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HQ 2.1/9.5 AIRFOIL (hq2195-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.1/9.5 AIRFOIL (hq2195-il)
Reynolds number: 50,000
Max Cl/Cd: 37.27 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq2195-il-50000.txt
Download as CSV file: xf-hq2195-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.1/9.5 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4187   0.09996   0.09292  -0.0123   1.0000   0.3133
  -8.250  -0.4201   0.09764   0.09068  -0.0117   1.0000   0.3302
  -8.000  -0.4270   0.09587   0.08901  -0.0110   1.0000   0.3466
  -7.750  -0.4019   0.09119   0.08429  -0.0095   1.0000   0.3656
  -7.500  -0.4015   0.08924   0.08242  -0.0078   1.0000   0.3898
  -7.250  -0.3832   0.08569   0.07888  -0.0060   1.0000   0.4162
  -7.000  -0.3750   0.08291   0.07615  -0.0042   1.0000   0.4405
  -6.750  -0.3876   0.08216   0.07554  -0.0012   1.0000   0.4654
  -6.500  -0.3638   0.07775   0.07112  -0.0006   1.0000   0.4838
  -5.750  -0.4898   0.05953   0.05324  -0.0295   1.0000   0.2564
  -5.500  -0.4699   0.04714   0.03934  -0.0425   1.0000   0.1547
  -5.250  -0.4531   0.04374   0.03570  -0.0421   1.0000   0.1494
  -5.000  -0.4344   0.04020   0.03168  -0.0421   1.0000   0.1439
  -4.750  -0.4113   0.03726   0.02792  -0.0422   1.0000   0.1374
  -4.500  -0.3895   0.03454   0.02484  -0.0416   1.0000   0.1354
  -4.250  -0.3663   0.03223   0.02213  -0.0410   1.0000   0.1347
  -4.000  -0.3422   0.03047   0.01985  -0.0403   1.0000   0.1379
  -3.750  -0.3195   0.02855   0.01787  -0.0395   1.0000   0.1437
  -3.500  -0.2955   0.02706   0.01616  -0.0385   1.0000   0.1489
  -3.250  -0.2719   0.02564   0.01461  -0.0373   1.0000   0.1575
  -3.000  -0.2491   0.02443   0.01343  -0.0361   1.0000   0.1732
  -2.750  -0.2265   0.02324   0.01228  -0.0347   1.0000   0.1950
  -2.500  -0.2023   0.02162   0.01108  -0.0342   1.0000   0.2496
  -2.250  -0.2047   0.01926   0.01153  -0.0253   1.0000   0.7353
  -2.000  -0.2167   0.01923   0.01172  -0.0136   1.0000   0.8584
  -1.750  -0.1379   0.01929   0.01121  -0.0197   1.0000   1.0000
  -1.500  -0.1348   0.01894   0.01066  -0.0177   1.0000   1.0000
  -1.250  -0.1188   0.01883   0.01026  -0.0175   1.0000   1.0000
  -1.000  -0.0974   0.01887   0.01000  -0.0181   1.0000   1.0000
  -0.750  -0.0742   0.01901   0.00986  -0.0187   1.0000   1.0000
  -0.500  -0.0504   0.01922   0.00983  -0.0193   1.0000   1.0000
  -0.250  -0.0267   0.01948   0.00987  -0.0199   1.0000   1.0000
   0.000  -0.0033   0.01979   0.00997  -0.0203   1.0000   1.0000
   0.250   0.0197   0.02014   0.01016  -0.0206   1.0000   1.0000
   0.500   0.0423   0.02053   0.01041  -0.0209   1.0000   1.0000
   0.750   0.0645   0.02097   0.01072  -0.0211   1.0000   1.0000
   1.000   0.0863   0.02145   0.01110  -0.0213   1.0000   1.0000
   1.250   0.1075   0.02198   0.01156  -0.0215   1.0000   1.0000
   1.500   0.1282   0.02258   0.01209  -0.0217   1.0000   1.0000
   1.750   0.1482   0.02323   0.01271  -0.0219   1.0000   1.0000
   2.000   0.1675   0.02397   0.01342  -0.0221   1.0000   1.0000
   2.250   0.2067   0.02520   0.01468  -0.0263   0.9893   1.0000
   2.500   0.2678   0.02678   0.01631  -0.0341   0.9659   1.0000
   2.750   0.3288   0.02812   0.01775  -0.0413   0.9416   1.0000
   3.000   0.3807   0.02908   0.01885  -0.0466   0.9149   1.0000
   3.250   0.4322   0.02987   0.01979  -0.0513   0.8876   1.0000
   3.500   0.4845   0.03042   0.02054  -0.0555   0.8597   1.0000
   3.750   0.5383   0.03067   0.02103  -0.0593   0.8320   1.0000
   4.000   0.5999   0.03038   0.02103  -0.0635   0.8055   1.0000
   4.250   0.6513   0.02984   0.02078  -0.0653   0.7777   1.0000
   4.500   0.6965   0.02907   0.02026  -0.0656   0.7488   1.0000
   4.750   0.7412   0.02783   0.01926  -0.0649   0.7186   1.0000
   5.000   0.7773   0.02663   0.01826  -0.0628   0.6837   1.0000
   5.250   0.8130   0.02515   0.01689  -0.0601   0.6451   1.0000
   5.500   0.8424   0.02407   0.01582  -0.0571   0.6007   1.0000
   5.750   0.8670   0.02361   0.01528  -0.0540   0.5487   1.0000
   6.000   0.8896   0.02387   0.01523  -0.0510   0.4884   1.0000
   6.250   0.9073   0.02496   0.01588  -0.0480   0.4174   1.0000
   6.500   0.9218   0.02662   0.01705  -0.0450   0.3410   1.0000
   6.750   0.9370   0.02836   0.01839  -0.0424   0.2792   1.0000
   7.000   0.9551   0.03006   0.01981  -0.0406   0.2348   1.0000
   7.250   0.9759   0.03192   0.02137  -0.0393   0.2013   1.0000
   7.500   0.9980   0.03402   0.02338  -0.0382   0.1755   1.0000
   7.750   1.0211   0.03656   0.02601  -0.0373   0.1571   1.0000
   8.000   1.0449   0.03941   0.02899  -0.0366   0.1445   1.0000
   8.250   1.0654   0.04208   0.03188  -0.0356   0.1339   1.0000
   8.500   1.0815   0.04573   0.03604  -0.0341   0.1290   1.0000
   8.750   1.0929   0.04952   0.04040  -0.0323   0.1257   1.0000
   9.000   1.1068   0.05293   0.04405  -0.0311   0.1211   1.0000
   9.250   1.1174   0.05705   0.04839  -0.0299   0.1175   1.0000
   9.500   1.1154   0.06146   0.05334  -0.0278   0.1169   1.0000
   9.750   1.1114   0.06628   0.05859  -0.0260   0.1172   1.0000
  10.000   1.1059   0.07134   0.06396  -0.0246   0.1178   1.0000
  10.250   1.0720   0.07628   0.06937  -0.0224   0.1204   1.0000
  10.500   1.0094   0.08279   0.07615  -0.0218   0.1250   1.0000
  10.750   0.9758   0.09022   0.08364  -0.0246   0.1286   1.0000
  11.000   0.9592   0.09734   0.09078  -0.0275   0.1308   1.0000
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