XFOIL         Version 6.94
  
 Calculated polar for: E176  (8.83%)                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3331   0.00791   0.00254  -0.0652  0.6898  1.0000
   0.500   0.3871   0.00808   0.00255  -0.0647  0.6770  1.0000
   1.000   0.4414   0.00828   0.00263  -0.0644  0.6648  1.0000
   1.500   0.4959   0.00850   0.00275  -0.0640  0.6532  1.0000
   2.000   0.5505   0.00875   0.00287  -0.0637  0.6415  1.0000
   2.500   0.6051   0.00893   0.00305  -0.0634  0.6276  1.0000
   3.000   0.6596   0.00911   0.00322  -0.0630  0.6125  1.0000
   3.500   0.7139   0.00929   0.00342  -0.0626  0.5962  1.0000
   4.000   0.7680   0.00946   0.00362  -0.0621  0.5792  1.0000
   4.500   0.8219   0.00963   0.00385  -0.0616  0.5595  1.0000
   5.000   0.8748   0.00974   0.00405  -0.0609  0.5308  1.0000
   5.500   0.9257   0.00992   0.00421  -0.0598  0.4781  1.0000
   6.000   0.9722   0.01069   0.00468  -0.0583  0.3761  1.0000
   6.500   1.0072   0.01298   0.00606  -0.0559  0.1988  1.0000
   7.000   1.0330   0.01636   0.00832  -0.0527  0.0290  1.0000
   7.500   1.0691   0.01842   0.01052  -0.0498  0.0169  1.0000
   8.000   1.1030   0.02042   0.01282  -0.0467  0.0154  1.0000
   8.500   1.1286   0.02292   0.01554  -0.0427  0.0146  1.0000
   9.000   1.1482   0.02584   0.01867  -0.0380  0.0144  1.0000
   9.500   1.1710   0.02918   0.02225  -0.0338  0.0142  1.0000
  10.000   1.1968   0.03247   0.02578  -0.0306  0.0133  1.0000
  10.500   1.2211   0.03659   0.03028  -0.0277  0.0128  1.0000
  11.000   1.2361   0.04292   0.03723  -0.0242  0.0133  1.0000
  11.500   1.2188   0.05154   0.04662  -0.0200  0.0146  1.0000
  12.000   1.1903   0.06039   0.05600  -0.0184  0.0155  1.0000
  12.500   1.1549   0.07072   0.06674  -0.0197  0.0162  1.0000