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


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 100,000
Max Cl/Cd: 36.83 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq09-il-100000.txt
Download as CSV file: xf-hq09-il-100000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5705   0.10293   0.09855   0.0068   1.0000   0.1215
  -9.500  -0.5902   0.09764   0.09333   0.0021   1.0000   0.1224
  -9.250  -0.6040   0.09160   0.08736  -0.0015   1.0000   0.1235
  -9.000  -0.5702   0.08900   0.08471   0.0044   1.0000   0.1401
  -8.750  -0.5786   0.08439   0.08013   0.0022   1.0000   0.1473
  -8.500  -0.6119   0.07837   0.07422  -0.0044   1.0000   0.1489
  -8.250  -0.5885   0.07463   0.07042  -0.0004   1.0000   0.1594
  -7.250  -0.7210   0.04714   0.04064  -0.0147   1.0000   0.0562
  -7.000  -0.7085   0.04240   0.03552  -0.0138   1.0000   0.0558
  -6.750  -0.6954   0.03848   0.03094  -0.0125   1.0000   0.0576
  -6.500  -0.6772   0.03427   0.02624  -0.0113   1.0000   0.0578
  -6.250  -0.6566   0.03053   0.02220  -0.0104   1.0000   0.0599
  -6.000  -0.6340   0.02857   0.01997  -0.0095   1.0000   0.0671
  -5.750  -0.6090   0.02609   0.01695  -0.0083   1.0000   0.0690
  -5.500  -0.5837   0.02334   0.01401  -0.0075   1.0000   0.0726
  -5.250  -0.5590   0.02206   0.01256  -0.0068   1.0000   0.0819
  -5.000  -0.5346   0.02007   0.01067  -0.0059   1.0000   0.0893
  -4.750  -0.5115   0.01849   0.00913  -0.0048   1.0000   0.0998
  -4.500  -0.4905   0.01707   0.00784  -0.0034   1.0000   0.1222
  -4.250  -0.4735   0.01517   0.00664  -0.0018   1.0000   0.1995
  -4.000  -0.4668   0.01267   0.00596   0.0016   1.0000   0.4994
  -3.750  -0.4522   0.01235   0.00616   0.0055   1.0000   0.6488
  -3.500  -0.4353   0.01237   0.00622   0.0090   1.0000   0.7185
  -3.250  -0.4190   0.01246   0.00633   0.0127   1.0000   0.7671
  -3.000  -0.4033   0.01253   0.00637   0.0165   1.0000   0.8057
  -2.750  -0.3886   0.01263   0.00646   0.0207   1.0000   0.8421
  -2.500  -0.3702   0.01280   0.00659   0.0245   1.0000   0.8776
  -2.250  -0.3390   0.01295   0.00657   0.0251   1.0000   0.9048
  -2.000  -0.2994   0.01299   0.00644   0.0232   1.0000   0.9240
  -1.750  -0.2539   0.01302   0.00630   0.0198   1.0000   0.9409
  -1.500  -0.1933   0.01307   0.00614   0.0135   1.0000   0.9548
  -1.250  -0.1138   0.01306   0.00593   0.0036   1.0000   0.9706
  -1.000  -0.0506   0.01280   0.00557  -0.0038   1.0000   0.9848
  -0.750   0.0025   0.01245   0.00517  -0.0097   1.0000   0.9964
  -0.500   0.0257   0.01222   0.00495  -0.0104   1.0000   1.0000
  -0.250   0.0238   0.01213   0.00493  -0.0068   1.0000   1.0000
   0.000   0.0000   0.01215   0.00498   0.0000   1.0000   1.0000
   0.250  -0.0238   0.01213   0.00493   0.0068   1.0000   1.0000
   0.500  -0.0257   0.01221   0.00495   0.0104   1.0000   1.0000
   0.750  -0.0025   0.01245   0.00516   0.0097   0.9965   1.0000
   1.000   0.0506   0.01280   0.00557   0.0038   0.9848   1.0000
   1.250   0.1138   0.01306   0.00593  -0.0037   0.9706   1.0000
   1.500   0.1934   0.01307   0.00613  -0.0135   0.9548   1.0000
   1.750   0.2539   0.01301   0.00629  -0.0198   0.9409   1.0000
   2.000   0.2993   0.01299   0.00644  -0.0232   0.9240   1.0000
   2.250   0.3389   0.01295   0.00656  -0.0251   0.9048   1.0000
   2.500   0.3699   0.01280   0.00659  -0.0244   0.8776   1.0000
   2.750   0.3883   0.01263   0.00646  -0.0206   0.8421   1.0000
   3.000   0.4030   0.01253   0.00637  -0.0164   0.8058   1.0000
   3.250   0.4187   0.01246   0.00633  -0.0127   0.7672   1.0000
   3.500   0.4351   0.01237   0.00622  -0.0089   0.7190   1.0000
   3.750   0.4519   0.01235   0.00616  -0.0054   0.6491   1.0000
   4.000   0.4666   0.01267   0.00596  -0.0016   0.5001   1.0000
   4.250   0.4734   0.01517   0.00664   0.0018   0.1996   1.0000
   4.500   0.4904   0.01708   0.00784   0.0035   0.1224   1.0000
   4.750   0.5115   0.01848   0.00912   0.0048   0.0999   1.0000
   5.000   0.5346   0.02007   0.01067   0.0059   0.0894   1.0000
   5.250   0.5590   0.02206   0.01255   0.0068   0.0820   1.0000
   5.500   0.5837   0.02334   0.01401   0.0075   0.0726   1.0000
   5.750   0.6090   0.02609   0.01696   0.0083   0.0690   1.0000
   6.000   0.6341   0.02858   0.01997   0.0095   0.0670   1.0000
   6.250   0.6568   0.03055   0.02223   0.0104   0.0600   1.0000
   6.500   0.6773   0.03430   0.02626   0.0112   0.0578   1.0000
   6.750   0.6961   0.03824   0.03073   0.0125   0.0578   1.0000
   7.000   0.7087   0.04257   0.03565   0.0137   0.0560   1.0000
   7.250   0.7212   0.04720   0.04071   0.0147   0.0562   1.0000
   8.500   0.7293   0.08324   0.07868   0.0103   0.1227   1.0000
   8.750   0.6939   0.08757   0.08312   0.0044   0.1219   1.0000
   9.000   0.6646   0.09368   0.08910  -0.0047   0.1195   1.0000
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