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HQ 1.5/8.5 AIRFOIL (hq1585-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/8.5 AIRFOIL (hq1585-il)
Reynolds number: 50,000
Max Cl/Cd: 35.57 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1585-il-50000.txt
Download as CSV file: xf-hq1585-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8.5 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5009   0.11434   0.10734   0.0017   1.0000   0.2522
  -9.250  -0.5025   0.11137   0.10444   0.0008   1.0000   0.2621
  -9.000  -0.5148   0.11006   0.10324  -0.0007   1.0000   0.2734
  -8.750  -0.5000   0.10558   0.09876   0.0000   1.0000   0.2868
  -8.500  -0.4876   0.10145   0.09464   0.0005   1.0000   0.2994
  -8.250  -0.4851   0.09839   0.09163   0.0006   1.0000   0.3144
  -8.000  -0.4900   0.09605   0.08939   0.0006   1.0000   0.3301
  -7.750  -0.4671   0.09132   0.08463   0.0021   1.0000   0.3497
  -7.500  -0.4601   0.08811   0.08146   0.0028   1.0000   0.3676
  -7.250  -0.4658   0.08601   0.07947   0.0035   1.0000   0.3864
  -7.000  -0.4638   0.08292   0.07646   0.0043   1.0000   0.4025
  -6.750  -0.4558   0.07953   0.07307   0.0051   1.0000   0.4172
  -6.500  -0.5252   0.06135   0.05481  -0.0338   1.0000   0.1901
  -6.250  -0.5165   0.05228   0.04504  -0.0400   1.0000   0.1445
  -6.000  -0.5050   0.04678   0.03896  -0.0414   1.0000   0.1341
  -5.750  -0.4884   0.04290   0.03467  -0.0411   1.0000   0.1308
  -5.500  -0.4707   0.03910   0.03050  -0.0406   1.0000   0.1273
  -5.250  -0.4511   0.03553   0.02637  -0.0399   1.0000   0.1240
  -5.000  -0.4298   0.03257   0.02289  -0.0389   1.0000   0.1234
  -4.750  -0.4084   0.03037   0.02031  -0.0378   1.0000   0.1292
  -4.500  -0.3863   0.02822   0.01780  -0.0367   1.0000   0.1351
  -4.250  -0.3635   0.02631   0.01571  -0.0353   1.0000   0.1413
  -4.000  -0.3407   0.02459   0.01387  -0.0338   1.0000   0.1510
  -3.750  -0.3193   0.02323   0.01251  -0.0323   1.0000   0.1720
  -3.500  -0.2973   0.02168   0.01108  -0.0306   1.0000   0.2010
  -3.250  -0.2783   0.01922   0.00963  -0.0293   1.0000   0.3171
  -3.000  -0.2883   0.01767   0.01033  -0.0182   1.0000   0.7137
  -2.750  -0.2907   0.01772   0.01045  -0.0090   1.0000   0.8156
  -2.500  -0.0801   0.01802   0.00929  -0.0336   1.0000   1.0000
  -2.250  -0.0886   0.01768   0.00892  -0.0298   1.0000   1.0000
  -2.000  -0.1012   0.01738   0.00858  -0.0251   1.0000   1.0000
  -1.750  -0.1131   0.01708   0.00822  -0.0205   1.0000   1.0000
  -1.500  -0.1158   0.01687   0.00786  -0.0171   1.0000   1.0000
  -1.250  -0.1053   0.01680   0.00758  -0.0157   1.0000   1.0000
  -1.000  -0.0884   0.01684   0.00740  -0.0151   1.0000   1.0000
  -0.750  -0.0690   0.01696   0.00727  -0.0148   1.0000   1.0000
  -0.500  -0.0487   0.01713   0.00724  -0.0146   1.0000   1.0000
  -0.250  -0.0279   0.01734   0.00729  -0.0145   1.0000   1.0000
