USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 200,000 Max Cl/Cd: 65.11 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa50-il-200000.txt Download as CSV file: xf-usa50-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 50 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4933 0.08713 0.08370 -0.0197 1.0000 0.0365
-7.750 -0.5011 0.08393 0.08057 -0.0218 1.0000 0.0372
-7.500 -0.5034 0.07973 0.07639 -0.0268 1.0000 0.0379
-7.250 -0.5014 0.07552 0.07211 -0.0309 1.0000 0.0383
-7.000 -0.4981 0.07176 0.06823 -0.0329 1.0000 0.0385
-6.750 -0.4941 0.06814 0.06446 -0.0335 1.0000 0.0387
-6.500 -0.4969 0.06129 0.05760 -0.0337 1.0000 0.0395
-6.250 -0.4911 0.05765 0.05405 -0.0322 1.0000 0.0407
-6.000 -0.4832 0.05473 0.05111 -0.0311 1.0000 0.0422
-5.750 -0.4737 0.05153 0.04782 -0.0305 1.0000 0.0446
-5.500 -0.4562 0.04966 0.04536 -0.0303 1.0000 0.0508
-5.250 -0.4504 0.04400 0.03952 -0.0295 1.0000 0.0524
-5.000 -0.4388 0.04118 0.03685 -0.0284 1.0000 0.0558
-4.750 -0.4239 0.03851 0.03397 -0.0271 1.0000 0.0598
-4.500 -0.4087 0.03554 0.03039 -0.0253 1.0000 0.0659
-4.250 -0.3930 0.03244 0.02727 -0.0241 1.0000 0.0678
-4.000 -0.3703 0.02620 0.02010 -0.0208 1.0000 0.0410
-3.750 -0.3490 0.02268 0.01617 -0.0188 1.0000 0.0333
-3.500 -0.3260 0.01946 0.01234 -0.0168 1.0000 0.0308
-3.250 -0.3026 0.01764 0.01021 -0.0153 1.0000 0.0311
-3.000 -0.2798 0.01683 0.00925 -0.0142 1.0000 0.0353
-2.750 -0.2563 0.01583 0.00805 -0.0129 1.0000 0.0378
-2.500 -0.2341 0.01437 0.00651 -0.0114 1.0000 0.0409
-2.250 -0.2126 0.01371 0.00581 -0.0101 1.0000 0.0488
-2.000 -0.1917 0.01302 0.00520 -0.0087 1.0000 0.0738
-1.750 -0.1714 0.01244 0.00465 -0.0074 1.0000 0.1049
-1.500 -0.1513 0.01207 0.00444 -0.0063 1.0000 0.1387
-1.250 -0.0621 0.00913 0.00438 -0.0189 1.0000 1.0000
-1.000 -0.0436 0.00925 0.00434 -0.0173 1.0000 1.0000
-0.750 -0.0249 0.00940 0.00436 -0.0158 0.9999 1.0000
-0.500 0.0239 0.00959 0.00440 -0.0206 0.9936 1.0000
-0.250 0.0709 0.00969 0.00436 -0.0249 0.9860 1.0000
0.000 0.1170 0.00973 0.00432 -0.0290 0.9779 1.0000
0.250 0.1686 0.00968 0.00422 -0.0341 0.9694 1.0000
0.500 0.2185 0.00947 0.00399 -0.0386 0.9576 1.0000
0.750 0.2649 0.00921 0.00372 -0.0423 0.9446 1.0000
1.000 0.3086 0.00894 0.00346 -0.0454 0.9314 1.0000
1.250 0.3551 0.00869 0.00323 -0.0491 0.9152 1.0000
1.500 0.4051 0.00844 0.00304 -0.0535 0.8945 1.0000
1.750 0.4560 0.00825 0.00285 -0.0581 0.8653 1.0000
2.000 0.4981 0.00820 0.00275 -0.0607 0.8231 1.0000
2.250 0.5297 0.00833 0.00276 -0.0611 0.7745 1.0000
2.500 0.5551 0.00858 0.00282 -0.0602 0.7293 1.0000
2.750 0.5775 0.00887 0.00296 -0.0587 0.6898 1.0000
3.000 0.5969 0.00919 0.00309 -0.0566 0.6415 1.0000
3.250 0.6158 0.00953 0.00323 -0.0545 0.5939 1.0000
3.500 0.6361 0.00985 0.00343 -0.0528 0.5574 1.0000
3.750 0.6542 0.01026 0.00369 -0.0506 0.5052 1.0000
4.000 0.6716 0.01066 0.00388 -0.0483 0.4285 1.0000
4.250 0.6827 0.01167 0.00417 -0.0451 0.2824 1.0000
4.500 0.6884 0.01381 0.00511 -0.0415 0.0634 1.0000
4.750 0.7072 0.01472 0.00590 -0.0395 0.0355 1.0000
5.000 0.7257 0.01564 0.00692 -0.0375 0.0300 1.0000
5.250 0.7430 0.01668 0.00809 -0.0352 0.0285 1.0000
5.500 0.7602 0.01783 0.00939 -0.0330 0.0276 1.0000
5.750 0.7781 0.01923 0.01088 -0.0308 0.0272 1.0000
6.000 0.7985 0.02082 0.01257 -0.0291 0.0267 1.0000
6.250 0.8202 0.02215 0.01401 -0.0277 0.0244 1.0000
6.500 0.8431 0.02428 0.01635 -0.0263 0.0241 1.0000
6.750 0.8659 0.02725 0.01966 -0.0246 0.0255 1.0000
7.000 0.8848 0.03103 0.02382 -0.0226 0.0278 1.0000
7.250 0.9040 0.03644 0.02988 -0.0194 0.0385 1.0000
12.250 0.6735 0.13006 0.12678 -0.0180 0.0520 1.0000
12.500 0.6652 0.13503 0.13174 -0.0216 0.0518 1.0000
12.750 0.6526 0.13958 0.13627 -0.0260 0.0513 1.0000
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