.0035
-3.750 -0.4172 0.01074 0.00546 0.0012 0.9907 0.0030
-3.500 -0.3887 0.00935 0.00390 0.0012 0.9875 0.0027
-3.250 -0.3603 0.00857 0.00299 0.0012 0.9843 0.0025
-3.000 -0.3314 0.00808 0.00240 0.0010 0.9805 0.0024
-2.750 -0.3000 0.00771 0.00193 0.0002 0.9775 0.0024
-2.500 -0.2699 0.00748 0.00163 -0.0003 0.9732 0.0024
-2.250 -0.2383 0.00731 0.00134 -0.0012 0.9686 0.0022
-2.000 -0.2057 0.00715 0.00115 -0.0023 0.9639 0.0025
-1.750 -0.1740 0.00704 0.00104 -0.0032 0.9565 0.0040
-1.500 -0.1411 0.00691 0.00096 -0.0045 0.9473 0.0054
-1.250 -0.1133 0.00576 0.00069 -0.0051 0.9234 0.2787
-1.000 -0.0925 0.00484 0.00054 -0.0041 0.8840 0.5291
-0.750 -0.0690 0.00472 0.00049 -0.0032 0.8471 0.5869
-0.500 -0.0460 0.00458 0.00045 -0.0021 0.8160 0.6452
-0.250 -0.0225 0.00450 0.00042 -0.0012 0.7852 0.6916
0.000 0.0001 0.00444 0.00042 0.0000 0.7462 0.7456
0.250 0.0226 0.00450 0.00042 0.0012 0.6917 0.7855
0.500 0.0459 0.00459 0.00045 0.0021 0.6417 0.8166
0.750 0.0691 0.00472 0.00049 0.0031 0.5876 0.8476
1.000 0.0929 0.00482 0.00054 0.0040 0.5351 0.8839
1.250 0.1125 0.00588 0.00071 0.0052 0.2490 0.9234
1.500 0.1412 0.00691 0.00096 0.0044 0.0051 0.9473
1.750 0.1742 0.00704 0.00104 0.0032 0.0039 0.9565
2.000 0.2059 0.00715 0.00115 0.0023 0.0026 0.9638
2.250 0.2385 0.00731 0.00134 0.0011 0.0022 0.9687
2.500 0.2699 0.00748 0.00162 0.0003 0.0024 0.9735
2.750 0.3001 0.00771 0.00193 -0.0002 0.0024 0.9777
3.000 0.3313 0.00808 0.00240 -0.0010 0.0024 0.9807
3.250 0.3600 0.00854 0.00296 -0.0012 0.0025 0.9847
3.500 0.3889 0.00937 0.00392 -0.0013 0.0027 0.9878
3.750 0.4176 0.01075 0.00547 -0.0013 0.0030 0.9909
4.000 0.4456 0.01267 0.00757 -0.0008 0.0035 0.9939
4.250 0.4755 0.01489 0.01005 -0.0007 0.0041 0.9960
4.500 0.5047 0.01772 0.01321 -0.0006 0.0050 0.9981
5.500 0.5728 0.03058 0.02730 0.0068 0.0057 1.0000
5.750 0.5857 0.03362 0.03060 0.0091 0.0057 1.0000
6.000 0.5977 0.03684 0.03406 0.0113 0.0057 1.0000
6.250 0.6084 0.04032 0.03776 0.0132 0.0057 1.0000
6.500 0.6174 0.04405 0.04171 0.0150 0.0057 1.0000
6.750 0.6255 0.04782 0.04565 0.0164 0.0056 1.0000
7.000 0.6315 0.05186 0.04986 0.0175 0.0056 1.0000
7.250 0.6361 0.05600 0.05414 0.0181 0.0056 1.0000
7.500 0.6381 0.06042 0.05868 0.0181 0.0056 1.0000
7.750 0.6389 0.06491 0.06327 0.0174 0.0056 1.0000
8.000 0.6367 0.06971 0.06814 0.0159 0.0056 1.0000
8.250 0.6282 0.07430 0.07278 0.0141 0.0056 1.0000
8.500 0.6229 0.08039 0.07886 0.0071 0.0056 1.0000
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Polar data table (+)
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