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HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/9 AIRFOIL (hq259-il)
Reynolds number: 200,000
Max Cl/Cd: 79.98 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq259-il-200000.txt
Download as CSV file: xf-hq259-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/9 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4073   0.08675   0.08344  -0.0406   1.0000   0.0364
  -8.000  -0.4150   0.08338   0.08015  -0.0423   1.0000   0.0365
  -7.750  -0.4217   0.07935   0.07615  -0.0459   1.0000   0.0365
  -7.500  -0.4279   0.07615   0.07294  -0.0467   1.0000   0.0366
  -7.250  -0.4371   0.07326   0.07002  -0.0463   1.0000   0.0366
  -7.000  -0.4457   0.07063   0.06733  -0.0450   1.0000   0.0367
  -6.750  -0.4609   0.06477   0.06150  -0.0442   1.0000   0.0373
  -6.500  -0.4634   0.06145   0.05826  -0.0421   1.0000   0.0381
  -6.250  -0.4638   0.05890   0.05571  -0.0405   1.0000   0.0388
  -6.000  -0.4622   0.05635   0.05314  -0.0393   1.0000   0.0398
  -5.750  -0.4380   0.05210   0.04874  -0.0434   0.9966   0.0423
  -5.500  -0.3934   0.04880   0.04461  -0.0505   0.9908   0.0492
  -5.250  -0.3733   0.04147   0.03737  -0.0543   0.9869   0.0524
  -5.000  -0.3357   0.03809   0.03339  -0.0588   0.9829   0.0639
  -4.750  -0.3078   0.03452   0.02985  -0.0607   0.9785   0.0671
  -4.500  -0.2673   0.02706   0.02129  -0.0612   0.9747   0.0344
  -4.250  -0.2315   0.02272   0.01633  -0.0627   0.9718   0.0299
  -4.000  -0.1934   0.01984   0.01287  -0.0641   0.9695   0.0286
  -3.750  -0.1612   0.01821   0.01102  -0.0647   0.9648   0.0299
  -3.500  -0.1265   0.01719   0.00983  -0.0658   0.9603   0.0350
  -3.250  -0.0896   0.01571   0.00833  -0.0676   0.9572   0.0429
  -3.000  -0.0536   0.01449   0.00713  -0.0693   0.9536   0.0668
  -2.750  -0.0223   0.01390   0.00660  -0.0703   0.9473   0.1010
  -2.500   0.0130   0.01216   0.00585  -0.0729   0.9439   0.3200
  -2.250   0.0437   0.01126   0.00602  -0.0734   0.9401   0.6424
  -2.000   0.0722   0.01121   0.00605  -0.0731   0.9331   0.7064
  -1.750   0.1090   0.01112   0.00595  -0.0743   0.9295   0.7486
  -1.500   0.1368   0.01107   0.00589  -0.0738   0.9224   0.7801
  -1.250   0.1676   0.01093   0.00578  -0.0737   0.9171   0.8154
  -1.000   0.1903   0.01079   0.00571  -0.0715   0.9101   0.8578
  -0.750   0.2156   0.01056   0.00553  -0.0701   0.9036   0.8889
  -0.500   0.2449   0.01040   0.00534  -0.0699   0.8973   0.9094
  -0.250   0.2792   0.01025   0.00516  -0.0709   0.8908   0.9304
   0.000   0.3232   0.01005   0.00492  -0.0739   0.8871   0.9519
   0.250   0.3675   0.00991   0.00474  -0.0774   0.8783   0.9773
   0.500   0.4073   0.00970   0.00447  -0.0798   0.8685   1.0000
   0.750   0.4326   0.00960   0.00427  -0.0792   0.8574   1.0000
   1.000   0.4594   0.00956   0.00416  -0.0790   0.8472   1.0000
   1.250   0.4854   0.00958   0.00414  -0.0786   0.8360   1.0000
