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AG12 (ag12-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG12 (ag12-il)
Reynolds number: 100,000
Max Cl/Cd: 46.97 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag12-il-100000-n5.txt
Download as CSV file: xf-ag12-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4633   0.10348   0.09885   0.0037   1.0000   0.0490
  -8.750  -0.4670   0.09967   0.09509   0.0008   1.0000   0.0495
  -8.500  -0.5614   0.10512   0.10029   0.0149   1.0000   0.0407
  -8.250  -0.5556   0.10176   0.09695   0.0137   1.0000   0.0428
  -8.000  -0.5516   0.09802   0.09325   0.0116   1.0000   0.0441
  -7.500  -0.4627   0.07292   0.06847  -0.0054   1.0000   0.0249
  -7.250  -0.5372   0.08012   0.07545  -0.0047   1.0000   0.0251
  -7.000  -0.5280   0.07538   0.07072  -0.0083   1.0000   0.0243
  -6.750  -0.5156   0.06988   0.06521  -0.0138   1.0000   0.0236
  -6.500  -0.4998   0.06361   0.05889  -0.0204   1.0000   0.0229
  -6.250  -0.4810   0.05661   0.05176  -0.0274   1.0000   0.0222
  -6.000  -0.4598   0.04923   0.04412  -0.0337   1.0000   0.0216
  -5.750  -0.4367   0.04232   0.03680  -0.0386   1.0000   0.0210
  -5.500  -0.4118   0.03624   0.03016  -0.0420   1.0000   0.0207
  -5.250  -0.3856   0.03132   0.02459  -0.0439   1.0000   0.0206
  -5.000  -0.3586   0.02749   0.02013  -0.0448   1.0000   0.0209
  -4.750  -0.3312   0.02451   0.01658  -0.0451   1.0000   0.0215
  -4.500  -0.3039   0.02229   0.01391  -0.0450   1.0000   0.0233
  -4.250  -0.2766   0.02065   0.01189  -0.0447   1.0000   0.0264
  -4.000  -0.2504   0.01883   0.00995  -0.0445   1.0000   0.0292
  -3.750  -0.2237   0.01748   0.00842  -0.0441   1.0000   0.0328
  -3.500  -0.1971   0.01635   0.00716  -0.0438   1.0000   0.0400
  -3.250  -0.1702   0.01535   0.00611  -0.0435   1.0000   0.0521
  -3.000  -0.1435   0.01443   0.00528  -0.0434   1.0000   0.0805
  -2.750  -0.1169   0.01355   0.00466  -0.0433   1.0000   0.1380
  -2.500  -0.0908   0.01284   0.00427  -0.0433   1.0000   0.2216
  -2.250  -0.0649   0.01229   0.00399  -0.0431   1.0000   0.3113
  -2.000  -0.0394   0.01184   0.00380  -0.0427   1.0000   0.3987
  -1.750  -0.0139   0.01143   0.00364  -0.0421   1.0000   0.4872
  -1.500   0.0109   0.01101   0.00351  -0.0414   1.0000   0.5812
  -1.250   0.0332   0.01049   0.00337  -0.0398   1.0000   0.7005
  -1.000   0.0566   0.00983   0.00317  -0.0380   1.0000   1.0000
  -0.750   0.0828   0.00989   0.00307  -0.0378   1.0000   1.0000
  -0.500   0.1088   0.00996   0.00303  -0.0377   1.0000   1.0000
  -0.250   0.1503   0.01005   0.00297  -0.0407   0.9837   1.0000
   0.000   0.1917   0.01011   0.00293  -0.0435   0.9610   1.0000
   0.250   0.2316   0.01017   0.00290  -0.0458   0.9347   1.0000
   0.500   0.2694   0.01023   0.00287  -0.0475   0.9040   1.0000
   0.750   0.3037   0.01030   0.00285  -0.0483   0.8692   1.0000
   1.000   0.3341   0.01041   0.00284  -0.0482   0.8315   1.0000
   1.250   0.3614   0.01055   0.00285  -0.0474   0.7922   1.0000
   1.500   0.3876   0.01074   0.00289  -0.0465   0.7525   1.0000
   1.750   0.4134   0.01095   0.00298  -0.0455   0.7128   1.0000
   2.000   0.4392   0.01120   0.00307  -0.0446   0.6734   1.0000
   2.250   0.4652   0.01146   0.00319  -0.0439   0.6345   1.0000
   2.500   0.4911   0.01176   0.00334  -0.0431   0.5964   1.0000
   2.750   0.5170   0.01207   0.00352  -0.0425   0.5585   1.0000
   3.000   0.5429   0.01240   0.00377  -0.0419   0.5211   1.0000
   3.250   0.5688   0.01275   0.00401  -0.0413   0.4846   1.0000
   3.500   0.5947   0.01312   0.00428  -0.0408   0.4484   1.0000
   3.750   0.6206   0.01351   0.00458  -0.0403   0.4123   1.0000
   4.000   0.6464   0.01394   0.00497  -0.0399   0.3768   1.0000
   4.250   0.6722   0.01439   0.00535  -0.0395   0.3411   1.0000
   4.500   0.6978   0.01487   0.00576  -0.0391   0.3058   1.0000
   4.750   0.7233   0.01540   0.00622  -0.0387   0.2708   1.0000
   5.000   0.7487   0.01596   0.00674  -0.0383   0.2371   1.0000
   5.250   0.7740   0.01659   0.00737  -0.0380   0.2045   1.0000
   5.500   0.7989   0.01729   0.00801  -0.0377   0.1739   1.0000
   5.750   0.8237   0.01804   0.00874  -0.0373   0.1464   1.0000
   6.000   0.8481   0.01888   0.00954  -0.0369   0.1202   1.0000
   6.250   0.8724   0.01978   0.01051  -0.0365   0.0987   1.0000
   6.500   0.8960   0.02081   0.01152  -0.0361   0.0789   1.0000
   6.750   0.9192   0.02199   0.01285  -0.0355   0.0640   1.0000
   7.000   0.9420   0.02323   0.01418  -0.0349   0.0506   1.0000
   7.250   0.9633   0.02482   0.01587  -0.0341   0.0421   1.0000
   7.500   0.9846   0.02627   0.01743  -0.0336   0.0338   1.0000
   7.750   1.0050   0.02830   0.01975  -0.0324   0.0296   1.0000
   8.000   1.0254   0.02992   0.02155  -0.0318   0.0250   1.0000
   8.250   1.0428   0.03237   0.02421  -0.0309   0.0219   1.0000
   8.500   1.0605   0.03517   0.02744  -0.0297   0.0203   1.0000
   8.750   1.0757   0.03842   0.03114  -0.0286   0.0190   1.0000
   9.000   1.0879   0.04184   0.03501  -0.0276   0.0179   1.0000
   9.250   1.0982   0.04475   0.03836  -0.0268   0.0166   1.0000
   9.500   1.1057   0.04749   0.04132  -0.0262   0.0153   1.0000
  10.000   1.0956   0.05649   0.05103  -0.0253   0.0143   1.0000
  10.250   1.0826   0.06084   0.05569  -0.0253   0.0142   1.0000
  10.500   1.0660   0.06600   0.06109  -0.0276   0.0142   1.0000
  10.750   1.0488   0.07258   0.06786  -0.0325   0.0143   1.0000
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