   0.000  -0.0070   0.01760   0.00740  -0.0144   1.0000   1.0000
   0.250   0.0139   0.01790   0.00757  -0.0142   1.0000   1.0000
   0.500   0.0348   0.01823   0.00778  -0.0141   1.0000   1.0000
   0.750   0.0555   0.01860   0.00806  -0.0140   1.0000   1.0000
   1.000   0.0760   0.01902   0.00840  -0.0140   1.0000   1.0000
   1.250   0.0962   0.01948   0.00881  -0.0139   1.0000   1.0000
   1.500   0.1162   0.01998   0.00928  -0.0138   1.0000   1.0000
   1.750   0.1358   0.02054   0.00981  -0.0138   1.0000   1.0000
   2.000   0.1550   0.02115   0.01042  -0.0139   1.0000   1.0000
   2.250   0.1738   0.02182   0.01111  -0.0139   1.0000   1.0000
   2.500   0.1984   0.02267   0.01201  -0.0153   0.9967   1.0000
   2.750   0.2635   0.02420   0.01370  -0.0241   0.9725   1.0000
   3.000   0.3282   0.02543   0.01513  -0.0320   0.9448   1.0000
   3.250   0.3924   0.02634   0.01632  -0.0391   0.9154   1.0000
   3.500   0.4552   0.02688   0.01718  -0.0451   0.8856   1.0000
   3.750   0.5167   0.02702   0.01773  -0.0501   0.8541   1.0000
   4.000   0.5737   0.02677   0.01787  -0.0532   0.8197   1.0000
   4.250   0.6318   0.02580   0.01731  -0.0546   0.7844   1.0000
   4.500   0.6720   0.02470   0.01656  -0.0525   0.7430   1.0000
   4.750   0.7011   0.02372   0.01578  -0.0489   0.6952   1.0000
   5.000   0.7305   0.02251   0.01465  -0.0447   0.6434   1.0000
   5.250   0.7517   0.02166   0.01364  -0.0398   0.5745   1.0000
   5.500   0.7668   0.02156   0.01313  -0.0350   0.4888   1.0000
   5.750   0.7807   0.02254   0.01358  -0.0315   0.3992   1.0000
   6.000   0.7927   0.02456   0.01486  -0.0284   0.2973   1.0000
   6.250   0.8119   0.02730   0.01685  -0.0264   0.2210   1.0000
   6.500   0.8365   0.02950   0.01899  -0.0253   0.1852   1.0000
   6.750   0.8629   0.03174   0.02113  -0.0246   0.1645   1.0000
   7.000   0.8904   0.03446   0.02396  -0.0240   0.1516   1.0000
   7.250   0.9128   0.03700   0.02674  -0.0230   0.1375   1.0000
   7.500   0.9337   0.04003   0.03026  -0.0217   0.1274   1.0000
   7.750   0.9493   0.04342   0.03415  -0.0203   0.1175   1.0000
   8.000   0.9671   0.04719   0.03804  -0.0193   0.1102   1.0000
   8.250   0.9732   0.05133   0.04295  -0.0174   0.1060   1.0000
   8.500   0.9780   0.05595   0.04810  -0.0160   0.1042   1.0000
   8.750   0.9762   0.06124   0.05388  -0.0148   0.1052   1.0000
   9.000   0.9699   0.06666   0.05969  -0.0139   0.1070   1.0000
   9.250   0.9622   0.07204   0.06531  -0.0135   0.1087   1.0000
   9.500   0.9561   0.07743   0.07085  -0.0133   0.1101   1.0000
   9.750   0.9555   0.08301   0.07653  -0.0134   0.1114   1.0000
  10.000   0.8759   0.09112   0.08481  -0.0196   0.1207   1.0000
  10.250   0.8662   0.09834   0.09203  -0.0231   0.1261   1.0000
  10.500   0.8344   0.11244   0.10603  -0.0346   0.1526   1.0000
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