AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 50,000 Max Cl/Cd: 35.89 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg12a-il-50000-n5.txt Download as CSV file: xf-rg12a-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL PROFILE12A 9.00%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4943 0.09080 0.08379 -0.0332 1.0000 0.0408
-8.750 -0.4943 0.08696 0.08001 -0.0340 1.0000 0.0403
-8.500 -0.4986 0.08268 0.07580 -0.0358 1.0000 0.0398
-8.250 -0.5056 0.07870 0.07189 -0.0371 1.0000 0.0393
-8.000 -0.5116 0.07458 0.06780 -0.0382 1.0000 0.0389
-7.750 -0.5175 0.07035 0.06356 -0.0391 1.0000 0.0384
-7.500 -0.5219 0.06619 0.05934 -0.0396 1.0000 0.0380
-7.250 -0.5246 0.06197 0.05501 -0.0398 1.0000 0.0376
-7.000 -0.5247 0.05779 0.05064 -0.0396 1.0000 0.0372
-6.750 -0.5218 0.05370 0.04631 -0.0393 1.0000 0.0369
-6.500 -0.5158 0.04965 0.04188 -0.0386 1.0000 0.0368
-6.250 -0.5065 0.04584 0.03769 -0.0378 1.0000 0.0368
-6.000 -0.4941 0.04223 0.03363 -0.0368 1.0000 0.0371
-5.750 -0.4788 0.03898 0.02987 -0.0358 1.0000 0.0387
-5.500 -0.4606 0.03609 0.02633 -0.0346 1.0000 0.0407
-5.250 -0.4422 0.03342 0.02339 -0.0336 1.0000 0.0431
-5.000 -0.4217 0.03123 0.02093 -0.0325 1.0000 0.0452
-4.750 -0.3997 0.02917 0.01855 -0.0313 1.0000 0.0479
-4.500 -0.3767 0.02750 0.01645 -0.0299 1.0000 0.0533
-4.250 -0.3558 0.02601 0.01495 -0.0287 1.0000 0.0609
-4.000 -0.3336 0.02453 0.01329 -0.0271 1.0000 0.0687
-3.750 -0.3124 0.02335 0.01211 -0.0259 1.0000 0.0864
-3.500 -0.2908 0.02210 0.01091 -0.0249 1.0000 0.1183
-3.250 -0.2705 0.02076 0.00993 -0.0240 1.0000 0.1791
-3.000 -0.2515 0.01916 0.00912 -0.0233 1.0000 0.3092
-2.750 -0.2415 0.01765 0.00904 -0.0194 1.0000 0.5694
-2.500 -0.2353 0.01756 0.00939 -0.0130 1.0000 0.7678
-2.250 -0.2228 0.01760 0.00938 -0.0083 1.0000 0.8578
-2.000 -0.1679 0.01770 0.00915 -0.0115 1.0000 0.9511
-1.750 -0.0934 0.01769 0.00859 -0.0207 1.0000 1.0000
-1.500 -0.0901 0.01759 0.00834 -0.0174 1.0000 1.0000
-1.250 -0.0834 0.01755 0.00814 -0.0147 1.0000 1.0000
-1.000 -0.0723 0.01759 0.00801 -0.0129 1.0000 1.0000
-0.750 -0.0438 0.01781 0.00802 -0.0142 0.9946 1.0000
-0.500 0.0000 0.01813 0.00809 -0.0182 0.9829 1.0000
-0.250 0.0426 0.01839 0.00818 -0.0219 0.9704 1.0000
0.000 0.0838 0.01860 0.00824 -0.0252 0.9573 1.0000
0.250 0.1240 0.01877 0.00828 -0.0282 0.9436 1.0000
0.500 0.1636 0.01890 0.00833 -0.0309 0.9294 1.0000
0.750 0.2025 0.01899 0.00838 -0.0333 0.9148 1.0000
