XFOIL         Version 6.94
  
 Calculated polar for: GRUMMAN K-3 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.5105   0.01455   0.00956  -0.1362  0.8694  0.5977
   0.500   0.5711   0.01405   0.00915  -0.1360  0.8555  0.6047
   1.000   0.6396   0.01340   0.00855  -0.1373  0.8384  0.6126
   1.500   0.7015   0.01293   0.00809  -0.1374  0.8104  0.6185
   2.000   0.7630   0.01241   0.00750  -0.1373  0.7569  0.6253
   2.500   0.7835   0.01364   0.00753  -0.1287  0.5183  0.6313
   3.000   0.7788   0.01594   0.00874  -0.1170  0.3122  0.6366
   3.500   0.8029   0.01749   0.00971  -0.1111  0.2145  0.6426
   4.000   0.8440   0.01859   0.01057  -0.1084  0.1775  0.6492
   4.500   0.8877   0.01965   0.01147  -0.1062  0.1557  0.6551
   5.000   0.9341   0.02037   0.01218  -0.1045  0.1395  0.6622
   5.500   0.9778   0.02126   0.01305  -0.1023  0.1260  0.6688
   6.000   1.0207   0.02225   0.01400  -0.1000  0.1136  0.6757
   6.500   1.0655   0.02313   0.01492  -0.0980  0.1010  0.6833
   7.500   1.1515   0.02506   0.01686  -0.0938  0.0735  0.6988
   8.000   1.1897   0.02637   0.01817  -0.0909  0.0599  0.7068
   8.500   1.2242   0.02805   0.01984  -0.0876  0.0506  0.7153
   9.000   1.2561   0.02997   0.02173  -0.0842  0.0446  0.7225
  10.000   1.3127   0.03408   0.02609  -0.0767  0.0374  0.7442
  10.500   1.3378   0.03663   0.02872  -0.0730  0.0353  0.7540
  11.000   1.3611   0.03912   0.03138  -0.0692  0.0335  0.7662
  11.500   1.3819   0.04211   0.03448  -0.0655  0.0322  0.7791
  12.000   1.4016   0.04514   0.03777  -0.0619  0.0310  0.7932
  12.500   1.4194   0.04840   0.04119  -0.0584  0.0300  0.8105
  13.000   1.4367   0.05188   0.04477  -0.0550  0.0293  0.8323
  13.500   1.4410   0.05605   0.04939  -0.0509  0.0287  0.8629
  14.000   1.4271   0.05963   0.05338  -0.0439  0.0282  0.9810
  14.500   1.4311   0.06533   0.05932  -0.0429  0.0276  1.0001
  15.000   1.4297   0.07174   0.06597  -0.0422  0.0271  1.0001
  15.500   1.4253   0.07876   0.07321  -0.0423  0.0267  1.0001
  16.000   1.4209   0.08606   0.08067  -0.0431  0.0263  1.0001
  16.500   1.4098   0.09473   0.08957  -0.0451  0.0261  1.0001
  17.000   1.3787   0.10735   0.10262  -0.0510  0.0260  1.0001
  17.500   1.3416   0.12244   0.11815  -0.0603  0.0260  1.0001