USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 50,000 Max Cl/Cd: 35.22 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa51-il-50000-n5.txt Download as CSV file: xf-usa51-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 51 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4361 0.10069 0.09385 -0.0273 1.0000 0.0610
-8.750 -0.4282 0.09691 0.09010 -0.0263 1.0000 0.0579
-8.500 -0.4296 0.09293 0.08619 -0.0274 1.0000 0.0552
-8.250 -0.4354 0.08881 0.08218 -0.0294 1.0000 0.0529
-8.000 -0.4573 0.08281 0.07627 -0.0364 1.0000 0.0487
-7.750 -0.4551 0.07932 0.07284 -0.0357 1.0000 0.0480
-7.500 -0.4559 0.07558 0.06912 -0.0359 1.0000 0.0473
-7.250 -0.4572 0.07170 0.06525 -0.0362 1.0000 0.0467
-7.000 -0.4579 0.06777 0.06128 -0.0365 1.0000 0.0460
-6.750 -0.4577 0.06370 0.05714 -0.0365 1.0000 0.0455
-6.500 -0.4558 0.05970 0.05302 -0.0363 1.0000 0.0450
-6.250 -0.4524 0.05562 0.04874 -0.0359 1.0000 0.0449
-6.000 -0.4468 0.05169 0.04455 -0.0352 1.0000 0.0453
-5.750 -0.4390 0.04787 0.04036 -0.0343 1.0000 0.0463
-5.500 -0.4286 0.04425 0.03628 -0.0330 1.0000 0.0473
-5.250 -0.4157 0.04086 0.03238 -0.0316 1.0000 0.0482
-5.000 -0.4005 0.03776 0.02875 -0.0300 1.0000 0.0487
-4.750 -0.3842 0.03481 0.02545 -0.0286 1.0000 0.0498
-4.500 -0.3669 0.03295 0.02346 -0.0274 1.0000 0.0532
-4.250 -0.3475 0.03109 0.02120 -0.0260 1.0000 0.0577
-4.000 -0.3258 0.02903 0.01865 -0.0246 1.0000 0.0605
-3.750 -0.3044 0.02732 0.01673 -0.0234 1.0000 0.0644
-3.500 -0.2835 0.02623 0.01546 -0.0222 1.0000 0.0727
-3.250 -0.2610 0.02492 0.01399 -0.0209 1.0000 0.0782
-3.000 -0.2389 0.02400 0.01291 -0.0197 1.0000 0.0874
-2.750 -0.2164 0.02323 0.01204 -0.0189 1.0000 0.1012
-2.500 -0.1950 0.02241 0.01116 -0.0178 1.0000 0.1145
-2.250 -0.1725 0.02160 0.01042 -0.0171 0.9993 0.1416
-2.000 -0.1326 0.02022 0.00974 -0.0204 0.9890 0.2771
-1.750 -0.0621 0.01789 0.00968 -0.0264 0.9906 1.0000
-1.500 -0.0158 0.01809 0.00938 -0.0305 0.9773 1.0000
-1.250 0.0278 0.01821 0.00909 -0.0340 0.9628 1.0000
-1.000 0.0692 0.01829 0.00887 -0.0370 0.9469 1.0000
-0.750 0.1097 0.01833 0.00863 -0.0397 0.9304 1.0000
-0.500 0.1500 0.01833 0.00841 -0.0422 0.9131 1.0000
-0.250 0.1912 0.01827 0.00816 -0.0447 0.8960 1.0000
0.000 0.2278 0.01819 0.00795 -0.0463 0.8759 1.0000
0.250 0.2661 0.01809 0.00772 -0.0480 0.8563 1.0000
0.500 0.3026 0.01800 0.00751 -0.0493 0.8349 1.0000
0.750 0.3402 0.01791 0.00730 -0.0506 0.8132 1.0000
1.000 0.3739 0.01789 0.00717 -0.0513 0.7889 1.0000
1.250 0.4068 0.01791 0.00707 -0.0517 0.7643 1.0000
1.500 0.4374 0.01799 0.00702 -0.0517 0.7396 1.0000
1.750 0.4642 0.01817 0.00710 -0.0511 0.7138 1.0000
2.000 0.4902 0.01838 0.00721 -0.0504 0.6894 1.0000
2.250 0.5159 0.01863 0.00736 -0.0496 0.6673 1.0000
2.500 0.5401 0.01893 0.00761 -0.0487 0.6457 1.0000
