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HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/9 AIRFOIL (hq109-il)
Reynolds number: 100,000
Max Cl/Cd: 47.98 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq109-il-100000.txt
Download as CSV file: xf-hq109-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/9 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5433   0.10314   0.09828  -0.0083   1.0000   0.1042
  -9.250  -0.5590   0.09918   0.09442  -0.0141   1.0000   0.1085
  -9.000  -0.5870   0.09446   0.08985  -0.0230   1.0000   0.1094
  -8.750  -0.5462   0.09098   0.08629  -0.0139   1.0000   0.1154
  -8.500  -0.5525   0.08703   0.08241  -0.0171   1.0000   0.1201
  -8.250  -0.5922   0.08174   0.07720  -0.0275   1.0000   0.1225
  -8.000  -0.5691   0.07789   0.07343  -0.0229   1.0000   0.1270
  -7.750  -0.5726   0.07377   0.06931  -0.0255   1.0000   0.1330
  -6.750  -0.5864   0.04553   0.03933  -0.0355   1.0000   0.0707
  -6.500  -0.5704   0.04014   0.03322  -0.0336   1.0000   0.0560
  -6.250  -0.5563   0.03562   0.02847  -0.0325   1.0000   0.0542
  -6.000  -0.5400   0.03212   0.02454  -0.0310   1.0000   0.0531
  -5.750  -0.5216   0.02938   0.02131  -0.0293   1.0000   0.0542
  -5.500  -0.5029   0.02683   0.01830  -0.0277   1.0000   0.0581
  -5.250  -0.4824   0.02456   0.01586  -0.0264   1.0000   0.0611
  -5.000  -0.4608   0.02266   0.01375  -0.0248   1.0000   0.0658
  -4.750  -0.4406   0.02095   0.01194  -0.0233   1.0000   0.0753
  -4.500  -0.4208   0.01936   0.01036  -0.0216   1.0000   0.0889
  -4.250  -0.4021   0.01811   0.00915  -0.0200   1.0000   0.1111
  -4.000  -0.3845   0.01695   0.00819  -0.0183   1.0000   0.1398
  -3.750  -0.3673   0.01575   0.00730  -0.0167   1.0000   0.1837
  -3.500  -0.3566   0.01331   0.00646  -0.0145   1.0000   0.4262
  -3.250  -0.3467   0.01285   0.00678  -0.0102   1.0000   0.6371
  -3.000  -0.3337   0.01295   0.00695  -0.0064   1.0000   0.7100
  -2.750  -0.3206   0.01308   0.00707  -0.0028   1.0000   0.7600
  -2.500  -0.3082   0.01317   0.00714   0.0009   1.0000   0.7998
  -2.250  -0.2967   0.01321   0.00717   0.0048   1.0000   0.8370
  -2.000  -0.2853   0.01321   0.00717   0.0088   1.0000   0.8772
  -1.750  -0.2616   0.01330   0.00722   0.0106   1.0000   0.9229
  -1.500  -0.2020   0.01360   0.00729   0.0049   1.0000   0.9603
  -1.250  -0.1206   0.01390   0.00733  -0.0055   1.0000   0.9859
  -1.000  -0.0835   0.01384   0.00712  -0.0093   1.0000   1.0000
  -0.750  -0.1021   0.01350   0.00674  -0.0040   1.0000   1.0000
  -0.500  -0.0685   0.01371   0.00681  -0.0070   0.9940   1.0000
  -0.250  -0.0225   0.01400   0.00699  -0.0118   0.9845   1.0000
   0.000   0.0223   0.01429   0.00716  -0.0163   0.9745   1.0000
   0.250   0.0680   0.01458   0.00738  -0.0207   0.9653   1.0000
   0.500   0.1137   0.01484   0.00759  -0.0250   0.9562   1.0000
   0.750   0.1527   0.01505   0.00779  -0.0279   0.9456   1.0000
   1.000   0.1972   0.01523   0.00797  -0.0316   0.9350   1.0000
   1.250   0.2546   0.01516   0.00795  -0.0372   0.9221   1.0000
   1.500   0.3129   0.01489   0.00777  -0.0424   0.9077   1.0000
   1.750   0.3561   0.01476   0.00771  -0.0450   0.8946   1.0000
   2.000   0.3922   0.01470   0.00772  -0.0462   0.8809   1.0000
   2.250   0.4238   0.01465   0.00774  -0.0463   0.8660   1.0000
   2.500   0.4527   0.01460   0.00778  -0.0458   0.8500   1.0000
   2.750   0.4799   0.01452   0.00776  -0.0447   0.8332   1.0000
   3.000   0.5052   0.01442   0.00773  -0.0432   0.8152   1.0000
   3.250   0.5276   0.01435   0.00773  -0.0412   0.7935   1.0000
   3.500   0.5507   0.01420   0.00767  -0.0390   0.7704   1.0000
   3.750   0.5736   0.01402   0.00754  -0.0368   0.7442   1.0000
   4.000   0.5958   0.01386   0.00742  -0.0345   0.7125   1.0000
   4.250   0.6179   0.01370   0.00727  -0.0322   0.6743   1.0000
   4.500   0.6395   0.01364   0.00717  -0.0299   0.6247   1.0000
   4.750   0.6602   0.01376   0.00720  -0.0276   0.5588   1.0000
   5.000   0.6791   0.01424   0.00733  -0.0252   0.4734   1.0000
   5.250   0.6965   0.01512   0.00774  -0.0232   0.3839   1.0000
   5.500   0.7134   0.01626   0.00840  -0.0214   0.3049   1.0000
   5.750   0.7300   0.01765   0.00936  -0.0198   0.2326   1.0000
   6.000   0.7474   0.01904   0.01039  -0.0183   0.1700   1.0000
   6.250   0.7657   0.02055   0.01165  -0.0169   0.1286   1.0000
   6.500   0.7848   0.02233   0.01322  -0.0156   0.1012   1.0000
   6.750   0.8059   0.02407   0.01499  -0.0144   0.0804   1.0000
   7.000   0.8281   0.02615   0.01709  -0.0133   0.0667   1.0000
   7.250   0.8509   0.02859   0.01964  -0.0123   0.0572   1.0000
   7.500   0.8733   0.03097   0.02211  -0.0115   0.0509   1.0000
   7.750   0.8929   0.03382   0.02541  -0.0102   0.0461   1.0000
   8.000   0.9108   0.03708   0.02918  -0.0086   0.0441   1.0000
   8.250   0.9244   0.04086   0.03351  -0.0068   0.0437   1.0000
   8.500   0.9327   0.04520   0.03846  -0.0048   0.0441   1.0000
   8.750   0.9359   0.04986   0.04363  -0.0029   0.0451   1.0000
   9.000   0.9352   0.05469   0.04885  -0.0012   0.0464   1.0000
   9.250   0.9320   0.05966   0.05410   0.0002   0.0477   1.0000
   9.500   0.9338   0.06407   0.05877   0.0014   0.0508   1.0000
   9.750   0.8856   0.07008   0.06527   0.0032   0.0543   1.0000
  10.000   0.8586   0.07526   0.07060   0.0021   0.0556   1.0000
  10.250   0.8328   0.08174   0.07716  -0.0016   0.0568   1.0000
  10.500   0.8139   0.08904   0.08450  -0.0065   0.0586   1.0000
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