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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 50,000
Max Cl/Cd: 33.05 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa49-il-50000.txt
Download as CSV file: xf-usa49-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 49 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.5628   0.09495   0.08846   0.0123   1.0000   0.2580
  -7.250  -0.5604   0.09154   0.08512   0.0122   1.0000   0.2728
  -7.000  -0.5631   0.08855   0.08222   0.0117   1.0000   0.2907
  -6.750  -0.5543   0.08483   0.07855   0.0133   1.0000   0.3125
  -6.500  -0.5506   0.08162   0.07540   0.0143   1.0000   0.3390
  -6.250  -0.5405   0.07832   0.07214   0.0170   1.0000   0.3694
  -6.000  -0.5404   0.07574   0.06963   0.0187   1.0000   0.4052
  -5.750  -0.5224   0.07202   0.06593   0.0231   1.0000   0.4395
  -5.000  -0.4572   0.04381   0.03522  -0.0269   1.0000   0.1234
  -4.750  -0.4350   0.03949   0.03055  -0.0267   1.0000   0.1151
  -4.500  -0.4099   0.03592   0.02622  -0.0261   1.0000   0.1086
  -4.250  -0.3854   0.03276   0.02263  -0.0253   1.0000   0.1069
  -4.000  -0.3596   0.03007   0.01944  -0.0244   1.0000   0.1079
  -3.750  -0.3346   0.02780   0.01691  -0.0235   1.0000   0.1151
  -3.500  -0.3080   0.02584   0.01453  -0.0224   1.0000   0.1231
  -3.250  -0.2811   0.02376   0.01240  -0.0214   1.0000   0.1344
  -3.000  -0.2537   0.02187   0.01051  -0.0203   1.0000   0.1655
  -2.750  -0.1569   0.01580   0.00775  -0.0271   1.0000   1.0000
  -2.500  -0.1386   0.01560   0.00704  -0.0257   1.0000   1.0000
  -2.250  -0.1208   0.01544   0.00651  -0.0242   1.0000   1.0000
  -2.000  -0.1031   0.01532   0.00608  -0.0227   1.0000   1.0000
  -1.750  -0.0857   0.01524   0.00573  -0.0210   1.0000   1.0000
  -1.500  -0.0682   0.01519   0.00544  -0.0194   1.0000   1.0000
  -1.250  -0.0503   0.01517   0.00518  -0.0178   1.0000   1.0000
  -1.000  -0.0318   0.01518   0.00500  -0.0164   1.0000   1.0000
  -0.750  -0.0129   0.01521   0.00488  -0.0151   1.0000   1.0000
  -0.500   0.0064   0.01527   0.00480  -0.0138   1.0000   1.0000
  -0.250   0.0260   0.01537   0.00478  -0.0127   1.0000   1.0000
   0.000   0.0459   0.01549   0.00479  -0.0117   1.0000   1.0000
   0.250   0.0659   0.01565   0.00488  -0.0107   1.0000   1.0000
   0.500   0.0859   0.01586   0.00503  -0.0099   1.0000   1.0000
   0.750   0.1058   0.01611   0.00525  -0.0092   1.0000   1.0000
   1.000   0.1254   0.01642   0.00555  -0.0086   1.0000   1.0000
   1.250   0.1445   0.01681   0.00594  -0.0082   1.0000   1.0000
   1.500   0.1631   0.01729   0.00644  -0.0078   1.0000   1.0000
   1.750   0.1808   0.01789   0.00707  -0.0077   1.0000   1.0000
   2.000   0.1978   0.01861   0.00784  -0.0078   1.0000   1.0000
   2.250   0.2141   0.01947   0.00875  -0.0082   1.0000   1.0000
   2.500   0.2724   0.02054   0.01001  -0.0164   0.9807   1.0000
   2.750   0.3496   0.02135   0.01111  -0.0270   0.9465   1.0000
   3.000   0.4258   0.02168   0.01186  -0.0363   0.9114   1.0000
   3.250   0.4903   0.02147   0.01202  -0.0416   0.8730   1.0000
   3.500   0.5347   0.02062   0.01142  -0.0409   0.8237   1.0000
   3.750   0.5644   0.01996   0.01094  -0.0375   0.7771   1.0000
   4.000   0.5880   0.01969   0.01072  -0.0338   0.7332   1.0000
   4.250   0.6101   0.01975   0.01083  -0.0307   0.6912   1.0000
   4.500   0.6304   0.01984   0.01093  -0.0272   0.6427   1.0000
   4.750   0.6475   0.01985   0.01086  -0.0230   0.5802   1.0000
   5.000   0.6634   0.02007   0.01091  -0.0191   0.5012   1.0000
   5.250   0.6738   0.02109   0.01116  -0.0146   0.3427   1.0000
   5.500   0.6911   0.02367   0.01271  -0.0126   0.2322   1.0000
   5.750   0.7154   0.02547   0.01422  -0.0115   0.1999   1.0000
   6.000   0.7407   0.02712   0.01594  -0.0106   0.1772   1.0000
   6.250   0.7684   0.02920   0.01812  -0.0098   0.1645   1.0000
   6.500   0.7938   0.03132   0.02035  -0.0090   0.1500   1.0000
   6.750   0.8190   0.03399   0.02333  -0.0081   0.1398   1.0000
   7.000   0.8416   0.03675   0.02637  -0.0071   0.1287   1.0000
   7.250   0.8638   0.03999   0.02999  -0.0061   0.1221   1.0000
   7.500   0.8831   0.04408   0.03445  -0.0051   0.1185   1.0000
   7.750   0.8953   0.04824   0.03945  -0.0038   0.1152   1.0000
   8.000   0.9045   0.05305   0.04488  -0.0028   0.1146   1.0000
   8.250   0.9112   0.05850   0.05078  -0.0022   0.1171   1.0000
   8.500   0.9134   0.06429   0.05704  -0.0021   0.1213   1.0000
   8.750   0.8939   0.07194   0.06521  -0.0039   0.1286   1.0000
   9.000   0.8766   0.07927   0.07272  -0.0069   0.1362   1.0000
   9.250   0.8693   0.08686   0.08039  -0.0096   0.1470   1.0000
   9.500   0.8152   0.09999   0.09339  -0.0244   0.1701   1.0000
   9.750   0.7617   0.11967   0.11284  -0.0491   0.3129   1.0000
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