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(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il)
Reynolds number: 50,000
Max Cl/Cd: 25.38 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651212a06-il-50000.txt
Download as CSV file: xf-naca651212a06-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-212 a=0.6                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.4949   0.14356   0.13638  -0.0102   1.0000   0.2185
 -12.000  -0.4797   0.13863   0.13144  -0.0104   1.0000   0.2263
 -11.750  -0.4989   0.13825   0.13117  -0.0129   1.0000   0.2348
 -11.500  -0.4715   0.13229   0.12517  -0.0120   1.0000   0.2464
 -11.250  -0.4646   0.12848   0.12138  -0.0126   1.0000   0.2559
 -11.000  -0.4768   0.12680   0.11977  -0.0142   1.0000   0.2661
 -10.750  -0.4723   0.12384   0.11683  -0.0144   1.0000   0.2799
 -10.500  -0.4652   0.12044   0.11346  -0.0144   1.0000   0.2949
 -10.250  -0.4643   0.11767   0.11074  -0.0144   1.0000   0.3104
  -9.500  -0.4259   0.10611   0.09922  -0.0133   1.0000   0.3589
  -9.250  -0.4162   0.10277   0.09590  -0.0130   1.0000   0.3749
  -9.000  -0.4068   0.09946   0.09262  -0.0127   1.0000   0.3911
  -8.750  -0.3974   0.09626   0.08942  -0.0123   1.0000   0.4081
  -8.500  -0.3905   0.09341   0.08661  -0.0118   1.0000   0.4265
  -8.000  -0.3769   0.08712   0.08042  -0.0113   1.0000   0.4533
  -7.750  -0.4730   0.07799   0.07176  -0.0262   1.0000   0.3019
  -7.500  -0.5374   0.07404   0.06811  -0.0271   1.0000   0.2771
  -7.000  -0.6310   0.06475   0.05869  -0.0264   1.0000   0.2157
  -6.750  -0.6672   0.05952   0.05254  -0.0258   1.0000   0.1751
  -6.500  -0.6604   0.05483   0.04754  -0.0243   1.0000   0.1585
  -6.250  -0.6581   0.05121   0.04322  -0.0222   1.0000   0.1447
  -6.000  -0.6460   0.04751   0.03943  -0.0207   1.0000   0.1390
  -5.750  -0.6367   0.04475   0.03570  -0.0184   1.0000   0.1299
  -5.500  -0.6212   0.04173   0.03245  -0.0170   1.0000   0.1275
  -5.250  -0.6048   0.03926   0.02960  -0.0155   1.0000   0.1259
  -5.000  -0.5870   0.03722   0.02712  -0.0141   1.0000   0.1272
  -4.750  -0.5672   0.03539   0.02486  -0.0127   1.0000   0.1288
  -4.500  -0.5453   0.03367   0.02276  -0.0115   1.0000   0.1295
  -4.250  -0.5224   0.03201   0.02085  -0.0104   1.0000   0.1312
  -4.000  -0.4997   0.03049   0.01935  -0.0095   1.0000   0.1367
  -3.750  -0.4774   0.02952   0.01817  -0.0082   1.0000   0.1449
  -3.500  -0.4558   0.02832   0.01712  -0.0068   1.0000   0.1541
  -3.250  -0.1219   0.02937   0.02057  -0.0333   1.0000   0.9972
  -3.000  -0.1138   0.02908   0.02010  -0.0326   1.0000   1.0000
  -2.750  -0.1165   0.02888   0.01981  -0.0298   1.0000   1.0000
  -2.500  -0.1190   0.02867   0.01949  -0.0271   1.0000   1.0000
  -2.250  -0.1213   0.02844   0.01916  -0.0243   1.0000   1.0000
  -2.000  -0.1238   0.02819   0.01880  -0.0214   1.0000   1.0000
  -1.750  -0.1266   0.02791   0.01843  -0.0185   1.0000   1.0000
  -1.500  -0.1299   0.02759   0.01803  -0.0154   1.0000   1.0000
  -1.250  -0.1340   0.02724   0.01759  -0.0122   1.0000   1.0000
  -1.000  -0.1390   0.02683   0.01711  -0.0089   1.0000   1.0000
  -0.750  -0.1445   0.02638   0.01658  -0.0053   1.0000   1.0000
  -0.500  -0.1489   0.02594   0.01604  -0.0018   1.0000   1.0000
