XFOIL         Version 6.94
  
 Calculated polar for: agora10                                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.035 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0358   0.03532   0.02594  -0.0285  0.9066  0.4007
  -2.750  -0.0134   0.03548   0.02610  -0.0262  0.8783  0.4132
  -2.500   0.0085   0.03607   0.02651  -0.0238  0.8523  0.4273
  -2.250   0.0276   0.03655   0.02703  -0.0206  0.8245  0.4394
  -2.000   0.0483   0.03724   0.02768  -0.0177  0.7979  0.4518
  -1.750   0.0707   0.03787   0.02836  -0.0153  0.7708  0.4652
  -1.500   0.0923   0.03860   0.02928  -0.0128  0.7366  0.4842
  -1.250   0.1071   0.04016   0.03109  -0.0081  0.6955  0.5081
  -1.000   0.1212   0.04280   0.03415  -0.0037  0.6329  0.5474
  -0.750   0.1314   0.04653   0.03891   0.0008  0.5520  0.6663
  -0.500   0.1618   0.05047   0.04276   0.0001  0.4769  1.0008
  -0.250   0.1903   0.05332   0.04513  -0.0006  0.4256  1.0008
   0.000   0.2276   0.05463   0.04620  -0.0036  0.3968  1.0008
   0.250   0.2679   0.05511   0.04654  -0.0071  0.3767  1.0008
   0.500   0.2984   0.05515   0.04636  -0.0067  0.3603  1.0008
   0.750   0.3295   0.05603   0.04708  -0.0072  0.3478  1.0008
   1.000   0.3588   0.05703   0.04792  -0.0072  0.3380  1.0008
   1.250   0.3873   0.05788   0.04862  -0.0068  0.3277  1.0008
   1.500   0.4180   0.05961   0.05027  -0.0083  0.3176  1.0008
   1.750   0.4503   0.06290   0.05354  -0.0118  0.3123  1.0008
   2.000   0.4824   0.06707   0.05773  -0.0167  0.3076  1.0008
   2.250   0.5131   0.07230   0.06303  -0.0236  0.3033  1.0008
   2.500   0.5396   0.07831   0.06910  -0.0310  0.3037  1.0008
   2.750   0.5595   0.08588   0.07676  -0.0403  0.3130  1.0008