XFOIL         Version 6.94
  
 Calculated polar for: rt                                              
  
 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  -0.2775   0.04643   0.03488   0.0512  0.9991  0.7487
  -2.500  -0.2203   0.04781   0.03641   0.0470  0.9991  0.7773
  -2.250  -0.1864   0.04861   0.03741   0.0437  0.9991  0.7993
  -2.000  -0.1441   0.04970   0.03879   0.0381  0.9991  0.8266
  -1.500   0.2113   0.05148   0.04096  -0.0255  0.9000  1.0009
  -1.250   0.4323   0.04208   0.03144  -0.0543  0.7805  1.0009
  -1.000   0.5493   0.03657   0.02493  -0.0638  0.6444  1.0009
  -0.750   0.5860   0.03672   0.02426  -0.0642  0.5841  1.0009
  -0.500   0.6157   0.03738   0.02433  -0.0643  0.5493  1.0009
  -0.250   0.6421   0.03827   0.02484  -0.0641  0.5244  1.0009
   0.000   0.6712   0.03914   0.02543  -0.0646  0.5073  1.0009
   0.250   0.6962   0.04033   0.02653  -0.0646  0.4933  1.0009
   0.500   0.7286   0.04115   0.02706  -0.0657  0.4835  1.0009
   0.750   0.7495   0.04271   0.02870  -0.0653  0.4741  1.0009
   1.000   0.7752   0.04400   0.02997  -0.0656  0.4674  1.0009
   1.250   0.8061   0.04505   0.03083  -0.0666  0.4616  1.0009
   1.500   0.8205   0.04711   0.03305  -0.0653  0.4569  1.0009
   1.750   0.8339   0.04926   0.03535  -0.0640  0.4518  1.0009
   2.000   0.8492   0.05135   0.03749  -0.0629  0.4478  1.0009
   2.250   0.8667   0.05330   0.03946  -0.0621  0.4441  1.0009
   2.500   0.8868   0.05521   0.04135  -0.0617  0.4410  1.0009
   3.000   0.8922   0.06193   0.04834  -0.0574  0.4391  1.0009
   3.250   0.8841   0.06634   0.05289  -0.0545  0.4396  1.0009
   3.500   0.6594   0.09008   0.07698  -0.0395  0.4556  1.0009
   3.750   0.6265   0.09847   0.08536  -0.0399  0.4644  1.0009
   4.000   0.5884   0.10736   0.09422  -0.0407  0.4795  1.0009