  -0.250  -0.1494   0.02563   0.01560   0.0011   1.0000   1.0000
   0.000  -0.1449   0.02548   0.01530   0.0035   1.0000   1.0000
   0.250  -0.1356   0.02551   0.01518   0.0051   1.0000   1.0000
   0.500  -0.1231   0.02569   0.01521   0.0063   1.0000   1.0000
   0.750  -0.1084   0.02596   0.01534   0.0071   1.0000   1.0000
   1.000  -0.0924   0.02630   0.01556   0.0077   1.0000   1.0000
   1.250  -0.0757   0.02671   0.01586   0.0082   1.0000   1.0000
   1.500  -0.0585   0.02717   0.01622   0.0086   1.0000   1.0000
   1.750  -0.0409   0.02768   0.01664   0.0089   1.0000   1.0000
   2.000  -0.0231   0.02823   0.01713   0.0092   1.0000   1.0000
   2.250  -0.0053   0.02883   0.01767   0.0094   1.0000   1.0000
   2.500   0.0125   0.02947   0.01827   0.0095   1.0000   1.0000
   2.750   0.0304   0.03015   0.01893   0.0097   1.0000   1.0000
   3.000   0.0481   0.03088   0.01964   0.0098   1.0000   1.0000
   3.250   0.0657   0.03165   0.02042   0.0098   1.0000   1.0000
   3.500   0.0832   0.03247   0.02124   0.0098   1.0000   1.0000
   3.750   0.1005   0.03333   0.02213   0.0098   1.0000   1.0000
   4.000   0.1175   0.03426   0.02309   0.0098   1.0000   1.0000
   4.250   0.1343   0.03523   0.02411   0.0097   1.0000   1.0000
   4.500   0.1508   0.03627   0.02522   0.0095   1.0000   1.0000
   4.750   0.1670   0.03737   0.02639   0.0093   1.0000   1.0000
   5.000   0.1827   0.03854   0.02764   0.0091   1.0000   1.0000
   5.250   0.1981   0.03978   0.02898   0.0088   1.0000   1.0000
   5.500   0.2130   0.04111   0.03042   0.0084   1.0000   1.0000
   5.750   0.2275   0.04253   0.03198   0.0080   1.0000   1.0000
   6.000   0.4599   0.04856   0.03907  -0.0222   0.8022   1.0000
   6.250   0.5043   0.04925   0.04012  -0.0240   0.7694   1.0000
   6.500   0.7008   0.02761   0.01722  -0.0064   0.2213   1.0000
   6.750   0.6995   0.02983   0.01883  -0.0026   0.1826   1.0000
   7.000   0.7232   0.03174   0.02038  -0.0012   0.1554   1.0000
   7.250   0.7847   0.03378   0.02239  -0.0040   0.1316   1.0000
   7.500   0.8423   0.03653   0.02524  -0.0067   0.1211   1.0000
   7.750   0.8798   0.03936   0.02804  -0.0078   0.1135   1.0000
   8.000   0.9102   0.04258   0.03155  -0.0077   0.1118   1.0000
   8.250   0.9336   0.04573   0.03511  -0.0068   0.1116   1.0000
   8.500   0.9521   0.04900   0.03883  -0.0054   0.1116   1.0000
   8.750   0.9648   0.05224   0.04252  -0.0036   0.1113   1.0000
   9.000   0.9736   0.05565   0.04636  -0.0017   0.1111   1.0000
   9.250   0.9833   0.05965   0.05066  -0.0001   0.1119   1.0000
   9.500   0.9713   0.06257   0.05434   0.0037   0.1159   1.0000
   9.750   0.9554   0.06677   0.05902   0.0066   0.1202   1.0000
  10.000   0.9513   0.07140   0.06388   0.0081   0.1241   1.0000
  10.250   0.9369   0.07552   0.06829   0.0100   0.1289   1.0000
  10.500   0.8918   0.07931   0.07230   0.0125   0.1314   1.0000
  10.750   0.8542   0.08478   0.07787   0.0119   0.1344   1.0000
  11.000   0.8486   0.09099   0.08416   0.0107   0.1424   1.0000
  11.250   0.7614   0.10566   0.09876  -0.0016   0.1595   1.0000
<< Back to NACA 65(1)-212 a=0.6 (naca651212a06-il)

Polar data table (+)

Polar graphs


<< Back to NACA 65(1)-212 a=0.6 (naca651212a06-il)