1.000 0.2410 0.01905 0.00841 -0.0356 0.8998 1.0000
1.250 0.2796 0.01908 0.00843 -0.0377 0.8845 1.0000
1.500 0.3180 0.01908 0.00846 -0.0397 0.8687 1.0000
1.750 0.3547 0.01907 0.00849 -0.0412 0.8521 1.0000
2.000 0.3914 0.01903 0.00850 -0.0426 0.8351 1.0000
2.250 0.4283 0.01897 0.00849 -0.0439 0.8178 1.0000
2.500 0.4585 0.01901 0.00861 -0.0440 0.7966 1.0000
2.750 0.4925 0.01897 0.00863 -0.0445 0.7765 1.0000
3.000 0.5218 0.01903 0.00874 -0.0443 0.7536 1.0000
3.250 0.5528 0.01907 0.00883 -0.0443 0.7307 1.0000
3.500 0.5818 0.01917 0.00902 -0.0439 0.7061 1.0000
3.750 0.6089 0.01934 0.00924 -0.0432 0.6798 1.0000
4.000 0.6354 0.01954 0.00948 -0.0424 0.6526 1.0000
4.250 0.6610 0.01978 0.00977 -0.0415 0.6242 1.0000
4.500 0.6858 0.02007 0.01013 -0.0404 0.5945 1.0000
4.750 0.7099 0.02039 0.01048 -0.0393 0.5639 1.0000
5.000 0.7327 0.02076 0.01089 -0.0380 0.5313 1.0000
5.250 0.7547 0.02118 0.01134 -0.0366 0.4970 1.0000
5.500 0.7762 0.02165 0.01185 -0.0351 0.4618 1.0000
5.750 0.7965 0.02219 0.01240 -0.0336 0.4243 1.0000
6.000 0.8160 0.02280 0.01301 -0.0320 0.3854 1.0000
6.250 0.8345 0.02351 0.01369 -0.0304 0.3450 1.0000
6.500 0.8519 0.02434 0.01447 -0.0288 0.3031 1.0000
6.750 0.8683 0.02531 0.01536 -0.0271 0.2603 1.0000
7.000 0.8835 0.02647 0.01647 -0.0255 0.2176 1.0000
7.250 0.8977 0.02784 0.01771 -0.0238 0.1765 1.0000
7.500 0.9105 0.02944 0.01912 -0.0222 0.1395 1.0000
7.750 0.9231 0.03117 0.02074 -0.0206 0.1084 1.0000
8.000 0.9351 0.03304 0.02249 -0.0190 0.0887 1.0000
8.250 0.9487 0.03494 0.02446 -0.0174 0.0742 1.0000
8.500 0.9635 0.03697 0.02659 -0.0159 0.0660 1.0000
8.750 0.9819 0.03925 0.02917 -0.0146 0.0599 1.0000
9.000 0.9966 0.04128 0.03133 -0.0134 0.0544 1.0000
9.250 1.0122 0.04375 0.03403 -0.0123 0.0501 1.0000
9.500 1.0276 0.04660 0.03730 -0.0111 0.0472 1.0000
9.750 1.0393 0.04964 0.04069 -0.0097 0.0453 1.0000
10.000 1.0469 0.05277 0.04414 -0.0083 0.0439 1.0000
10.250 1.0505 0.05605 0.04769 -0.0067 0.0428 1.0000
10.500 1.0496 0.05937 0.05125 -0.0050 0.0419 1.0000
10.750 1.0454 0.06306 0.05508 -0.0035 0.0411 1.0000
11.000 1.0313 0.06677 0.05912 -0.0020 0.0408 1.0000
11.250 1.0145 0.07094 0.06358 -0.0015 0.0407 1.0000
11.500 0.9965 0.07568 0.06856 -0.0023 0.0406 1.0000
11.750 0.9787 0.08099 0.07406 -0.0042 0.0408 1.0000
12.000 0.9589 0.08718 0.08041 -0.0075 0.0409 1.0000
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Polar data table (+)
Polar graphs
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