Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HQ 3.5/10 AIRFOIL (hq3510-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 3.5/10 AIRFOIL (hq3510-il)
Reynolds number: 50,000
Max Cl/Cd: 37.42 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq3510-il-50000.txt
Download as CSV file: xf-hq3510-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.5/10 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3523   0.11557   0.10889  -0.0342   1.0000   0.1762
  -9.000  -0.3374   0.11088   0.10422  -0.0326   1.0000   0.1842
  -8.750  -0.3536   0.11038   0.10389  -0.0331   1.0000   0.1905
  -8.500  -0.3408   0.10628   0.09982  -0.0312   1.0000   0.2017
  -8.250  -0.3419   0.10365   0.09729  -0.0301   1.0000   0.2104
  -8.000  -0.3577   0.10297   0.09677  -0.0291   1.0000   0.2185
  -7.750  -0.3533   0.09998   0.09381  -0.0269   1.0000   0.2313
  -7.500  -0.3524   0.09735   0.09126  -0.0247   1.0000   0.2424
  -7.250  -0.3830   0.09771   0.09182  -0.0220   1.0000   0.2473
  -7.000  -0.3832   0.09524   0.08943  -0.0190   1.0000   0.2608
  -6.750  -0.3881   0.09316   0.08745  -0.0158   1.0000   0.2743
  -6.500  -0.3929   0.09112   0.08550  -0.0126   1.0000   0.2874
  -6.250  -0.4004   0.08930   0.08378  -0.0093   1.0000   0.3009
  -6.000  -0.4105   0.08769   0.08227  -0.0063   1.0000   0.3163
  -5.750  -0.4277   0.08653   0.08124  -0.0039   1.0000   0.3327
  -5.500  -0.4164   0.08355   0.07830   0.0006   1.0000   0.3571
  -5.250  -0.4175   0.08144   0.07626   0.0045   1.0000   0.3827
  -5.000  -0.4180   0.07951   0.07440   0.0088   1.0000   0.4121
  -4.750  -0.4239   0.07806   0.07304   0.0132   1.0000   0.4447
  -4.500  -0.4140   0.07559   0.07061   0.0185   1.0000   0.4852
  -3.750  -0.3060   0.04583   0.03841  -0.0524   1.0000   0.1481
  -3.500  -0.2745   0.04158   0.03372  -0.0542   1.0000   0.1293
  -3.250  -0.2394   0.03837   0.02963  -0.0560   1.0000   0.1188
  -3.000  -0.2097   0.03604   0.02669  -0.0568   1.0000   0.1198
  -2.750  -0.1815   0.03374   0.02396  -0.0571   1.0000   0.1213
  -2.500  -0.1551   0.03170   0.02166  -0.0569   1.0000   0.1241
  -2.250  -0.1290   0.03029   0.01991  -0.0565   1.0000   0.1361
  -2.000  -0.1050   0.02884   0.01844  -0.0557   1.0000   0.1531
  -1.750  -0.0810   0.02766   0.01722  -0.0548   1.0000   0.1788
  -1.500  -0.0545   0.02656   0.01621  -0.0545   1.0000   0.2211
  -1.250  -0.0229   0.02386   0.01588  -0.0547   1.0000   0.5940
  -1.000  -0.0355   0.02257   0.01544  -0.0438   1.0000   1.0000
  -0.750  -0.0101   0.02303   0.01516  -0.0443   1.0000   1.0000
  -0.500   0.0143   0.02354   0.01520  -0.0449   1.0000   1.0000
  -0.250   0.0379   0.02409   0.01536  -0.0455   1.0000   1.0000
   0.000   0.0608   0.02469   0.01564  -0.0461   1.0000   1.0000
   0.250   0.0829   0.02533   0.01597  -0.0465   1.0000   1.0000
   0.500   0.1045   0.02602   0.01642  -0.0469   1.0000   1.0000
   0.750   0.1255   0.02675   0.01694  -0.0472   1.0000   1.0000
   1.000   0.1460   0.02753   0.01754  -0.0475   1.0000   1.0000
   1.250   0.1660   0.02836   0.01819  -0.0478   1.0000   1.0000
   1.500   0.1990   0.02965   0.01931  -0.0506   0.9942   1.0000
   1.750   0.2451   0.03120   0.02071  -0.0558   0.9785   1.0000
