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NACA 66(1)-212 (naca661212-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 66(1)-212 (naca661212-il)
Reynolds number: 50,000
Max Cl/Cd: 22.02 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca661212-il-50000.txt
Download as CSV file: xf-naca661212-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66(1)-212                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4323   0.11298   0.10575  -0.0249   1.0000   0.3357
 -10.000  -0.4362   0.11079   0.10364  -0.0241   1.0000   0.3543
  -9.750  -0.4467   0.10913   0.10207  -0.0230   1.0000   0.3709
  -9.500  -0.4186   0.10422   0.09713  -0.0220   1.0000   0.3901
  -9.250  -0.4106   0.10119   0.09413  -0.0209   1.0000   0.4077
  -9.000  -0.4071   0.09856   0.09155  -0.0195   1.0000   0.4256
  -8.750  -0.4193   0.09710   0.09021  -0.0171   1.0000   0.4461
  -8.500  -0.4123   0.09401   0.08713  -0.0157   1.0000   0.4615
  -7.500  -0.5260   0.08060   0.07440  -0.0141   1.0000   0.3618
  -7.250  -0.5802   0.07794   0.07195  -0.0110   1.0000   0.3462
  -7.000  -0.6321   0.07527   0.06945  -0.0074   1.0000   0.3435
  -6.500  -0.7203   0.05813   0.05087  -0.0150   1.0000   0.1744
  -6.250  -0.7145   0.05380   0.04603  -0.0133   1.0000   0.1572
  -6.000  -0.7059   0.05021   0.04178  -0.0113   1.0000   0.1458
  -5.750  -0.6930   0.04678   0.03794  -0.0096   1.0000   0.1399
  -5.500  -0.6770   0.04399   0.03438  -0.0076   1.0000   0.1322
  -5.250  -0.6585   0.04114   0.03116  -0.0062   1.0000   0.1286
  -5.000  -0.6377   0.03861   0.02816  -0.0048   1.0000   0.1257
  -4.750  -0.6156   0.03652   0.02567  -0.0035   1.0000   0.1255
  -4.500  -0.5934   0.03482   0.02364  -0.0023   1.0000   0.1293
  -4.250  -0.5694   0.03325   0.02172  -0.0012   1.0000   0.1317
  -4.000  -0.5441   0.03188   0.02001   0.0000   1.0000   0.1339
  -3.750  -0.1493   0.03122   0.02198  -0.0384   1.0000   1.0000
  -3.500  -0.1464   0.03094   0.02151  -0.0358   1.0000   1.0000
  -3.250  -0.1427   0.03069   0.02109  -0.0333   1.0000   1.0000
  -3.000  -0.1384   0.03046   0.02070  -0.0308   1.0000   1.0000
  -2.750  -0.1335   0.03026   0.02034  -0.0282   1.0000   1.0000
  -2.500  -0.1282   0.03008   0.02002  -0.0258   1.0000   1.0000
  -2.250  -0.1225   0.02992   0.01970  -0.0233   1.0000   1.0000
  -2.000  -0.1166   0.02977   0.01943  -0.0208   1.0000   1.0000
  -1.750  -0.1103   0.02964   0.01918  -0.0183   1.0000   1.0000
  -1.500  -0.1039   0.02952   0.01895  -0.0159   1.0000   1.0000
  -1.250  -0.0974   0.02941   0.01874  -0.0134   1.0000   1.0000
  -1.000  -0.0908   0.02930   0.01855  -0.0109   1.0000   1.0000
  -0.750  -0.0841   0.02921   0.01838  -0.0085   1.0000   1.0000
  -0.500  -0.0776   0.02912   0.01820  -0.0060   1.0000   1.0000
  -0.250  -0.0711   0.02904   0.01805  -0.0034   1.0000   1.0000
   0.000  -0.0647   0.02896   0.01792  -0.0009   1.0000   1.0000
   0.250  -0.0584   0.02888   0.01779   0.0017   1.0000   1.0000
   0.500  -0.0524   0.02880   0.01766   0.0043   1.0000   1.0000
   0.750  -0.0466   0.02872   0.01756   0.0069   1.0000   1.0000
   1.000  -0.0410   0.02864   0.01745   0.0095   1.0000   1.0000
   1.250  -0.0357   0.02857   0.01735   0.0122   1.0000   1.0000
   1.500  -0.0303   0.02849   0.01727   0.0149   1.0000   1.0000
   1.750  -0.0249   0.02843   0.01720   0.0175   1.0000   1.0000
   2.000  -0.0189   0.02840   0.01716   0.0199   1.0000   1.0000
   2.250  -0.0120   0.02842   0.01717   0.0222   1.0000   1.0000
   2.500  -0.0028   0.02854   0.01729   0.0241   1.0000   1.0000
   2.750   0.0086   0.02877   0.01752   0.0254   1.0000   1.0000
   3.000   0.0219   0.02912   0.01788   0.0264   1.0000   1.0000
   3.250   0.0366   0.02956   0.01835   0.0270   1.0000   1.0000
   3.500   0.0521   0.03009   0.01892   0.0274   1.0000   1.0000
   3.750   0.0682   0.03070   0.01958   0.0276   1.0000   1.0000
   4.000   0.0844   0.03139   0.02033   0.0278   1.0000   1.0000
   4.250   0.1007   0.03216   0.02117   0.0278   1.0000   1.0000
   4.500   0.1168   0.03301   0.02212   0.0277   1.0000   1.0000
   4.750   0.1328   0.03394   0.02314   0.0276   1.0000   1.0000
   5.000   0.1484   0.03495   0.02427   0.0273   1.0000   1.0000
   5.250   0.2386   0.03914   0.02888   0.0131   0.9557   1.0000
   5.500   0.2967   0.04120   0.03129   0.0061   0.9171   1.0000
   5.750   0.3499   0.04279   0.03326   0.0009   0.8809   1.0000
   6.000   0.5924   0.02722   0.01628   0.0037   0.2100   1.0000
   6.250   0.6018   0.02883   0.01751   0.0067   0.1747   1.0000
   6.500   0.6934   0.03183   0.02016  -0.0017   0.1385   1.0000
   6.750   0.7730   0.03510   0.02354  -0.0090   0.1236   1.0000
   7.000   0.8061   0.03741   0.02623  -0.0090   0.1206   1.0000
   7.250   0.8282   0.03962   0.02879  -0.0077   0.1169   1.0000
   7.500   0.8489   0.04206   0.03145  -0.0065   0.1138   1.0000
   7.750   0.8648   0.04479   0.03463  -0.0043   0.1145   1.0000
   8.000   0.8762   0.04778   0.03814  -0.0016   0.1169   1.0000
   8.250   0.8856   0.05113   0.04192   0.0010   0.1199   1.0000
   8.500   0.8964   0.05488   0.04597   0.0029   0.1228   1.0000
   8.750   0.9092   0.05889   0.05023   0.0044   0.1253   1.0000
   9.000   0.8877   0.06190   0.05405   0.0097   0.1326   1.0000
   9.250   0.8896   0.06628   0.05864   0.0114   0.1367   1.0000
   9.500   0.8635   0.07018   0.06302   0.0151   0.1454   1.0000
   9.750   0.8546   0.07509   0.06815   0.0165   0.1556   1.0000
  10.000   0.8082   0.07948   0.07270   0.0191   0.1630   1.0000
  10.250   0.7926   0.08889   0.08228   0.0162   0.1966   1.0000
  10.500   0.7100   0.09701   0.09029   0.0096   0.2000   1.0000
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