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HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 100,000
Max Cl/Cd: 33.41 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq09-il-100000-n5.txt
Download as CSV file: xf-hq09-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6829   0.08853   0.08378   0.0012   1.0000   0.0160
  -9.500  -0.6932   0.08038   0.07579  -0.0045   1.0000   0.0151
  -9.250  -0.7368   0.06447   0.05972  -0.0154   1.0000   0.0117
  -9.000  -0.7546   0.05976   0.05485  -0.0163   1.0000   0.0115
  -8.750  -0.7664   0.05528   0.05014  -0.0164   1.0000   0.0116
  -8.500  -0.7720   0.05094   0.04551  -0.0161   1.0000   0.0117
  -8.250  -0.7720   0.04697   0.04123  -0.0155   1.0000   0.0119
  -8.000  -0.7616   0.04539   0.03950  -0.0152   1.0000   0.0136
  -7.750  -0.7516   0.04254   0.03634  -0.0145   1.0000   0.0151
  -7.500  -0.7409   0.03894   0.03224  -0.0135   1.0000   0.0168
  -7.250  -0.7273   0.03509   0.02776  -0.0121   1.0000   0.0186
  -7.000  -0.7096   0.03174   0.02365  -0.0107   1.0000   0.0201
  -6.750  -0.6924   0.02893   0.02064  -0.0100   1.0000   0.0236
  -6.500  -0.6705   0.02736   0.01870  -0.0092   1.0000   0.0294
  -6.250  -0.6475   0.02497   0.01595  -0.0082   1.0000   0.0325
  -6.000  -0.6254   0.02372   0.01464  -0.0077   1.0000   0.0381
  -5.750  -0.6017   0.02211   0.01284  -0.0067   1.0000   0.0406
  -5.500  -0.5786   0.02078   0.01133  -0.0058   1.0000   0.0432
  -5.250  -0.5575   0.01943   0.00996  -0.0048   1.0000   0.0467
  -5.000  -0.5352   0.01846   0.00890  -0.0039   1.0000   0.0524
  -4.750  -0.5129   0.01757   0.00798  -0.0031   1.0000   0.0636
  -4.500  -0.4906   0.01667   0.00699  -0.0022   1.0000   0.0820
  -4.250  -0.4695   0.01560   0.00614  -0.0013   1.0000   0.1169
  -4.000  -0.4497   0.01445   0.00555  -0.0004   1.0000   0.2203
  -3.750  -0.4318   0.01327   0.00507   0.0008   1.0000   0.3512
  -3.500  -0.4145   0.01240   0.00486   0.0027   1.0000   0.4942
  -3.250  -0.3956   0.01201   0.00483   0.0048   1.0000   0.5959
  -3.000  -0.3757   0.01184   0.00478   0.0068   1.0000   0.6603
  -2.750  -0.3568   0.01174   0.00478   0.0091   1.0000   0.7126
  -2.500  -0.3377   0.01168   0.00469   0.0113   1.0000   0.7499
  -2.250  -0.3172   0.01162   0.00460   0.0130   1.0000   0.7732
  -2.000  -0.2964   0.01156   0.00448   0.0143   1.0000   0.7927
  -1.750  -0.2762   0.01152   0.00441   0.0158   1.0000   0.8103
  -1.500  -0.2413   0.01153   0.00436   0.0143   0.9895   0.8303
  -1.250  -0.2045   0.01158   0.00434   0.0127   0.9802   0.8534
  -1.000  -0.1674   0.01162   0.00437   0.0112   0.9708   0.8764
  -0.750  -0.1263   0.01167   0.00439   0.0087   0.9617   0.8941
  -0.500  -0.0833   0.01169   0.00439   0.0057   0.9519   0.9078
  -0.250  -0.0405   0.01170   0.00436   0.0026   0.9413   0.9190
   0.000   0.0000   0.01169   0.00435   0.0000   0.9299   0.9299
   0.250   0.0406   0.01170   0.00436  -0.0026   0.9191   0.9413
   0.500   0.0833   0.01169   0.00439  -0.0057   0.9079   0.9519
   0.750   0.1263   0.01167   0.00439  -0.0087   0.8941   0.9617
   1.000   0.1673   0.01162   0.00437  -0.0112   0.8765   0.9708
   1.250   0.2045   0.01158   0.00434  -0.0127   0.8535   0.9802
   1.500   0.2413   0.01153   0.00436  -0.0143   0.8304   0.9895
   1.750   0.2761   0.01152   0.00440  -0.0158   0.8103   1.0000
   2.000   0.2963   0.01155   0.00447  -0.0143   0.7927   1.0000
   2.250   0.3170   0.01161   0.00460  -0.0129   0.7732   1.0000
   2.500   0.3375   0.01168   0.00469  -0.0113   0.7500   1.0000
   2.750   0.3566   0.01174   0.00478  -0.0091   0.7127   1.0000
   3.000   0.3755   0.01184   0.00478  -0.0068   0.6604   1.0000
   3.250   0.3955   0.01201   0.00483  -0.0048   0.5959   1.0000
   3.500   0.4143   0.01240   0.00486  -0.0027   0.4947   1.0000
   3.750   0.4317   0.01327   0.00507  -0.0008   0.3513   1.0000
   4.000   0.4497   0.01444   0.00555   0.0005   0.2211   1.0000
   4.250   0.4694   0.01560   0.00615   0.0013   0.1167   1.0000
   4.500   0.4905   0.01667   0.00699   0.0022   0.0819   1.0000
   4.750   0.5128   0.01757   0.00799   0.0031   0.0636   1.0000
   5.000   0.5352   0.01846   0.00889   0.0039   0.0524   1.0000
   5.250   0.5575   0.01942   0.00995   0.0048   0.0469   1.0000
   5.500   0.5786   0.02078   0.01132   0.0057   0.0432   1.0000
   5.750   0.6018   0.02210   0.01283   0.0067   0.0407   1.0000
   6.000   0.6254   0.02372   0.01465   0.0077   0.0381   1.0000
   6.250   0.6476   0.02498   0.01596   0.0082   0.0325   1.0000
   6.500   0.6706   0.02737   0.01871   0.0092   0.0294   1.0000
   6.750   0.6925   0.02891   0.02061   0.0100   0.0235   1.0000
   7.000   0.7097   0.03175   0.02366   0.0107   0.0201   1.0000
   7.250   0.7275   0.03514   0.02782   0.0121   0.0186   1.0000
   7.500   0.7412   0.03894   0.03225   0.0134   0.0168   1.0000
   7.750   0.7519   0.04259   0.03639   0.0144   0.0152   1.0000
   8.000   0.7609   0.04575   0.03990   0.0151   0.0138   1.0000
   8.250   0.7703   0.04780   0.04214   0.0155   0.0123   1.0000
   8.500   0.7722   0.05089   0.04544   0.0159   0.0115   1.0000
   8.750   0.7666   0.05530   0.05015   0.0162   0.0115   1.0000
   9.000   0.7546   0.05978   0.05487   0.0162   0.0114   1.0000
   9.250   0.7344   0.06536   0.06065   0.0149   0.0119   1.0000
   9.500   0.6937   0.08053   0.07593   0.0043   0.0150   1.0000
   9.750   0.6835   0.08859   0.08384  -0.0013   0.0163   1.0000
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