Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 500,000
Max Cl/Cd: 80.42 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa50-il-500000.txt
Download as CSV file: xf-usa50-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5022   0.08315   0.08097  -0.0178   1.0000   0.0135
  -7.750  -0.5098   0.07994   0.07781  -0.0185   1.0000   0.0136
  -7.500  -0.5157   0.07621   0.07410  -0.0202   1.0000   0.0138
  -7.250  -0.5163   0.07195   0.06985  -0.0229   1.0000   0.0138
  -7.000  -0.5151   0.06746   0.06534  -0.0253   1.0000   0.0140
  -6.750  -0.5117   0.06286   0.06069  -0.0274   1.0000   0.0146
  -6.500  -0.5065   0.05838   0.05613  -0.0286   1.0000   0.0148
  -6.250  -0.4983   0.05383   0.05146  -0.0294   1.0000   0.0157
  -6.000  -0.4797   0.05079   0.04826  -0.0291   1.0000   0.0170
  -5.750  -0.4692   0.04713   0.04443  -0.0282   1.0000   0.0172
  -5.500  -0.4594   0.04345   0.04056  -0.0268   1.0000   0.0172
  -5.250  -0.4490   0.03987   0.03676  -0.0252   1.0000   0.0173
  -4.250  -0.3787   0.02156   0.01684  -0.0217   0.9961   0.0177
  -4.000  -0.3474   0.01878   0.01365  -0.0225   0.9942   0.0178
  -3.750  -0.3200   0.01546   0.00992  -0.0220   0.9922   0.0160
  -3.500  -0.2900   0.01350   0.00767  -0.0221   0.9900   0.0157
  -3.250  -0.2583   0.01226   0.00624  -0.0227   0.9876   0.0162
  -3.000  -0.2242   0.01145   0.00531  -0.0239   0.9852   0.0175
  -2.750  -0.1901   0.01039   0.00412  -0.0252   0.9831   0.0214
  -2.500  -0.1599   0.00994   0.00360  -0.0256   0.9785   0.0261
  -2.250  -0.1265   0.00941   0.00318  -0.0266   0.9747   0.0599
  -2.000  -0.0896   0.00924   0.00302  -0.0287   0.9719   0.0789
  -1.750  -0.0557   0.00893   0.00276  -0.0301   0.9682   0.0955
  -1.500  -0.0255   0.00843   0.00249  -0.0307   0.9623   0.1554
  -1.250   0.0023   0.00702   0.00235  -0.0315   0.9583   0.5199
  -1.000   0.0228   0.00616   0.00224  -0.0297   0.9502   0.7304
  -0.750   0.1256   0.00587   0.00254  -0.0459   0.9614   0.9919
  -0.500   0.1764   0.00578   0.00238  -0.0510   0.9574   1.0000
  -0.250   0.2118   0.00559   0.00213  -0.0524   0.9441   1.0000
   0.000   0.2576   0.00543   0.00194  -0.0563   0.9307   1.0000
   0.250   0.3148   0.00530   0.00174  -0.0628   0.9131   1.0000
   0.500   0.3712   0.00523   0.00156  -0.0690   0.8815   1.0000
   0.750   0.4058   0.00531   0.00146  -0.0703   0.8320   1.0000
   1.000   0.4286   0.00550   0.00142  -0.0690   0.7722   1.0000
   1.250   0.4477   0.00576   0.00141  -0.0670   0.7177   1.0000
   1.500   0.4670   0.00602   0.00144  -0.0651   0.6697   1.0000
   1.750   0.4872   0.00624   0.00149  -0.0634   0.6276   1.0000
   2.000   0.5081   0.00645   0.00155  -0.0619   0.5932   1.0000
   2.250   0.5286   0.00671   0.00166  -0.0603   0.5561   1.0000
   2.500   0.5504   0.00691   0.00176  -0.0590   0.5295   1.0000
   2.750   0.5726   0.00712   0.00187  -0.0578   0.5015   1.0000
   3.000   0.5935   0.00738   0.00196  -0.0564   0.4501   1.0000
   3.250   0.6125   0.00782   0.00207  -0.0547   0.3634   1.0000
   3.500   0.6286   0.00864   0.00237  -0.0525   0.2521   1.0000
   3.750   0.6402   0.01003   0.00293  -0.0498   0.0741   1.0000
   4.000   0.6590   0.01078   0.00341  -0.0479   0.0207   1.0000
   4.250   0.6815   0.01114   0.00388  -0.0467   0.0185   1.0000
   4.500   0.7031   0.01162   0.00447  -0.0452   0.0170   1.0000
   4.750   0.7237   0.01220   0.00517  -0.0436   0.0161   1.0000
   5.000   0.7430   0.01291   0.00598  -0.0417   0.0152   1.0000
   5.250   0.7620   0.01364   0.00679  -0.0399   0.0135   1.0000
   5.500   0.7792   0.01462   0.00787  -0.0377   0.0128   1.0000
   5.750   0.7942   0.01614   0.00949  -0.0351   0.0119   1.0000
   6.000   0.8123   0.01772   0.01118  -0.0330   0.0117   1.0000
   6.250   0.8335   0.01895   0.01252  -0.0314   0.0121   1.0000
   6.500   0.8572   0.02270   0.01672  -0.0290   0.0177   1.0000
   6.750   0.8491   0.01371   0.00842  -0.0224   0.0249   1.0000
   7.000   0.8654   0.01518   0.01005  -0.0207   0.0226   1.0000
   7.250   0.8803   0.01721   0.01218  -0.0191   0.0211   1.0000
  19.000   0.7050   0.22203   0.22006  -0.0746   0.0072   1.0000
  19.250   0.7096   0.22545   0.22349  -0.0763   0.0071   1.0000
<< Back to USA 50 AIRFOIL (usa50-il)

Polar data table (+)

Polar graphs


<< Back to USA 50 AIRFOIL (usa50-il)