XFOIL         Version 6.94
  
 Calculated polar for: LISSAMAN 7769 AIRFOIL                           
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.040 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0279   0.05470   0.04756  -0.0738  0.7898  0.2949
  -2.750  -0.0033   0.05291   0.04520  -0.0775  0.7885  0.2552
  -2.500   0.0143   0.05184   0.04369  -0.0788  0.7885  0.2373
  -2.250   0.0266   0.05102   0.04260  -0.0789  0.7902  0.2300
  -2.000   0.0398   0.05051   0.04174  -0.0789  0.7921  0.2271
  -1.750   0.0527   0.05009   0.04108  -0.0785  0.7937  0.2283
  -1.500   0.0673   0.04974   0.04048  -0.0782  0.7953  0.2300
  -1.250   0.0802   0.04951   0.04000  -0.0775  0.7977  0.2310
  -1.000   0.0588   0.05004   0.04046  -0.0730  0.8102  0.2305
  -0.750   0.0755   0.04995   0.04009  -0.0731  0.8176  0.2339
  -0.500   0.0733   0.05030   0.04025  -0.0710  0.8317  0.2374
  -0.250   0.0698   0.05065   0.04035  -0.0687  0.8501  0.2423