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

NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 63-212 AIRFOIL (n63212-il)
Reynolds number: 500,000
Max Cl/Cd: 81.69 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63212-il-500000.txt
Download as CSV file: xf-n63212-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-212 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
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.5662   0.12264   0.12023  -0.0084   1.0000   0.0260
 -12.250  -0.5642   0.11840   0.11600  -0.0103   1.0000   0.0266
  -9.500  -0.8082   0.03374   0.02917  -0.0420   1.0000   0.0192
  -9.250  -0.7950   0.03166   0.02686  -0.0409   1.0000   0.0191
  -9.000  -0.7810   0.02945   0.02443  -0.0397   1.0000   0.0191
  -8.750  -0.7657   0.02753   0.02230  -0.0385   1.0000   0.0191
  -8.500  -0.7516   0.02507   0.01966  -0.0370   1.0000   0.0189
  -8.000  -0.7041   0.02157   0.01578  -0.0372   0.9934   0.0190
  -7.750  -0.6723   0.01929   0.01333  -0.0388   0.9874   0.0192
  -7.500  -0.6393   0.01756   0.01151  -0.0405   0.9813   0.0195
  -7.250  -0.6061   0.01634   0.01025  -0.0423   0.9734   0.0200
  -7.000  -0.5738   0.01543   0.00928  -0.0437   0.9638   0.0205
  -6.750  -0.5432   0.01470   0.00852  -0.0447   0.9528   0.0212
  -6.500  -0.5164   0.01414   0.00790  -0.0447   0.9402   0.0220
  -6.250  -0.4926   0.01364   0.00730  -0.0440   0.9270   0.0228
  -6.000  -0.4692   0.01315   0.00673  -0.0433   0.9143   0.0236
  -5.750  -0.4462   0.01253   0.00603  -0.0425   0.9026   0.0246
  -5.500  -0.4228   0.01194   0.00538  -0.0418   0.8920   0.0265
  -5.250  -0.3978   0.01157   0.00494  -0.0413   0.8822   0.0290
  -5.000  -0.3727   0.01105   0.00439  -0.0409   0.8720   0.0348
  -4.750  -0.3474   0.01058   0.00391  -0.0406   0.8631   0.0497
  -4.500  -0.3218   0.01009   0.00354  -0.0404   0.8538   0.0790
  -4.250  -0.2969   0.00943   0.00317  -0.0403   0.8449   0.1499
  -4.000  -0.2729   0.00850   0.00274  -0.0403   0.8364   0.2774
  -3.750  -0.2480   0.00773   0.00248  -0.0403   0.8277   0.4097
  -3.500  -0.2212   0.00748   0.00234  -0.0402   0.8199   0.4679
  -3.250  -0.1935   0.00732   0.00228  -0.0402   0.8113   0.5106
  -3.000  -0.1659   0.00725   0.00222  -0.0401   0.8039   0.5423
  -2.750  -0.1376   0.00719   0.00215  -0.0402   0.7954   0.5635
  -2.500  -0.1099   0.00715   0.00211  -0.0400   0.7882   0.5836
  -2.250  -0.0817   0.00710   0.00209  -0.0400   0.7797   0.6037
  -2.000  -0.0537   0.00712   0.00207  -0.0399   0.7725   0.6234
  -1.750  -0.0253   0.00711   0.00207  -0.0400   0.7642   0.6395
  -1.500   0.0027   0.00712   0.00206  -0.0399   0.7571   0.6513
  -1.250   0.0313   0.00709   0.00204  -0.0400   0.7488   0.6609
  -1.000   0.0597   0.00713   0.00200  -0.0400   0.7417   0.6709
  -0.750   0.0881   0.00710   0.00201  -0.0400   0.7335   0.6793
  -0.500   0.1166   0.00713   0.00198  -0.0401   0.7265   0.6883
  -0.250   0.1451   0.00711   0.00199  -0.0402   0.7183   0.6966
   0.000   0.1734   0.00715   0.00198  -0.0402   0.7112   0.7051
   0.250   0.2020   0.00714   0.00201  -0.0403   0.7030   0.7135
   0.500   0.2303   0.00718   0.00202  -0.0403   0.6958   0.7217
   0.750   0.2589   0.00718   0.00205  -0.0404   0.6875   0.7306
   1.000   0.2870   0.00722   0.00208  -0.0404   0.6801   0.7386
   1.250   0.3156   0.00723   0.00212  -0.0405   0.6714   0.7478
