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

USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 500,000
Max Cl/Cd: 59.67 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa49-il-500000.txt
Download as CSV file: xf-usa49-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.6120   0.12516   0.12287   0.0222   1.0000   0.0117
 -10.250  -0.6087   0.12126   0.11898   0.0206   1.0000   0.0120
  -7.000  -0.5752   0.05088   0.04828  -0.0267   1.0000   0.0161
  -6.750  -0.5646   0.04804   0.04536  -0.0273   1.0000   0.0167
  -6.500  -0.5502   0.04523   0.04244  -0.0278   1.0000   0.0174
  -6.250  -0.5345   0.04163   0.03867  -0.0282   1.0000   0.0181
  -6.000  -0.5178   0.03781   0.03462  -0.0282   1.0000   0.0190
  -5.750  -0.4998   0.03397   0.03050  -0.0277   1.0000   0.0202
  -5.500  -0.4774   0.03159   0.02781  -0.0266   1.0000   0.0224
  -4.250  -0.3812   0.01561   0.00976  -0.0183   1.0000   0.0162
  -4.000  -0.3586   0.01420   0.00819  -0.0171   1.0000   0.0163
  -3.750  -0.3362   0.01294   0.00686  -0.0161   1.0000   0.0174
  -3.500  -0.3132   0.01222   0.00611  -0.0151   1.0000   0.0187
  -3.250  -0.2895   0.01136   0.00515  -0.0142   1.0000   0.0188
  -3.000  -0.2612   0.01062   0.00436  -0.0143   0.9992   0.0198
  -2.750  -0.2248   0.00999   0.00366  -0.0161   0.9966   0.0215
  -2.500  -0.1875   0.00957   0.00320  -0.0181   0.9942   0.0234
  -2.250  -0.1518   0.00899   0.00253  -0.0197   0.9904   0.0303
  -2.000  -0.1151   0.00823   0.00219  -0.0218   0.9869   0.1306
  -1.750  -0.0754   0.00741   0.00198  -0.0250   0.9836   0.2891
  -1.500  -0.0439   0.00612   0.00182  -0.0264   0.9763   0.6023
  -1.000   0.0074   0.00515   0.00187  -0.0241   0.9590   0.9157
  -0.750   0.0376   0.00512   0.00185  -0.0239   0.9488   0.9521
  -0.500   0.0725   0.00512   0.00179  -0.0251   0.9381   0.9723
  -0.250   0.1085   0.00511   0.00173  -0.0265   0.9251   0.9858
   0.000   0.1486   0.00510   0.00164  -0.0289   0.9026   0.9969
   0.250   0.1764   0.00510   0.00150  -0.0285   0.8647   1.0000
   0.500   0.1996   0.00515   0.00137  -0.0272   0.8202   1.0000
   0.750   0.2235   0.00530   0.00127  -0.0261   0.7667   1.0000
   1.000   0.2481   0.00549   0.00123  -0.0253   0.7167   1.0000
   1.250   0.2729   0.00571   0.00122  -0.0247   0.6672   1.0000
   1.500   0.2977   0.00596   0.00124  -0.0241   0.6160   1.0000
   1.750   0.3229   0.00620   0.00128  -0.0236   0.5717   1.0000
   2.000   0.3484   0.00642   0.00133  -0.0231   0.5357   1.0000
   2.250   0.3742   0.00662   0.00140  -0.0228   0.5006   1.0000
   2.500   0.4000   0.00685   0.00150  -0.0224   0.4599   1.0000
   2.750   0.4253   0.00722   0.00159  -0.0221   0.3905   1.0000
   3.000   0.4500   0.00777   0.00175  -0.0218   0.3047   1.0000
   3.250   0.4739   0.00854   0.00201  -0.0214   0.1904   1.0000
   3.500   0.4974   0.00951   0.00244  -0.0211   0.0753   1.0000
   3.750   0.5238   0.00984   0.00277  -0.0208   0.0668   1.0000
   4.000   0.5503   0.01013   0.00310  -0.0206   0.0636   1.0000
   4.250   0.5766   0.01048   0.00348  -0.0204   0.0608   1.0000
   4.500   0.6026   0.01089   0.00392  -0.0201   0.0586   1.0000
   4.750   0.6282   0.01141   0.00450  -0.0198   0.0566   1.0000
   5.000   0.6530   0.01208   0.00524  -0.0193   0.0548   1.0000
   5.250   0.6796   0.01229   0.00549  -0.0192   0.0528   1.0000
   5.500   0.7058   0.01262   0.00586  -0.0190   0.0495   1.0000
   5.750   0.7313   0.01306   0.00635  -0.0187   0.0461   1.0000
   6.000   0.7532   0.01431   0.00764  -0.0180   0.0419   1.0000
   6.250   0.7815   0.01411   0.00751  -0.0181   0.0393   1.0000
   6.500   0.8087   0.01416   0.00760  -0.0181   0.0348   1.0000
   6.750   0.8338   0.01465   0.00811  -0.0178   0.0295   1.0000
   7.000   0.8623   0.01445   0.00790  -0.0179   0.0242   1.0000
   7.250   0.8866   0.01512   0.00862  -0.0175   0.0202   1.0000
   7.500   0.9119   0.01558   0.00912  -0.0171   0.0177   1.0000
   7.750   0.9320   0.01707   0.01072  -0.0160   0.0154   1.0000
   8.000   0.9540   0.01821   0.01202  -0.0152   0.0145   1.0000
   8.250   0.9767   0.01918   0.01312  -0.0145   0.0134   1.0000
   8.500   0.9998   0.01993   0.01398  -0.0140   0.0122   1.0000
   8.750   1.0213   0.02103   0.01520  -0.0132   0.0114   1.0000
   9.000   1.0392   0.02297   0.01735  -0.0121   0.0108   1.0000
   9.250   1.0489   0.02709   0.02195  -0.0100   0.0103   1.0000
   9.500   1.0599   0.03047   0.02575  -0.0083   0.0100   1.0000
   9.750   1.0705   0.03343   0.02909  -0.0067   0.0098   1.0000
  10.000   1.0723   0.03768   0.03380  -0.0045   0.0097   1.0000
  10.250   1.0631   0.04303   0.03961  -0.0021   0.0097   1.0000
  10.500   1.0447   0.04848   0.04544   0.0002   0.0098   1.0000
  10.750   1.0211   0.05305   0.05025   0.0022   0.0099   1.0000
  11.000   0.9970   0.05847   0.05587   0.0009   0.0100   1.0000
<< Back to USA 49 AIRFOIL (usa49-il)

Polar data table (+)

Polar graphs


<< Back to USA 49 AIRFOIL (usa49-il)