   1.500   0.5121   0.00959   0.00410  -0.0782   0.8251   1.0000
   1.750   0.5390   0.00958   0.00405  -0.0778   0.8140   1.0000
   2.000   0.5658   0.00957   0.00400  -0.0774   0.8020   1.0000
   2.250   0.5924   0.00957   0.00397  -0.0769   0.7893   1.0000
   2.500   0.6189   0.00958   0.00399  -0.0763   0.7757   1.0000
   2.750   0.6451   0.00961   0.00400  -0.0757   0.7609   1.0000
   3.000   0.6711   0.00963   0.00401  -0.0751   0.7444   1.0000
   3.250   0.6972   0.00966   0.00402  -0.0744   0.7264   1.0000
   3.500   0.7230   0.00972   0.00406  -0.0738   0.7063   1.0000
   3.750   0.7486   0.00980   0.00415  -0.0730   0.6842   1.0000
   4.000   0.7737   0.00991   0.00424  -0.0723   0.6583   1.0000
   4.250   0.7984   0.01006   0.00434  -0.0714   0.6283   1.0000
   4.500   0.8222   0.01028   0.00449  -0.0704   0.5902   1.0000
   4.750   0.8449   0.01060   0.00470  -0.0692   0.5429   1.0000
   5.000   0.8660   0.01107   0.00495  -0.0679   0.4864   1.0000
   5.250   0.8861   0.01169   0.00532  -0.0665   0.4285   1.0000
   5.500   0.9062   0.01237   0.00577  -0.0652   0.3759   1.0000
   5.750   0.9265   0.01307   0.00628  -0.0640   0.3301   1.0000
   6.000   0.9462   0.01382   0.00683  -0.0629   0.2871   1.0000
   6.250   0.9658   0.01457   0.00735  -0.0618   0.2379   1.0000
   6.500   0.9866   0.01525   0.00794  -0.0609   0.1992   1.0000
   6.750   1.0064   0.01603   0.00849  -0.0599   0.1536   1.0000
   7.000   1.0252   0.01701   0.00918  -0.0588   0.1005   1.0000
   7.250   1.0342   0.01936   0.01092  -0.0561   0.0261   1.0000
   7.500   1.0519   0.02058   0.01228  -0.0544   0.0214   1.0000
   7.750   1.0642   0.02231   0.01418  -0.0520   0.0187   1.0000
   8.000   1.0802   0.02358   0.01557  -0.0502   0.0171   1.0000
   8.250   1.0965   0.02482   0.01694  -0.0486   0.0152   1.0000
   8.500   1.1106   0.02648   0.01873  -0.0466   0.0145   1.0000
   8.750   1.1255   0.02832   0.02070  -0.0448   0.0140   1.0000
   9.000   1.1415   0.03041   0.02302  -0.0431   0.0136   1.0000
   9.250   1.1585   0.03285   0.02569  -0.0417   0.0136   1.0000
   9.500   1.1742   0.03570   0.02884  -0.0401   0.0137   1.0000
   9.750   1.1859   0.03900   0.03254  -0.0381   0.0141   1.0000
  10.000   1.1913   0.04271   0.03668  -0.0355   0.0147   1.0000
  10.250   1.1888   0.04643   0.04081  -0.0324   0.0152   1.0000
  10.500   1.1797   0.04992   0.04464  -0.0288   0.0157   1.0000
  10.750   1.1671   0.05345   0.04846  -0.0257   0.0161   1.0000
  11.000   1.1513   0.05727   0.05256  -0.0232   0.0164   1.0000
  11.250   1.1345   0.06133   0.05685  -0.0218   0.0167   1.0000
  11.500   1.1157   0.06588   0.06163  -0.0214   0.0169   1.0000
  11.750   1.0975   0.07076   0.06671  -0.0220   0.0171   1.0000
  12.000   1.0767   0.07650   0.07264  -0.0239   0.0172   1.0000
  12.250   1.0572   0.08266   0.07897  -0.0269   0.0174   1.0000
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