Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NPL 9615 AIRFOIL (npl9615-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NPL 9615 AIRFOIL (npl9615-il)
Reynolds number: 1,000,000
Max Cl/Cd: 77.44 at α=8.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-npl9615-il-1000000-n5.txt
Download as CSV file: xf-npl9615-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NPL 9615 AIRFOIL                                
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -15.000  -1.0509   0.06001   0.05733  -0.0395   1.0000   0.0092
 -14.750  -1.0775   0.05198   0.04910  -0.0449   1.0000   0.0092
 -14.500  -1.0943   0.04655   0.04349  -0.0477   1.0000   0.0092
 -14.250  -1.1063   0.04241   0.03918  -0.0489   1.0000   0.0092
 -14.000  -1.1146   0.03915   0.03577  -0.0489   1.0000   0.0092
 -13.750  -1.1200   0.03654   0.03301  -0.0479   1.0000   0.0092
 -13.500  -1.1231   0.03436   0.03068  -0.0463   1.0000   0.0092
 -13.250  -1.1237   0.03255   0.02874  -0.0442   1.0000   0.0092
 -13.000  -1.1227   0.03099   0.02704  -0.0417   1.0000   0.0092
 -12.750  -1.1201   0.02966   0.02558  -0.0389   1.0000   0.0093
 -12.500  -1.1148   0.02855   0.02437  -0.0360   1.0000   0.0093
 -12.250  -1.1049   0.02755   0.02326  -0.0338   1.0000   0.0094
 -12.000  -1.0928   0.02665   0.02226  -0.0318   1.0000   0.0094
 -11.750  -1.0785   0.02592   0.02145  -0.0300   1.0000   0.0095
 -11.500  -1.0660   0.02497   0.02039  -0.0279   1.0000   0.0096
 -11.250  -1.0590   0.02335   0.01857  -0.0251   1.0000   0.0097
 -11.000  -1.0466   0.02225   0.01734  -0.0229   1.0000   0.0098
 -10.750  -1.0325   0.02130   0.01629  -0.0208   1.0000   0.0098
 -10.500  -1.0173   0.02047   0.01536  -0.0188   1.0000   0.0098
 -10.250  -1.0010   0.01975   0.01456  -0.0170   1.0000   0.0098
 -10.000  -0.9842   0.01908   0.01382  -0.0151   1.0000   0.0098
  -9.750  -0.9577   0.01834   0.01300  -0.0154   0.9987   0.0098
  -9.500  -0.9303   0.01763   0.01222  -0.0158   0.9965   0.0098
  -9.250  -0.9031   0.01696   0.01148  -0.0161   0.9932   0.0098
  -9.000  -0.8769   0.01634   0.01080  -0.0162   0.9886   0.0098
  -8.750  -0.8527   0.01577   0.01017  -0.0157   0.9834   0.0098
  -8.500  -0.8259   0.01521   0.00956  -0.0158   0.9774   0.0098
  -8.250  -0.7950   0.01464   0.00894  -0.0168   0.9709   0.0098
  -8.000  -0.7602   0.01407   0.00833  -0.0186   0.9655   0.0098
  -7.750  -0.7233   0.01352   0.00773  -0.0209   0.9592   0.0097
  -7.500  -0.6890   0.01302   0.00718  -0.0225   0.9512   0.0097
  -7.250  -0.6564   0.01256   0.00667  -0.0237   0.9410   0.0097
  -7.000  -0.6276   0.01218   0.00623  -0.0241   0.9269   0.0097
  -6.750  -0.6007   0.01184   0.00582  -0.0240   0.9117   0.0097
  -6.500  -0.5748   0.01154   0.00545  -0.0236   0.8966   0.0097
  -6.250  -0.5496   0.01126   0.00510  -0.0231   0.8835   0.0097
  -6.000  -0.5245   0.01101   0.00478  -0.0225   0.8711   0.0097
  -5.750  -0.4993   0.01076   0.00448  -0.0219   0.8585   0.0097
  -5.500  -0.4741   0.01053   0.00419  -0.0214   0.8407   0.0097
  -5.250  -0.4510   0.01040   0.00391  -0.0203   0.7944   0.0097
  -5.000  -0.4266   0.01025   0.00365  -0.0195   0.7757   0.0098
