XFOIL         Version 6.94
  
 Calculated polar for: NYU/GRUMMAN K-1 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3004   0.00804   0.00339  -0.0875  0.8608  0.6433
   0.500   0.3557   0.00786   0.00329  -0.0866  0.8448  0.6554
   1.000   0.4124   0.00781   0.00326  -0.0860  0.8279  0.6667
   1.500   0.4687   0.00769   0.00316  -0.0854  0.8042  0.6767
   2.000   0.5244   0.00769   0.00317  -0.0846  0.7679  0.6862
   2.500   0.5767   0.00806   0.00314  -0.0831  0.6653  0.6960
   3.000   0.6140   0.01036   0.00389  -0.0802  0.3237  0.7046
   3.500   0.6625   0.01160   0.00454  -0.0792  0.2001  0.7138
   4.000   0.7157   0.01229   0.00507  -0.0788  0.1650  0.7226
   4.500   0.7693   0.01277   0.00550  -0.0784  0.1429  0.7316
   5.000   0.8224   0.01328   0.00598  -0.0778  0.1218  0.7404
   5.500   0.8756   0.01386   0.00648  -0.0774  0.0966  0.7490
   6.000   0.9251   0.01483   0.00718  -0.0765  0.0498  0.7581
   6.500   0.9736   0.01593   0.00824  -0.0752  0.0359  0.7669
   7.000   1.0227   0.01701   0.00935  -0.0741  0.0309  0.7759
   7.500   1.0705   0.01805   0.01048  -0.0728  0.0277  0.7857
   8.000   1.1169   0.01920   0.01172  -0.0713  0.0251  0.7953
   9.500   1.2431   0.02372   0.01663  -0.0652  0.0196  0.8291
  11.500   1.3522   0.03285   0.02690  -0.0497  0.0159  1.0000
  12.500   1.3781   0.03966   0.03424  -0.0414  0.0148  1.0000
  13.000   1.3821   0.04403   0.03882  -0.0376  0.0144  1.0000
  13.500   1.3714   0.05027   0.04539  -0.0340  0.0142  1.0000
  14.000   1.3475   0.05788   0.05348  -0.0322  0.0141  1.0000
  14.500   1.3152   0.06785   0.06392  -0.0336  0.0140  1.0000
  15.000   1.2779   0.08058   0.07703  -0.0389  0.0141  1.0000
  15.500   1.2360   0.09700   0.09382  -0.0488  0.0141  1.0000