   2.000   0.2889   0.03263   0.02203  -0.0604   0.9623   1.0000
   2.250   0.3398   0.03410   0.02340  -0.0658   0.9431   1.0000
   2.500   0.3845   0.03515   0.02440  -0.0697   0.9201   1.0000
   2.750   0.4352   0.03613   0.02536  -0.0742   0.8977   1.0000
   3.000   0.4764   0.03689   0.02612  -0.0770   0.8763   1.0000
   3.250   0.5227   0.03759   0.02685  -0.0803   0.8569   1.0000
   3.500   0.5568   0.03818   0.02752  -0.0817   0.8359   1.0000
   3.750   0.6074   0.03844   0.02788  -0.0849   0.8169   1.0000
   4.000   0.6377   0.03885   0.02837  -0.0853   0.7945   1.0000
   4.250   0.6909   0.03854   0.02826  -0.0880   0.7754   1.0000
   4.500   0.7235   0.03856   0.02842  -0.0880   0.7517   1.0000
   4.750   0.7847   0.03722   0.02731  -0.0906   0.7332   1.0000
   5.000   0.8220   0.03655   0.02683  -0.0902   0.7086   1.0000
   5.250   0.8683   0.03523   0.02577  -0.0904   0.6853   1.0000
   5.500   0.9301   0.03273   0.02353  -0.0914   0.6640   1.0000
   5.750   0.9743   0.03125   0.02227  -0.0909   0.6365   1.0000
   6.000   1.0174   0.02996   0.02109  -0.0904   0.6073   1.0000
   6.250   1.0473   0.02970   0.02091  -0.0890   0.5759   1.0000
   6.500   1.0783   0.02959   0.02083  -0.0879   0.5460   1.0000
   6.750   1.1073   0.02981   0.02112  -0.0868   0.5175   1.0000
   7.000   1.1343   0.03031   0.02163  -0.0856   0.4904   1.0000
   7.250   1.1605   0.03101   0.02232  -0.0845   0.4641   1.0000
   7.500   1.1864   0.03184   0.02314  -0.0835   0.4382   1.0000
   7.750   1.2097   0.03287   0.02420  -0.0822   0.4124   1.0000
   8.000   1.2284   0.03416   0.02567  -0.0805   0.3870   1.0000
   8.250   1.2475   0.03548   0.02707  -0.0788   0.3611   1.0000
   8.500   1.2663   0.03690   0.02855  -0.0770   0.3349   1.0000
   8.750   1.2840   0.03841   0.03011  -0.0751   0.3078   1.0000
   9.000   1.2991   0.03997   0.03170  -0.0729   0.2793   1.0000
   9.250   1.3111   0.04153   0.03326  -0.0703   0.2492   1.0000
   9.500   1.3226   0.04265   0.03410  -0.0677   0.2156   1.0000
   9.750   1.3213   0.04402   0.03554  -0.0636   0.1864   1.0000
  10.000   1.3174   0.04569   0.03721  -0.0592   0.1578   1.0000
  10.250   1.3101   0.04708   0.03845  -0.0549   0.1329   1.0000
  10.500   1.3008   0.04917   0.04059  -0.0507   0.1144   1.0000
  10.750   1.2964   0.05156   0.04289  -0.0474   0.0990   1.0000
  11.000   1.2988   0.05504   0.04654  -0.0449   0.0879   1.0000
  11.250   1.3044   0.05909   0.05076  -0.0431   0.0807   1.0000
  11.500   1.2999   0.06267   0.05470  -0.0406   0.0764   1.0000
  11.750   1.3078   0.06664   0.05860  -0.0396   0.0709   1.0000
  12.000   1.2918   0.07058   0.06297  -0.0372   0.0701   1.0000
  12.250   1.2741   0.07499   0.06774  -0.0356   0.0696   1.0000
  12.500   1.2550   0.07983   0.07291  -0.0349   0.0695   1.0000
  12.750   1.2344   0.08515   0.07851  -0.0350   0.0697   1.0000
  13.000   1.2127   0.09100   0.08459  -0.0361   0.0700   1.0000
  13.250   1.1919   0.09734   0.09113  -0.0380   0.0706   1.0000
  13.500   1.1709   0.10423   0.09818  -0.0407   0.0711   1.0000
<< Back to HQ 3.5/10 AIRFOIL (hq3510-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.5/10 AIRFOIL (hq3510-il)