XFOIL         Version 6.94
  
 Calculated polar for: FX 67-K-150/17 AIRFOIL                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5368   0.01101   0.00614  -0.0976  0.6832  0.8792
   1.000   0.5915   0.01103   0.00609  -0.0978  0.6795  0.8903
   1.500   0.6452   0.01099   0.00601  -0.0977  0.6763  0.9005
   2.000   0.7016   0.01117   0.00613  -0.0984  0.6723  0.9096
   2.500   0.7492   0.01127   0.00633  -0.0975  0.6680  0.9195
   3.000   0.8004   0.01134   0.00645  -0.0974  0.6629  0.9295
   3.500   0.8535   0.01125   0.00640  -0.0973  0.6584  0.9397
   4.000   0.9112   0.01116   0.00636  -0.0983  0.6543  0.9501
   4.500   0.9663   0.01101   0.00632  -0.0988  0.6456  0.9631
   5.000   1.0332   0.00993   0.00517  -0.1010  0.6232  0.9742
   5.500   1.0984   0.00956   0.00473  -0.1037  0.5840  0.9926
   6.000   1.1441   0.00983   0.00503  -0.1030  0.5569  1.0000
   6.500   1.1366   0.01124   0.00593  -0.0925  0.4610  1.0000
   7.000   1.0970   0.01417   0.00849  -0.0786  0.3857  1.0000
   7.500   1.0585   0.01841   0.01226  -0.0673  0.2973  1.0000
   8.000   1.0184   0.02340   0.01659  -0.0572  0.1801  1.0000
   8.500   0.9758   0.02930   0.02160  -0.0482  0.0286  1.0000
   9.000   0.9903   0.03208   0.02437  -0.0451  0.0032  1.0000
   9.500   1.0118   0.03454   0.02698  -0.0429  0.0026  1.0000
  10.000   1.0305   0.03731   0.02993  -0.0406  0.0025  1.0000
  10.500   1.0439   0.04068   0.03347  -0.0381  0.0025  1.0000
  11.000   1.0552   0.04440   0.03733  -0.0358  0.0026  1.0000
  11.500   1.0651   0.04842   0.04150  -0.0337  0.0026  1.0000
  12.000   1.0729   0.05283   0.04607  -0.0316  0.0029  1.0000
  12.500   1.0864   0.05675   0.05015  -0.0294  0.0033  1.0000
  13.000   1.1298   0.05806   0.05148  -0.0272  0.0041  1.0000