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 =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0129   0.05302   0.04610  -0.0710  0.7556  0.2954
  -2.750   0.0225   0.05055   0.04287  -0.0767  0.7540  0.2335
  -2.500   0.0462   0.04927   0.04112  -0.0786  0.7537  0.2156
  -2.250   0.0696   0.04881   0.03993  -0.0800  0.7540  0.2037
  -2.000   0.0855   0.04797   0.03894  -0.0803  0.7544  0.2049
  -1.750   0.0996   0.04743   0.03834  -0.0802  0.7548  0.2092
  -1.500   0.1147   0.04723   0.03794  -0.0800  0.7555  0.2131
  -1.250   0.1315   0.04712   0.03754  -0.0799  0.7571  0.2162
  -1.000   0.1497   0.04726   0.03730  -0.0800  0.7604  0.2220
  -0.750   0.1277   0.04829   0.03831  -0.0765  0.7746  0.2220
  -0.500   0.1504   0.04816   0.03811  -0.0772  0.7780  0.2369
  -0.250   0.1088   0.04924   0.03925  -0.0717  0.8087  0.2297
   0.000   0.0890   0.05002   0.04001  -0.0688  0.8517  0.2307