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HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/9 AIRFOIL (hq109-il)
Reynolds number: 50,000
Max Cl/Cd: 32.8 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq109-il-50000.txt
Download as CSV file: xf-hq109-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/9 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5085   0.09612   0.08946   0.0049   1.0000   0.3272
  -8.000  -0.4998   0.09256   0.08593   0.0060   1.0000   0.3487
  -7.750  -0.5078   0.09061   0.08408   0.0068   1.0000   0.3713
  -7.000  -0.5830   0.06164   0.05503  -0.0311   1.0000   0.1444
  -6.750  -0.5853   0.05449   0.04713  -0.0342   1.0000   0.1219
  -6.500  -0.5726   0.04954   0.04197  -0.0340   1.0000   0.1154
  -6.250  -0.5613   0.04476   0.03637  -0.0336   1.0000   0.1082
  -6.000  -0.5452   0.04089   0.03207  -0.0327   1.0000   0.1073
  -5.750  -0.5285   0.03753   0.02817  -0.0315   1.0000   0.1108
  -5.500  -0.5090   0.03468   0.02519  -0.0303   1.0000   0.1173
  -5.250  -0.4871   0.03170   0.02161  -0.0289   1.0000   0.1212
  -5.000  -0.4658   0.02931   0.01897  -0.0275   1.0000   0.1341
  -4.750  -0.4430   0.02710   0.01648  -0.0261   1.0000   0.1504
  -4.500  -0.4214   0.02513   0.01459  -0.0246   1.0000   0.1784
  -4.250  -0.3990   0.02307   0.01276  -0.0229   1.0000   0.2117
  -4.000  -0.3787   0.02095   0.01104  -0.0212   1.0000   0.2761
  -3.750  -0.3729   0.01807   0.01019  -0.0166   1.0000   0.4990
  -3.500  -0.3784   0.01829   0.01120  -0.0054   1.0000   0.7207
  -3.250  -0.3765   0.01869   0.01158   0.0036   1.0000   0.8038
  -3.000  -0.2027   0.02060   0.01207  -0.0111   1.0000   0.9502
  -2.750  -0.0633   0.01924   0.00974  -0.0331   1.0000   1.0000
  -2.500  -0.0545   0.01878   0.00920  -0.0317   1.0000   1.0000
  -2.250  -0.0484   0.01840   0.00877  -0.0298   1.0000   1.0000
  -2.000  -0.0460   0.01810   0.00844  -0.0273   1.0000   1.0000
  -1.750  -0.0482   0.01786   0.00820  -0.0240   1.0000   1.0000
  -1.500  -0.0561   0.01766   0.00799  -0.0198   1.0000   1.0000
  -1.250  -0.0682   0.01747   0.00776  -0.0149   1.0000   1.0000
  -1.000  -0.0789   0.01729   0.00752  -0.0102   1.0000   1.0000
  -0.750  -0.0810   0.01719   0.00732  -0.0067   1.0000   1.0000
  -0.500  -0.0725   0.01722   0.00720  -0.0048   1.0000   1.0000
  -0.250  -0.0587   0.01735   0.00719  -0.0038   1.0000   1.0000
   0.000  -0.0424   0.01754   0.00725  -0.0030   1.0000   1.0000
   0.250  -0.0249   0.01779   0.00738  -0.0025   1.0000   1.0000
   0.500  -0.0067   0.01809   0.00756  -0.0020   1.0000   1.0000
   0.750   0.0118   0.01843   0.00782  -0.0017   1.0000   1.0000
   1.000   0.0304   0.01882   0.00814  -0.0014   1.0000   1.0000
   1.250   0.0490   0.01925   0.00852  -0.0012   1.0000   1.0000
   1.500   0.0676   0.01972   0.00897  -0.0010   1.0000   1.0000
   1.750   0.0860   0.02025   0.00948  -0.0009   1.0000   1.0000
   2.000   0.1041   0.02083   0.01006  -0.0008   1.0000   1.0000
   2.250   0.1220   0.02147   0.01071  -0.0008   1.0000   1.0000
   2.500   0.1538   0.02245   0.01174  -0.0036   0.9938   1.0000
   2.750   0.2256   0.02396   0.01347  -0.0134   0.9662   1.0000
   3.000   0.2906   0.02509   0.01481  -0.0214   0.9400   1.0000
   3.250   0.3515   0.02591   0.01589  -0.0280   0.9137   1.0000
   3.500   0.4114   0.02643   0.01678  -0.0337   0.8859   1.0000
   3.750   0.4682   0.02661   0.01733  -0.0380   0.8547   1.0000
   4.000   0.5317   0.02619   0.01741  -0.0420   0.8201   1.0000
   4.250   0.5830   0.02535   0.01697  -0.0424   0.7816   1.0000
   4.500   0.6247   0.02411   0.01606  -0.0402   0.7378   1.0000
   4.750   0.6556   0.02285   0.01506  -0.0359   0.6866   1.0000
   5.000   0.6802   0.02179   0.01401  -0.0309   0.6241   1.0000
   5.250   0.6992   0.02138   0.01340  -0.0261   0.5451   1.0000
   5.500   0.7160   0.02183   0.01331  -0.0220   0.4564   1.0000
   5.750   0.7308   0.02324   0.01413  -0.0187   0.3652   1.0000
   6.000   0.7475   0.02529   0.01557  -0.0163   0.2866   1.0000
   6.250   0.7672   0.02759   0.01753  -0.0145   0.2277   1.0000
   6.500   0.7874   0.02967   0.01950  -0.0130   0.1850   1.0000
   6.750   0.8107   0.03212   0.02189  -0.0118   0.1559   1.0000
   7.000   0.8323   0.03478   0.02475  -0.0105   0.1343   1.0000
   7.250   0.8538   0.03802   0.02829  -0.0092   0.1209   1.0000
   7.500   0.8718   0.04112   0.03168  -0.0079   0.1090   1.0000
   7.750   0.8880   0.04507   0.03614  -0.0064   0.1044   1.0000
   8.000   0.9027   0.04916   0.04042  -0.0053   0.0991   1.0000
   8.250   0.9067   0.05341   0.04536  -0.0035   0.0967   1.0000
   8.500   0.9073   0.05802   0.05048  -0.0021   0.0955   1.0000
   8.750   0.9040   0.06300   0.05587  -0.0011   0.0961   1.0000
   9.000   0.8990   0.06813   0.06128  -0.0005   0.0977   1.0000
   9.250   0.8967   0.07342   0.06670  -0.0001   0.0993   1.0000
   9.500   0.8578   0.07880   0.07244  -0.0010   0.1039   1.0000
   9.750   0.8266   0.08538   0.07906  -0.0039   0.1094   1.0000
  10.000   0.8189   0.09131   0.08499  -0.0060   0.1120   1.0000
  10.250   0.7839   0.10521   0.09881  -0.0176   0.1392   1.0000
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