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

NACA 642-015 AIRFOIL (n64015-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 642-015 AIRFOIL (n64015-il)
Reynolds number: 50,000
Max Cl/Cd: 27.05 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n64015-il-50000-n5.txt
Download as CSV file: xf-n64015-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 642-015 AIRFOIL                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.500  -0.7502   0.10087   0.09324  -0.0223   1.0000   0.0574
 -13.250  -0.7786   0.09252   0.08476  -0.0273   1.0000   0.0571
 -13.000  -0.8031   0.08564   0.07772  -0.0309   1.0000   0.0568
 -12.750  -0.8279   0.07946   0.07134  -0.0337   1.0000   0.0566
 -12.500  -0.8491   0.07424   0.06591  -0.0354   1.0000   0.0567
 -12.250  -0.8673   0.06970   0.06113  -0.0362   1.0000   0.0568
 -12.000  -0.8831   0.06568   0.05682  -0.0364   1.0000   0.0572
 -11.750  -0.8966   0.06209   0.05292  -0.0358   1.0000   0.0577
 -11.500  -0.9071   0.05893   0.04942  -0.0346   1.0000   0.0584
 -11.250  -0.8955   0.05634   0.04680  -0.0343   1.0000   0.0600
 -11.000  -0.8875   0.05392   0.04424  -0.0336   1.0000   0.0614
 -10.750  -0.8778   0.05146   0.04158  -0.0328   1.0000   0.0630
 -10.500  -0.8633   0.04900   0.03887  -0.0322   1.0000   0.0646
 -10.250  -0.8464   0.04671   0.03626  -0.0317   1.0000   0.0671
 -10.000  -0.8228   0.04464   0.03398  -0.0316   1.0000   0.0704
  -9.750  -0.7945   0.04300   0.03234  -0.0317   1.0000   0.0743
  -9.500  -0.7611   0.04155   0.03067  -0.0317   1.0000   0.0790
  -9.250  -0.7318   0.04032   0.02944  -0.0314   1.0000   0.0850
  -9.000  -0.7067   0.03921   0.02824  -0.0307   1.0000   0.0914
  -8.750  -0.6869   0.03800   0.02706  -0.0297   1.0000   0.0981
  -8.500  -0.6714   0.03675   0.02577  -0.0287   1.0000   0.1073
  -8.250  -0.6611   0.03530   0.02446  -0.0274   1.0000   0.1162
  -8.000  -0.6536   0.03379   0.02304  -0.0260   1.0000   0.1280
  -7.750  -0.6501   0.03222   0.02163  -0.0242   1.0000   0.1436
  -7.500  -0.6511   0.03064   0.02026  -0.0217   1.0000   0.1638
  -7.250  -0.6576   0.02911   0.01905  -0.0184   1.0000   0.1905
  -7.000  -0.6676   0.02769   0.01800  -0.0143   1.0000   0.2279
  -6.750  -0.6834   0.02655   0.01732  -0.0088   1.0000   0.2731
  -6.500  -0.6977   0.02580   0.01718  -0.0028   1.0000   0.3405
  -6.250  -0.7003   0.02597   0.01792   0.0028   1.0000   0.4308
  -6.000  -0.6752   0.02629   0.01816   0.0029   0.9867   0.5054
  -5.750  -0.6415   0.02717   0.01889   0.0026   0.9748   0.5529
  -5.500  -0.6056   0.02837   0.01992   0.0025   0.9645   0.5881
  -5.250  -0.5597   0.03017   0.02156   0.0023   0.9577   0.6151
  -5.000  -0.5180   0.03112   0.02229   0.0013   0.9496   0.6355
  -4.750  -0.4769   0.03171   0.02266   0.0001   0.9413   0.6487
  -4.500  -0.4386   0.03172   0.02246  -0.0017   0.9331   0.6602
  -4.250  -0.4156   0.03127   0.02182  -0.0020   0.9210   0.6721
  -4.000  -0.3747   0.03142   0.02178  -0.0037   0.9131   0.6774
  -3.750  -0.3481   0.03101   0.02120  -0.0044   0.9031   0.6862
  -3.500  -0.3209   0.03081   0.02087  -0.0046   0.8929   0.6924
  -3.250  -0.2887   0.03060   0.02050  -0.0057   0.8849   0.6988
  -3.000  -0.2723   0.03019   0.01998  -0.0047   0.8737   0.7069
  -2.750  -0.2423   0.03011   0.01980  -0.0051   0.8655   0.7121
  -2.500  -0.2268   0.02964   0.01921  -0.0040   0.8558   0.7210
  -2.250  -0.1995   0.02960   0.01910  -0.0040   0.8473   0.7255
  -2.000  -0.1737   0.02942   0.01884  -0.0040   0.8397   0.7313
  -1.750  -0.1600   0.02909   0.01843  -0.0026   0.8303   0.7391
  -1.500  -0.1308   0.02902   0.01830  -0.0029   0.8235   0.7434
  -1.250  -0.1111   0.02889   0.01812  -0.0021   0.8154   0.7499
  -1.000  -0.0915   0.02867   0.01786  -0.0015   0.8080   0.7565
  -0.750  -0.0643   0.02865   0.01780  -0.0016   0.8018   0.7613
