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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 200,000
Max Cl/Cd: 52.14 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa49-il-200000.txt
Download as CSV file: xf-usa49-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5817   0.09213   0.08876   0.0022   1.0000   0.0408
  -8.000  -0.5819   0.08765   0.08432  -0.0016   1.0000   0.0421
  -7.750  -0.5826   0.08238   0.07907  -0.0083   1.0000   0.0433
  -7.500  -0.5771   0.07535   0.07189  -0.0203   1.0000   0.0450
  -7.250  -0.5706   0.07069   0.06697  -0.0246   1.0000   0.0454
  -7.000  -0.5692   0.06373   0.06012  -0.0251   1.0000   0.0468
  -6.750  -0.5577   0.06114   0.05757  -0.0242   1.0000   0.0485
  -6.500  -0.5450   0.05786   0.05423  -0.0249   1.0000   0.0511
  -6.250  -0.5271   0.05492   0.05049  -0.0282   1.0000   0.0577
  -6.000  -0.5182   0.04817   0.04399  -0.0285   1.0000   0.0599
  -5.750  -0.5015   0.04567   0.04148  -0.0281   1.0000   0.0631
  -5.500  -0.4850   0.04176   0.03706  -0.0286   1.0000   0.0721
  -5.250  -0.4668   0.03923   0.03460  -0.0280   1.0000   0.0757
  -5.000  -0.4486   0.03628   0.03128  -0.0276   1.0000   0.0860
  -4.750  -0.4224   0.02730   0.02116  -0.0248   1.0000   0.0436
  -4.500  -0.4014   0.02374   0.01719  -0.0235   1.0000   0.0408
  -4.250  -0.3769   0.02037   0.01309  -0.0215   1.0000   0.0355
  -4.000  -0.3520   0.01948   0.01195  -0.0202   1.0000   0.0340
  -3.750  -0.3281   0.01779   0.01005  -0.0191   1.0000   0.0335
  -3.500  -0.3040   0.01631   0.00843  -0.0180   1.0000   0.0337
  -3.250  -0.2807   0.01454   0.00659  -0.0169   1.0000   0.0348
  -3.000  -0.2576   0.01351   0.00558  -0.0161   1.0000   0.0395
  -2.750  -0.2337   0.01281   0.00484  -0.0152   1.0000   0.0433
  -2.500  -0.2093   0.01210   0.00408  -0.0144   1.0000   0.0496
  -2.250  -0.1845   0.01112   0.00357  -0.0139   1.0000   0.1381
  -2.000  -0.1607   0.00981   0.00328  -0.0141   1.0000   0.3736
  -1.750  -0.1444   0.00823   0.00338  -0.0112   1.0000   0.7760
  -1.500  -0.0718   0.00812   0.00345  -0.0185   1.0000   1.0000
  -1.250  -0.0574   0.00814   0.00334  -0.0162   1.0000   1.0000
  -1.000  -0.0401   0.00818   0.00326  -0.0145   1.0000   1.0000
  -0.750  -0.0207   0.00826   0.00324  -0.0133   1.0000   1.0000
  -0.500  -0.0001   0.00837   0.00327  -0.0123   1.0000   1.0000
  -0.250   0.0404   0.00846   0.00328  -0.0153   0.9948   1.0000
   0.000   0.0904   0.00849   0.00323  -0.0201   0.9855   1.0000
   0.250   0.1395   0.00848   0.00319  -0.0247   0.9758   1.0000
   0.500   0.1891   0.00842   0.00312  -0.0292   0.9657   1.0000
   0.750   0.2326   0.00833   0.00304  -0.0323   0.9514   1.0000
   1.000   0.2700   0.00820   0.00291  -0.0337   0.9291   1.0000
   1.250   0.2959   0.00807   0.00273  -0.0324   0.8931   1.0000
   1.500   0.3180   0.00798   0.00255  -0.0302   0.8494   1.0000
   1.750   0.3404   0.00801   0.00244  -0.0284   0.8048   1.0000
   2.000   0.3637   0.00812   0.00239  -0.0269   0.7613   1.0000
   2.250   0.3874   0.00831   0.00240  -0.0256   0.7165   1.0000
   2.500   0.4113   0.00856   0.00247  -0.0245   0.6699   1.0000
   2.750   0.4353   0.00886   0.00255  -0.0235   0.6216   1.0000
   3.000   0.4594   0.00919   0.00266  -0.0226   0.5720   1.0000
   3.250   0.4842   0.00950   0.00282  -0.0220   0.5293   1.0000
   3.500   0.5090   0.00983   0.00303  -0.0214   0.4826   1.0000
   3.750   0.5334   0.01023   0.00320  -0.0207   0.4158   1.0000
   4.000   0.5567   0.01085   0.00343  -0.0200   0.3259   1.0000
   4.250   0.5788   0.01189   0.00384  -0.0194   0.1938   1.0000
   4.500   0.6010   0.01305   0.00443  -0.0189   0.1008   1.0000
   4.750   0.6261   0.01366   0.00505  -0.0184   0.0906   1.0000
   5.000   0.6511   0.01426   0.00571  -0.0180   0.0855   1.0000
   5.250   0.6744   0.01516   0.00662  -0.0173   0.0813   1.0000
   5.500   0.6990   0.01585   0.00739  -0.0167   0.0779   1.0000
   5.750   0.7225   0.01674   0.00831  -0.0161   0.0724   1.0000
   6.000   0.7448   0.01818   0.00977  -0.0153   0.0667   1.0000
   6.250   0.7691   0.01913   0.01081  -0.0148   0.0603   1.0000
   6.500   0.7917   0.02116   0.01287  -0.0141   0.0532   1.0000
   6.750   0.8160   0.02231   0.01418  -0.0134   0.0464   1.0000
   7.000   0.8383   0.02458   0.01658  -0.0127   0.0404   1.0000
   7.250   0.8623   0.02616   0.01851  -0.0118   0.0366   1.0000
   7.500   0.8850   0.02794   0.02047  -0.0110   0.0337   1.0000
   7.750   0.9022   0.03185   0.02460  -0.0103   0.0307   1.0000
   8.000   0.9227   0.03326   0.02650  -0.0089   0.0289   1.0000
   8.250   0.9384   0.03676   0.03050  -0.0072   0.0282   1.0000
   8.500   0.9495   0.04092   0.03518  -0.0055   0.0282   1.0000
   8.750   0.9559   0.04565   0.04038  -0.0038   0.0287   1.0000
   9.000   0.9573   0.05094   0.04603  -0.0024   0.0295   1.0000
  11.000   0.8154   0.11397   0.11062  -0.0315   0.0472   1.0000
  11.250   0.8117   0.11958   0.11620  -0.0341   0.0465   1.0000
  11.500   0.8100   0.12468   0.12128  -0.0362   0.0458   1.0000
  11.750   0.8103   0.12924   0.12584  -0.0377   0.0449   1.0000
  12.000   0.8130   0.13323   0.12983  -0.0384   0.0441   1.0000
  12.250   0.6418   0.13011   0.12679  -0.0248   0.0496   1.0000
  12.500   0.6375   0.13413   0.13080  -0.0262   0.0485   1.0000
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