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AG19 (ag19-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG19 (ag19-il)
Reynolds number: 200,000
Max Cl/Cd: 66.09 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag19-il-200000.txt
Download as CSV file: xf-ag19-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG19                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5241   0.10298   0.09958   0.0053   1.0000   0.0272
  -7.750  -0.5208   0.09779   0.09447   0.0023   1.0000   0.0274
  -7.500  -0.5191   0.09234   0.08902   0.0042   1.0000   0.0285
  -7.250  -0.5106   0.08888   0.08558   0.0042   1.0000   0.0294
  -7.000  -0.5006   0.08530   0.08202   0.0021   1.0000   0.0304
  -6.750  -0.4890   0.08147   0.07820  -0.0013   1.0000   0.0316
  -6.500  -0.4751   0.07732   0.07406  -0.0058   1.0000   0.0330
  -6.250  -0.4580   0.07276   0.06950  -0.0116   1.0000   0.0348
  -6.000  -0.4326   0.06729   0.06399  -0.0207   1.0000   0.0374
  -5.750  -0.3900   0.06090   0.05731  -0.0346   1.0000   0.0391
  -5.500  -0.3758   0.05343   0.04979  -0.0387   1.0000   0.0403
  -5.250  -0.3626   0.05070   0.04713  -0.0380   1.0000   0.0422
  -5.000  -0.3395   0.04759   0.04390  -0.0402   1.0000   0.0483
  -4.750  -0.3087   0.04164   0.03763  -0.0458   1.0000   0.0546
  -4.500  -0.2862   0.03889   0.03482  -0.0466   1.0000   0.0587
  -4.250  -0.2542   0.03417   0.02960  -0.0502   1.0000   0.0674
  -4.000  -0.2308   0.03168   0.02710  -0.0505   1.0000   0.0714
  -3.750  -0.2024   0.02858   0.02363  -0.0518   1.0000   0.0819
  -3.500  -0.1609   0.02007   0.01385  -0.0520   1.0000   0.0334
  -3.250  -0.1307   0.01691   0.01000  -0.0514   1.0000   0.0295
  -3.000  -0.1031   0.01511   0.00792  -0.0509   1.0000   0.0297
  -2.750  -0.0762   0.01377   0.00646  -0.0504   1.0000   0.0322
  -2.500  -0.0498   0.01329   0.00583  -0.0498   1.0000   0.0391
  -2.250  -0.0224   0.01172   0.00430  -0.0496   1.0000   0.0476
  -2.000   0.0053   0.01080   0.00344  -0.0495   1.0000   0.0747
  -1.750   0.0337   0.00960   0.00287  -0.0499   1.0000   0.2093
  -1.500   0.0607   0.00869   0.00278  -0.0503   1.0000   0.4314
  -1.250   0.0833   0.00760   0.00271  -0.0491   1.0000   0.7182
  -1.000   0.1061   0.00711   0.00255  -0.0474   1.0000   1.0000
  -0.750   0.1328   0.00719   0.00248  -0.0473   1.0000   1.0000
  -0.500   0.1593   0.00729   0.00246  -0.0472   1.0000   1.0000
  -0.250   0.1855   0.00742   0.00250  -0.0472   1.0000   1.0000
   0.000   0.2274   0.00751   0.00248  -0.0503   0.9908   1.0000
   0.250   0.2747   0.00753   0.00244  -0.0544   0.9770   1.0000
   0.500   0.3180   0.00751   0.00238  -0.0573   0.9600   1.0000
   0.750   0.3573   0.00749   0.00232  -0.0593   0.9395   1.0000
   1.000   0.3907   0.00747   0.00226  -0.0599   0.9141   1.0000
   1.250   0.4183   0.00750   0.00223  -0.0591   0.8838   1.0000
   1.500   0.4434   0.00758   0.00222  -0.0577   0.8508   1.0000
   1.750   0.4678   0.00770   0.00224  -0.0563   0.8159   1.0000
   2.000   0.4922   0.00787   0.00227  -0.0549   0.7781   1.0000
   2.250   0.5168   0.00808   0.00232  -0.0537   0.7371   1.0000
   2.500   0.5416   0.00832   0.00239  -0.0526   0.6921   1.0000
   2.750   0.5666   0.00860   0.00248  -0.0517   0.6423   1.0000
   3.000   0.5915   0.00895   0.00264  -0.0508   0.5867   1.0000
   3.250   0.6165   0.00935   0.00280  -0.0500   0.5276   1.0000
   3.500   0.6415   0.00981   0.00301  -0.0494   0.4694   1.0000
   3.750   0.6668   0.01030   0.00327  -0.0488   0.4158   1.0000
   4.000   0.6921   0.01081   0.00358  -0.0484   0.3681   1.0000
   4.250   0.7177   0.01132   0.00396  -0.0481   0.3253   1.0000
   4.500   0.7433   0.01184   0.00434  -0.0477   0.2861   1.0000
   4.750   0.7690   0.01236   0.00475  -0.0474   0.2488   1.0000
   5.000   0.7944   0.01293   0.00521  -0.0471   0.2152   1.0000
   5.250   0.8199   0.01353   0.00572  -0.0468   0.1828   1.0000
   5.500   0.8450   0.01420   0.00636  -0.0464   0.1555   1.0000
   5.750   0.8700   0.01492   0.00702  -0.0460   0.1312   1.0000
   6.000   0.8943   0.01580   0.00779  -0.0456   0.1108   1.0000
   6.250   0.9193   0.01657   0.00863  -0.0451   0.0930   1.0000
   6.500   0.9437   0.01741   0.00949  -0.0447   0.0767   1.0000
   6.750   0.9675   0.01845   0.01057  -0.0441   0.0630   1.0000
   7.000   0.9909   0.01958   0.01176  -0.0434   0.0506   1.0000
   7.250   1.0138   0.02093   0.01327  -0.0426   0.0405   1.0000
   7.500   1.0350   0.02292   0.01540  -0.0416   0.0333   1.0000
   7.750   1.0572   0.02441   0.01702  -0.0408   0.0278   1.0000
   8.000   1.0767   0.02707   0.01998  -0.0397   0.0237   1.0000
   8.250   1.0967   0.02967   0.02296  -0.0384   0.0215   1.0000
   8.500   1.1143   0.03272   0.02643  -0.0371   0.0201   1.0000
   8.750   1.1284   0.03638   0.03056  -0.0358   0.0194   1.0000
   9.000   1.1373   0.04094   0.03566  -0.0343   0.0192   1.0000
   9.250   1.1434   0.04497   0.04012  -0.0331   0.0186   1.0000
   9.500   1.1473   0.04849   0.04387  -0.0323   0.0175   1.0000
   9.750   1.1389   0.05381   0.04957  -0.0314   0.0171   1.0000
  10.000   1.1260   0.05885   0.05494  -0.0308   0.0170   1.0000
  10.250   1.1097   0.06305   0.05938  -0.0304   0.0170   1.0000
  10.500   1.0943   0.06780   0.06435  -0.0326   0.0172   1.0000
  10.750   1.0180   0.09452   0.09160  -0.0562   0.0225   1.0000
  11.000   0.9960   0.10668   0.10371  -0.0633   0.0234   1.0000
  11.250   0.9814   0.11550   0.11248  -0.0677   0.0240   1.0000
  11.500   0.9699   0.12315   0.12009  -0.0711   0.0247   1.0000
  11.750   0.9466   0.13849   0.13535  -0.0786   0.0318   1.0000
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