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=500,000 Ncrit=5


Details Polar file
Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il)
Reynolds number: 500,000
Max Cl/Cd: 69.31 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca651212a06-il-500000-n5.txt
Download as CSV file: xf-naca651212a06-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.5868   0.09144   0.08919  -0.0309   1.0000   0.0098
 -11.500  -0.7120   0.05718   0.05474  -0.0539   1.0000   0.0093
 -11.250  -0.7773   0.04653   0.04375  -0.0562   1.0000   0.0093
 -11.000  -0.8039   0.03980   0.03656  -0.0568   0.9411   0.0096
 -10.750  -0.8262   0.03565   0.03197  -0.0525   0.9154   0.0099
 -10.500  -0.8370   0.03165   0.02749  -0.0490   0.8999   0.0103
 -10.250  -0.8360   0.02861   0.02400  -0.0464   0.8884   0.0107
 -10.000  -0.8257   0.02652   0.02154  -0.0446   0.8785   0.0111
  -9.750  -0.8116   0.02491   0.01962  -0.0431   0.8702   0.0115
  -9.500  -0.7943   0.02390   0.01848  -0.0421   0.8624   0.0119
  -9.250  -0.7743   0.02328   0.01776  -0.0413   0.8555   0.0122
  -9.000  -0.7539   0.02253   0.01689  -0.0406   0.8484   0.0125
  -8.750  -0.7332   0.02172   0.01593  -0.0398   0.8421   0.0128
  -8.500  -0.7117   0.02080   0.01485  -0.0392   0.8357   0.0133
  -8.250  -0.6899   0.01991   0.01378  -0.0385   0.8298   0.0138
  -8.000  -0.6674   0.01902   0.01273  -0.0379   0.8242   0.0143
  -7.750  -0.6440   0.01832   0.01188  -0.0374   0.8183   0.0148
  -7.500  -0.6200   0.01782   0.01123  -0.0370   0.8131   0.0152
  -7.250  -0.5982   0.01666   0.00997  -0.0363   0.8080   0.0157
  -7.000  -0.5748   0.01600   0.00924  -0.0359   0.8026   0.0162
  -6.750  -0.5508   0.01548   0.00865  -0.0354   0.7977   0.0166
  -6.500  -0.5263   0.01498   0.00810  -0.0351   0.7929   0.0171
  -6.250  -0.5016   0.01452   0.00757  -0.0347   0.7881   0.0178
  -6.000  -0.4767   0.01411   0.00708  -0.0344   0.7837   0.0186
  -5.750  -0.4518   0.01369   0.00658  -0.0341   0.7794   0.0194
  -5.500  -0.4268   0.01324   0.00607  -0.0337   0.7747   0.0198
  -5.250  -0.4018   0.01283   0.00559  -0.0334   0.7702   0.0203
  -5.000  -0.3780   0.01226   0.00495  -0.0329   0.7663   0.0214
  -4.750  -0.3522   0.01190   0.00457  -0.0327   0.7623   0.0224
  -4.500  -0.3261   0.01160   0.00424  -0.0326   0.7581   0.0235
  -4.250  -0.2998   0.01135   0.00394  -0.0324   0.7541   0.0250
  -4.000  -0.2733   0.01114   0.00366  -0.0323   0.7504   0.0265
  -3.750  -0.2468   0.01083   0.00332  -0.0322   0.7464   0.0286
  -3.500  -0.2202   0.01057   0.00306  -0.0321   0.7426   0.0314
  -3.250  -0.1932   0.01038   0.00282  -0.0321   0.7390   0.0345
  -3.000  -0.1662   0.01019   0.00261  -0.0320   0.7357   0.0404
  -2.750  -0.1394   0.00994   0.00243  -0.0320   0.7320   0.0572
  -2.500  -0.1140   0.00947   0.00221  -0.0319   0.7281   0.1257
  -2.250  -0.0902   0.00878   0.00196  -0.0316   0.7244   0.2520
  -2.000  -0.0690   0.00780   0.00169  -0.0310   0.7213   0.4483
  -1.750  -0.0450   0.00736   0.00168  -0.0305   0.7180   0.5707
  -1.500  -0.0186   0.00722   0.00169  -0.0303   0.7142   0.6221
  -1.250   0.0080   0.00713   0.00172  -0.0300   0.7106   0.6587
  -1.000   0.0357   0.00710   0.00170  -0.0300   0.7073   0.6744
  -0.750   0.0637   0.00710   0.00168  -0.0300   0.7043   0.6865
  -0.500   0.0920   0.00707   0.00167  -0.0302   0.7007   0.6956
  -0.250   0.1204   0.00706   0.00166  -0.0304   0.6970   0.7015
   0.000   0.1486   0.00706   0.00166  -0.0305   0.6936   0.7077
   0.500   0.2051   0.00708   0.00169  -0.0308   0.6870   0.7206
   0.750   0.2335   0.00709   0.00173  -0.0310   0.6833   0.7277
   1.000   0.2617   0.00711   0.00177  -0.0311   0.6796   0.7342
   1.250   0.2900   0.00714   0.00180  -0.0312   0.6763   0.7414
   1.500   0.3183   0.00717   0.00185  -0.0314   0.6731   0.7480
