XFOIL         Version 6.96
  
 Calculated polar for: AH 79-100 B AIRFOIL
                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -4.000  -0.0898   0.07032   0.06606  -0.0907   0.8995   0.1085
  -2.500   0.2462   0.04067   0.03422  -0.1577   0.8754   0.1314
  -2.000   0.3779   0.03354   0.02543  -0.1724   0.8717   0.0973
  -1.500   0.4486   0.03207   0.02332  -0.1754   0.8608   0.1026
  -1.000   0.5391   0.02974   0.02104  -0.1816   0.8552   0.1244
   1.000   0.7851   0.02764   0.02049  -0.1843   0.8082   1.0000
   1.500   0.8568   0.02778   0.02038  -0.1872   0.7968   1.0000
   2.000   0.9187   0.02820   0.02066  -0.1884   0.7829   1.0000
   3.500   1.1051   0.02881   0.02121  -0.1909   0.7408   1.0000
   4.500   1.2325   0.02846   0.02105  -0.1921   0.7113   1.0000
   5.500   1.3622   0.02754   0.02045  -0.1931   0.6787   1.0000
   6.000   1.4317   0.02668   0.01979  -0.1940   0.6613   1.0000
   6.500   1.4760   0.02706   0.02044  -0.1914   0.6372   1.0000
   7.000   1.5351   0.02591   0.01949  -0.1900   0.6084   1.0000
   7.500   1.5925   0.02461   0.01824  -0.1881   0.5723   1.0000
   8.000   1.6264   0.02501   0.01893  -0.1834   0.5344   1.0000
   8.500   1.6544   0.02541   0.01942  -0.1774   0.4859   1.0000
   9.000   1.6613   0.02666   0.02069  -0.1684   0.4273   1.0000
   9.500   1.6461   0.02923   0.02307  -0.1569   0.3539   1.0000
  10.000   1.6117   0.03429   0.02753  -0.1455   0.2479   1.0000