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HQ 0/10 AIRFOIL (hq010-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 0/10 AIRFOIL (hq010-il)
Reynolds number: 50,000
Max Cl/Cd: 27.99 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq010-il-50000.txt
Download as CSV file: xf-hq010-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/10 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7339   0.07540   0.06886  -0.0155   1.0000   0.1425
  -8.750  -0.7801   0.06813   0.06125  -0.0185   1.0000   0.1335
  -8.500  -0.7761   0.06295   0.05596  -0.0185   1.0000   0.1305
  -8.250  -0.7806   0.05758   0.05025  -0.0183   1.0000   0.1270
  -8.000  -0.7828   0.05243   0.04458  -0.0177   1.0000   0.1247
  -7.750  -0.7734   0.04882   0.04063  -0.0167   1.0000   0.1282
  -7.500  -0.7640   0.04503   0.03631  -0.0155   1.0000   0.1309
  -7.250  -0.7501   0.04130   0.03199  -0.0143   1.0000   0.1324
  -7.000  -0.7325   0.03789   0.02793  -0.0130   1.0000   0.1342
  -6.750  -0.7091   0.03489   0.02487  -0.0123   1.0000   0.1393
  -6.500  -0.6871   0.03247   0.02201  -0.0112   1.0000   0.1485
  -6.250  -0.6640   0.03052   0.02006  -0.0103   1.0000   0.1633
  -6.000  -0.6384   0.02842   0.01788  -0.0094   1.0000   0.1790
  -5.750  -0.6141   0.02649   0.01600  -0.0083   1.0000   0.2054
  -5.500  -0.5931   0.02453   0.01437  -0.0069   1.0000   0.2502
  -5.250  -0.5773   0.02210   0.01273  -0.0045   1.0000   0.3340
  -5.000  -0.5705   0.02038   0.01221   0.0004   1.0000   0.4779
  -4.750  -0.5603   0.02059   0.01298   0.0066   1.0000   0.5920
  -4.500  -0.5464   0.02147   0.01394   0.0128   1.0000   0.6638
  -4.250  -0.5299   0.02234   0.01470   0.0187   1.0000   0.7183
  -4.000  -0.5043   0.02355   0.01577   0.0241   1.0000   0.7670
  -3.750  -0.4590   0.02491   0.01680   0.0267   1.0000   0.8180
  -3.500  -0.4024   0.02525   0.01672   0.0242   1.0000   0.8566
  -3.250  -0.3211   0.02517   0.01608   0.0158   1.0000   0.8869
  -3.000  -0.2579   0.02462   0.01518   0.0093   1.0000   0.9139
  -2.750  -0.1958   0.02389   0.01413   0.0024   1.0000   0.9404
  -2.500  -0.1161   0.02275   0.01271  -0.0082   1.0000   0.9699
  -2.250  -0.0311   0.02114   0.01084  -0.0207   1.0000   1.0000
  -2.000  -0.0164   0.02062   0.01031  -0.0203   1.0000   1.0000
  -1.750  -0.0022   0.02018   0.00987  -0.0197   1.0000   1.0000
  -1.500   0.0109   0.01980   0.00951  -0.0189   1.0000   1.0000
  -1.250   0.0219   0.01951   0.00926  -0.0177   1.0000   1.0000
  -1.000   0.0291   0.01932   0.00911  -0.0158   1.0000   1.0000
  -0.750   0.0298   0.01925   0.00909  -0.0130   1.0000   1.0000
  -0.500   0.0237   0.01928   0.00917  -0.0092   1.0000   1.0000
  -0.250   0.0126   0.01935   0.00925  -0.0047   1.0000   1.0000
   0.000   0.0000   0.01938   0.00928   0.0000   1.0000   1.0000
   0.250  -0.0126   0.01935   0.00925   0.0047   1.0000   1.0000
   0.500  -0.0237   0.01928   0.00917   0.0092   1.0000   1.0000
   0.750  -0.0298   0.01925   0.00909   0.0130   1.0000   1.0000
   1.000  -0.0290   0.01931   0.00911   0.0158   1.0000   1.0000
   1.250  -0.0219   0.01951   0.00925   0.0177   1.0000   1.0000
   1.500  -0.0108   0.01980   0.00951   0.0189   1.0000   1.0000
   1.750   0.0022   0.02017   0.00986   0.0197   1.0000   1.0000
   2.000   0.0164   0.02062   0.01030   0.0203   1.0000   1.0000
   2.250   0.0312   0.02114   0.01083   0.0207   1.0000   1.0000
   2.500   0.1158   0.02274   0.01270   0.0082   0.9701   1.0000
   2.750   0.1963   0.02389   0.01413  -0.0025   0.9403   1.0000
   3.000   0.2578   0.02462   0.01518  -0.0093   0.9139   1.0000
   3.250   0.3210   0.02516   0.01607  -0.0158   0.8869   1.0000
   3.500   0.4024   0.02524   0.01672  -0.0242   0.8566   1.0000
   3.750   0.4595   0.02489   0.01679  -0.0267   0.8179   1.0000
   4.000   0.5043   0.02355   0.01577  -0.0241   0.7670   1.0000
   4.250   0.5299   0.02235   0.01471  -0.0187   0.7184   1.0000
   4.500   0.5464   0.02147   0.01394  -0.0128   0.6638   1.0000
   4.750   0.5603   0.02059   0.01298  -0.0066   0.5921   1.0000
   5.000   0.5705   0.02038   0.01220  -0.0004   0.4783   1.0000
   5.250   0.5773   0.02210   0.01273   0.0045   0.3340   1.0000
   5.500   0.5931   0.02453   0.01437   0.0069   0.2501   1.0000
   5.750   0.6141   0.02649   0.01599   0.0083   0.2053   1.0000
   6.000   0.6384   0.02843   0.01789   0.0094   0.1789   1.0000
   6.250   0.6640   0.03052   0.02005   0.0103   0.1632   1.0000
   6.500   0.6871   0.03249   0.02200   0.0112   0.1482   1.0000
   6.750   0.7091   0.03489   0.02487   0.0123   0.1393   1.0000
   7.000   0.7325   0.03789   0.02794   0.0130   0.1342   1.0000
   7.250   0.7500   0.04131   0.03201   0.0143   0.1323   1.0000
   7.500   0.7640   0.04504   0.03633   0.0155   0.1309   1.0000
   7.750   0.7734   0.04881   0.04062   0.0167   0.1281   1.0000
   8.000   0.7827   0.05245   0.04460   0.0177   0.1248   1.0000
   8.250   0.7805   0.05759   0.05026   0.0183   0.1270   1.0000
   8.500   0.7762   0.06294   0.05594   0.0185   0.1305   1.0000
   8.750   0.7805   0.06812   0.06124   0.0185   0.1336   1.0000
   9.000   0.7338   0.07545   0.06892   0.0155   0.1426   1.0000
   9.250   0.6151   0.07835   0.07213   0.0123   0.1629   1.0000
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