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

RAE6-9CK AIRFOIL (rae69ck-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAE6-9CK AIRFOIL (rae69ck-il)
Reynolds number: 50,000
Max Cl/Cd: 25.44 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae69ck-il-50000.txt
Download as CSV file: xf-rae69ck-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE6-9CK AIRFOIL                                
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4902   0.11830   0.11070  -0.0136   1.0000   0.3440
 -10.250  -0.6088   0.09011   0.08292  -0.0438   1.0000   0.1519
 -10.000  -0.6758   0.08073   0.07368  -0.0508   1.0000   0.1402
  -9.750  -0.6833   0.07609   0.06902  -0.0507   1.0000   0.1365
  -9.500  -0.7035   0.07176   0.06471  -0.0499   1.0000   0.1331
  -9.250  -0.7651   0.06740   0.05995  -0.0483   1.0000   0.1245
  -9.000  -0.7670   0.06326   0.05570  -0.0469   1.0000   0.1235
  -8.750  -0.7710   0.05937   0.05161  -0.0452   1.0000   0.1224
  -8.500  -0.7735   0.05557   0.04750  -0.0433   1.0000   0.1213
  -8.250  -0.7729   0.05186   0.04346  -0.0414   1.0000   0.1199
  -8.000  -0.7687   0.04831   0.03950  -0.0395   1.0000   0.1189
  -7.750  -0.7600   0.04502   0.03579  -0.0376   1.0000   0.1191
  -7.500  -0.7493   0.04222   0.03251  -0.0358   1.0000   0.1224
  -7.250  -0.7354   0.03948   0.02931  -0.0341   1.0000   0.1262
  -7.000  -0.7155   0.03687   0.02666  -0.0328   1.0000   0.1312
  -6.750  -0.6961   0.03472   0.02410  -0.0311   1.0000   0.1381
  -6.500  -0.6751   0.03265   0.02206  -0.0297   1.0000   0.1496
  -6.250  -0.6521   0.03076   0.02020  -0.0279   1.0000   0.1621
  -6.000  -0.6324   0.02918   0.01863  -0.0258   1.0000   0.1828
  -5.750  -0.6161   0.02743   0.01718  -0.0233   1.0000   0.2106
  -5.500  -0.6070   0.02516   0.01554  -0.0207   1.0000   0.2590
  -5.250  -0.6117   0.02395   0.01667  -0.0132   1.0000   0.5060
  -5.000  -0.6084   0.02851   0.02117  -0.0003   1.0000   0.6138
  -4.750  -0.5985   0.03140   0.02389   0.0096   1.0000   0.6568
  -4.500  -0.5866   0.03325   0.02554   0.0176   1.0000   0.6928
  -4.250  -0.5702   0.03471   0.02680   0.0246   1.0000   0.7261
  -4.000  -0.5341   0.03647   0.02824   0.0298   1.0000   0.7657
  -3.750  -0.4217   0.03811   0.02921   0.0229   1.0000   0.8242
  -3.250  -0.3497   0.03655   0.02715   0.0192   1.0000   0.8612
  -3.000  -0.3504   0.03577   0.02632   0.0227   1.0000   0.8708
  -2.750  -0.3240   0.03489   0.02526   0.0215   1.0000   0.8793
  -2.500  -0.3393   0.03412   0.02452   0.0272   1.0000   0.8858
  -2.250  -0.3144   0.03331   0.02356   0.0261   1.0000   0.8920
  -2.000  -0.3184   0.03254   0.02277   0.0298   1.0000   0.8982
  -1.750  -0.3100   0.03179   0.02195   0.0314   1.0000   0.9039
  -1.500  -0.2989   0.03110   0.02120   0.0326   1.0000   0.9101
  -1.250  -0.3005   0.03035   0.02043   0.0359   1.0000   0.9164
