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NLR-7223-62 AIRFOIL (nl722362-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NLR-7223-62 AIRFOIL (nl722362-il)
Reynolds number: 50,000
Max Cl/Cd: 30.01 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nl722362-il-50000.txt
Download as CSV file: xf-nl722362-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NLR-7223-62 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4739   0.10236   0.09544  -0.0097   1.0000   0.2864
  -8.500  -0.4727   0.09952   0.09268  -0.0089   1.0000   0.3025
  -8.250  -0.4776   0.09717   0.09041  -0.0080   1.0000   0.3195
  -8.000  -0.4635   0.09324   0.08652  -0.0061   1.0000   0.3406
  -7.750  -0.4615   0.09074   0.08408  -0.0041   1.0000   0.3647
  -7.500  -0.4487   0.08732   0.08070  -0.0018   1.0000   0.3913
  -7.250  -0.4418   0.08458   0.07802   0.0008   1.0000   0.4204
  -7.000  -0.4346   0.08212   0.07562   0.0041   1.0000   0.4577
  -6.750  -0.4136   0.07890   0.07240   0.0072   1.0000   0.5007
  -4.750  -0.5459   0.04714   0.03903  -0.0046   1.0000   0.1647
  -4.500  -0.5285   0.04377   0.03470  -0.0022   1.0000   0.1362
  -4.250  -0.5105   0.04027   0.03113  -0.0007   1.0000   0.1310
  -4.000  -0.4919   0.03761   0.02776   0.0017   1.0000   0.1227
  -3.750  -0.4712   0.03547   0.02515   0.0037   1.0000   0.1191
  -3.500  -0.4490   0.03307   0.02246   0.0053   1.0000   0.1161
  -3.250  -0.4247   0.03094   0.01995   0.0068   1.0000   0.1137
  -3.000  -0.3989   0.02910   0.01776   0.0081   1.0000   0.1136
  -2.750  -0.3725   0.02764   0.01602   0.0092   1.0000   0.1177
  -2.500  -0.1426   0.01972   0.01157  -0.0218   1.0000   1.0000
  -2.250  -0.1313   0.01961   0.01089  -0.0188   1.0000   1.0000
  -2.000  -0.1198   0.01952   0.01040  -0.0161   1.0000   1.0000
  -1.750  -0.1074   0.01947   0.01004  -0.0136   1.0000   1.0000
  -1.500  -0.0941   0.01945   0.00974  -0.0113   1.0000   1.0000
  -1.250  -0.0801   0.01945   0.00951  -0.0090   1.0000   1.0000
  -1.000  -0.0655   0.01948   0.00930  -0.0069   1.0000   1.0000
  -0.750  -0.0504   0.01953   0.00916  -0.0048   1.0000   1.0000
  -0.500  -0.0349   0.01961   0.00907  -0.0029   1.0000   1.0000
  -0.250  -0.0191   0.01971   0.00902  -0.0010   1.0000   1.0000
   0.000  -0.0030   0.01983   0.00901   0.0009   1.0000   1.0000
   0.250   0.0132   0.01997   0.00904   0.0026   1.0000   1.0000
   0.500   0.0297   0.02014   0.00911   0.0043   1.0000   1.0000
   0.750   0.0461   0.02033   0.00924   0.0060   1.0000   1.0000
   1.000   0.0627   0.02056   0.00940   0.0076   1.0000   1.0000
   1.250   0.0792   0.02081   0.00961   0.0091   1.0000   1.0000
   1.500   0.0955   0.02109   0.00988   0.0106   1.0000   1.0000
   1.750   0.1116   0.02142   0.01021   0.0121   1.0000   1.0000
   2.000   0.1274   0.02179   0.01059   0.0135   1.0000   1.0000
   2.250   0.1428   0.02222   0.01105   0.0148   1.0000   1.0000
   2.500   0.1576   0.02271   0.01159   0.0161   1.0000   1.0000
   2.750   0.1717   0.02329   0.01224   0.0173   1.0000   1.0000
   3.000   0.1963   0.02414   0.01321   0.0161   0.9956   1.0000
   3.250   0.3306   0.02546   0.01505  -0.0050   0.9408   1.0000
   3.500   0.4707   0.02420   0.01455  -0.0229   0.8724   1.0000
   3.750   0.5816   0.01938   0.00998  -0.0277   0.6514   1.0000
   4.000   0.6168   0.02071   0.00990  -0.0263   0.5033   1.0000
   4.250   0.6431   0.02196   0.01078  -0.0256   0.4531   1.0000
   4.500   0.6715   0.02309   0.01177  -0.0255   0.4213   1.0000
   4.750   0.6995   0.02421   0.01278  -0.0252   0.3973   1.0000
   5.000   0.7245   0.02531   0.01398  -0.0246   0.3770   1.0000
   5.250   0.7484   0.02647   0.01521  -0.0237   0.3590   1.0000
   5.500   0.7714   0.02773   0.01659  -0.0228   0.3425   1.0000
   5.750   0.7934   0.02906   0.01805  -0.0217   0.3266   1.0000
   6.000   0.8144   0.03049   0.01963  -0.0204   0.3108   1.0000
   6.250   0.8342   0.03206   0.02138  -0.0190   0.2951   1.0000
   6.500   0.8529   0.03369   0.02319  -0.0173   0.2783   1.0000
   6.750   0.8707   0.03547   0.02512  -0.0156   0.2615   1.0000
   7.000   0.8885   0.03737   0.02707  -0.0138   0.2436   1.0000
   7.250   0.8996   0.03942   0.02943  -0.0112   0.2251   1.0000
   7.500   0.9094   0.04145   0.03175  -0.0084   0.2055   1.0000
   7.750   0.9242   0.04311   0.03339  -0.0062   0.1844   1.0000
   8.000   0.9362   0.04540   0.03572  -0.0038   0.1669   1.0000
   8.500   0.9351   0.05183   0.04306   0.0033   0.1469   1.0000
   8.750   0.9485   0.05442   0.04560   0.0051   0.1372   1.0000
   9.000   0.9340   0.05889   0.05061   0.0091   0.1356   1.0000
   9.250   0.9176   0.06345   0.05553   0.0124   0.1345   1.0000
   9.500   0.8988   0.06808   0.06041   0.0152   0.1343   1.0000
   9.750   0.8773   0.07282   0.06530   0.0175   0.1348   1.0000
  10.000   0.8559   0.07747   0.07001   0.0194   0.1359   1.0000
  10.250   0.8369   0.08224   0.07480   0.0205   0.1369   1.0000
  10.500   0.8213   0.08755   0.08011   0.0203   0.1377   1.0000
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