  -4.750  -0.4014   0.01010   0.00343  -0.0189   0.7616   0.0098
  -4.500  -0.3761   0.00995   0.00322  -0.0184   0.7461   0.0098
  -4.250  -0.3509   0.00983   0.00301  -0.0178   0.7227   0.0098
  -4.000  -0.3263   0.00976   0.00282  -0.0170   0.6920   0.0098
  -3.750  -0.3015   0.00969   0.00265  -0.0164   0.6637   0.0099
  -3.500  -0.2761   0.00961   0.00248  -0.0158   0.6404   0.0099
  -3.250  -0.2503   0.00953   0.00233  -0.0153   0.6208   0.0099
  -3.000  -0.2243   0.00945   0.00219  -0.0149   0.6060   0.0100
  -2.750  -0.1980   0.00936   0.00206  -0.0145   0.5924   0.0101
  -2.500  -0.1717   0.00928   0.00194  -0.0141   0.5796   0.0102
  -2.250  -0.1456   0.00923   0.00183  -0.0137   0.5624   0.0103
  -2.000  -0.1196   0.00920   0.00173  -0.0133   0.5428   0.0105
  -1.750  -0.0935   0.00917   0.00164  -0.0129   0.5243   0.0108
  -1.500  -0.0672   0.00910   0.00155  -0.0125   0.5120   0.0116
  -1.250  -0.0410   0.00904   0.00147  -0.0121   0.4986   0.0170
  -1.000  -0.0144   0.00901   0.00142  -0.0118   0.4854   0.0189
  -0.750   0.0117   0.00896   0.00137  -0.0114   0.4709   0.0278
  -0.500   0.0376   0.00899   0.00135  -0.0110   0.4466   0.0318
  -0.250   0.0632   0.00906   0.00133  -0.0105   0.4161   0.0351
   0.000   0.0888   0.00912   0.00132  -0.0100   0.3890   0.0398
   0.250   0.1147   0.00914   0.00132  -0.0096   0.3678   0.0519
   0.500   0.1325   0.00834   0.00119  -0.0079   0.3503   0.2978
   0.750   0.1481   0.00754   0.00112  -0.0057   0.3315   0.5317
   1.000   0.1678   0.00728   0.00115  -0.0041   0.3063   0.6390
   1.250   0.1901   0.00720   0.00120  -0.0029   0.2862   0.6973
   1.500   0.2124   0.00717   0.00126  -0.0017   0.2610   0.7529
   1.750   0.2362   0.00716   0.00132  -0.0007   0.2461   0.7851
   2.000   0.2593   0.00713   0.00143   0.0004   0.2363   0.8322
   2.250   0.2851   0.00718   0.00151   0.0010   0.2239   0.8614
   2.500   0.3111   0.00731   0.00160   0.0014   0.2071   0.8760
   2.750   0.3378   0.00743   0.00169   0.0017   0.1940   0.8862
   3.000   0.3649   0.00756   0.00179   0.0019   0.1801   0.8954
   3.250   0.3918   0.00771   0.00190   0.0021   0.1661   0.9055
   3.500   0.4197   0.00788   0.00203   0.0021   0.1496   0.9156
   3.750   0.4472   0.00806   0.00216   0.0021   0.1367   0.9267
   4.000   0.4777   0.00823   0.00230   0.0015   0.1271   0.9353
   4.250   0.5080   0.00838   0.00244   0.0009   0.1201   0.9441
   4.750   0.5742   0.00881   0.00278  -0.0016   0.1030   0.9553
   5.000   0.6084   0.00904   0.00297  -0.0032   0.0960   0.9596
   5.250   0.6417   0.00923   0.00313  -0.0045   0.0897   0.9637
   5.500   0.6735   0.00944   0.00330  -0.0055   0.0836   0.9677
   5.750   0.7070   0.00962   0.00348  -0.0068   0.0807   0.9704
   6.000   0.7397   0.00982   0.00367  -0.0080   0.0739   0.9743
   6.250   0.7701   0.01006   0.00391  -0.0087   0.0677   0.9796
   6.500   0.7974   0.01039   0.00420  -0.0088   0.0582   0.9848
   6.750   0.8247   0.01075   0.00449  -0.0089   0.0472   0.9885
   7.000   0.8546   0.01116   0.00483  -0.0096   0.0357   0.9913
   7.250   0.8840   0.01156   0.00518  -0.0103   0.0281   0.9940
