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NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il)
Reynolds number: 100,000
Max Cl/Cd: 46.71 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651212a06-il-100000.txt
Download as CSV file: xf-naca651212a06-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-212 a=0.6                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4851   0.09496   0.09021  -0.0337   1.0000   0.1349
  -9.500  -0.4912   0.09128   0.08660  -0.0364   1.0000   0.1427
  -9.250  -0.5285   0.08513   0.08059  -0.0451   1.0000   0.1443
  -9.000  -0.4954   0.08349   0.07892  -0.0395   1.0000   0.1566
  -8.750  -0.5330   0.07803   0.07360  -0.0457   1.0000   0.1584
  -8.500  -0.5131   0.07540   0.07100  -0.0432   1.0000   0.1698
  -8.250  -0.5505   0.07277   0.06850  -0.0423   1.0000   0.1717
  -8.000  -0.5822   0.07145   0.06725  -0.0376   1.0000   0.1715
  -7.750  -0.6158   0.07021   0.06602  -0.0329   1.0000   0.1722
  -7.500  -0.6582   0.06917   0.06484  -0.0287   1.0000   0.1741
  -6.000  -0.5979   0.03949   0.03208  -0.0287   0.9808   0.0853
  -5.750  -0.5721   0.03547   0.02766  -0.0290   0.9762   0.0785
  -5.500  -0.5399   0.03284   0.02452  -0.0300   0.9721   0.0779
  -5.250  -0.5106   0.03072   0.02191  -0.0302   0.9674   0.0768
  -5.000  -0.4801   0.02889   0.01969  -0.0305   0.9627   0.0758
  -4.750  -0.4450   0.02725   0.01786  -0.0318   0.9592   0.0768
  -4.500  -0.4120   0.02609   0.01659  -0.0330   0.9554   0.0808
  -4.250  -0.3861   0.02524   0.01562  -0.0327   0.9500   0.0850
  -4.000  -0.3531   0.02404   0.01443  -0.0337   0.9460   0.0887
  -3.750  -0.3192   0.02308   0.01358  -0.0352   0.9427   0.0967
  -3.500  -0.3056   0.02251   0.01305  -0.0330   0.9365   0.1063
  -3.250  -0.2790   0.02178   0.01237  -0.0331   0.9317   0.1257
  -3.000  -0.2709   0.01896   0.01203  -0.0303   0.9276   0.5723
  -2.750  -0.2699   0.01936   0.01284  -0.0233   0.9209   0.7220
  -2.500  -0.2507   0.01977   0.01320  -0.0202   0.9163   0.7831
  -2.250  -0.2284   0.02053   0.01413  -0.0159   0.9128   0.8362
  -2.000  -0.2140   0.02232   0.01592  -0.0092   0.9076   0.8856
  -1.750  -0.1561   0.02335   0.01672  -0.0127   0.9057   0.9207
  -1.500  -0.1088   0.02345   0.01664  -0.0171   0.9025   0.9299
  -1.250  -0.0457   0.02358   0.01658  -0.0246   0.9006   0.9342
  -1.000  -0.0063   0.02366   0.01653  -0.0279   0.8965   0.9426
  -0.750   0.0407   0.02375   0.01653  -0.0327   0.8926   0.9486
  -0.500   0.0834   0.02382   0.01651  -0.0366   0.8887   0.9564
  -0.250   0.1437   0.02384   0.01646  -0.0438   0.8864   0.9603
   0.000   0.2009   0.02385   0.01641  -0.0502   0.8843   0.9653
   0.250   0.2298   0.02406   0.01662  -0.0522   0.8781   0.9753
   0.500   0.2753   0.02415   0.01671  -0.0568   0.8741   0.9830
   0.750   0.3313   0.02409   0.01667  -0.0632   0.8714   0.9880
   1.000   0.3854   0.02402   0.01664  -0.0690   0.8690   0.9935
   1.250   0.3953   0.02451   0.01716  -0.0681   0.8605   1.0000
   1.500   0.4177   0.02467   0.01737  -0.0682   0.8555   1.0000
   1.750   0.3792   0.02541   0.01806  -0.0585   0.8443   1.0000
   2.000   0.4093   0.02557   0.01827  -0.0597   0.8399   1.0000
   2.250   0.3666   0.02626   0.01887  -0.0488   0.8282   1.0000
   2.500   0.4121   0.02645   0.01913  -0.0522   0.8243   1.0000
   2.750   0.3876   0.02714   0.01976  -0.0440   0.8129   1.0000
   3.000   0.4379   0.02730   0.02003  -0.0480   0.8089   1.0000
   3.250   0.4215   0.02806   0.02075  -0.0412   0.7977   1.0000
   3.750   0.4675   0.02888   0.02173  -0.0401   0.7818   1.0000
   4.000   0.4881   0.02929   0.02223  -0.0390   0.7723   1.0000
   4.250   0.5379   0.02905   0.02220  -0.0421   0.7648   1.0000
   4.500   0.6282   0.02571   0.01920  -0.0470   0.7463   1.0000
   4.750   0.7072   0.01894   0.01256  -0.0448   0.6992   1.0000
   5.000   0.7308   0.01742   0.01113  -0.0412   0.6676   1.0000
   5.250   0.7455   0.01596   0.00966  -0.0358   0.5953   1.0000
   5.500   0.7244   0.01733   0.00909  -0.0259   0.2696   1.0000
   5.750   0.7091   0.01995   0.01053  -0.0197   0.1300   1.0000
   6.000   0.7146   0.02131   0.01168  -0.0162   0.1051   1.0000
   6.250   0.7234   0.02254   0.01284  -0.0132   0.0944   1.0000
   6.500   0.7352   0.02385   0.01400  -0.0108   0.0857   1.0000
   6.750   0.7550   0.02500   0.01517  -0.0095   0.0787   1.0000
   7.000   0.7817   0.02652   0.01658  -0.0092   0.0740   1.0000
   7.250   0.8202   0.02893   0.01893  -0.0105   0.0707   1.0000
   7.500   0.8484   0.03045   0.02067  -0.0101   0.0676   1.0000
   7.750   0.8765   0.03232   0.02271  -0.0100   0.0648   1.0000
   8.000   0.9043   0.03471   0.02537  -0.0096   0.0642   1.0000
   8.250   0.9282   0.03743   0.02849  -0.0087   0.0647   1.0000
   8.500   0.9484   0.04057   0.03205  -0.0074   0.0663   1.0000
   8.750   0.9646   0.04391   0.03578  -0.0058   0.0677   1.0000
   9.000   0.9768   0.04728   0.03952  -0.0040   0.0684   1.0000
   9.250   0.9873   0.05109   0.04363  -0.0023   0.0696   1.0000
   9.500   0.9840   0.05590   0.04927   0.0018   0.0806   1.0000
  11.500   0.7208   0.11706   0.11214  -0.0122   0.1745   1.0000
  11.750   0.7390   0.12058   0.11569  -0.0105   0.1686   1.0000
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