XFOIL         Version 6.96
  
 Calculated polar for: BOEING VERTOL V43012-1.58 AIRFOIL
              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -4.500  -0.4155   0.04412   0.03835  -0.0053   0.9885   0.1462
  -3.000  -0.1746   0.02767   0.02108  -0.0230   0.9155   0.2729
  -2.500  -0.0469   0.02606   0.01789  -0.0288   0.8939   0.1581
  -2.000   0.0505   0.02355   0.01480  -0.0314   0.8630   0.1105
  -1.500   0.1273   0.02139   0.01232  -0.0321   0.8225   0.1017
  -1.000   0.1938   0.01967   0.01047  -0.0315   0.7732   0.0997
  -0.500   0.2471   0.01830   0.00905  -0.0289   0.7140   0.1008
   0.000   0.2897   0.01741   0.00788  -0.0247   0.6481   0.1100
   0.500   0.3332   0.01710   0.00702  -0.0208   0.5744   0.1269
   4.000   0.7738   0.02119   0.01096  -0.0289   0.3279   1.0000