   1.750   0.3467   0.00720   0.00192  -0.0316   0.6692   0.7554
   2.000   0.3751   0.00723   0.00198  -0.0318   0.6652   0.7618
   2.250   0.4032   0.00727   0.00202  -0.0319   0.6604   0.7681
   2.500   0.4311   0.00729   0.00207  -0.0320   0.6503   0.7741
   2.750   0.4584   0.00729   0.00207  -0.0319   0.6361   0.7800
   3.000   0.4852   0.00736   0.00207  -0.0317   0.6136   0.7859
   3.250   0.5109   0.00742   0.00208  -0.0314   0.5794   0.7910
   3.500   0.5344   0.00771   0.00215  -0.0307   0.5161   0.7968
   3.750   0.5523   0.00848   0.00244  -0.0293   0.3996   0.8028
   4.000   0.5675   0.00954   0.00296  -0.0276   0.2689   0.8101
   4.500   0.5980   0.01120   0.00404  -0.0238   0.0869   0.8310
   4.750   0.6117   0.01176   0.00455  -0.0209   0.0510   0.8464
   5.000   0.6390   0.01227   0.00492  -0.0214   0.0370   0.8942
   5.250   0.6619   0.01266   0.00527  -0.0208   0.0304   0.9007
   5.500   0.6850   0.01295   0.00559  -0.0201   0.0273   0.9060
   5.750   0.7069   0.01331   0.00595  -0.0193   0.0243   0.9102
   6.000   0.7278   0.01368   0.00635  -0.0182   0.0222   0.9141
   6.250   0.7496   0.01400   0.00672  -0.0173   0.0209   0.9178
   6.500   0.7704   0.01435   0.00712  -0.0163   0.0196   0.9212
   6.750   0.7902   0.01472   0.00753  -0.0151   0.0186   0.9248
   7.000   0.8086   0.01519   0.00802  -0.0137   0.0175   0.9283
   7.250   0.8249   0.01579   0.00867  -0.0120   0.0167   0.9314
   7.500   0.8426   0.01625   0.00920  -0.0105   0.0164   0.9344
   7.750   0.8584   0.01673   0.00975  -0.0087   0.0160   0.9377
   8.000   0.8725   0.01726   0.01035  -0.0066   0.0155   0.9411
   8.250   0.8870   0.01783   0.01098  -0.0047   0.0151   0.9444
   8.500   0.9033   0.01839   0.01159  -0.0033   0.0144   0.9474
   8.750   0.9202   0.01900   0.01226  -0.0020   0.0140   0.9505
   9.000   0.9379   0.01969   0.01302  -0.0011   0.0137   0.9539
   9.250   0.9571   0.02049   0.01388  -0.0006   0.0134   0.9576
   9.750   0.9986   0.02261   0.01614  -0.0008   0.0127   0.9704
  10.000   1.0146   0.02393   0.01755   0.0000   0.0125   1.0000
  10.250   1.0299   0.02487   0.01860   0.0012   0.0124   1.0000
  10.500   1.0449   0.02583   0.01966   0.0024   0.0122   1.0000
  10.750   1.0595   0.02695   0.02089   0.0036   0.0120   1.0000
  11.000   1.0737   0.02804   0.02211   0.0047   0.0117   1.0000
  11.250   1.0872   0.02911   0.02328   0.0058   0.0115   1.0000
  11.500   1.0996   0.03028   0.02458   0.0070   0.0111   1.0000
  11.750   1.1110   0.03167   0.02610   0.0082   0.0109   1.0000
  12.000   1.1210   0.03318   0.02777   0.0094   0.0107   1.0000
  12.250   1.1297   0.03472   0.02945   0.0106   0.0105   1.0000
  12.500   1.1374   0.03625   0.03110   0.0117   0.0103   1.0000
  12.750   1.1430   0.03809   0.03309   0.0128   0.0102   1.0000
  13.000   1.1483   0.03981   0.03493   0.0138   0.0100   1.0000
  13.250   1.1520   0.04174   0.03699   0.0147   0.0099   1.0000
  13.500   1.1541   0.04390   0.03928   0.0155   0.0098   1.0000
  13.750   1.1556   0.04604   0.04153   0.0161   0.0096   1.0000
  14.000   1.1551   0.04852   0.04413   0.0166   0.0095   1.0000
  14.250   1.1508   0.05157   0.04733   0.0169   0.0094   1.0000
  14.500   1.1437   0.05498   0.05089   0.0169   0.0092   1.0000
  14.750   1.1283   0.05974   0.05587   0.0165   0.0092   1.0000
  15.000   1.1195   0.06381   0.06013   0.0155   0.0090   1.0000
  15.250   1.1034   0.06929   0.06582   0.0139   0.0090   1.0000
  15.500   1.0834   0.07578   0.07254   0.0114   0.0090   1.0000
  15.750   1.0681   0.08210   0.07905   0.0083   0.0089   1.0000
  16.000   1.0475   0.08993   0.08706   0.0041   0.0089   1.0000
  16.250   1.0195   0.10025   0.09760  -0.0021   0.0089   1.0000
  16.500   0.9863   0.11312   0.11066  -0.0101   0.0090   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)