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USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: USA 5 AIRFOIL (usa5-il)
Reynolds number: 1,000,000
Max Cl/Cd: 112.54 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa5-il-1000000.txt
Download as CSV file: xf-usa5-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 5 AIRFOIL                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3411   0.09987   0.09829  -0.0222   1.0000   0.0079
  -8.500  -0.3394   0.09720   0.09564  -0.0227   1.0000   0.0080
  -8.250  -0.3387   0.09460   0.09307  -0.0227   1.0000   0.0080
  -6.500  -0.2633   0.06544   0.06395  -0.0506   0.9848   0.0085
  -6.250  -0.2357   0.06146   0.05993  -0.0566   0.9827   0.0088
  -6.000  -0.2135   0.05950   0.05795  -0.0591   0.9782   0.0102
  -5.750  -0.1823   0.05483   0.05322  -0.0657   0.9740   0.0114
  -5.500  -0.1414   0.05043   0.04873  -0.0729   0.9710   0.0126
  -5.250  -0.1095   0.04583   0.04403  -0.0780   0.9655   0.0126
  -4.500  -0.0700   0.01346   0.00965  -0.0804   0.9308   0.0087
  -4.250  -0.0486   0.01157   0.00735  -0.0789   0.9210   0.0087
  -4.000  -0.0246   0.01041   0.00594  -0.0779   0.9116   0.0095
  -3.750   0.0008   0.00991   0.00533  -0.0772   0.9014   0.0104
  -3.500   0.0266   0.00954   0.00485  -0.0765   0.8906   0.0116
  -3.250   0.0519   0.00901   0.00417  -0.0757   0.8774   0.0122
  -3.000   0.0774   0.00867   0.00368  -0.0750   0.8615   0.0126
  -2.750   0.1005   0.00786   0.00266  -0.0737   0.8434   0.0151
  -2.500   0.1258   0.00772   0.00244  -0.0729   0.8242   0.0176
  -2.250   0.1506   0.00760   0.00219  -0.0720   0.8026   0.0192
  -2.000   0.1741   0.00726   0.00171  -0.0709   0.7791   0.0248
  -1.750   0.1988   0.00725   0.00159  -0.0701   0.7554   0.0289
  -1.500   0.2223   0.00710   0.00134  -0.0689   0.7286   0.0388
  -1.250   0.2462   0.00706   0.00120  -0.0680   0.7015   0.0485
  -1.000   0.2701   0.00707   0.00117  -0.0670   0.6733   0.0703
  -0.750   0.2950   0.00717   0.00119  -0.0663   0.6487   0.0838
  -0.500   0.3204   0.00730   0.00123  -0.0657   0.6251   0.0910
  -0.250   0.3453   0.00735   0.00119  -0.0650   0.6017   0.0959
   0.000   0.3702   0.00744   0.00119  -0.0643   0.5758   0.1011
   0.250   0.3949   0.00757   0.00118  -0.0636   0.5433   0.1044
   0.500   0.4185   0.00773   0.00116  -0.0627   0.5003   0.1075
   0.750   0.4416   0.00791   0.00116  -0.0618   0.4565   0.1117
   1.000   0.4660   0.00804   0.00118  -0.0611   0.4295   0.1155
   1.250   0.4911   0.00816   0.00122  -0.0605   0.4107   0.1204
   1.500   0.5162   0.00821   0.00123  -0.0599   0.3965   0.1264
   2.000   0.5672   0.00832   0.00130  -0.0589   0.3749   0.1404
   2.250   0.5923   0.00832   0.00137  -0.0584   0.3668   0.1774
   2.500   0.7282   0.00692   0.00175  -0.0842   0.3506   1.0000
   2.750   0.7534   0.00705   0.00183  -0.0836   0.3432   1.0000
   3.000   0.7785   0.00718   0.00191  -0.0831   0.3313   1.0000
   3.250   0.8037   0.00731   0.00199  -0.0825   0.3205   1.0000
   3.500   0.8288   0.00744   0.00209  -0.0820   0.3121   1.0000
   3.750   0.8536   0.00759   0.00219  -0.0814   0.2981   1.0000
   4.000   0.8778   0.00780   0.00229  -0.0807   0.2760   1.0000
   4.250   0.8972   0.00846   0.00256  -0.0793   0.1952   1.0000
   4.500   0.9170   0.00911   0.00295  -0.0780   0.1512   1.0000
   4.750   0.9301   0.01052   0.00370  -0.0755   0.0242   1.0000
   5.000   0.9533   0.01086   0.00402  -0.0745   0.0147   1.0000
   5.250   0.9770   0.01112   0.00434  -0.0737   0.0132   1.0000
   5.500   1.0004   0.01142   0.00471  -0.0728   0.0123   1.0000
   5.750   1.0231   0.01180   0.00515  -0.0718   0.0112   1.0000
   6.000   1.0448   0.01227   0.00570  -0.0706   0.0100   1.0000
   6.250   1.0641   0.01300   0.00653  -0.0690   0.0087   1.0000
   6.500   1.0801   0.01404   0.00771  -0.0668   0.0082   1.0000
   6.750   1.0990   0.01471   0.00845  -0.0651   0.0079   1.0000
   7.000   1.1194   0.01521   0.00901  -0.0638   0.0076   1.0000
   7.250   1.1389   0.01577   0.00963  -0.0622   0.0071   1.0000
   7.500   1.1553   0.01659   0.01052  -0.0602   0.0068   1.0000
   7.750   1.1706   0.01748   0.01148  -0.0580   0.0064   1.0000
   8.000   1.1838   0.01855   0.01263  -0.0553   0.0061   1.0000
   8.250   1.1949   0.01988   0.01405  -0.0524   0.0061   1.0000
   8.500   1.2072   0.02121   0.01545  -0.0497   0.0061   1.0000
   8.750   1.2213   0.02235   0.01668  -0.0475   0.0059   1.0000
   9.000   1.2355   0.02356   0.01795  -0.0454   0.0057   1.0000
   9.250   1.2501   0.02432   0.01872  -0.0436   0.0053   1.0000
   9.500   1.2641   0.02703   0.02165  -0.0411   0.0058   1.0000
   9.750   1.2786   0.02939   0.02419  -0.0390   0.0061   1.0000
  10.000   1.2915   0.03218   0.02719  -0.0368   0.0063   1.0000
  10.250   1.3011   0.03358   0.02871  -0.0343   0.0059   1.0000
  11.250   1.2965   0.04136   0.03702  -0.0201   0.0042   1.0000
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