   1.500   0.3435   0.00727   0.00217  -0.0404   0.6635   0.7557
   1.750   0.3720   0.00730   0.00222  -0.0405   0.6544   0.7648
   2.000   0.3998   0.00732   0.00229  -0.0404   0.6456   0.7728
   2.250   0.4279   0.00738   0.00233  -0.0404   0.6358   0.7817
   2.500   0.4555   0.00740   0.00240  -0.0403   0.6232   0.7901
   2.750   0.4832   0.00744   0.00245  -0.0402   0.6102   0.7987
   3.000   0.5107   0.00749   0.00252  -0.0400   0.5959   0.8077
   3.250   0.5376   0.00755   0.00257  -0.0397   0.5770   0.8161
   3.500   0.5647   0.00765   0.00265  -0.0395   0.5561   0.8258
   3.750   0.5913   0.00772   0.00276  -0.0392   0.5382   0.8342
   4.000   0.6182   0.00783   0.00287  -0.0390   0.5187   0.8437
   4.250   0.6442   0.00797   0.00300  -0.0386   0.4934   0.8533
   4.500   0.6690   0.00819   0.00315  -0.0380   0.4560   0.8629
   4.750   0.6926   0.00859   0.00335  -0.0373   0.3981   0.8736
   5.000   0.7124   0.00938   0.00373  -0.0362   0.3018   0.8853
   5.250   0.7297   0.01037   0.00428  -0.0348   0.1996   0.8976
   5.500   0.7474   0.01124   0.00480  -0.0333   0.1233   0.9113
   5.750   0.7648   0.01199   0.00530  -0.0316   0.0722   0.9277
   6.000   0.7827   0.01252   0.00574  -0.0298   0.0518   0.9483
   6.250   0.8096   0.01302   0.00625  -0.0300   0.0421   0.9771
   6.500   0.8385   0.01353   0.00679  -0.0307   0.0371   1.0000
   6.750   0.8614   0.01439   0.00764  -0.0303   0.0326   1.0000
   7.000   0.8868   0.01487   0.00816  -0.0302   0.0305   1.0000
   7.250   0.9110   0.01542   0.00873  -0.0300   0.0283   1.0000
   7.500   0.9309   0.01637   0.00968  -0.0292   0.0263   1.0000
   7.750   0.9515   0.01723   0.01060  -0.0284   0.0252   1.0000
   8.000   0.9734   0.01791   0.01134  -0.0277   0.0243   1.0000
   8.250   0.9943   0.01867   0.01216  -0.0269   0.0233   1.0000
   8.500   1.0147   0.01946   0.01299  -0.0261   0.0225   1.0000
   8.750   1.0343   0.02032   0.01387  -0.0252   0.0218   1.0000
   9.000   1.0526   0.02135   0.01493  -0.0242   0.0211   1.0000
   9.250   1.0680   0.02350   0.01713  -0.0230   0.0202   1.0000
   9.500   1.0883   0.02409   0.01783  -0.0221   0.0198   1.0000
   9.750   1.1076   0.02513   0.01897  -0.0212   0.0194   1.0000
  10.000   1.1264   0.02635   0.02030  -0.0203   0.0189   1.0000
  10.250   1.1444   0.02765   0.02173  -0.0193   0.0185   1.0000
  10.500   1.1613   0.02908   0.02329  -0.0183   0.0182   1.0000
  10.750   1.1761   0.03051   0.02485  -0.0170   0.0178   1.0000
  11.000   1.1890   0.03206   0.02652  -0.0155   0.0175   1.0000
  11.250   1.1997   0.03368   0.02829  -0.0139   0.0173   1.0000
  11.500   1.2073   0.03521   0.02994  -0.0120   0.0170   1.0000
  11.750   1.2133   0.03693   0.03179  -0.0102   0.0168   1.0000
  12.000   1.2189   0.03851   0.03344  -0.0087   0.0164   1.0000
  12.250   1.2204   0.04138   0.03646  -0.0073   0.0161   1.0000
  12.500   1.2104   0.04566   0.04103  -0.0055   0.0159   1.0000
  12.750   1.1999   0.04925   0.04486  -0.0041   0.0159   1.0000
  13.000   1.1849   0.05281   0.04866  -0.0032   0.0158   1.0000
  13.250   1.1714   0.05670   0.05276  -0.0030   0.0157   1.0000
  13.500   1.1566   0.06084   0.05711  -0.0035   0.0157   1.0000
  13.750   1.1378   0.06624   0.06272  -0.0047   0.0158   1.0000
<< Back to NACA 63-212 AIRFOIL (n63212-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63-212 AIRFOIL (n63212-il)