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

HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 1,000,000
Max Cl/Cd: 64.9 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq09-il-1000000.txt
Download as CSV file: xf-hq09-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/9 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.8237   0.05180   0.05008  -0.0198   1.0000   0.0045
  -9.750  -0.8347   0.04981   0.04801  -0.0193   1.0000   0.0047
  -9.500  -0.8462   0.04677   0.04483  -0.0177   1.0000   0.0047
  -9.250  -0.8429   0.04488   0.04282  -0.0169   1.0000   0.0049
  -9.000  -0.8783   0.03333   0.03060  -0.0134   1.0000   0.0038
  -8.750  -0.8865   0.02638   0.02295  -0.0106   1.0000   0.0036
  -8.500  -0.8780   0.02219   0.01824  -0.0087   1.0000   0.0035
  -8.250  -0.8630   0.01923   0.01489  -0.0071   1.0000   0.0035
  -8.000  -0.8444   0.01710   0.01245  -0.0059   1.0000   0.0035
  -7.750  -0.8239   0.01549   0.01060  -0.0049   1.0000   0.0036
  -7.500  -0.8020   0.01426   0.00918  -0.0040   1.0000   0.0037
  -7.250  -0.7791   0.01328   0.00806  -0.0032   1.0000   0.0039
  -7.000  -0.7551   0.01254   0.00719  -0.0026   1.0000   0.0042
  -6.750  -0.7328   0.01146   0.00598  -0.0018   1.0000   0.0058
  -6.500  -0.7072   0.01110   0.00560  -0.0014   1.0000   0.0070
  -6.250  -0.6827   0.01050   0.00498  -0.0009   1.0000   0.0099
  -6.000  -0.6561   0.01042   0.00493  -0.0008   1.0000   0.0123
  -5.750  -0.6292   0.01040   0.00495  -0.0007   1.0000   0.0130
  -5.500  -0.6056   0.00967   0.00413   0.0000   1.0000   0.0166
  -5.250  -0.5798   0.00948   0.00394   0.0003   1.0000   0.0191
  -5.000  -0.5543   0.00925   0.00369   0.0007   1.0000   0.0206
  -4.750  -0.5287   0.00911   0.00355   0.0011   1.0000   0.0218
  -4.500  -0.5052   0.00857   0.00294   0.0019   1.0000   0.0260
  -4.250  -0.4802   0.00827   0.00265   0.0024   0.9996   0.0306
  -4.000  -0.4447   0.00791   0.00233   0.0006   0.9922   0.0454
  -3.750  -0.4086   0.00750   0.00206  -0.0014   0.9828   0.0794
  -3.500  -0.3735   0.00695   0.00178  -0.0034   0.9698   0.1454
  -3.250  -0.3418   0.00654   0.00160  -0.0045   0.9514   0.2139
  -2.750  -0.2919   0.00587   0.00125  -0.0034   0.9015   0.3365
  -2.500  -0.2683   0.00542   0.00109  -0.0027   0.8803   0.4413
  -2.250  -0.2437   0.00507   0.00098  -0.0021   0.8638   0.5266
  -2.000  -0.2180   0.00487   0.00093  -0.0017   0.8495   0.5889
  -1.750  -0.1913   0.00479   0.00089  -0.0013   0.8363   0.6262
  -1.500  -0.1644   0.00472   0.00085  -0.0011   0.8242   0.6560
  -1.250  -0.1372   0.00468   0.00083  -0.0009   0.8134   0.6795
  -1.000  -0.1102   0.00464   0.00082  -0.0006   0.8021   0.7034
  -0.750  -0.0828   0.00463   0.00080  -0.0004   0.7891   0.7167
  -0.500  -0.0552   0.00463   0.00077  -0.0003   0.7744   0.7271
  -0.250  -0.0276   0.00464   0.00076  -0.0001   0.7613   0.7381
   0.250   0.0277   0.00464   0.00076   0.0001   0.7381   0.7613
   0.500   0.0553   0.00463   0.00077   0.0003   0.7271   0.7744
   0.750   0.0828   0.00463   0.00080   0.0004   0.7167   0.7891
   1.000   0.1102   0.00464   0.00082   0.0006   0.7034   0.8020
   1.250   0.1372   0.00468   0.00083   0.0009   0.6795   0.8133
   1.500   0.1644   0.00472   0.00085   0.0011   0.6560   0.8242
   1.750   0.1913   0.00479   0.00089   0.0013   0.6260   0.8363
   2.000   0.2180   0.00487   0.00093   0.0017   0.5894   0.8496
   2.250   0.2437   0.00507   0.00098   0.0021   0.5254   0.8639
   2.500   0.2683   0.00542   0.00109   0.0027   0.4412   0.8803
   2.750   0.2918   0.00588   0.00126   0.0034   0.3333   0.9017
   3.250   0.3418   0.00654   0.00160   0.0045   0.2139   0.9515
   3.500   0.3735   0.00695   0.00178   0.0034   0.1454   0.9699
   3.750   0.4087   0.00750   0.00206   0.0014   0.0793   0.9830
   4.000   0.4448   0.00791   0.00233  -0.0006   0.0456   0.9922
   4.250   0.4803   0.00827   0.00265  -0.0025   0.0307   0.9998
   4.500   0.5049   0.00857   0.00293  -0.0019   0.0260   1.0000
   4.750   0.5285   0.00911   0.00355  -0.0010   0.0218   1.0000
   5.000   0.5540   0.00927   0.00371  -0.0007   0.0207   1.0000
   5.250   0.5796   0.00948   0.00394  -0.0003   0.0191   1.0000
   5.500   0.6055   0.00967   0.00413   0.0000   0.0166   1.0000
   5.750   0.6291   0.01041   0.00495   0.0007   0.0130   1.0000
   6.000   0.6561   0.01042   0.00493   0.0008   0.0123   1.0000
   6.250   0.6827   0.01052   0.00501   0.0009   0.0100   1.0000
   6.500   0.7073   0.01110   0.00560   0.0014   0.0070   1.0000
   6.750   0.7329   0.01145   0.00598   0.0017   0.0058   1.0000
   7.000   0.7564   0.01226   0.00685   0.0024   0.0044   1.0000
   7.250   0.7791   0.01331   0.00809   0.0032   0.0039   1.0000
   7.500   0.8021   0.01425   0.00918   0.0039   0.0037   1.0000
   7.750   0.8240   0.01550   0.01061   0.0048   0.0036   1.0000
   8.000   0.8446   0.01710   0.01245   0.0059   0.0035   1.0000
   8.250   0.8631   0.01927   0.01492   0.0071   0.0035   1.0000
   8.500   0.8782   0.02220   0.01825   0.0086   0.0035   1.0000
   8.750   0.8865   0.02643   0.02301   0.0106   0.0036   1.0000
   9.000   0.8775   0.03357   0.03085   0.0134   0.0038   1.0000
   9.250   0.8498   0.04174   0.03955   0.0161   0.0042   1.0000
   9.500   0.8522   0.04434   0.04231   0.0171   0.0043   1.0000
   9.750   0.8481   0.04704   0.04515   0.0187   0.0044   1.0000
  10.000   0.8253   0.05161   0.04989   0.0198   0.0045   1.0000
<< Back to HQ 0/9 AIRFOIL (hq09-il)

Polar data table (+)

Polar graphs


<< Back to HQ 0/9 AIRFOIL (hq09-il)