XFOIL         Version 6.94
  
 Calculated polar for: GOE 415 AIRFOIL                                 
  
 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   0.1940   0.00772   0.00235  -0.0299  0.5758  0.9850
   0.500   0.2590   0.00806   0.00231  -0.0317  0.5356  1.0000
   1.000   0.3115   0.00836   0.00235  -0.0307  0.5028  1.0000
   1.500   0.3640   0.00868   0.00244  -0.0298  0.4724  1.0000
   2.000   0.4167   0.00897   0.00256  -0.0289  0.4422  1.0000
   2.500   0.4695   0.00929   0.00271  -0.0280  0.4130  1.0000
   3.000   0.5224   0.00963   0.00291  -0.0272  0.3829  1.0000
   3.500   0.5752   0.01002   0.00316  -0.0264  0.3546  1.0000
   4.000   0.6280   0.01046   0.00348  -0.0256  0.3288  1.0000
   4.500   0.6808   0.01092   0.00386  -0.0248  0.3075  1.0000
   5.000   0.7335   0.01143   0.00430  -0.0241  0.2893  1.0000
   5.500   0.7864   0.01192   0.00481  -0.0234  0.2730  1.0000
   6.000   0.8389   0.01245   0.00536  -0.0227  0.2575  1.0000
   6.500   0.8912   0.01299   0.00593  -0.0220  0.2412  1.0000
   7.000   0.9432   0.01351   0.00653  -0.0213  0.2219  1.0000
   7.500   0.9946   0.01409   0.00720  -0.0205  0.1977  1.0000
   8.000   1.0444   0.01487   0.00796  -0.0196  0.1658  1.0000
   8.500   1.0930   0.01583   0.00889  -0.0186  0.1381  1.0000
   9.000   1.1401   0.01693   0.01000  -0.0176  0.1183  1.0000
   9.500   1.1851   0.01822   0.01135  -0.0163  0.1045  1.0000
  10.000   1.2293   0.01951   0.01278  -0.0150  0.0908  1.0000
  10.500   1.2726   0.02085   0.01424  -0.0137  0.0703  1.0000
  11.000   1.3061   0.02324   0.01653  -0.0115  0.0388  1.0000
  11.500   1.3310   0.02609   0.01951  -0.0086  0.0306  1.0000
  12.000   1.3439   0.02935   0.02297  -0.0049  0.0268  1.0000
  12.500   1.3429   0.03342   0.02725  -0.0015  0.0246  1.0000
  13.000   1.3358   0.03940   0.03345  -0.0017  0.0227  1.0000
  13.500   1.3273   0.04735   0.04171  -0.0052  0.0215  1.0000
  14.000   1.3145   0.05647   0.05109  -0.0095  0.0204  1.0000
  14.500   1.2976   0.06572   0.06049  -0.0131  0.0197  1.0000
  15.000   1.2768   0.07564   0.07065  -0.0169  0.0186  1.0000
  15.500   1.2581   0.08548   0.08072  -0.0209  0.0178  1.0000
  16.000   1.2462   0.09416   0.08944  -0.0240  0.0166  1.0000
  17.000   1.2054   0.11740   0.11323  -0.0351  0.0147  1.0000
  17.500   0.8734   0.10436   0.10094  -0.0109  0.0214  1.0000