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NACA 66 (p51htip-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 66 (p51htip-il)
Reynolds number: 50,000
Max Cl/Cd: 22.24 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-p51htip-il-50000.txt
Download as CSV file: xf-p51htip-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66                                         
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.4477   0.12821   0.12105  -0.0272   1.0000   0.2497
 -11.250  -0.4662   0.12740   0.12035  -0.0288   1.0000   0.2605
 -11.000  -0.4499   0.12299   0.11594  -0.0282   1.0000   0.2745
 -10.750  -0.4440   0.11971   0.11269  -0.0279   1.0000   0.2893
 -10.500  -0.4433   0.11695   0.10998  -0.0277   1.0000   0.3047
 -10.250  -0.4449   0.11442   0.10752  -0.0274   1.0000   0.3202
 -10.000  -0.4215   0.10966   0.10274  -0.0263   1.0000   0.3384
  -9.750  -0.4113   0.10630   0.09940  -0.0254   1.0000   0.3561
  -9.500  -0.4055   0.10346   0.09658  -0.0245   1.0000   0.3734
  -9.250  -0.4096   0.10144   0.09464  -0.0231   1.0000   0.3933
  -9.000  -0.4136   0.09915   0.09245  -0.0213   1.0000   0.4114
  -7.000  -0.5214   0.08095   0.07499  -0.0018   1.0000   0.4222
  -6.500  -0.6485   0.07372   0.06813   0.0012   1.0000   0.3625
  -6.250  -0.7096   0.06240   0.05605  -0.0086   1.0000   0.2375
  -6.000  -0.7100   0.05537   0.04774  -0.0096   1.0000   0.1696
  -5.750  -0.6976   0.05111   0.04299  -0.0081   1.0000   0.1521
  -5.500  -0.6840   0.04745   0.03863  -0.0061   1.0000   0.1389
  -5.250  -0.6682   0.04414   0.03472  -0.0043   1.0000   0.1309
  -5.000  -0.6493   0.04180   0.03171  -0.0024   1.0000   0.1261
  -4.750  -0.6296   0.03953   0.02899  -0.0009   1.0000   0.1262
  -4.500  -0.6085   0.03763   0.02660   0.0005   1.0000   0.1274
  -4.250  -0.5851   0.03566   0.02427   0.0017   1.0000   0.1277
  -4.000  -0.5604   0.03363   0.02204   0.0025   1.0000   0.1290
  -3.750  -0.5360   0.03205   0.02043   0.0034   1.0000   0.1345
  -3.500  -0.1706   0.03100   0.02206  -0.0347   1.0000   1.0000
  -3.250  -0.1666   0.03072   0.02155  -0.0322   1.0000   1.0000
  -3.000  -0.1619   0.03046   0.02108  -0.0297   1.0000   1.0000
  -2.750  -0.1568   0.03023   0.02066  -0.0272   1.0000   1.0000
  -2.500  -0.1511   0.03003   0.02028  -0.0247   1.0000   1.0000
  -2.250  -0.1452   0.02984   0.01991  -0.0223   1.0000   1.0000
  -2.000  -0.1390   0.02968   0.01959  -0.0198   1.0000   1.0000
  -1.750  -0.1327   0.02953   0.01929  -0.0173   1.0000   1.0000
  -1.500  -0.1262   0.02939   0.01902  -0.0148   1.0000   1.0000
  -1.250  -0.1196   0.02926   0.01877  -0.0123   1.0000   1.0000
  -1.000  -0.1131   0.02914   0.01854  -0.0098   1.0000   1.0000
  -0.750  -0.1065   0.02903   0.01831  -0.0072   1.0000   1.0000
  -0.500  -0.1000   0.02892   0.01810  -0.0046   1.0000   1.0000
  -0.250  -0.0935   0.02882   0.01792  -0.0020   1.0000   1.0000
   0.000  -0.0873   0.02872   0.01774   0.0006   1.0000   1.0000
