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

NACA 642-015A AIRFOIL (n64015a-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 642-015A AIRFOIL (n64015a-il)
Reynolds number: 500,000
Max Cl/Cd: 64.38 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n64015a-il-500000.txt
Download as CSV file: xf-n64015a-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 642-015A AIRFOIL                           
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -16.750  -0.9958   0.09366   0.09041  -0.0242   1.0000   0.0187
 -16.500  -1.0245   0.08506   0.08157  -0.0290   1.0000   0.0186
 -16.250  -1.0456   0.07837   0.07469  -0.0323   1.0000   0.0186
 -16.000  -1.0642   0.07248   0.06861  -0.0347   1.0000   0.0186
 -15.750  -1.0800   0.06734   0.06331  -0.0364   1.0000   0.0187
 -15.500  -1.0929   0.06281   0.05860  -0.0375   1.0000   0.0187
 -15.250  -1.1033   0.05876   0.05439  -0.0382   1.0000   0.0188
 -15.000  -1.1104   0.05521   0.05070  -0.0385   1.0000   0.0189
 -14.750  -1.1151   0.05204   0.04737  -0.0386   1.0000   0.0190
 -14.500  -1.1176   0.04919   0.04439  -0.0384   1.0000   0.0192
 -14.250  -1.1181   0.04660   0.04167  -0.0381   1.0000   0.0193
 -14.000  -1.1170   0.04423   0.03917  -0.0376   1.0000   0.0195
 -13.750  -1.1143   0.04202   0.03683  -0.0369   1.0000   0.0197
 -13.500  -1.1094   0.04003   0.03472  -0.0362   1.0000   0.0199
 -13.250  -1.1032   0.03816   0.03274  -0.0353   1.0000   0.0201
 -13.000  -1.0957   0.03640   0.03085  -0.0344   1.0000   0.0204
 -12.750  -1.0868   0.03474   0.02906  -0.0334   1.0000   0.0207
 -12.500  -1.0766   0.03321   0.02740  -0.0324   1.0000   0.0210
 -12.250  -1.0655   0.03184   0.02590  -0.0314   1.0000   0.0214
 -12.000  -1.0536   0.03060   0.02454  -0.0303   1.0000   0.0218
 -11.750  -1.0410   0.02951   0.02332  -0.0292   1.0000   0.0221
 -11.500  -1.0259   0.02798   0.02170  -0.0282   1.0000   0.0225
 -11.250  -1.0112   0.02649   0.02017  -0.0271   1.0000   0.0230
 -11.000  -0.9971   0.02538   0.01904  -0.0260   1.0000   0.0234
 -10.750  -0.9826   0.02441   0.01804  -0.0248   1.0000   0.0239
 -10.500  -0.9680   0.02349   0.01708  -0.0236   1.0000   0.0244
 -10.250  -0.9533   0.02261   0.01615  -0.0223   1.0000   0.0250
 -10.000  -0.9387   0.02178   0.01526  -0.0209   1.0000   0.0256
  -9.750  -0.9242   0.02101   0.01444  -0.0194   1.0000   0.0262
  -9.500  -0.9090   0.02040   0.01376  -0.0178   1.0000   0.0268
  -9.250  -0.9031   0.01931   0.01265  -0.0147   1.0000   0.0276
  -9.000  -0.8946   0.01863   0.01198  -0.0118   1.0000   0.0285
  -8.750  -0.8837   0.01810   0.01144  -0.0092   1.0000   0.0295
  -8.500  -0.8755   0.01764   0.01096  -0.0060   1.0000   0.0304
  -8.000  -0.8336   0.01629   0.00955  -0.0049   0.9943   0.0334
  -7.750  -0.8003   0.01558   0.00883  -0.0069   0.9881   0.0361
  -7.500  -0.7650   0.01495   0.00817  -0.0091   0.9828   0.0394
  -7.250  -0.7342   0.01430   0.00754  -0.0104   0.9728   0.0443
  -7.000  -0.7046   0.01368   0.00694  -0.0113   0.9613   0.0509
