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NACA 63012A AIRFOIL (n63012a-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 63012A AIRFOIL (n63012a-il)
Reynolds number: 100,000
Max Cl/Cd: 41.25 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63012a-il-100000.txt
Download as CSV file: xf-n63012a-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63012A AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.6451   0.10393   0.09898   0.0008   1.0000   0.1695
 -10.000  -0.7116   0.06091   0.05593  -0.0347   1.0000   0.0882
  -9.750  -0.7390   0.05556   0.05044  -0.0336   1.0000   0.0844
  -9.500  -0.7820   0.04933   0.04370  -0.0300   1.0000   0.0758
  -9.250  -0.8475   0.05545   0.04925  -0.0237   1.0000   0.0739
  -9.000  -0.8659   0.05060   0.04324  -0.0191   1.0000   0.0659
  -8.750  -0.8555   0.04768   0.03993  -0.0173   1.0000   0.0656
  -8.500  -0.8453   0.04389   0.03581  -0.0156   1.0000   0.0658
  -8.250  -0.8314   0.04062   0.03218  -0.0140   1.0000   0.0659
  -8.000  -0.8145   0.03794   0.02913  -0.0124   1.0000   0.0656
  -7.750  -0.7959   0.03474   0.02560  -0.0112   1.0000   0.0659
  -7.500  -0.7740   0.03158   0.02225  -0.0104   1.0000   0.0670
  -7.250  -0.7514   0.02947   0.02008  -0.0097   1.0000   0.0698
  -7.000  -0.7290   0.02793   0.01843  -0.0087   1.0000   0.0740
  -6.750  -0.7053   0.02642   0.01669  -0.0076   1.0000   0.0775
  -6.500  -0.6811   0.02436   0.01475  -0.0069   1.0000   0.0822
  -6.250  -0.6602   0.02319   0.01353  -0.0056   1.0000   0.0900
  -6.000  -0.6425   0.02176   0.01228  -0.0039   1.0000   0.1002
  -5.750  -0.6277   0.02048   0.01113  -0.0017   1.0000   0.1139
  -5.500  -0.6162   0.01916   0.01003   0.0010   1.0000   0.1381
  -5.250  -0.6115   0.01722   0.00876   0.0045   1.0000   0.2086
  -5.000  -0.6175   0.01497   0.00823   0.0099   1.0000   0.4738
  -4.750  -0.6054   0.01487   0.00841   0.0135   1.0000   0.5696
  -4.500  -0.5896   0.01502   0.00858   0.0163   1.0000   0.6188
  -4.250  -0.5728   0.01525   0.00882   0.0191   1.0000   0.6545
  -4.000  -0.5558   0.01551   0.00907   0.0217   1.0000   0.6846
  -3.750  -0.5391   0.01574   0.00925   0.0243   1.0000   0.7119
  -3.500  -0.5221   0.01607   0.00960   0.0272   1.0000   0.7353
  -3.250  -0.5062   0.01632   0.00983   0.0300   1.0000   0.7594
  -3.000  -0.4900   0.01662   0.01013   0.0330   1.0000   0.7808
  -2.750  -0.4743   0.01676   0.01022   0.0355   1.0000   0.8022
  -2.500  -0.4580   0.01691   0.01033   0.0379   1.0000   0.8201
  -2.250  -0.4413   0.01701   0.01040   0.0400   1.0000   0.8368
  -2.000  -0.4049   0.01724   0.01054   0.0382   0.9942   0.8522
  -1.750  -0.3622   0.01744   0.01062   0.0350   0.9854   0.8660
  -1.500  -0.3200   0.01761   0.01071   0.0318   0.9764   0.8791
  -1.250  -0.2686   0.01791   0.01091   0.0271   0.9700   0.8891
  -1.000  -0.2248   0.01805   0.01099   0.0237   0.9607   0.9001
