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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.000   0.1686   0.00244  -0.00463  -0.0210   0.3791   0.0292
  -0.500   0.2150   0.00225  -0.00501  -0.0190   0.3362   0.0294
   0.000   0.2623   0.00218  -0.00530  -0.0173   0.2906   0.0300
   0.500   0.3120   0.00215  -0.00544  -0.0159   0.2683   0.0309
   1.000   0.3627   0.00215  -0.00552  -0.0148   0.2537   0.0322
   1.500   0.4144   0.00214  -0.00556  -0.0139   0.2430   0.0341
   2.000   0.4660   0.00214  -0.00552  -0.0130   0.2328   0.0511
   4.500   0.7577   0.00150  -0.00401  -0.0165   0.1962   0.9858
   5.000   0.8277   0.00158  -0.00398  -0.0200   0.1902   0.9908
   5.500   0.9014   0.00162  -0.00393  -0.0242   0.1858   0.9956
   6.000   0.9679   0.00171  -0.00388  -0.0271   0.1793   0.9988
   6.500   1.0186   0.00180  -0.00378  -0.0267   0.1741   1.0000
   7.000   1.0580   0.00189  -0.00366  -0.0239   0.1695   1.0000
   7.500   1.0951   0.00204  -0.00353  -0.0207   0.1627   1.0000
   8.000   1.1343   0.00217  -0.00336  -0.0178   0.1571   1.0000
   8.500   1.1715   0.00236  -0.00320  -0.0147   0.1486   1.0000
   9.000   1.2099   0.00255  -0.00297  -0.0119   0.1421   1.0000
   9.500   1.2448   0.00285  -0.00270  -0.0086   0.1347   1.0000
  10.000   1.2827   0.00307  -0.00242  -0.0059   0.1300   1.0000