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

NACA 65-209 (naca65209-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA 65-209 (naca65209-il)
Reynolds number: 500,000
Max Cl/Cd: 61.19 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca65209-il-500000-n5.txt
Download as CSV file: xf-naca65209-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65-209                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5759   0.08054   0.07830  -0.0276   1.0000   0.0052
  -9.500  -0.5872   0.07333   0.07116  -0.0333   1.0000   0.0052
  -9.250  -0.6048   0.06598   0.06380  -0.0403   1.0000   0.0051
  -9.000  -0.6236   0.06139   0.05915  -0.0418   1.0000   0.0051
  -8.750  -0.6373   0.05584   0.05347  -0.0428   1.0000   0.0050
  -8.500  -0.6506   0.04927   0.04668  -0.0427   1.0000   0.0051
  -8.250  -0.6700   0.03981   0.03678  -0.0407   1.0000   0.0052
  -8.000  -0.6884   0.02568   0.02155  -0.0378   0.9943   0.0060
  -7.750  -0.6584   0.02512   0.02087  -0.0392   0.9893   0.0063
  -7.500  -0.6289   0.02371   0.01925  -0.0406   0.9843   0.0070
  -7.250  -0.6027   0.02040   0.01544  -0.0413   0.9776   0.0082
  -7.000  -0.5718   0.01957   0.01438  -0.0424   0.9720   0.0090
  -6.750  -0.5433   0.01800   0.01258  -0.0434   0.9653   0.0100
  -6.500  -0.5132   0.01778   0.01232  -0.0445   0.9585   0.0108
  -6.250  -0.4841   0.01682   0.01120  -0.0451   0.9512   0.0114
  -6.000  -0.4568   0.01608   0.01033  -0.0453   0.9425   0.0124
  -5.750  -0.4300   0.01514   0.00923  -0.0453   0.9342   0.0132
  -5.500  -0.4046   0.01421   0.00812  -0.0449   0.9251   0.0138
  -5.250  -0.3792   0.01353   0.00732  -0.0446   0.9162   0.0142
  -5.000  -0.3536   0.01300   0.00668  -0.0442   0.9078   0.0146
  -4.750  -0.3302   0.01195   0.00552  -0.0436   0.8986   0.0153
  -4.500  -0.3059   0.01123   0.00471  -0.0431   0.8900   0.0162
  -4.250  -0.2803   0.01085   0.00428  -0.0428   0.8820   0.0175
  -4.000  -0.2542   0.01052   0.00388  -0.0426   0.8738   0.0189
  -3.750  -0.2281   0.01016   0.00344  -0.0423   0.8660   0.0203
  -3.500  -0.2017   0.00986   0.00307  -0.0421   0.8578   0.0218
  -3.250  -0.1750   0.00963   0.00277  -0.0420   0.8504   0.0231
  -3.000  -0.1483   0.00933   0.00239  -0.0419   0.8428   0.0258
  -2.750  -0.1213   0.00910   0.00212  -0.0418   0.8354   0.0309
  -2.500  -0.0942   0.00894   0.00192  -0.0417   0.8278   0.0377
  -2.250  -0.0672   0.00867   0.00173  -0.0417   0.8206   0.0637
  -2.000  -0.0416   0.00805   0.00150  -0.0418   0.8135   0.1902
  -1.750  -0.0183   0.00679   0.00121  -0.0420   0.8061   0.4721
  -1.500   0.0061   0.00624   0.00124  -0.0415   0.7991   0.6442
  -1.250   0.0329   0.00614   0.00126  -0.0412   0.7921   0.6897
  -1.000   0.0598   0.00610   0.00126  -0.0410   0.7855   0.7193
  -0.750   0.0873   0.00608   0.00127  -0.0409   0.7786   0.7392
  -0.500   0.1151   0.00609   0.00125  -0.0409   0.7719   0.7475
  -0.250   0.1432   0.00609   0.00123  -0.0410   0.7653   0.7552
   0.000   0.1709   0.00610   0.00124  -0.0409   0.7585   0.7619