  -0.500  -0.0471   0.02855   0.01768  -0.0006   0.7936   0.7689
  -0.250  -0.0195   0.02845   0.01755  -0.0007   0.7885   0.7745
   0.000   0.0000   0.02856   0.01769   0.0000   0.7805   0.7805
   0.250   0.0195   0.02845   0.01755   0.0007   0.7745   0.7885
   0.500   0.0471   0.02855   0.01768   0.0006   0.7689   0.7936
   0.750   0.0643   0.02865   0.01780   0.0016   0.7613   0.8018
   1.000   0.0915   0.02867   0.01785   0.0015   0.7565   0.8081
   1.250   0.1111   0.02888   0.01812   0.0021   0.7499   0.8155
   1.500   0.1308   0.02902   0.01830   0.0029   0.7434   0.8235
   1.750   0.1600   0.02908   0.01842   0.0026   0.7391   0.8304
   2.000   0.1737   0.02942   0.01884   0.0040   0.7313   0.8397
   2.250   0.1995   0.02960   0.01910   0.0040   0.7256   0.8474
   2.500   0.2269   0.02964   0.01921   0.0040   0.7211   0.8558
   2.750   0.2423   0.03011   0.01980   0.0051   0.7122   0.8655
   3.000   0.2722   0.03018   0.01998   0.0047   0.7069   0.8737
   3.250   0.2888   0.03060   0.02050   0.0056   0.6988   0.8850
   3.500   0.3209   0.03081   0.02086   0.0046   0.6924   0.8929
   3.750   0.3483   0.03100   0.02119   0.0044   0.6862   0.9031
   4.000   0.3748   0.03141   0.02177   0.0037   0.6774   0.9131
   4.250   0.4157   0.03126   0.02181   0.0020   0.6721   0.9210
   4.500   0.4388   0.03172   0.02245   0.0017   0.6602   0.9331
   4.750   0.4771   0.03169   0.02265  -0.0002   0.6487   0.9414
   5.000   0.5183   0.03111   0.02228  -0.0014   0.6355   0.9497
   5.250   0.5598   0.03017   0.02155  -0.0023   0.6151   0.9578
   5.500   0.6056   0.02836   0.01991  -0.0025   0.5881   0.9646
   5.750   0.6416   0.02716   0.01888  -0.0026   0.5527   0.9749
   6.000   0.6753   0.02628   0.01815  -0.0030   0.5054   0.9867
   6.500   0.6976   0.02579   0.01718   0.0028   0.3410   1.0000
   6.750   0.6833   0.02654   0.01731   0.0088   0.2731   1.0000
   7.000   0.6678   0.02768   0.01799   0.0142   0.2280   1.0000
   7.250   0.6578   0.02911   0.01904   0.0183   0.1902   1.0000
   7.500   0.6513   0.03064   0.02026   0.0217   0.1635   1.0000
   7.750   0.6504   0.03222   0.02163   0.0241   0.1434   1.0000
   8.000   0.6539   0.03380   0.02304   0.0259   0.1276   1.0000
   8.250   0.6617   0.03530   0.02445   0.0274   0.1161   1.0000
   8.500   0.6720   0.03676   0.02577   0.0286   0.1069   1.0000
   8.750   0.6876   0.03800   0.02705   0.0296   0.0980   1.0000
   9.000   0.7075   0.03921   0.02824   0.0306   0.0913   1.0000
   9.250   0.7321   0.04034   0.02945   0.0313   0.0847   1.0000
   9.500   0.7623   0.04155   0.03068   0.0316   0.0789   1.0000
   9.750   0.7953   0.04302   0.03235   0.0316   0.0741   1.0000
  10.000   0.8238   0.04466   0.03398   0.0314   0.0702   1.0000
  10.250   0.8473   0.04675   0.03631   0.0315   0.0670   1.0000
  10.500   0.8642   0.04902   0.03889   0.0321   0.0645   1.0000
  10.750   0.8788   0.05150   0.04162   0.0327   0.0629   1.0000
  11.000   0.8885   0.05394   0.04426   0.0335   0.0613   1.0000
  11.250   0.8970   0.05635   0.04679   0.0341   0.0598   1.0000
  11.500   0.9075   0.05899   0.04945   0.0345   0.0583   1.0000
  11.750   0.8970   0.06218   0.05299   0.0356   0.0577   1.0000
  12.000   0.8847   0.06570   0.05684   0.0362   0.0572   1.0000
  12.250   0.8676   0.06978   0.06122   0.0360   0.0567   1.0000
  12.500   0.8502   0.07429   0.06597   0.0352   0.0566   1.0000
  12.750   0.8288   0.07959   0.07147   0.0334   0.0567   1.0000
  13.000   0.8045   0.08571   0.07779   0.0307   0.0568   1.0000
  13.250   0.7778   0.09289   0.08513   0.0268   0.0570   1.0000
  13.500   0.7522   0.10088   0.09325   0.0221   0.0575   1.0000
<< Back to NACA 642-015 AIRFOIL (n64015-il)

Polar data table (+)

Polar graphs


<< Back to NACA 642-015 AIRFOIL (n64015-il)