USA 25 AIRFOIL (usa25-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: USA 25 AIRFOIL (usa25-il) Reynolds number: 50,000 Max Cl/Cd: 35.86 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa25-il-50000-n5.txt Download as CSV file: xf-usa25-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 25 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3379 0.10816 0.10164 -0.0256 1.0000 0.1038
-7.750 -0.3518 0.10718 0.10079 -0.0248 1.0000 0.1042
-7.500 -0.3604 0.10566 0.09936 -0.0251 1.0000 0.1044
-7.250 -0.3669 0.10402 0.09779 -0.0256 1.0000 0.1047
-7.000 -0.3607 0.09936 0.09320 -0.0235 1.0000 0.1054
-6.750 -0.3548 0.09538 0.08925 -0.0212 1.0000 0.1068
-6.500 -0.3528 0.09249 0.08641 -0.0195 1.0000 0.1089
-6.250 -0.3532 0.09011 0.08408 -0.0184 1.0000 0.1121
-6.000 -0.3554 0.08823 0.08223 -0.0185 1.0000 0.1160
-5.750 -0.3574 0.08738 0.08134 -0.0212 1.0000 0.1186
-5.250 -0.3503 0.08078 0.07481 -0.0190 1.0000 0.1206
-5.000 -0.3455 0.07782 0.07188 -0.0177 1.0000 0.1220
-4.750 -0.3304 0.07461 0.06863 -0.0191 0.9971 0.1235
-4.500 -0.2798 0.06857 0.06195 -0.0314 0.9871 0.0899
-4.250 -0.2532 0.06436 0.05771 -0.0337 0.9789 0.0885
-4.000 -0.2222 0.06052 0.05371 -0.0372 0.9697 0.0868
-3.750 -0.1883 0.05662 0.04961 -0.0411 0.9605 0.0827
-3.250 -0.1148 0.04830 0.04037 -0.0488 0.9388 0.0750
-3.000 -0.0827 0.04532 0.03712 -0.0508 0.9273 0.0746
-2.750 -0.0497 0.04282 0.03446 -0.0529 0.9156 0.0770
-2.500 -0.0152 0.04077 0.03222 -0.0548 0.9034 0.0795
-2.250 0.0179 0.03824 0.02937 -0.0561 0.8895 0.0800
-2.000 0.0518 0.03572 0.02647 -0.0572 0.8754 0.0803
-1.750 0.0867 0.03340 0.02378 -0.0583 0.8608 0.0813
-1.500 0.1227 0.03120 0.02120 -0.0593 0.8458 0.0831
-1.250 0.1601 0.02907 0.01862 -0.0603 0.8301 0.0860
-1.000 0.1969 0.02740 0.01659 -0.0613 0.8116 0.0919
-0.750 0.2366 0.02636 0.01537 -0.0629 0.7911 0.1015
-0.500 0.2846 0.02514 0.01383 -0.0657 0.7705 0.1152
-0.250 0.3269 0.02426 0.01271 -0.0675 0.7458 0.1365
0.000 0.3646 0.02400 0.01229 -0.0689 0.7213 0.1679
0.250 0.3980 0.02384 0.01199 -0.0697 0.6965 0.1989
0.500 0.4344 0.02346 0.01136 -0.0708 0.6752 0.2180
0.750 0.4667 0.02316 0.01082 -0.0711 0.6557 0.2277
1.000 0.4967 0.02300 0.01045 -0.0712 0.6394 0.2398
1.250 0.5246 0.02288 0.01020 -0.0710 0.6256 0.2551
1.500 0.5518 0.02278 0.00998 -0.0708 0.6137 0.2708
1.750 0.5771 0.02273 0.00988 -0.0702 0.6023 0.2888
2.000 0.6016 0.02264 0.00983 -0.0697 0.5924 0.3177
2.250 0.6234 0.02193 0.00973 -0.0688 0.5844 0.4517
