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

NACA 65(2)-415 a=0.5 (naca652415a05-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(2)-415 a=0.5 (naca652415a05-il)
Reynolds number: 100,000
Max Cl/Cd: 33.64 at α=9.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca652415a05-il-100000.txt
Download as CSV file: xf-naca652415a05-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(2)-415 a=0.5                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.3453   0.11736   0.11281  -0.0504   1.0000   0.1179
 -11.250  -0.3718   0.11837   0.11394  -0.0452   1.0000   0.1188
 -11.000  -0.3942   0.11478   0.11044  -0.0535   0.9938   0.1246
 -10.750  -0.3711   0.10935   0.10497  -0.0551   0.9895   0.1277
 -10.500  -0.3446   0.10560   0.10117  -0.0573   0.9856   0.1341
 -10.250  -0.3811   0.09869   0.09436  -0.0724   0.9767   0.1411
 -10.000  -0.3225   0.09627   0.09182  -0.0670   0.9758   0.1475
  -9.750  -0.3832   0.08814   0.08376  -0.0876   0.9634   0.1560
  -9.500  -0.3056   0.08704   0.08260  -0.0784   0.9656   0.1689
  -9.250  -0.2859   0.08293   0.07846  -0.0814   0.9614   0.1789
  -9.000  -0.2950   0.07775   0.07332  -0.0881   0.9527   0.1908
  -8.750  -0.3132   0.07326   0.06882  -0.0950   0.9420   0.2021
  -8.500  -0.2737   0.07114   0.06667  -0.0926   0.9393   0.2141
  -8.250  -0.2727   0.06746   0.06299  -0.0951   0.9322   0.2272
  -8.000  -0.3136   0.06496   0.06048  -0.0947   0.9190   0.2383
  -7.250  -0.4115   0.04578   0.03874  -0.0881   0.8883   0.0939
  -7.000  -0.3909   0.04205   0.03473  -0.0880   0.8853   0.0888
  -6.750  -0.3994   0.04078   0.03311  -0.0825   0.8771   0.0859
  -6.500  -0.3802   0.03871   0.03062  -0.0814   0.8730   0.0857
  -6.250  -0.3520   0.03665   0.02816  -0.0817   0.8704   0.0860
  -6.000  -0.3577   0.03623   0.02750  -0.0765   0.8636   0.0854
  -5.750  -0.3397   0.03499   0.02598  -0.0750   0.8597   0.0853
  -5.500  -0.3101   0.03365   0.02446  -0.0754   0.8569   0.0872
  -5.250  -0.2756   0.03251   0.02314  -0.0765   0.8547   0.0913
  -5.000  -0.2824   0.03263   0.02317  -0.0711   0.8483   0.0924
  -4.750  -0.2679   0.03217   0.02257  -0.0690   0.8447   0.0941
  -4.500  -0.2410   0.03124   0.02161  -0.0689   0.8421   0.0971
  -4.250  -0.2088   0.03040   0.02089  -0.0699   0.8399   0.1046
  -4.000  -0.1996   0.03039   0.02080  -0.0670   0.8360   0.1095
  -3.750  -0.2082   0.03049   0.02097  -0.0615   0.8315   0.1123
  -3.500  -0.1973   0.03031   0.02086  -0.0592   0.8286   0.1195
  -3.250  -0.1774   0.02992   0.02054  -0.0581   0.8259   0.1357
  -3.000  -0.1563   0.02851   0.01987  -0.0576   0.8235   0.2361
  -2.750  -0.1708   0.02747   0.02085  -0.0495   0.8207   0.6648
  -2.500  -0.1765   0.02828   0.02178  -0.0432   0.8182   0.7278
  -2.250  -0.1736   0.02901   0.02253  -0.0381   0.8153   0.7716
  -2.000  -0.1648   0.02968   0.02319  -0.0339   0.8126   0.8060
  -1.750  -0.1575   0.03041   0.02391  -0.0293   0.8109   0.8400
  -1.500  -0.1502   0.03120   0.02471  -0.0237   0.8091   0.8815
  -1.250  -0.0131   0.03365   0.02684  -0.0374   0.8065   0.9345
  -1.000   0.0036   0.03420   0.02731  -0.0379   0.8028   0.9468
  -0.750   0.0320   0.03473   0.02774  -0.0400   0.8002   0.9577
  -0.500   0.0799   0.03527   0.02815  -0.0455   0.7975   0.9639
  -0.250   0.1241   0.03567   0.02842  -0.0499   0.7944   0.9720
   0.000   0.1851   0.03598   0.02862  -0.0568   0.7917   0.9766
   0.250   0.2267   0.03655   0.02911  -0.0611   0.7890   0.9847
   0.500   0.2535   0.03736   0.02989  -0.0642   0.7854   0.9954
   0.750   0.2625   0.03793   0.03043  -0.0636   0.7820   1.0000
   1.000   0.2421   0.03832   0.03079  -0.0577   0.7796   1.0000