  -1.000  -0.2743   0.02982   0.01980   0.0343   1.0000   0.9224
  -0.750  -0.2737   0.02915   0.01910   0.0372   1.0000   0.9295
  -0.500  -0.2386   0.02879   0.01867   0.0340   1.0000   0.9353
  -0.250  -0.2271   0.02831   0.01816   0.0349   1.0000   0.9432
   0.000  -0.1879   0.02810   0.01789   0.0308   1.0000   0.9493
   0.250  -0.1661   0.02782   0.01759   0.0298   1.0000   0.9573
   0.500  -0.1256   0.02773   0.01749   0.0254   1.0000   0.9643
   0.750  -0.0939   0.02764   0.01741   0.0224   1.0000   0.9724
   1.000  -0.0565   0.02764   0.01743   0.0183   1.0000   0.9810
   1.250  -0.0187   0.02769   0.01753   0.0140   1.0000   0.9902
   1.500   0.0134   0.02780   0.01769   0.0106   1.0000   1.0000
   1.750   0.0088   0.02745   0.01738   0.0136   1.0000   1.0000
   2.000   0.0031   0.02702   0.01700   0.0167   1.0000   1.0000
   2.250  -0.0029   0.02653   0.01656   0.0198   1.0000   1.0000
   2.500  -0.0056   0.02609   0.01615   0.0222   1.0000   1.0000
   2.750   0.0038   0.02601   0.01610   0.0225   1.0000   1.0000
   3.000   0.0220   0.02625   0.01639   0.0213   1.0000   1.0000
   3.250   0.0443   0.02671   0.01692   0.0194   1.0000   1.0000
   3.500   0.0682   0.02734   0.01761   0.0172   1.0000   1.0000
   3.750   0.0923   0.02809   0.01844   0.0150   1.0000   1.0000
   4.000   0.1233   0.02921   0.01967   0.0115   0.9962   1.0000
   4.250   0.1959   0.03172   0.02245   0.0007   0.9700   1.0000
   4.500   0.2596   0.03346   0.02448  -0.0078   0.9414   1.0000
   4.750   0.3174   0.03459   0.02595  -0.0143   0.9088   1.0000
   5.000   0.3957   0.03488   0.02673  -0.0217   0.8616   1.0000
   5.250   0.5766   0.02598   0.01929  -0.0281   0.7430   1.0000
   5.500   0.6025   0.02368   0.01436  -0.0147   0.2838   1.0000
   5.750   0.6231   0.02637   0.01620  -0.0139   0.2097   1.0000
   6.000   0.6692   0.02884   0.01836  -0.0163   0.1700   1.0000
   6.250   0.7147   0.03137   0.02075  -0.0189   0.1486   1.0000
   6.500   0.7522   0.03399   0.02360  -0.0200   0.1378   1.0000
   6.750   0.7835   0.03686   0.02662  -0.0206   0.1299   1.0000
   7.000   0.8083   0.03948   0.02967  -0.0201   0.1238   1.0000
   7.250   0.8321   0.04272   0.03334  -0.0194   0.1222   1.0000
   7.500   0.8520   0.04610   0.03707  -0.0185   0.1208   1.0000
   7.750   0.8720   0.05023   0.04129  -0.0182   0.1179   1.0000
   8.000   0.8832   0.05391   0.04545  -0.0164   0.1178   1.0000
   8.250   0.8779   0.05760   0.05009  -0.0129   0.1234   1.0000
   8.500   0.8785   0.06243   0.05533  -0.0112   0.1286   1.0000
   8.750   0.8719   0.06734   0.06069  -0.0094   0.1364   1.0000
   9.000   0.8773   0.07355   0.06710  -0.0092   0.1476   1.0000
   9.250   0.8267   0.07856   0.07250  -0.0076   0.1563   1.0000
   9.500   0.8113   0.08804   0.08213  -0.0104   0.1896   1.0000
<< Back to RAE6-9CK AIRFOIL (rae69ck-il)

Polar data table (+)

Polar graphs


<< Back to RAE6-9CK AIRFOIL (rae69ck-il)