2.750 0.7402 0.02221 0.01034 -0.0807 0.5645 1.0000
3.000 0.7651 0.02262 0.01063 -0.0802 0.5571 1.0000
3.250 0.7906 0.02304 0.01094 -0.0797 0.5505 1.0000
3.500 0.8136 0.02351 0.01138 -0.0789 0.5435 1.0000
4.000 0.8606 0.02448 0.01229 -0.0775 0.5308 1.0000
4.250 0.8850 0.02497 0.01277 -0.0769 0.5247 1.0000
4.500 0.9070 0.02553 0.01335 -0.0759 0.5184 1.0000
4.750 0.9284 0.02609 0.01396 -0.0749 0.5116 1.0000
5.000 0.9527 0.02662 0.01446 -0.0742 0.5052 1.0000
5.250 0.9707 0.02722 0.01517 -0.0726 0.4966 1.0000
5.500 0.9929 0.02772 0.01568 -0.0715 0.4882 1.0000
5.750 1.0123 0.02826 0.01628 -0.0700 0.4788 1.0000
6.000 1.0311 0.02885 0.01697 -0.0685 0.4700 1.0000
6.250 1.0532 0.02937 0.01753 -0.0674 0.4617 1.0000
6.500 1.0699 0.03008 0.01839 -0.0656 0.4535 1.0000
6.750 1.0921 0.03066 0.01905 -0.0646 0.4461 1.0000
7.000 1.1079 0.03144 0.02004 -0.0628 0.4379 1.0000
7.250 1.1295 0.03203 0.02074 -0.0617 0.4302 1.0000
7.500 1.1436 0.03288 0.02180 -0.0596 0.4216 1.0000
7.750 1.1661 0.03343 0.02244 -0.0586 0.4136 1.0000
8.000 1.1765 0.03441 0.02368 -0.0561 0.4044 1.0000
8.250 1.2013 0.03488 0.02427 -0.0554 0.3967 1.0000
8.500 1.2068 0.03605 0.02576 -0.0523 0.3870 1.0000
8.750 1.2228 0.03668 0.02655 -0.0504 0.3774 1.0000
9.000 1.2369 0.03676 0.02674 -0.0478 0.3641 1.0000
9.250 1.2437 0.03687 0.02697 -0.0443 0.3491 1.0000
9.500 1.2498 0.03709 0.02739 -0.0408 0.3352 1.0000
9.750 1.2504 0.03804 0.02859 -0.0371 0.3233 1.0000
10.000 1.2520 0.03895 0.02971 -0.0335 0.3123 1.0000
10.250 1.2520 0.03963 0.03053 -0.0296 0.3003 1.0000
10.500 1.2502 0.04039 0.03139 -0.0258 0.2868 1.0000
10.750 1.2460 0.04140 0.03249 -0.0221 0.2716 1.0000
11.000 1.2446 0.04264 0.03381 -0.0192 0.2587 1.0000
11.250 1.2423 0.04414 0.03539 -0.0166 0.2458 1.0000
11.500 1.2381 0.04600 0.03730 -0.0144 0.2325 1.0000
11.750 1.2305 0.04839 0.03972 -0.0126 0.2181 1.0000
12.000 1.2202 0.05137 0.04276 -0.0113 0.2033 1.0000
12.250 1.2077 0.05495 0.04636 -0.0107 0.1882 1.0000
12.500 1.1953 0.05890 0.05032 -0.0107 0.1744 1.0000
12.750 1.1825 0.06329 0.05476 -0.0112 0.1606 1.0000
13.000 1.1696 0.06800 0.05954 -0.0121 0.1466 1.0000
13.250 1.1562 0.07309 0.06477 -0.0135 0.1307 1.0000
13.500 1.1431 0.07824 0.06998 -0.0148 0.1099 1.0000
13.750 1.1310 0.08308 0.07471 -0.0160 0.0919 1.0000
14.000 1.1178 0.08818 0.07966 -0.0174 0.0796 1.0000
14.250 1.1053 0.09340 0.08479 -0.0189 0.0708 1.0000
14.500 1.0946 0.09852 0.08990 -0.0204 0.0635 1.0000
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