   1.250   0.2421   0.03857   0.03099  -0.0544   0.7766   1.0000
   1.500   0.2490   0.03898   0.03134  -0.0520   0.7736   1.0000
   1.750   0.2078   0.03939   0.03169  -0.0431   0.7719   1.0000
   2.000   0.2027   0.04012   0.03235  -0.0397   0.7700   1.0000
   2.250   0.2093   0.04096   0.03312  -0.0380   0.7675   1.0000
   2.500   0.2240   0.04182   0.03393  -0.0373   0.7645   1.0000
   2.750   0.2589   0.04241   0.03447  -0.0385   0.7553   1.0000
   3.000   0.2648   0.04338   0.03541  -0.0369   0.7519   1.0000
   3.250   0.3054   0.04394   0.03594  -0.0388   0.7441   1.0000
   3.500   0.3108   0.04486   0.03684  -0.0369   0.7371   1.0000
   3.750   0.3391   0.04552   0.03749  -0.0374   0.7298   1.0000
   4.000   0.3560   0.04632   0.03828  -0.0367   0.7210   1.0000
   4.250   0.3858   0.04695   0.03892  -0.0374   0.7137   1.0000
   4.500   0.4007   0.04777   0.03975  -0.0364   0.7042   1.0000
   4.750   0.4387   0.04821   0.04022  -0.0377   0.6970   1.0000
   5.000   0.4476   0.04910   0.04114  -0.0362   0.6864   1.0000
   5.250   0.4961   0.04924   0.04134  -0.0383   0.6802   1.0000
   5.500   0.4996   0.05023   0.04237  -0.0363   0.6685   1.0000
   5.750   0.5252   0.05078   0.04297  -0.0362   0.6592   1.0000
   6.000   0.5577   0.05101   0.04327  -0.0367   0.6509   1.0000
   6.250   0.5713   0.05183   0.04416  -0.0356   0.6394   1.0000
   6.500   0.6193   0.05140   0.04385  -0.0371   0.6339   1.0000
   6.750   0.6293   0.05225   0.04477  -0.0357   0.6212   1.0000
   7.000   0.6493   0.05274   0.04536  -0.0349   0.6100   1.0000
   7.250   0.6947   0.05193   0.04470  -0.0358   0.6039   1.0000
   7.500   0.7105   0.05239   0.04527  -0.0346   0.5910   1.0000
   7.750   0.7702   0.05008   0.04318  -0.0357   0.5862   1.0000
   8.000   0.8137   0.04745   0.04074  -0.0349   0.5731   1.0000
   8.750   0.9763   0.03204   0.02605  -0.0297   0.5183   1.0000
   9.000   0.9893   0.03122   0.02532  -0.0268   0.4791   1.0000
   9.250   0.9970   0.02964   0.02225  -0.0206   0.2801   1.0000
   9.500   0.9703   0.03324   0.02508  -0.0164   0.1939   1.0000
   9.750   0.9520   0.03661   0.02776  -0.0131   0.1366   1.0000
  10.000   0.9480   0.03902   0.02986  -0.0108   0.1112   1.0000
  10.250   0.9520   0.04087   0.03151  -0.0088   0.0969   1.0000
  10.500   0.9642   0.04219   0.03280  -0.0073   0.0870   1.0000
  10.750   0.9850   0.04314   0.03365  -0.0059   0.0801   1.0000
  11.000   1.0092   0.04407   0.03445  -0.0051   0.0734   1.0000
  11.250   1.0476   0.04487   0.03528  -0.0048   0.0687   1.0000
  11.500   1.0866   0.04603   0.03653  -0.0048   0.0651   1.0000
  11.750   1.1339   0.04766   0.03813  -0.0059   0.0619   1.0000
  12.000   1.1724   0.05040   0.04111  -0.0067   0.0593   1.0000
  12.250   1.1903   0.05291   0.04395  -0.0059   0.0585   1.0000
  12.500   1.2045   0.05586   0.04725  -0.0050   0.0581   1.0000
  12.750   1.2132   0.05904   0.05076  -0.0038   0.0581   1.0000
  13.000   1.2157   0.06249   0.05454  -0.0023   0.0582   1.0000
  13.250   1.2137   0.06616   0.05853  -0.0008   0.0585   1.0000
  13.500   1.2080   0.06998   0.06263   0.0006   0.0589   1.0000
  13.750   1.1983   0.07399   0.06690   0.0019   0.0592   1.0000
  14.000   1.1858   0.07818   0.07134   0.0030   0.0595   1.0000
  14.250   1.1716   0.08272   0.07611   0.0037   0.0598   1.0000
  14.500   1.1621   0.08785   0.08142   0.0038   0.0602   1.0000
  14.750   1.1522   0.09326   0.08703   0.0038   0.0606   1.0000
  15.000   1.1211   0.09664   0.09070   0.0037   0.0613   1.0000
  15.250   0.8163   0.10959   0.10489  -0.0018   0.0661   1.0000
<< Back to NACA 65(2)-415 a=0.5 (naca652415a05-il)

Polar data table (+)

Polar graphs


<< Back to NACA 65(2)-415 a=0.5 (naca652415a05-il)