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RAE 5214 AIRFOIL (rae5214-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAE 5214 AIRFOIL (rae5214-il)
Reynolds number: 50,000
Max Cl/Cd: 27.54 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rae5214-il-50000-n5.txt
Download as CSV file: xf-rae5214-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 5214 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5226   0.09849   0.09071  -0.0243   1.0000   0.0574
  -9.500  -0.5446   0.08858   0.08086  -0.0329   1.0000   0.0539
  -9.250  -0.5469   0.08394   0.07626  -0.0350   1.0000   0.0534
  -9.000  -0.5544   0.07908   0.07143  -0.0377   1.0000   0.0528
  -8.750  -0.5673   0.07442   0.06679  -0.0402   1.0000   0.0522
  -8.500  -0.5798   0.06990   0.06223  -0.0423   1.0000   0.0516
  -8.250  -0.5895   0.06548   0.05768  -0.0437   1.0000   0.0511
  -8.000  -0.5954   0.06121   0.05323  -0.0444   1.0000   0.0508
  -7.750  -0.5969   0.05724   0.04902  -0.0446   1.0000   0.0508
  -7.500  -0.5945   0.05354   0.04503  -0.0445   1.0000   0.0513
  -7.250  -0.5890   0.04999   0.04110  -0.0441   1.0000   0.0522
  -7.000  -0.5804   0.04659   0.03724  -0.0435   1.0000   0.0531
  -6.750  -0.5684   0.04336   0.03349  -0.0427   1.0000   0.0538
  -6.500  -0.5528   0.04059   0.03040  -0.0418   1.0000   0.0544
  -6.250  -0.5352   0.03831   0.02797  -0.0409   1.0000   0.0554
  -6.000  -0.5169   0.03635   0.02582  -0.0399   1.0000   0.0575
  -5.750  -0.4974   0.03439   0.02353  -0.0388   1.0000   0.0600
  -5.500  -0.4766   0.03248   0.02118  -0.0376   1.0000   0.0620
  -5.250  -0.4561   0.03082   0.01953  -0.0364   1.0000   0.0640
  -5.000  -0.4354   0.02950   0.01814  -0.0351   1.0000   0.0674
  -4.750  -0.4139   0.02826   0.01667  -0.0336   1.0000   0.0710
  -4.500  -0.3938   0.02705   0.01553  -0.0319   1.0000   0.0738
  -4.250  -0.3739   0.02608   0.01454  -0.0303   1.0000   0.0782
  -4.000  -0.3545   0.02515   0.01356  -0.0287   1.0000   0.0837
  -3.750  -0.3352   0.02425   0.01266  -0.0272   1.0000   0.0899
  -3.500  -0.3156   0.02336   0.01179  -0.0260   1.0000   0.0984
  -3.250  -0.2949   0.02243   0.01090  -0.0252   1.0000   0.1123
  -3.000  -0.2733   0.02133   0.00996  -0.0248   1.0000   0.1401
  -2.750  -0.2534   0.01881   0.00919  -0.0253   1.0000   0.3875
  -2.500  -0.2431   0.01897   0.01006  -0.0204   1.0000   0.6045
  -2.250  -0.2311   0.01938   0.01053  -0.0161   1.0000   0.6764
  -2.000  -0.2257   0.01972   0.01101  -0.0099   1.0000   0.7331
  -1.750  -0.2218   0.01985   0.01124  -0.0034   1.0000   0.7819
  -1.500  -0.2117   0.01975   0.01112   0.0010   1.0000   0.8141
  -1.250  -0.1920   0.01959   0.01082   0.0023   1.0000   0.8282
  -1.000  -0.1709   0.01943   0.01055   0.0032   1.0000   0.8401
  -0.750  -0.1491   0.01932   0.01033   0.0038   1.0000   0.8526
  -0.500  -0.1264   0.01924   0.01018   0.0042   0.9997   0.8658
