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HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 200,000
Max Cl/Cd: 38.62 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq09-il-200000-n5.txt
Download as CSV file: xf-hq09-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/9 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.7192   0.08439   0.08111   0.0026   1.0000   0.0057
 -10.000  -0.7327   0.07470   0.07143  -0.0036   1.0000   0.0056
  -9.750  -0.7502   0.06419   0.06085  -0.0121   1.0000   0.0055
  -9.500  -0.7707   0.05790   0.05442  -0.0164   1.0000   0.0054
  -9.250  -0.7898   0.05351   0.04988  -0.0168   1.0000   0.0054
  -9.000  -0.8048   0.04874   0.04485  -0.0162   1.0000   0.0054
  -8.750  -0.8137   0.04363   0.03936  -0.0152   1.0000   0.0053
  -8.500  -0.8155   0.03880   0.03408  -0.0140   1.0000   0.0053
  -8.250  -0.8108   0.03446   0.02923  -0.0127   1.0000   0.0053
  -8.000  -0.8009   0.03055   0.02478  -0.0112   1.0000   0.0054
  -7.750  -0.7861   0.02739   0.02114  -0.0100   1.0000   0.0055
  -7.500  -0.7693   0.02452   0.01786  -0.0087   1.0000   0.0057
  -7.250  -0.7505   0.02232   0.01536  -0.0077   1.0000   0.0063
  -7.000  -0.7285   0.02107   0.01394  -0.0070   1.0000   0.0071
  -6.750  -0.7064   0.01974   0.01243  -0.0062   1.0000   0.0085
  -6.500  -0.6836   0.01858   0.01097  -0.0053   1.0000   0.0110
  -6.250  -0.6600   0.01791   0.01029  -0.0050   1.0000   0.0163
  -6.000  -0.6349   0.01762   0.00992  -0.0048   1.0000   0.0224
  -5.750  -0.6108   0.01705   0.00929  -0.0044   1.0000   0.0266
  -5.500  -0.5872   0.01624   0.00836  -0.0037   1.0000   0.0292
  -5.250  -0.5640   0.01541   0.00741  -0.0029   1.0000   0.0306
  -5.000  -0.5412   0.01454   0.00645  -0.0021   1.0000   0.0333
  -4.750  -0.5177   0.01387   0.00573  -0.0015   1.0000   0.0380
  -4.500  -0.4936   0.01335   0.00507  -0.0008   1.0000   0.0447
  -4.250  -0.4700   0.01279   0.00455  -0.0002   1.0000   0.0587
  -4.000  -0.4472   0.01218   0.00410   0.0005   1.0000   0.0902
  -3.750  -0.4259   0.01140   0.00368   0.0012   1.0000   0.1570
  -3.500  -0.4043   0.01082   0.00336   0.0020   1.0000   0.2297
  -3.250  -0.3847   0.01006   0.00306   0.0030   1.0000   0.3248
  -3.000  -0.3588   0.00930   0.00285   0.0028   0.9928   0.4565
  -2.750  -0.3267   0.00883   0.00274   0.0016   0.9817   0.5593
  -2.500  -0.2939   0.00862   0.00273   0.0006   0.9705   0.6297
  -2.250  -0.2613   0.00849   0.00273  -0.0002   0.9593   0.6821
  -2.000  -0.2283   0.00843   0.00269  -0.0011   0.9479   0.7123
  -1.750  -0.1951   0.00838   0.00261  -0.0021   0.9353   0.7279
  -1.500  -0.1633   0.00834   0.00251  -0.0027   0.9200   0.7431
  -1.250  -0.1334   0.00830   0.00246  -0.0028   0.9037   0.7587
  -1.000  -0.1051   0.00828   0.00242  -0.0026   0.8887   0.7743
  -0.750  -0.0781   0.00826   0.00241  -0.0021   0.8757   0.7912
