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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 100,000
Max Cl/Cd: 42.13 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa49-il-100000-n5.txt
Download as CSV file: xf-usa49-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4358   0.08726   0.08276  -0.0107   1.0000   0.0495
  -8.250  -0.4380   0.08294   0.07848  -0.0121   1.0000   0.0492
  -8.000  -0.4413   0.07844   0.07402  -0.0137   1.0000   0.0482
  -7.500  -0.4573   0.06613   0.06178  -0.0193   1.0000   0.0264
  -7.000  -0.5287   0.06644   0.06162  -0.0261   1.0000   0.0217
  -6.750  -0.5209   0.06179   0.05688  -0.0280   1.0000   0.0212
  -6.500  -0.5108   0.05682   0.05176  -0.0299   1.0000   0.0209
  -6.250  -0.4988   0.05182   0.04654  -0.0314   1.0000   0.0209
  -6.000  -0.4850   0.04698   0.04141  -0.0322   1.0000   0.0209
  -5.750  -0.4696   0.04249   0.03656  -0.0325   1.0000   0.0211
  -5.500  -0.4530   0.03838   0.03204  -0.0322   1.0000   0.0211
  -5.250  -0.4350   0.03462   0.02783  -0.0316   1.0000   0.0208
  -5.000  -0.4155   0.03123   0.02395  -0.0307   1.0000   0.0206
  -4.750  -0.3944   0.02826   0.02043  -0.0296   1.0000   0.0206
  -4.500  -0.3721   0.02572   0.01742  -0.0285   1.0000   0.0207
  -4.250  -0.3489   0.02359   0.01486  -0.0274   1.0000   0.0210
  -4.000  -0.3250   0.02197   0.01290  -0.0263   1.0000   0.0220
  -3.750  -0.3019   0.02015   0.01089  -0.0253   1.0000   0.0239
  -3.500  -0.2788   0.01877   0.00941  -0.0244   1.0000   0.0256
  -3.250  -0.2555   0.01766   0.00817  -0.0234   1.0000   0.0274
  -3.000  -0.2320   0.01677   0.00714  -0.0224   1.0000   0.0302
  -2.750  -0.2084   0.01595   0.00624  -0.0217   1.0000   0.0363
  -2.500  -0.1842   0.01522   0.00545  -0.0209   1.0000   0.0452
  -2.250  -0.1601   0.01426   0.00496  -0.0204   1.0000   0.1380
  -2.000  -0.1363   0.01320   0.00465  -0.0203   1.0000   0.2993
  -1.750  -0.1208   0.01135   0.00474  -0.0170   1.0000   0.7646
  -1.500  -0.0737   0.01115   0.00459  -0.0197   1.0000   0.9709
  -1.250  -0.0447   0.01116   0.00435  -0.0203   1.0000   1.0000
  -1.000  -0.0243   0.01123   0.00424  -0.0192   1.0000   1.0000
  -0.750   0.0166   0.01130   0.00410  -0.0223   0.9911   1.0000
  -0.500   0.0581   0.01136   0.00398  -0.0254   0.9816   1.0000
  -0.250   0.0999   0.01140   0.00390  -0.0285   0.9713   1.0000
   0.000   0.1413   0.01143   0.00382  -0.0314   0.9592   1.0000
   0.250   0.1810   0.01142   0.00375  -0.0339   0.9440   1.0000
   0.500   0.2186   0.01140   0.00368  -0.0357   0.9267   1.0000
   0.750   0.2506   0.01139   0.00363  -0.0364   0.9057   1.0000
   1.000   0.2823   0.01137   0.00358  -0.0369   0.8847   1.0000
   1.250   0.3110   0.01137   0.00356  -0.0367   0.8602   1.0000
   1.500   0.3389   0.01138   0.00355  -0.0363   0.8339   1.0000
   1.750   0.3660   0.01139   0.00351  -0.0356   0.8026   1.0000
   2.000   0.3916   0.01144   0.00343  -0.0344   0.7600   1.0000
   2.250   0.4161   0.01158   0.00340  -0.0330   0.7102   1.0000
   2.500   0.4405   0.01181   0.00343  -0.0319   0.6620   1.0000
   2.750   0.4646   0.01209   0.00352  -0.0307   0.6139   1.0000
   3.000   0.4882   0.01245   0.00367  -0.0297   0.5606   1.0000
   3.250   0.5128   0.01279   0.00387  -0.0289   0.5195   1.0000
   3.500   0.5378   0.01311   0.00412  -0.0283   0.4844   1.0000
   3.750   0.5626   0.01346   0.00439  -0.0277   0.4420   1.0000
   4.000   0.5865   0.01392   0.00469  -0.0270   0.3809   1.0000
   4.250   0.6098   0.01453   0.00502  -0.0263   0.3150   1.0000
   4.500   0.6329   0.01527   0.00546  -0.0257   0.2398   1.0000
   4.750   0.6548   0.01632   0.00604  -0.0252   0.1424   1.0000
   5.000   0.6775   0.01730   0.00677  -0.0246   0.0966   1.0000
   5.250   0.7015   0.01803   0.00752  -0.0241   0.0879   1.0000
   5.500   0.7256   0.01872   0.00834  -0.0235   0.0832   1.0000
   5.750   0.7490   0.01950   0.00927  -0.0229   0.0797   1.0000
   6.000   0.7711   0.02049   0.01042  -0.0220   0.0770   1.0000
   6.250   0.7940   0.02135   0.01148  -0.0212   0.0740   1.0000
   6.500   0.8155   0.02244   0.01268  -0.0205   0.0666   1.0000
   6.750   0.8367   0.02361   0.01400  -0.0198   0.0556   1.0000
   7.000   0.8573   0.02499   0.01555  -0.0189   0.0455   1.0000
   7.250   0.8782   0.02639   0.01712  -0.0180   0.0357   1.0000
   7.500   0.8994   0.02762   0.01848  -0.0172   0.0291   1.0000
   7.750   0.9208   0.02938   0.02054  -0.0161   0.0240   1.0000
   8.000   0.9405   0.03093   0.02232  -0.0152   0.0216   1.0000
   8.250   0.9591   0.03361   0.02532  -0.0140   0.0197   1.0000
   8.500   0.9760   0.03681   0.02903  -0.0125   0.0177   1.0000
   8.750   0.9884   0.04062   0.03336  -0.0108   0.0169   1.0000
   9.000   0.9959   0.04476   0.03802  -0.0090   0.0163   1.0000
   9.250   0.9979   0.04915   0.04292  -0.0072   0.0160   1.0000
   9.500   0.9945   0.05365   0.04787  -0.0056   0.0159   1.0000
   9.750   0.9855   0.05819   0.05278  -0.0041   0.0158   1.0000
  10.000   0.9706   0.06249   0.05736  -0.0028   0.0159   1.0000
  10.250   0.9518   0.06705   0.06214  -0.0029   0.0159   1.0000
  10.500   0.9318   0.07263   0.06790  -0.0055   0.0161   1.0000
  10.750   0.9112   0.07989   0.07531  -0.0111   0.0163   1.0000
  11.000   0.8908   0.08919   0.08471  -0.0187   0.0167   1.0000
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