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NACA 0010-34 (naca001034-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA 0010-34 (naca001034-il)
Reynolds number: 1,000,000
Max Cl/Cd: 51.14 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca001034-il-1000000-n5.txt
Download as CSV file: xf-naca001034-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.6704   0.08564   0.08413  -0.0192   1.0000   0.0019
 -10.250  -0.6855   0.07448   0.07293  -0.0300   1.0000   0.0018
 -10.000  -0.7052   0.06779   0.06614  -0.0346   1.0000   0.0018
  -9.750  -0.7254   0.06288   0.06114  -0.0356   1.0000   0.0018
  -9.500  -0.7452   0.05894   0.05708  -0.0337   1.0000   0.0018
  -9.250  -0.7617   0.05491   0.05291  -0.0308   1.0000   0.0018
  -9.000  -0.7744   0.05037   0.04818  -0.0277   1.0000   0.0017
  -8.750  -0.7858   0.04587   0.04346  -0.0236   1.0000   0.0017
  -8.500  -0.7975   0.04126   0.03861  -0.0184   1.0000   0.0017
  -8.250  -0.8062   0.03452   0.03142  -0.0134   0.9987   0.0017
  -8.000  -0.8156   0.02046   0.01602  -0.0071   0.9951   0.0020
  -7.750  -0.7932   0.01831   0.01355  -0.0066   0.9930   0.0021
  -7.500  -0.7679   0.01692   0.01195  -0.0066   0.9910   0.0022
  -7.250  -0.7408   0.01586   0.01072  -0.0069   0.9893   0.0023
  -7.000  -0.7151   0.01442   0.00908  -0.0070   0.9878   0.0024
  -6.750  -0.6879   0.01351   0.00806  -0.0074   0.9861   0.0026
  -6.500  -0.6613   0.01309   0.00759  -0.0076   0.9828   0.0029
  -6.250  -0.6335   0.01261   0.00707  -0.0080   0.9800   0.0031
  -6.000  -0.6056   0.01202   0.00640  -0.0084   0.9777   0.0035
  -5.750  -0.5804   0.01142   0.00570  -0.0081   0.9737   0.0037
  -5.500  -0.5545   0.01093   0.00514  -0.0080   0.9694   0.0040
  -5.250  -0.5280   0.01033   0.00445  -0.0079   0.9660   0.0045
  -5.000  -0.5032   0.00998   0.00407  -0.0075   0.9606   0.0050
  -4.750  -0.4768   0.00967   0.00371  -0.0074   0.9558   0.0057
  -4.500  -0.4501   0.00943   0.00344  -0.0074   0.9512   0.0065
  -4.250  -0.4251   0.00908   0.00303  -0.0069   0.9449   0.0083
  -4.000  -0.3987   0.00884   0.00276  -0.0067   0.9395   0.0103
  -3.750  -0.3726   0.00863   0.00251  -0.0065   0.9331   0.0146
  -3.500  -0.3469   0.00837   0.00229  -0.0062   0.9266   0.0288
  -3.250  -0.3217   0.00806   0.00207  -0.0059   0.9194   0.0581
  -3.000  -0.2989   0.00748   0.00179  -0.0052   0.9117   0.1474
  -2.750  -0.2774   0.00678   0.00152  -0.0043   0.9028   0.2689
  -2.500  -0.2563   0.00608   0.00129  -0.0034   0.8938   0.4012
  -2.250  -0.2344   0.00555   0.00112  -0.0025   0.8839   0.5076
  -2.000  -0.2095   0.00530   0.00102  -0.0020   0.8733   0.5603
  -1.750  -0.1847   0.00508   0.00092  -0.0015   0.8626   0.6117
  -1.500  -0.1598   0.00488   0.00085  -0.0010   0.8523   0.6609
  -1.250  -0.1338   0.00477   0.00080  -0.0007   0.8421   0.6913
  -1.000  -0.1072   0.00471   0.00076  -0.0005   0.8314   0.7146