  -0.750  -0.1739   0.01821   0.01108   0.0187   0.9538   0.9110
  -0.500  -0.1188   0.01839   0.01121   0.0132   0.9462   0.9188
  -0.250  -0.0581   0.01847   0.01126   0.0063   0.9411   0.9269
   0.000   0.0000   0.01853   0.01132   0.0000   0.9333   0.9333
   0.250   0.0580   0.01847   0.01126  -0.0063   0.9269   0.9411
   0.500   0.1188   0.01838   0.01121  -0.0132   0.9188   0.9462
   0.750   0.1738   0.01821   0.01107  -0.0187   0.9110   0.9538
   1.000   0.2248   0.01805   0.01099  -0.0237   0.9001   0.9607
   1.250   0.2686   0.01790   0.01091  -0.0271   0.8891   0.9700
   1.500   0.3200   0.01761   0.01070  -0.0318   0.8791   0.9765
   1.750   0.3621   0.01743   0.01062  -0.0350   0.8660   0.9855
   2.000   0.4048   0.01724   0.01054  -0.0382   0.8522   0.9942
   2.250   0.4409   0.01701   0.01039  -0.0400   0.8368   1.0000
   2.500   0.4576   0.01691   0.01033  -0.0379   0.8201   1.0000
   2.750   0.4739   0.01676   0.01021  -0.0354   0.8023   1.0000
   3.000   0.4896   0.01661   0.01013  -0.0329   0.7809   1.0000
   3.250   0.5058   0.01632   0.00983  -0.0300   0.7595   1.0000
   3.500   0.5217   0.01607   0.00960  -0.0272   0.7355   1.0000
   3.750   0.5386   0.01574   0.00925  -0.0243   0.7120   1.0000
   4.000   0.5553   0.01551   0.00907  -0.0217   0.6848   1.0000
   4.250   0.5723   0.01525   0.00882  -0.0190   0.6548   1.0000
   4.500   0.5891   0.01502   0.00857  -0.0162   0.6192   1.0000
   4.750   0.6049   0.01487   0.00841  -0.0134   0.5702   1.0000
   5.000   0.6171   0.01496   0.00823  -0.0098   0.4753   1.0000
   5.250   0.6113   0.01719   0.00875  -0.0044   0.2099   1.0000
   5.500   0.6160   0.01915   0.01002  -0.0009   0.1383   1.0000
   5.750   0.6275   0.02047   0.01112   0.0017   0.1140   1.0000
   6.000   0.6423   0.02175   0.01227   0.0040   0.1002   1.0000
   6.250   0.6600   0.02318   0.01352   0.0056   0.0901   1.0000
   6.500   0.6809   0.02436   0.01475   0.0070   0.0822   1.0000
   6.750   0.7052   0.02641   0.01668   0.0077   0.0776   1.0000
   7.000   0.7289   0.02793   0.01843   0.0088   0.0740   1.0000
   7.250   0.7514   0.02946   0.02008   0.0097   0.0698   1.0000
   7.500   0.7740   0.03158   0.02225   0.0104   0.0670   1.0000
   7.750   0.7959   0.03474   0.02561   0.0112   0.0659   1.0000
   8.000   0.8146   0.03795   0.02912   0.0124   0.0656   1.0000
   8.250   0.8316   0.04063   0.03219   0.0140   0.0659   1.0000
   8.500   0.8455   0.04389   0.03581   0.0156   0.0658   1.0000
   8.750   0.8558   0.04771   0.03997   0.0172   0.0656   1.0000
   9.000   0.8663   0.05061   0.04326   0.0190   0.0659   1.0000
   9.250   0.8480   0.05552   0.04932   0.0236   0.0739   1.0000
   9.500   0.8552   0.06512   0.05914   0.0239   0.0912   1.0000
   9.750   0.7387   0.05553   0.05042   0.0335   0.0845   1.0000
  10.000   0.7115   0.06090   0.05592   0.0346   0.0882   1.0000
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