   0.250   0.1990   0.00612   0.00125  -0.0410   0.7519   0.7695
   0.500   0.2267   0.00613   0.00127  -0.0410   0.7451   0.7762
   0.750   0.2548   0.00615   0.00130  -0.0411   0.7384   0.7838
   1.000   0.2824   0.00617   0.00134  -0.0411   0.7314   0.7908
   1.500   0.3370   0.00625   0.00143  -0.0408   0.7051   0.8056
   1.750   0.3638   0.00632   0.00147  -0.0406   0.6853   0.8134
   2.000   0.3896   0.00643   0.00150  -0.0401   0.6507   0.8211
   2.250   0.4130   0.00675   0.00154  -0.0392   0.5732   0.8289
   2.500   0.4327   0.00759   0.00176  -0.0379   0.4144   0.8378
   2.750   0.4510   0.00879   0.00221  -0.0368   0.2231   0.8465
   3.000   0.4719   0.00972   0.00261  -0.0360   0.0876   0.8555
   3.250   0.4958   0.01019   0.00290  -0.0355   0.0414   0.8644
   3.500   0.5207   0.01047   0.00318  -0.0350   0.0298   0.8734
   3.750   0.5461   0.01068   0.00346  -0.0346   0.0256   0.8831
   4.000   0.5707   0.01095   0.00376  -0.0340   0.0219   0.8931
   4.250   0.5938   0.01140   0.00429  -0.0331   0.0186   0.9036
   4.500   0.6177   0.01167   0.00463  -0.0324   0.0179   0.9147
   4.750   0.6410   0.01200   0.00504  -0.0315   0.0169   0.9270
   5.000   0.6638   0.01239   0.00554  -0.0306   0.0160   0.9403
   5.250   0.6875   0.01286   0.00608  -0.0299   0.0151   0.9557
   5.750   0.7426   0.01415   0.00752  -0.0305   0.0140   1.0000
   6.000   0.7662   0.01488   0.00831  -0.0300   0.0137   1.0000
   6.250   0.7902   0.01552   0.00900  -0.0297   0.0131   1.0000
   6.500   0.8138   0.01627   0.00983  -0.0292   0.0126   1.0000
   6.750   0.8364   0.01735   0.01100  -0.0286   0.0122   1.0000
   7.000   0.8586   0.01876   0.01254  -0.0279   0.0118   1.0000
   7.250   0.8798   0.02113   0.01512  -0.0271   0.0114   1.0000
   7.500   0.9032   0.02221   0.01636  -0.0266   0.0112   1.0000
   7.750   0.9259   0.02344   0.01777  -0.0261   0.0110   1.0000
   8.000   0.9488   0.02430   0.01879  -0.0256   0.0105   1.0000
   8.250   0.9712   0.02518   0.01986  -0.0251   0.0097   1.0000
   8.500   0.9935   0.02567   0.02045  -0.0246   0.0088   1.0000
   8.750   1.0159   0.02594   0.02078  -0.0243   0.0082   1.0000
   9.000   1.0371   0.02646   0.02139  -0.0238   0.0078   1.0000
   9.250   1.0573   0.02713   0.02214  -0.0231   0.0075   1.0000
   9.500   1.0746   0.02825   0.02340  -0.0222   0.0070   1.0000
   9.750   1.0880   0.03049   0.02596  -0.0208   0.0066   1.0000
  10.000   1.0962   0.03392   0.02984  -0.0188   0.0061   1.0000
  10.250   1.0690   0.04389   0.04066  -0.0137   0.0049   1.0000
  10.500   1.0629   0.04797   0.04500  -0.0111   0.0046   1.0000
  10.750   1.0372   0.05315   0.05047  -0.0070   0.0045   1.0000
  11.000   1.0301   0.05555   0.05299  -0.0051   0.0044   1.0000
  11.250   0.9666   0.06601   0.06378  -0.0044   0.0045   1.0000
  11.500   0.9142   0.07742   0.07539  -0.0095   0.0047   1.0000
<< Back to NACA 65-209 (naca65209-il)

Polar data table (+)

Polar graphs


<< Back to NACA 65-209 (naca65209-il)