  -6.750  -0.6773   0.01314   0.00641  -0.0116   0.9479   0.0596
  -6.500  -0.6536   0.01263   0.00594  -0.0111   0.9326   0.0722
  -6.250  -0.6321   0.01214   0.00550  -0.0101   0.9170   0.0917
  -6.000  -0.6128   0.01154   0.00505  -0.0088   0.9017   0.1273
  -5.750  -0.5947   0.01082   0.00457  -0.0074   0.8874   0.1856
  -5.500  -0.5782   0.00996   0.00405  -0.0057   0.8742   0.2691
  -5.250  -0.5636   0.00896   0.00354  -0.0038   0.8613   0.3791
  -5.000  -0.5440   0.00845   0.00333  -0.0024   0.8502   0.4628
  -4.750  -0.5196   0.00830   0.00320  -0.0018   0.8404   0.4996
  -4.500  -0.4942   0.00818   0.00310  -0.0013   0.8300   0.5232
  -4.250  -0.4680   0.00812   0.00301  -0.0010   0.8210   0.5410
  -4.000  -0.4415   0.00807   0.00293  -0.0007   0.8120   0.5557
  -3.750  -0.4147   0.00804   0.00286  -0.0005   0.8037   0.5695
  -3.500  -0.3877   0.00802   0.00279  -0.0003   0.7951   0.5818
  -3.250  -0.3606   0.00799   0.00273  -0.0001   0.7870   0.5920
  -3.000  -0.3338   0.00797   0.00269   0.0001   0.7787   0.6045
  -2.750  -0.3069   0.00798   0.00268   0.0003   0.7711   0.6182
  -2.500  -0.2797   0.00800   0.00265   0.0005   0.7633   0.6291
  -2.250  -0.2522   0.00799   0.00263   0.0006   0.7564   0.6380
  -2.000  -0.2246   0.00797   0.00259   0.0006   0.7487   0.6464
  -1.500  -0.1687   0.00795   0.00251   0.0005   0.7343   0.6594
  -1.250  -0.1407   0.00792   0.00246   0.0005   0.7274   0.6646
  -1.000  -0.1125   0.00790   0.00244   0.0004   0.7206   0.6704
  -0.500  -0.0563   0.00789   0.00238   0.0002   0.7078   0.6822
  -0.250  -0.0281   0.00787   0.00238   0.0001   0.7008   0.6883
   0.000   0.0000   0.00788   0.00235   0.0000   0.6946   0.6946
   0.250   0.0281   0.00787   0.00238  -0.0001   0.6884   0.7008
   0.500   0.0563   0.00789   0.00238  -0.0002   0.6822   0.7078
   1.000   0.1125   0.00790   0.00244  -0.0004   0.6704   0.7206
   1.250   0.1407   0.00792   0.00246  -0.0005   0.6646   0.7274
   1.500   0.1687   0.00795   0.00251  -0.0005   0.6593   0.7343
   2.000   0.2246   0.00797   0.00259  -0.0006   0.6464   0.7487
   2.250   0.2522   0.00799   0.00263  -0.0006   0.6380   0.7564
   2.500   0.2797   0.00800   0.00265  -0.0005   0.6290   0.7633
   2.750   0.3069   0.00798   0.00268  -0.0003   0.6182   0.7711
   3.000   0.3338   0.00797   0.00269  -0.0001   0.6046   0.7787
   3.250   0.3606   0.00799   0.00273   0.0001   0.5920   0.7870
   3.500   0.3877   0.00802   0.00279   0.0003   0.5818   0.7952
   3.750   0.4148   0.00804   0.00286   0.0005   0.5695   0.8037
   4.000   0.4415   0.00807   0.00293   0.0007   0.5557   0.8120
   4.250   0.4681   0.00812   0.00301   0.0010   0.5410   0.8210
   4.500   0.4943   0.00818   0.00310   0.0013   0.5232   0.8300
   4.750   0.5196   0.00830   0.00320   0.0018   0.4997   0.8404
   5.000   0.5440   0.00845   0.00333   0.0024   0.4628   0.8502
   5.250   0.5637   0.00896   0.00354   0.0038   0.3796   0.8613
   5.500   0.5782   0.00996   0.00405   0.0057   0.2691   0.8742
   5.750   0.5948   0.01082   0.00457   0.0073   0.1856   0.8874