2.750 0.5652 0.01924 0.00786 -0.0480 0.6263 1.0000
3.000 0.5904 0.01955 0.00811 -0.0473 0.6083 1.0000
3.250 0.6139 0.01988 0.00841 -0.0463 0.5882 1.0000
3.500 0.6375 0.02020 0.00870 -0.0453 0.5684 1.0000
3.750 0.6606 0.02052 0.00895 -0.0441 0.5482 1.0000
4.000 0.6826 0.02085 0.00925 -0.0428 0.5262 1.0000
4.250 0.7048 0.02119 0.00955 -0.0415 0.5048 1.0000
4.500 0.7259 0.02156 0.00991 -0.0401 0.4821 1.0000
4.750 0.7472 0.02196 0.01026 -0.0388 0.4605 1.0000
5.000 0.7688 0.02239 0.01075 -0.0376 0.4410 1.0000
5.250 0.7911 0.02284 0.01133 -0.0366 0.4241 1.0000
5.500 0.8130 0.02331 0.01191 -0.0355 0.4070 1.0000
5.750 0.8344 0.02380 0.01249 -0.0343 0.3891 1.0000
6.000 0.8543 0.02430 0.01319 -0.0330 0.3678 1.0000
6.250 0.8738 0.02481 0.01385 -0.0315 0.3457 1.0000
6.500 0.8924 0.02538 0.01459 -0.0300 0.3205 1.0000
6.750 0.9110 0.02601 0.01533 -0.0285 0.2962 1.0000
7.000 0.9296 0.02672 0.01611 -0.0270 0.2755 1.0000
7.250 0.9461 0.02743 0.01685 -0.0254 0.2514 1.0000
7.500 0.9599 0.02815 0.01757 -0.0235 0.2240 1.0000
7.750 0.9727 0.02903 0.01835 -0.0217 0.1971 1.0000
8.000 0.9865 0.03000 0.01933 -0.0201 0.1774 1.0000
8.250 1.0001 0.03105 0.02043 -0.0185 0.1583 1.0000
8.500 1.0105 0.03234 0.02170 -0.0167 0.1355 1.0000
8.750 1.0222 0.03360 0.02307 -0.0149 0.1174 1.0000
9.000 1.0341 0.03489 0.02455 -0.0131 0.0858 1.0000
9.250 1.0366 0.03678 0.02622 -0.0107 0.0518 1.0000
9.500 1.0343 0.03898 0.02824 -0.0080 0.0462 1.0000
9.750 1.0319 0.04128 0.03061 -0.0057 0.0434 1.0000
10.000 1.0277 0.04385 0.03330 -0.0037 0.0411 1.0000
10.250 1.0219 0.04673 0.03630 -0.0024 0.0395 1.0000
10.500 1.0140 0.05008 0.03985 -0.0017 0.0382 1.0000
10.750 1.0084 0.05347 0.04339 -0.0016 0.0370 1.0000
11.000 1.0036 0.05703 0.04716 -0.0018 0.0359 1.0000
11.250 0.9995 0.06067 0.05098 -0.0023 0.0350 1.0000
11.500 0.9955 0.06446 0.05496 -0.0029 0.0343 1.0000
11.750 0.9924 0.06825 0.05892 -0.0036 0.0338 1.0000
12.000 0.9903 0.07199 0.06283 -0.0043 0.0333 1.0000
12.250 0.9886 0.07576 0.06676 -0.0049 0.0328 1.0000
12.500 0.9869 0.07960 0.07076 -0.0056 0.0322 1.0000
12.750 0.9850 0.08355 0.07485 -0.0065 0.0315 1.0000
13.000 0.9835 0.08754 0.07898 -0.0073 0.0310 1.0000
13.250 0.9815 0.09166 0.08321 -0.0081 0.0303 1.0000
13.500 0.9799 0.09583 0.08744 -0.0086 0.0295 1.0000
13.750 0.9737 0.10114 0.09299 -0.0110 0.0293 1.0000
14.000 0.9672 0.10664 0.09870 -0.0135 0.0292 1.0000
14.250 0.9589 0.11271 0.10498 -0.0165 0.0291 1.0000
14.500 0.9497 0.11921 0.11166 -0.0199 0.0291 1.0000
14.750 0.9393 0.12631 0.11894 -0.0238 0.0291 1.0000
15.000 0.9312 0.13311 0.12587 -0.0274 0.0293 1.0000
15.250 0.9209 0.14096 0.13390 -0.0318 0.0296 1.0000
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Polar data table (+)
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