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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 50,000
Max Cl/Cd: 36.13 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa50-il-50000-n5.txt
Download as CSV file: xf-usa50-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4931   0.10418   0.09738  -0.0195   1.0000   0.0870
  -8.250  -0.5040   0.10172   0.09506  -0.0226   1.0000   0.0877
  -8.000  -0.5091   0.09842   0.09185  -0.0255   1.0000   0.0879
  -7.750  -0.4873   0.09130   0.08466  -0.0226   1.0000   0.0584
  -7.500  -0.4871   0.08628   0.07969  -0.0256   1.0000   0.0459
  -7.250  -0.4851   0.08242   0.07587  -0.0265   1.0000   0.0447
  -6.750  -0.4814   0.07223   0.06553  -0.0338   1.0000   0.0373
  -6.500  -0.4759   0.06822   0.06148  -0.0340   1.0000   0.0371
  -6.250  -0.4693   0.06422   0.05740  -0.0343   1.0000   0.0369
  -6.000  -0.4614   0.06021   0.05325  -0.0345   1.0000   0.0367
  -5.750  -0.4519   0.05624   0.04911  -0.0345   1.0000   0.0365
  -5.500  -0.4409   0.05233   0.04498  -0.0343   1.0000   0.0361
  -5.250  -0.4281   0.04857   0.04097  -0.0338   1.0000   0.0356
  -5.000  -0.4137   0.04486   0.03693  -0.0331   1.0000   0.0351
  -4.750  -0.3974   0.04134   0.03300  -0.0323   1.0000   0.0348
  -4.500  -0.3794   0.03789   0.02907  -0.0312   1.0000   0.0348
  -4.250  -0.3597   0.03480   0.02547  -0.0299   1.0000   0.0352
  -4.000  -0.3378   0.03202   0.02199  -0.0285   1.0000   0.0370
  -3.750  -0.3173   0.02987   0.01963  -0.0275   1.0000   0.0410
  -3.500  -0.2937   0.02770   0.01703  -0.0262   1.0000   0.0438
  -3.250  -0.2686   0.02564   0.01450  -0.0248   1.0000   0.0470
  -3.000  -0.2454   0.02416   0.01289  -0.0237   1.0000   0.0565
  -2.750  -0.2197   0.02255   0.01105  -0.0227   1.0000   0.0685
  -2.500  -0.1947   0.02114   0.00944  -0.0219   1.0000   0.1014
  -2.250  -0.1710   0.01971   0.00821  -0.0214   1.0000   0.1610
  -2.000  -0.1480   0.01798   0.00734  -0.0210   1.0000   0.3055
  -1.750  -0.0813   0.01580   0.00672  -0.0269   1.0000   1.0000
  -1.500  -0.0613   0.01582   0.00629  -0.0255   1.0000   1.0000
  -1.250  -0.0414   0.01587   0.00599  -0.0243   1.0000   1.0000
  -1.000  -0.0214   0.01594   0.00577  -0.0230   1.0000   1.0000
  -0.750  -0.0014   0.01605   0.00558  -0.0218   1.0000   1.0000
  -0.500   0.0183   0.01618   0.00550  -0.0207   1.0000   1.0000
  -0.250   0.0379   0.01634   0.00549  -0.0195   1.0000   1.0000
   0.000   0.0572   0.01654   0.00553  -0.0184   1.0000   1.0000
   0.250   0.0760   0.01678   0.00565  -0.0173   1.0000   1.0000
   0.500   0.1006   0.01708   0.00583  -0.0176   0.9969   1.0000
   0.750   0.1458   0.01744   0.00611  -0.0219   0.9825   1.0000
   1.000   0.1901   0.01774   0.00637  -0.0259   0.9678   1.0000
   1.250   0.2301   0.01793   0.00656  -0.0290   0.9501   1.0000
   1.500   0.2716   0.01810   0.00676  -0.0321   0.9328   1.0000
   1.750   0.3150   0.01820   0.00696  -0.0354   0.9158   1.0000
   2.000   0.3568   0.01819   0.00707  -0.0381   0.8949   1.0000
   2.250   0.4029   0.01808   0.00712  -0.0414   0.8740   1.0000
   2.500   0.4423   0.01803   0.00727  -0.0433   0.8490   1.0000
   2.750   0.4798   0.01802   0.00744  -0.0448   0.8241   1.0000
   3.000   0.5169   0.01803   0.00764  -0.0461   0.7989   1.0000
   3.250   0.5533   0.01808   0.00793  -0.0471   0.7733   1.0000
   3.500   0.5877   0.01818   0.00824  -0.0477   0.7468   1.0000
   3.750   0.6184   0.01839   0.00866  -0.0476   0.7197   1.0000
   4.000   0.6470   0.01867   0.00924  -0.0472   0.6922   1.0000
   4.250   0.6741   0.01902   0.00986  -0.0464   0.6645   1.0000
   4.500   0.6947   0.01923   0.01005  -0.0433   0.6043   1.0000
   4.750   0.7067   0.01967   0.01014  -0.0387   0.5096   1.0000
   5.000   0.7101   0.02075   0.01025  -0.0334   0.3189   1.0000
   5.250   0.7152   0.02432   0.01191  -0.0310   0.0685   1.0000
   5.500   0.7336   0.02610   0.01383  -0.0293   0.0489   1.0000
   5.750   0.7510   0.02768   0.01556  -0.0275   0.0399   1.0000
   6.000   0.7671   0.02930   0.01731  -0.0256   0.0350   1.0000
   6.250   0.7839   0.03103   0.01930  -0.0234   0.0330   1.0000
   6.500   0.8032   0.03293   0.02148  -0.0215   0.0313   1.0000
   6.750   0.8244   0.03500   0.02378  -0.0200   0.0288   1.0000
   7.000   0.8452   0.03761   0.02645  -0.0190   0.0266   1.0000
   7.250   0.8685   0.04041   0.02973  -0.0176   0.0257   1.0000
   7.500   0.8880   0.04358   0.03346  -0.0161   0.0255   1.0000
   7.750   0.9031   0.04696   0.03732  -0.0143   0.0255   1.0000
   8.000   0.9140   0.05049   0.04130  -0.0123   0.0256   1.0000
   8.250   0.9209   0.05414   0.04538  -0.0101   0.0257   1.0000
   8.500   0.9242   0.05784   0.04946  -0.0080   0.0259   1.0000
   8.750   0.9239   0.06165   0.05361  -0.0060   0.0260   1.0000
   9.000   0.9207   0.06542   0.05765  -0.0041   0.0262   1.0000
   9.250   0.9144   0.06919   0.06166  -0.0023   0.0264   1.0000
   9.500   0.9038   0.07292   0.06558  -0.0006   0.0265   1.0000
   9.750   0.8908   0.07651   0.06931   0.0008   0.0266   1.0000
  10.000   0.8763   0.08048   0.07346   0.0011   0.0267   1.0000
  10.250   0.8508   0.08538   0.07860  -0.0015   0.0274   1.0000
  10.500   0.8306   0.09191   0.08523  -0.0062   0.0277   1.0000
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