   6.000   0.6128   0.01154   0.00505   0.0088   0.1273   0.9017
   6.250   0.6322   0.01214   0.00550   0.0101   0.0916   0.9169
   6.500   0.6537   0.01263   0.00594   0.0111   0.0722   0.9326
   6.750   0.6774   0.01314   0.00641   0.0116   0.0596   0.9479
   7.000   0.7048   0.01368   0.00694   0.0113   0.0508   0.9613
   7.250   0.7343   0.01430   0.00754   0.0104   0.0444   0.9728
   7.500   0.7650   0.01495   0.00817   0.0091   0.0394   0.9829
   7.750   0.8004   0.01558   0.00883   0.0069   0.0361   0.9882
   8.000   0.8337   0.01629   0.00955   0.0049   0.0334   0.9943
   8.250   0.8669   0.01722   0.01051   0.0028   0.0314   0.9994
   8.500   0.8754   0.01764   0.01096   0.0060   0.0304   1.0000
   8.750   0.8838   0.01810   0.01143   0.0092   0.0295   1.0000
   9.000   0.8948   0.01863   0.01197   0.0118   0.0285   1.0000
   9.250   0.9033   0.01931   0.01265   0.0147   0.0276   1.0000
   9.500   0.9094   0.02039   0.01375   0.0177   0.0268   1.0000
   9.750   0.9246   0.02101   0.01444   0.0193   0.0262   1.0000
  10.000   0.9392   0.02178   0.01526   0.0208   0.0256   1.0000
  10.250   0.9538   0.02261   0.01615   0.0222   0.0250   1.0000
  10.500   0.9685   0.02349   0.01708   0.0235   0.0244   1.0000
  10.750   0.9832   0.02440   0.01803   0.0247   0.0239   1.0000
  11.000   0.9977   0.02538   0.01904   0.0258   0.0234   1.0000
  11.250   1.0119   0.02648   0.02017   0.0270   0.0230   1.0000
  11.500   1.0266   0.02798   0.02170   0.0281   0.0225   1.0000
  11.750   1.0418   0.02951   0.02332   0.0291   0.0221   1.0000
  12.000   1.0544   0.03060   0.02454   0.0302   0.0218   1.0000
  12.250   1.0664   0.03183   0.02590   0.0312   0.0214   1.0000
  12.500   1.0776   0.03321   0.02740   0.0322   0.0210   1.0000
  12.750   1.0878   0.03474   0.02906   0.0332   0.0207   1.0000
  13.000   1.0967   0.03640   0.03085   0.0342   0.0204   1.0000
  13.250   1.1043   0.03816   0.03274   0.0351   0.0201   1.0000
  13.500   1.1106   0.04004   0.03473   0.0360   0.0199   1.0000
  13.750   1.1155   0.04202   0.03683   0.0367   0.0197   1.0000
  14.000   1.1183   0.04422   0.03916   0.0374   0.0195   1.0000
  14.250   1.1194   0.04660   0.04167   0.0379   0.0193   1.0000
  14.500   1.1190   0.04918   0.04439   0.0382   0.0191   1.0000
  14.750   1.1166   0.05203   0.04736   0.0383   0.0190   1.0000
  15.000   1.1119   0.05522   0.05070   0.0383   0.0189   1.0000
  15.250   1.1046   0.05880   0.05444   0.0379   0.0188   1.0000
  15.500   1.0944   0.06283   0.05863   0.0372   0.0187   1.0000
  15.750   1.0815   0.06738   0.06334   0.0361   0.0187   1.0000
  16.000   1.0656   0.07253   0.06867   0.0344   0.0186   1.0000
  16.250   1.0471   0.07840   0.07472   0.0319   0.0186   1.0000
  16.500   1.0257   0.08516   0.08168   0.0287   0.0186   1.0000
  16.750   0.9988   0.09346   0.09020   0.0240   0.0187   1.0000
<< Back to NACA 642-015A AIRFOIL (n64015a-il)

Polar data table (+)

Polar graphs


<< Back to NACA 642-015A AIRFOIL (n64015a-il)