  -0.500  -0.0519   0.00826   0.00239  -0.0014   0.8624   0.8072
  -0.250  -0.0258   0.00825   0.00238  -0.0007   0.8488   0.8212
   0.000   0.0000   0.00825   0.00237   0.0000   0.8349   0.8349
   0.250   0.0259   0.00825   0.00238   0.0007   0.8212   0.8488
   0.500   0.0519   0.00826   0.00239   0.0014   0.8072   0.8623
   0.750   0.0781   0.00826   0.00241   0.0021   0.7912   0.8756
   1.000   0.1051   0.00828   0.00242   0.0026   0.7743   0.8888
   1.250   0.1334   0.00830   0.00246   0.0028   0.7587   0.9037
   1.500   0.1633   0.00834   0.00251   0.0027   0.7431   0.9200
   1.750   0.1951   0.00838   0.00261   0.0021   0.7279   0.9353
   2.000   0.2282   0.00843   0.00269   0.0011   0.7123   0.9479
   2.250   0.2613   0.00849   0.00273   0.0002   0.6819   0.9594
   2.500   0.2938   0.00862   0.00272  -0.0006   0.6294   0.9706
   2.750   0.3266   0.00883   0.00274  -0.0016   0.5597   0.9818
   3.000   0.3588   0.00929   0.00285  -0.0028   0.4572   0.9929
   3.250   0.3845   0.01006   0.00306  -0.0030   0.3246   1.0000
   3.500   0.4042   0.01081   0.00336  -0.0020   0.2304   1.0000
   3.750   0.4258   0.01140   0.00368  -0.0012   0.1577   1.0000
   4.000   0.4471   0.01217   0.00410  -0.0004   0.0904   1.0000
   4.250   0.4699   0.01279   0.00455   0.0002   0.0584   1.0000
   4.500   0.4935   0.01335   0.00507   0.0008   0.0446   1.0000
   4.750   0.5176   0.01387   0.00573   0.0015   0.0380   1.0000
   5.000   0.5411   0.01454   0.00645   0.0021   0.0332   1.0000
   5.250   0.5639   0.01540   0.00740   0.0029   0.0307   1.0000
   5.500   0.5872   0.01624   0.00836   0.0037   0.0291   1.0000
   5.750   0.6108   0.01705   0.00929   0.0044   0.0266   1.0000
   6.000   0.6350   0.01761   0.00991   0.0048   0.0222   1.0000
   6.250   0.6600   0.01794   0.01033   0.0050   0.0164   1.0000
   6.500   0.6838   0.01856   0.01095   0.0053   0.0110   1.0000
   6.750   0.7066   0.01975   0.01244   0.0062   0.0084   1.0000
   7.000   0.7287   0.02107   0.01395   0.0070   0.0071   1.0000
   7.250   0.7507   0.02233   0.01537   0.0077   0.0063   1.0000
   7.500   0.7695   0.02453   0.01787   0.0087   0.0057   1.0000
   7.750   0.7864   0.02739   0.02114   0.0099   0.0055   1.0000
   8.000   0.8012   0.03057   0.02480   0.0111   0.0054   1.0000
   8.250   0.8114   0.03441   0.02917   0.0125   0.0054   1.0000
   8.500   0.8154   0.03896   0.03426   0.0140   0.0053   1.0000
   8.750   0.8139   0.04373   0.03948   0.0151   0.0053   1.0000
   9.000   0.8046   0.04893   0.04506   0.0160   0.0053   1.0000
   9.250   0.7908   0.05349   0.04986   0.0166   0.0054   1.0000
   9.500   0.7703   0.05812   0.05465   0.0161   0.0054   1.0000
   9.750   0.7514   0.06411   0.06077   0.0120   0.0055   1.0000
  10.000   0.7328   0.07508   0.07181   0.0032   0.0056   1.0000
  10.250   0.7202   0.08446   0.08118  -0.0028   0.0057   1.0000
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