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NACA 1408 (naca1408-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 1408 (naca1408-il)
Reynolds number: 200,000
Max Cl/Cd: 54.89 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca1408-il-200000.txt
Download as CSV file: xf-naca1408-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 1408                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6229   0.10337   0.09982  -0.0032   1.0000   0.0538
  -9.500  -0.6297   0.09724   0.09373  -0.0102   1.0000   0.0542
  -9.250  -0.6382   0.08993   0.08645  -0.0191   1.0000   0.0543
  -9.000  -0.6343   0.08641   0.08298  -0.0124   1.0000   0.0557
  -8.750  -0.6245   0.08455   0.08112  -0.0099   1.0000   0.0572
  -8.500  -0.6236   0.08060   0.07719  -0.0122   1.0000   0.0585
  -8.250  -0.6261   0.07518   0.07177  -0.0176   1.0000   0.0598
  -8.000  -0.6270   0.06908   0.06561  -0.0231   1.0000   0.0619
  -7.000  -0.6306   0.03607   0.03035  -0.0306   1.0000   0.0434
  -6.750  -0.6134   0.03231   0.02632  -0.0295   1.0000   0.0417
  -6.500  -0.5971   0.02860   0.02213  -0.0281   1.0000   0.0415
  -6.250  -0.5773   0.02627   0.01932  -0.0267   1.0000   0.0423
  -6.000  -0.5589   0.02298   0.01568  -0.0256   1.0000   0.0447
  -5.750  -0.5369   0.02141   0.01396  -0.0247   1.0000   0.0470
  -5.500  -0.5138   0.01995   0.01232  -0.0237   1.0000   0.0492
  -5.250  -0.4903   0.01893   0.01109  -0.0226   1.0000   0.0527
  -5.000  -0.4670   0.01747   0.00944  -0.0216   1.0000   0.0558
  -4.750  -0.4439   0.01633   0.00832  -0.0208   1.0000   0.0593
  -4.500  -0.4203   0.01569   0.00762  -0.0199   1.0000   0.0643
  -4.250  -0.3970   0.01479   0.00666  -0.0189   1.0000   0.0683
  -4.000  -0.3740   0.01400   0.00594  -0.0181   1.0000   0.0740
  -3.750  -0.3504   0.01344   0.00533  -0.0172   1.0000   0.0811
  -3.500  -0.3270   0.01277   0.00471  -0.0164   1.0000   0.0917
  -3.250  -0.3032   0.01198   0.00409  -0.0158   1.0000   0.1214
  -3.000  -0.2799   0.01083   0.00368  -0.0156   1.0000   0.2542
  -2.750  -0.2560   0.01017   0.00348  -0.0152   1.0000   0.3700
  -2.500  -0.2322   0.00958   0.00337  -0.0147   1.0000   0.4882
  -2.250  -0.2091   0.00903   0.00335  -0.0138   1.0000   0.6167
  -2.000  -0.1880   0.00856   0.00343  -0.0119   1.0000   0.7538
  -1.750  -0.1682   0.00838   0.00362  -0.0089   1.0000   0.8860
  -1.500  -0.1088   0.00843   0.00365  -0.0147   1.0000   0.9894
  -1.250  -0.0820   0.00845   0.00356  -0.0150   1.0000   1.0000
  -1.000  -0.0648   0.00852   0.00355  -0.0135   1.0000   1.0000
  -0.750  -0.0454   0.00864   0.00359  -0.0124   1.0000   1.0000
  -0.500  -0.0114   0.00876   0.00365  -0.0142   0.9965   1.0000
  -0.250   0.0358   0.00884   0.00366  -0.0185   0.9888   1.0000
   0.000   0.0840   0.00889   0.00368  -0.0229   0.9813   1.0000
   0.250   0.1316   0.00888   0.00366  -0.0272   0.9724   1.0000
   0.500   0.1792   0.00882   0.00361  -0.0312   0.9625   1.0000
   0.750   0.2253   0.00871   0.00353  -0.0349   0.9515   1.0000
   1.000   0.2658   0.00859   0.00343  -0.0371   0.9367   1.0000
   1.250   0.3004   0.00847   0.00334  -0.0381   0.9180   1.0000
   1.500   0.3311   0.00836   0.00323  -0.0380   0.8967   1.0000
   1.750   0.3569   0.00830   0.00316  -0.0369   0.8699   1.0000
   2.000   0.3813   0.00829   0.00310  -0.0355   0.8404   1.0000
   2.250   0.4050   0.00834   0.00308  -0.0340   0.8072   1.0000
   2.500   0.4280   0.00844   0.00305  -0.0323   0.7659   1.0000
   2.750   0.4508   0.00861   0.00307  -0.0307   0.7180   1.0000
   3.000   0.4735   0.00886   0.00310  -0.0291   0.6653   1.0000
   3.250   0.4962   0.00918   0.00319  -0.0277   0.6075   1.0000
   3.500   0.5195   0.00952   0.00333  -0.0266   0.5542   1.0000
   3.750   0.5429   0.00989   0.00354  -0.0256   0.5007   1.0000
   4.000   0.5658   0.01034   0.00376  -0.0246   0.4356   1.0000
   4.250   0.5872   0.01103   0.00404  -0.0235   0.3381   1.0000
   4.500   0.6037   0.01265   0.00465  -0.0221   0.1496   1.0000
   4.750   0.6237   0.01406   0.00562  -0.0209   0.0890   1.0000
   5.000   0.6462   0.01497   0.00648  -0.0200   0.0768   1.0000
   5.250   0.6704   0.01561   0.00718  -0.0193   0.0695   1.0000
   5.500   0.6920   0.01681   0.00831  -0.0183   0.0640   1.0000
   5.750   0.7163   0.01755   0.00916  -0.0175   0.0596   1.0000
   6.000   0.7401   0.01845   0.01008  -0.0168   0.0553   1.0000
   6.250   0.7622   0.02045   0.01203  -0.0159   0.0514   1.0000
   6.500   0.7873   0.02117   0.01295  -0.0152   0.0481   1.0000
   6.750   0.8117   0.02252   0.01444  -0.0144   0.0453   1.0000
   7.000   0.8354   0.02403   0.01605  -0.0137   0.0428   1.0000
   7.250   0.8563   0.02726   0.01949  -0.0129   0.0400   1.0000
   7.500   0.8785   0.02881   0.02142  -0.0118   0.0385   1.0000
   7.750   0.8982   0.03150   0.02454  -0.0104   0.0377   1.0000
   8.000   0.9148   0.03485   0.02835  -0.0088   0.0374   1.0000
   8.250   0.9287   0.03813   0.03209  -0.0071   0.0364   1.0000
   8.500   0.9386   0.04202   0.03647  -0.0053   0.0360   1.0000
   8.750   0.9409   0.04755   0.04247  -0.0034   0.0378   1.0000
  10.750   0.7903   0.11245   0.10898  -0.0295   0.0604   1.0000
  11.000   0.7841   0.11851   0.11502  -0.0332   0.0581   1.0000
  11.250   0.7830   0.12330   0.11979  -0.0354   0.0562   1.0000
  11.500   0.7859   0.12710   0.12358  -0.0362   0.0545   1.0000
  11.750   0.7937   0.12973   0.12623  -0.0353   0.0531   1.0000
  12.000   0.6220   0.12861   0.12520  -0.0233   0.0608   1.0000
  12.250   0.6156   0.13278   0.12937  -0.0253   0.0582   1.0000
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