  -0.250  -0.0896   0.01935   0.01020   0.0019   0.9917   0.8784
   0.000  -0.0506   0.01951   0.01030  -0.0009   0.9840   0.8916
   0.250  -0.0100   0.01966   0.01042  -0.0039   0.9756   0.9056
   0.500   0.0324   0.01981   0.01057  -0.0074   0.9665   0.9208
   0.750   0.0791   0.01999   0.01076  -0.0117   0.9573   0.9383
   1.250   0.1757   0.02033   0.01117  -0.0211   0.9363   1.0000
   1.500   0.2252   0.02042   0.01131  -0.0254   0.9192   1.0000
   1.750   0.2878   0.02001   0.01098  -0.0306   0.8877   1.0000
   2.000   0.3447   0.01928   0.01033  -0.0338   0.8506   1.0000
   2.250   0.3820   0.01881   0.00991  -0.0342   0.8157   1.0000
   2.500   0.4157   0.01843   0.00956  -0.0339   0.7769   1.0000
   2.750   0.4441   0.01820   0.00933  -0.0330   0.7282   1.0000
   3.000   0.4732   0.01801   0.00904  -0.0318   0.6551   1.0000
   3.250   0.5017   0.01822   0.00852  -0.0299   0.4957   1.0000
   3.500   0.5174   0.01954   0.00876  -0.0277   0.3431   1.0000
   3.750   0.5367   0.02084   0.00943  -0.0268   0.2732   1.0000
   4.000   0.5593   0.02193   0.01018  -0.0264   0.2388   1.0000
   4.250   0.5836   0.02292   0.01097  -0.0261   0.2157   1.0000
   4.500   0.6091   0.02388   0.01182  -0.0259   0.1984   1.0000
   4.750   0.6359   0.02483   0.01270  -0.0259   0.1840   1.0000
   5.000   0.6635   0.02583   0.01361  -0.0259   0.1719   1.0000
   5.250   0.6913   0.02681   0.01459  -0.0260   0.1609   1.0000
   5.500   0.7195   0.02787   0.01571  -0.0260   0.1508   1.0000
   5.750   0.7469   0.02904   0.01674  -0.0260   0.1426   1.0000
   6.000   0.7742   0.03024   0.01817  -0.0259   0.1340   1.0000
   6.250   0.8004   0.03152   0.01931  -0.0258   0.1273   1.0000
   6.500   0.8257   0.03293   0.02104  -0.0255   0.1197   1.0000
   6.750   0.8504   0.03424   0.02229  -0.0253   0.1139   1.0000
   7.000   0.8734   0.03599   0.02441  -0.0247   0.1074   1.0000
   7.250   0.8963   0.03742   0.02587  -0.0243   0.1022   1.0000
   7.500   0.9170   0.03949   0.02825  -0.0237   0.0974   1.0000
   7.750   0.9365   0.04150   0.03056  -0.0229   0.0927   1.0000
   8.000   0.9572   0.04320   0.03228  -0.0224   0.0894   1.0000
   8.250   0.9706   0.04619   0.03577  -0.0213   0.0861   1.0000
   8.500   0.9821   0.04909   0.03910  -0.0201   0.0829   1.0000
   8.750   0.9956   0.05155   0.04178  -0.0192   0.0805   1.0000
   9.000   1.0101   0.05388   0.04420  -0.0185   0.0787   1.0000
   9.250   1.0103   0.05786   0.04863  -0.0171   0.0772   1.0000
   9.500   1.0009   0.06253   0.05380  -0.0155   0.0760   1.0000
   9.750   0.9868   0.06724   0.05888  -0.0143   0.0750   1.0000
  10.000   0.9664   0.07202   0.06392  -0.0133   0.0745   1.0000
  10.250   0.9379   0.07744   0.06953  -0.0133   0.0745   1.0000
  10.500   0.9038   0.08473   0.07696  -0.0164   0.0750   1.0000
  10.750   0.8689   0.09451   0.08681  -0.0229   0.0756   1.0000
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