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
  -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   0.01709   0.00702  -0.0208   0.5744   0.1270
   1.500   0.5658   0.01673   0.00748  -0.0414   0.4193   1.0000
   2.000   0.6071   0.01758   0.00795  -0.0389   0.3938   1.0000
   2.500   0.6494   0.01849   0.00856  -0.0365   0.3738   1.0000
   3.000   0.6915   0.01937   0.00929  -0.0341   0.3575   1.0000
   3.500   0.7330   0.02027   0.01010  -0.0315   0.3424   1.0000
   4.000   0.7738   0.02119   0.01096  -0.0289   0.3279   1.0000
   4.500   0.8154   0.02227   0.01197  -0.0265   0.3159   1.0000
   5.000   0.8593   0.02345   0.01298  -0.0245   0.3051   1.0000
   5.500   0.8972   0.02471   0.01448  -0.0215   0.2939   1.0000
   6.000   0.9374   0.02621   0.01600  -0.0191   0.2836   1.0000
   6.500   0.9783   0.02767   0.01749  -0.0168   0.2736   1.0000
   7.000   1.0116   0.02980   0.01993  -0.0136   0.2643   1.0000
   7.500   1.0525   0.03171   0.02185  -0.0115   0.2559   1.0000
   8.000   1.0750   0.03471   0.02534  -0.0072   0.2483   1.0000
   8.500   1.1083   0.03711   0.02794  -0.0045   0.2412   1.0000
   9.000   1.1313   0.04057   0.03168  -0.0010   0.2353   1.0000
   9.500   1.1227   0.04574   0.03751   0.0054   0.2304   1.0000
  10.000   1.1039   0.05182   0.04404   0.0115   0.2270   1.0000