   0.250  -0.0810   0.02863   0.01758   0.0032   1.0000   1.0000
   0.500  -0.0749   0.02854   0.01743   0.0059   1.0000   1.0000
   0.750  -0.0687   0.02846   0.01729   0.0085   1.0000   1.0000
   1.000  -0.0622   0.02840   0.01717   0.0111   1.0000   1.0000
   1.250  -0.0553   0.02837   0.01710   0.0136   1.0000   1.0000
   1.500  -0.0475   0.02839   0.01707   0.0159   1.0000   1.0000
   1.750  -0.0387   0.02845   0.01709   0.0180   1.0000   1.0000
   2.000  -0.0285   0.02860   0.01719   0.0199   1.0000   1.0000
   2.250  -0.0168   0.02883   0.01738   0.0214   1.0000   1.0000
   2.500  -0.0037   0.02913   0.01767   0.0226   1.0000   1.0000
   2.750   0.0104   0.02952   0.01805   0.0236   1.0000   1.0000
   3.000   0.0253   0.02998   0.01851   0.0244   1.0000   1.0000
   3.250   0.0406   0.03051   0.01905   0.0251   1.0000   1.0000
   3.500   0.0561   0.03111   0.01967   0.0256   1.0000   1.0000
   3.750   0.0719   0.03177   0.02037   0.0260   1.0000   1.0000
   4.000   0.0877   0.03249   0.02114   0.0263   1.0000   1.0000
   4.250   0.1034   0.03328   0.02201   0.0266   1.0000   1.0000
   4.500   0.1189   0.03414   0.02294   0.0268   1.0000   1.0000
   4.750   0.1341   0.03507   0.02396   0.0268   1.0000   1.0000
   5.000   0.1491   0.03609   0.02507   0.0268   1.0000   1.0000
   5.250   0.1638   0.03718   0.02627   0.0267   1.0000   1.0000
   5.500   0.1780   0.03836   0.02760   0.0266   1.0000   1.0000
   5.750   0.1917   0.03965   0.02901   0.0263   1.0000   1.0000
   6.000   0.3837   0.04587   0.03624  -0.0018   0.8556   1.0000
   6.250   0.6126   0.02781   0.01692   0.0050   0.2175   1.0000
   6.500   0.6185   0.02964   0.01826   0.0084   0.1779   1.0000
   6.750   0.6564   0.03132   0.01970   0.0081   0.1509   1.0000
   7.000   0.7828   0.03520   0.02360  -0.0048   0.1252   1.0000
   7.250   0.8201   0.03755   0.02614  -0.0058   0.1177   1.0000
   7.500   0.8571   0.04102   0.02969  -0.0071   0.1142   1.0000
   7.750   0.8805   0.04393   0.03294  -0.0059   0.1141   1.0000
   8.000   0.8982   0.04667   0.03612  -0.0040   0.1146   1.0000
   8.250   0.9122   0.04966   0.03951  -0.0018   0.1147   1.0000
   8.500   0.9228   0.05284   0.04305   0.0005   0.1145   1.0000
   8.750   0.9166   0.05505   0.04604   0.0053   0.1176   1.0000
   9.000   0.9126   0.05858   0.05008   0.0089   0.1211   1.0000
   9.250   0.9138   0.06267   0.05447   0.0113   0.1253   1.0000
   9.500   0.9061   0.06639   0.05861   0.0144   0.1314   1.0000
   9.750   0.8838   0.07059   0.06316   0.0177   0.1376   1.0000
  10.000   0.8682   0.07485   0.06767   0.0201   0.1454   1.0000
  10.250   0.8270   0.07863   0.07159   0.0230   0.1487   1.0000
  10.500   0.8081   0.08432   0.07739   0.0229   0.1589   1.0000
  10.750   0.7476   0.09165   0.08470   0.0196   0.1639   1.0000
  11.000   0.6191   0.09395   0.08736   0.0211   0.1697   1.0000
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