  -0.750  -0.0807   0.00464   0.00073  -0.0003   0.8205   0.7374
  -0.500  -0.0539   0.00461   0.00071  -0.0002   0.8093   0.7559
  -0.250  -0.0269   0.00459   0.00070  -0.0001   0.7975   0.7713
   0.000   0.0000   0.00459   0.00070   0.0000   0.7849   0.7851
   0.250   0.0270   0.00459   0.00070   0.0001   0.7713   0.7976
   0.500   0.0540   0.00461   0.00071   0.0002   0.7560   0.8094
   0.750   0.0808   0.00464   0.00073   0.0003   0.7372   0.8206
   1.000   0.1073   0.00471   0.00076   0.0005   0.7147   0.8314
   1.250   0.1340   0.00477   0.00080   0.0007   0.6915   0.8421
   1.500   0.1600   0.00488   0.00085   0.0010   0.6615   0.8524
   1.750   0.1848   0.00507   0.00092   0.0015   0.6121   0.8626
   2.000   0.2097   0.00530   0.00101   0.0020   0.5610   0.8732
   2.500   0.2568   0.00605   0.00129   0.0033   0.4073   0.8937
   2.750   0.2777   0.00677   0.00152   0.0043   0.2709   0.9027
   3.000   0.2989   0.00749   0.00179   0.0051   0.1447   0.9117
   3.250   0.3218   0.00807   0.00207   0.0058   0.0568   0.9194
   3.500   0.3471   0.00837   0.00229   0.0062   0.0288   0.9266
   3.750   0.3728   0.00863   0.00251   0.0064   0.0145   0.9331
   4.000   0.3989   0.00884   0.00276   0.0067   0.0103   0.9396
   4.250   0.4252   0.00908   0.00303   0.0069   0.0084   0.9450
   4.500   0.4502   0.00944   0.00344   0.0073   0.0065   0.9511
   4.750   0.4770   0.00967   0.00370   0.0074   0.0057   0.9557
   5.000   0.5034   0.00998   0.00407   0.0075   0.0050   0.9605
   5.250   0.5283   0.01033   0.00445   0.0079   0.0045   0.9659
   5.500   0.5546   0.01094   0.00515   0.0079   0.0040   0.9694
   5.750   0.5805   0.01142   0.00570   0.0081   0.0037   0.9739
   6.000   0.6058   0.01203   0.00641   0.0083   0.0034   0.9778
   6.250   0.6336   0.01262   0.00707   0.0080   0.0031   0.9801
   6.500   0.6616   0.01306   0.00756   0.0075   0.0028   0.9828
   6.750   0.6882   0.01350   0.00805   0.0073   0.0026   0.9861
   7.000   0.7153   0.01443   0.00909   0.0070   0.0024   0.9878
   7.250   0.7411   0.01584   0.01070   0.0068   0.0023   0.9894
   7.500   0.7681   0.01692   0.01194   0.0065   0.0022   0.9911
   7.750   0.7934   0.01832   0.01356   0.0065   0.0021   0.9931
   8.000   0.8159   0.02047   0.01603   0.0071   0.0020   0.9951
   8.500   0.7973   0.04128   0.03863   0.0184   0.0017   1.0000
   8.750   0.7858   0.04590   0.04350   0.0236   0.0017   1.0000
   9.000   0.7745   0.05041   0.04822   0.0276   0.0017   1.0000
   9.250   0.7619   0.05496   0.05296   0.0307   0.0018   1.0000
   9.500   0.7455   0.05901   0.05715   0.0336   0.0018   1.0000
   9.750   0.7255   0.06301   0.06126   0.0354   0.0018   1.0000
  10.000   0.7055   0.06792   0.06628   0.0344   0.0018   1.0000
  10.250   0.6859   0.07470   0.07315   0.0296   0.0018   1.0000
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