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NACA 66 (p51htip-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 66 (p51htip-il)
Reynolds number: 100,000
Max Cl/Cd: 35.37 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-p51htip-il-100000.txt
Download as CSV file: xf-p51htip-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4602   0.09906   0.09424  -0.0455   1.0000   0.1205
 -10.000  -0.4776   0.09511   0.09040  -0.0493   1.0000   0.1257
  -9.750  -0.5193   0.09128   0.08675  -0.0530   1.0000   0.1268
  -9.500  -0.5479   0.08978   0.08533  -0.0498   1.0000   0.1267
  -9.250  -0.5806   0.08865   0.08427  -0.0455   1.0000   0.1267
  -9.000  -0.6123   0.08759   0.08327  -0.0407   1.0000   0.1266
  -8.750  -0.6465   0.08645   0.08214  -0.0358   1.0000   0.1267
  -8.500  -0.6802   0.08501   0.08064  -0.0316   1.0000   0.1270
  -8.250  -0.6469   0.08031   0.07611  -0.0304   1.0000   0.1338
  -8.000  -0.6675   0.07818   0.07396  -0.0271   1.0000   0.1359
  -7.750  -0.6930   0.07593   0.07164  -0.0239   1.0000   0.1390
  -7.500  -0.7463   0.07614   0.07131  -0.0187   1.0000   0.1422
  -7.250  -0.7129   0.07029   0.06592  -0.0182   1.0000   0.1497
  -7.000  -0.7452   0.06866   0.06384  -0.0146   1.0000   0.1575
  -6.750  -0.7268   0.06510   0.06055  -0.0129   1.0000   0.1666
  -6.500  -0.7293   0.06237   0.05777  -0.0104   1.0000   0.1780
  -6.250  -0.7295   0.05994   0.05526  -0.0077   1.0000   0.1937
  -6.000  -0.7281   0.05757   0.05282  -0.0052   1.0000   0.2103
  -5.000  -0.6610   0.03747   0.02931   0.0016   1.0000   0.0818
  -4.750  -0.6411   0.03462   0.02612   0.0031   1.0000   0.0778
  -4.500  -0.6208   0.03268   0.02381   0.0045   1.0000   0.0780
  -4.250  -0.5988   0.03075   0.02150   0.0059   1.0000   0.0769
  -4.000  -0.5760   0.02912   0.01953   0.0072   1.0000   0.0758
  -3.750  -0.5532   0.02775   0.01793   0.0083   1.0000   0.0762
  -3.500  -0.5313   0.02675   0.01681   0.0093   1.0000   0.0800
  -3.250  -0.5057   0.02594   0.01584   0.0098   0.9990   0.0832
  -3.000  -0.4794   0.02489   0.01476   0.0100   0.9977   0.0858
  -2.750  -0.4542   0.02403   0.01402   0.0101   0.9960   0.0918
  -2.500  -0.4280   0.02349   0.01349   0.0101   0.9940   0.1007
  -2.250  -0.4062   0.02292   0.01293   0.0108   0.9925   0.1110
  -2.000  -0.0917   0.02680   0.01971  -0.0289   0.9905   1.0000
  -1.750  -0.0729   0.02673   0.01951  -0.0289   0.9878   1.0000
  -1.500  -0.0520   0.02676   0.01943  -0.0292   0.9853   1.0000
  -1.250  -0.0348   0.02675   0.01931  -0.0287   0.9825   1.0000
  -1.000  -0.0206   0.02673   0.01921  -0.0276   0.9796   1.0000
  -0.750  -0.0028   0.02680   0.01920  -0.0272   0.9765   1.0000
  -0.500   0.0198   0.02699   0.01932  -0.0277   0.9734   1.0000
  -0.250   0.0341   0.02706   0.01934  -0.0266   0.9700   1.0000
   0.000   0.0480   0.02714   0.01938  -0.0253   0.9659   1.0000
   0.250   0.0693   0.02737   0.01957  -0.0255   0.9619   1.0000
   0.500   0.0889   0.02762   0.01979  -0.0253   0.9575   1.0000
   0.750   0.1003   0.02772   0.01987  -0.0236   0.9523   1.0000
   1.000   0.1266   0.02809   0.02023  -0.0246   0.9472   1.0000
   1.250   0.1381   0.02826   0.02040  -0.0228   0.9411   1.0000
   1.500   0.1592   0.02856   0.02070  -0.0228   0.9347   1.0000
   1.750   0.1771   0.02886   0.02102  -0.0221   0.9275   1.0000
   2.000   0.1990   0.02919   0.02137  -0.0222   0.9201   1.0000
   2.250   0.2138   0.02945   0.02167  -0.0208   0.9113   1.0000
   2.500   0.2496   0.02999   0.02226  -0.0232   0.9035   1.0000
   2.750   0.2567   0.03011   0.02242  -0.0204   0.8927   1.0000
   3.000   0.2767   0.03043   0.02280  -0.0197   0.8825   1.0000
   3.250   0.3186   0.03096   0.02343  -0.0228   0.8738   1.0000
   3.500   0.3289   0.03108   0.02360  -0.0203   0.8612   1.0000
   3.750   0.3465   0.03128   0.02390  -0.0189   0.8487   1.0000
   4.000   0.3704   0.03151   0.02423  -0.0185   0.8364   1.0000
   4.250   0.4025   0.03169   0.02454  -0.0193   0.8240   1.0000
   4.500   0.4452   0.03169   0.02476  -0.0215   0.8115   1.0000
   4.750   0.6929   0.01959   0.01027  -0.0355   0.1673   1.0000
   5.000   0.6881   0.02089   0.01109  -0.0301   0.1184   1.0000
   5.250   0.6878   0.02169   0.01179  -0.0253   0.1037   1.0000
   5.500   0.6865   0.02242   0.01251  -0.0204   0.0968   1.0000
   5.750   0.6868   0.02304   0.01316  -0.0157   0.0915   1.0000
   6.000   0.6897   0.02402   0.01400  -0.0115   0.0868   1.0000
   6.250   0.7019   0.02473   0.01477  -0.0087   0.0813   1.0000
   6.500   0.7254   0.02594   0.01589  -0.0079   0.0760   1.0000
   6.750   0.7788   0.02863   0.01852  -0.0120   0.0724   1.0000
   7.000   0.8049   0.03007   0.02019  -0.0113   0.0702   1.0000
   7.250   0.8270   0.03156   0.02187  -0.0101   0.0673   1.0000
   7.500   0.8513   0.03366   0.02424  -0.0092   0.0668   1.0000
   7.750   0.8720   0.03607   0.02698  -0.0076   0.0676   1.0000
   8.000   0.8900   0.03887   0.03015  -0.0057   0.0692   1.0000
   8.250   0.9058   0.04208   0.03367  -0.0038   0.0709   1.0000
   8.500   0.9193   0.04552   0.03737  -0.0018   0.0720   1.0000
   8.750   0.9368   0.05003   0.04206  -0.0009   0.0735   1.0000
  10.500   0.7530   0.09093   0.08615   0.0160   0.1986   1.0000
  10.750   0.6895   0.10023   0.09531   0.0073   0.1935   1.0000
  11.000   0.6982   0.10356   0.09864   0.0083   0.1827   1.0000
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