   7.500   0.9138   0.01190   0.00551  -0.0110   0.0243   0.9962
   7.750   0.9433   0.01224   0.00584  -0.0117   0.0219   0.9979
   8.000   0.9730   0.01257   0.00617  -0.0124   0.0203   0.9992
   8.250   1.0005   0.01292   0.00652  -0.0126   0.0191   1.0000
   8.500   1.0208   0.01322   0.00684  -0.0112   0.0181   1.0000
   8.750   1.0407   0.01355   0.00718  -0.0098   0.0172   1.0000
   9.000   1.0607   0.01389   0.00753  -0.0085   0.0165   1.0000
   9.250   1.0808   0.01421   0.00788  -0.0071   0.0158   1.0000
   9.500   1.1005   0.01457   0.00826  -0.0057   0.0152   1.0000
   9.750   1.1198   0.01498   0.00868  -0.0043   0.0146   1.0000
  10.000   1.1393   0.01536   0.00909  -0.0029   0.0141   1.0000
  10.250   1.1590   0.01573   0.00950  -0.0016   0.0136   1.0000
  10.500   1.1783   0.01613   0.00994  -0.0002   0.0131   1.0000
  10.750   1.1970   0.01656   0.01040   0.0013   0.0127   1.0000
  11.000   1.2148   0.01704   0.01090   0.0028   0.0123   1.0000
  11.250   1.2323   0.01752   0.01142   0.0044   0.0120   1.0000
  11.500   1.2499   0.01798   0.01193   0.0060   0.0117   1.0000
  11.750   1.2669   0.01848   0.01247   0.0076   0.0114   1.0000
  12.000   1.2833   0.01899   0.01304   0.0093   0.0112   1.0000
  12.250   1.2993   0.01954   0.01363   0.0109   0.0109   1.0000
  12.500   1.3136   0.02010   0.01424   0.0129   0.0107   1.0000
  12.750   1.3263   0.02070   0.01488   0.0150   0.0106   1.0000
  13.000   1.3382   0.02134   0.01558   0.0172   0.0104   1.0000
  13.250   1.3497   0.02207   0.01636   0.0192   0.0103   1.0000
  13.500   1.3601   0.02289   0.01724   0.0212   0.0101   1.0000
  13.750   1.3692   0.02385   0.01826   0.0232   0.0100   1.0000
  14.000   1.3785   0.02484   0.01933   0.0249   0.0099   1.0000
  14.250   1.3873   0.02592   0.02049   0.0265   0.0098   1.0000
  14.500   1.3952   0.02714   0.02178   0.0279   0.0097   1.0000
  14.750   1.4021   0.02850   0.02323   0.0291   0.0097   1.0000
  15.000   1.4084   0.03001   0.02483   0.0301   0.0096   1.0000
  15.250   1.4133   0.03175   0.02666   0.0308   0.0095   1.0000
  15.500   1.4174   0.03371   0.02870   0.0313   0.0094   1.0000
  15.750   1.4199   0.03593   0.03102   0.0314   0.0093   1.0000
  16.000   1.4213   0.03843   0.03362   0.0313   0.0092   1.0000
  16.250   1.4208   0.04126   0.03656   0.0309   0.0092   1.0000
  16.500   1.4189   0.04442   0.03982   0.0301   0.0091   1.0000
  16.750   1.4145   0.04804   0.04355   0.0289   0.0090   1.0000
  17.000   1.4081   0.05204   0.04766   0.0275   0.0090   1.0000
  17.250   1.3990   0.05657   0.05232   0.0257   0.0089   1.0000
  17.500   1.3866   0.06167   0.05753   0.0235   0.0089   1.0000
  17.750   1.3717   0.06734   0.06334   0.0210   0.0089   1.0000
  18.000   1.3525   0.07380   0.06993   0.0180   0.0088   1.0000
  18.250   1.3299   0.08102   0.07729   0.0146   0.0088   1.0000
  18.500   1.3043   0.08891   0.08532   0.0108   0.0089   1.0000
  18.750   1.2773   0.09728   0.09383   0.0068   0.0089   1.0000
  19.000   1.2474   0.10632   0.10302   0.0023   0.0089   1.0000
<< Back to NPL 9615 AIRFOIL (npl9615-il)

Polar data table (+)

Polar graphs


<< Back to NPL 9615 AIRFOIL (npl9615-il)