XFOIL         Version 6.94
  
 Calculated polar for: agora02                                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.030 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.1287   0.03987   0.02957  -0.0194  0.9988  0.4188
  -2.750  -0.1032   0.03918   0.02889  -0.0196  0.9988  0.4328
  -2.500  -0.0765   0.03879   0.02842  -0.0200  0.9988  0.4499
  -2.250  -0.0504   0.03830   0.02805  -0.0201  0.9988  0.4638
  -2.000  -0.0221   0.03807   0.02778  -0.0207  0.9988  0.4771
  -1.750   0.0057   0.03780   0.02770  -0.0211  0.9988  0.4919
  -1.500   0.1906   0.03314   0.02472  -0.0439  0.9074  0.6342
  -1.250   0.2050   0.03391   0.02636  -0.0351  0.8441  1.0012
  -1.000   0.2028   0.03709   0.02902  -0.0244  0.7770  1.0012
  -0.750   0.1952   0.04072   0.03231  -0.0126  0.6990  1.0012
  -0.500   0.1903   0.04508   0.03635  -0.0024  0.6111  1.0012
  -0.250   0.2509   0.04219   0.03349  -0.0106  0.4643  1.0012
   0.000   0.2683   0.04054   0.03139  -0.0037  0.4408  1.0012
   0.250   0.2858   0.03941   0.02972   0.0025  0.4206  1.0012
   0.500   0.3032   0.03884   0.02857   0.0083  0.4032  1.0012
   0.750   0.3318   0.03966   0.02911   0.0093  0.3862  1.0012
   1.000   0.3582   0.04044   0.02958   0.0110  0.3734  1.0012
   1.250   0.3819   0.04152   0.03027   0.0135  0.3636  1.0012
   1.500   0.4171   0.04335   0.03220   0.0111  0.3491  1.0012
   1.750   0.4426   0.04474   0.03333   0.0122  0.3322  1.0012
   2.000   0.4682   0.04698   0.03532   0.0129  0.3212  1.0012
   2.250   0.5061   0.04986   0.03849   0.0088  0.3165  1.0012
   2.500   0.5459   0.05417   0.04314   0.0022  0.3130  1.0012
   2.750   0.5856   0.06092   0.05030  -0.0082  0.3125  1.0012
   3.000   0.6166   0.06904   0.05869  -0.0189  0.3158  1.0012
   3.250   0.6384   0.07584   0.06557  -0.0250  0.3200  1.0012