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HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il)
Reynolds number: 50,000
Max Cl/Cd: 33.44 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq108-il-50000.txt
Download as CSV file: xf-hq108-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/8 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  -0.5417   0.11230   0.10550   0.0070   1.0000   0.2389
  -9.000  -0.5397   0.10885   0.10211   0.0065   1.0000   0.2504
  -8.500  -0.5350   0.10215   0.09553   0.0060   1.0000   0.2762
  -8.250  -0.5265   0.09849   0.09190   0.0065   1.0000   0.2922
  -8.000  -0.5264   0.09569   0.08917   0.0069   1.0000   0.3116
  -7.750  -0.5142   0.09195   0.08546   0.0083   1.0000   0.3341
  -7.500  -0.5061   0.08874   0.08229   0.0099   1.0000   0.3621
  -7.250  -0.5074   0.08605   0.07968   0.0109   1.0000   0.3840
  -6.250  -0.5480   0.05420   0.04715  -0.0318   1.0000   0.1328
  -6.000  -0.5347   0.04798   0.04016  -0.0335   1.0000   0.1124
  -5.750  -0.5167   0.04389   0.03561  -0.0332   1.0000   0.1077
  -5.500  -0.4984   0.04021   0.03144  -0.0327   1.0000   0.1078
  -5.250  -0.4781   0.03677   0.02746  -0.0318   1.0000   0.1080
  -5.000  -0.4556   0.03345   0.02365  -0.0308   1.0000   0.1075
  -4.750  -0.4313   0.03056   0.02021  -0.0295   1.0000   0.1090
  -4.500  -0.4077   0.02789   0.01740  -0.0283   1.0000   0.1170
  -4.250  -0.3841   0.02582   0.01503  -0.0270   1.0000   0.1354
  -4.000  -0.3599   0.02356   0.01286  -0.0253   1.0000   0.1589
  -3.750  -0.3374   0.02147   0.01100  -0.0236   1.0000   0.2070
  -3.500  -0.3210   0.01845   0.00919  -0.0216   1.0000   0.3425
  -3.250  -0.3341   0.01706   0.00999  -0.0087   1.0000   0.7462
  -3.000  -0.1055   0.01818   0.00907  -0.0322   1.0000   1.0000
  -2.750  -0.0945   0.01766   0.00839  -0.0312   1.0000   1.0000
  -2.500  -0.0849   0.01721   0.00782  -0.0298   1.0000   1.0000
  -2.250  -0.0774   0.01683   0.00734  -0.0279   1.0000   1.0000
  -2.000  -0.0731   0.01650   0.00696  -0.0254   1.0000   1.0000
  -1.750  -0.0726   0.01623   0.00659  -0.0221   1.0000   1.0000
  -1.500  -0.0756   0.01601   0.00632  -0.0183   1.0000   1.0000
  -1.250  -0.0778   0.01585   0.00608  -0.0144   1.0000   1.0000
  -1.000  -0.0733   0.01580   0.00588  -0.0116   1.0000   1.0000
  -0.750  -0.0616   0.01583   0.00575  -0.0098   1.0000   1.0000
  -0.500  -0.0463   0.01592   0.00569  -0.0086   1.0000   1.0000
  -0.250  -0.0292   0.01607   0.00569  -0.0077   1.0000   1.0000
   0.000  -0.0111   0.01625   0.00576  -0.0070   1.0000   1.0000
   0.250   0.0076   0.01648   0.00585  -0.0063   1.0000   1.0000
   0.500   0.0267   0.01674   0.00602  -0.0058   1.0000   1.0000
   0.750   0.0458   0.01704   0.00626  -0.0053   1.0000   1.0000
   1.000   0.0650   0.01738   0.00655  -0.0049   1.0000   1.0000
   1.250   0.0841   0.01776   0.00690  -0.0046   1.0000   1.0000
   1.500   0.1031   0.01819   0.00731  -0.0043   1.0000   1.0000
   1.750   0.1220   0.01865   0.00778  -0.0040   1.0000   1.0000
   2.000   0.1406   0.01917   0.00831  -0.0038   1.0000   1.0000
   2.250   0.1590   0.01974   0.00892  -0.0037   1.0000   1.0000
   2.500   0.1769   0.02038   0.00960  -0.0037   1.0000   1.0000
   2.750   0.1944   0.02108   0.01037  -0.0037   1.0000   1.0000
   3.000   0.2573   0.02250   0.01207  -0.0123   0.9768   1.0000
   3.250   0.3270   0.02370   0.01359  -0.0213   0.9468   1.0000
   3.500   0.3993   0.02451   0.01489  -0.0298   0.9152   1.0000
   3.750   0.4654   0.02480   0.01567  -0.0361   0.8792   1.0000
   4.000   0.5433   0.02418   0.01571  -0.0421   0.8363   1.0000
   4.250   0.5979   0.02292   0.01504  -0.0418   0.7853   1.0000
   4.500   0.6327   0.02144   0.01388  -0.0372   0.7258   1.0000
   4.750   0.6537   0.02038   0.01288  -0.0310   0.6492   1.0000
   5.000   0.6695   0.02002   0.01219  -0.0248   0.5394   1.0000
   5.250   0.6810   0.02122   0.01241  -0.0198   0.4000   1.0000
   5.500   0.6935   0.02374   0.01389  -0.0167   0.2766   1.0000
   5.750   0.7135   0.02642   0.01618  -0.0148   0.1998   1.0000
   6.000   0.7372   0.02896   0.01847  -0.0134   0.1557   1.0000
   6.250   0.7603   0.03154   0.02092  -0.0123   0.1259   1.0000
   6.500   0.7865   0.03495   0.02460  -0.0112   0.1124   1.0000
   6.750   0.8076   0.03794   0.02786  -0.0101   0.0992   1.0000
   7.000   0.8268   0.04172   0.03239  -0.0085   0.0948   1.0000
   7.250   0.8429   0.04606   0.03734  -0.0071   0.0939   1.0000
   7.500   0.8545   0.05075   0.04261  -0.0058   0.0947   1.0000
   7.750   0.8613   0.05558   0.04796  -0.0048   0.0953   1.0000
   8.000   0.8642   0.06059   0.05339  -0.0041   0.0963   1.0000
   8.250   0.8644   0.06566   0.05876  -0.0037   0.0975   1.0000
   8.500   0.8639   0.07082   0.06412  -0.0036   0.0989   1.0000
   8.750   0.8707   0.07621   0.06961  -0.0035   0.1009   1.0000
   9.000   0.8131   0.08505   0.07886  -0.0091   0.1231   1.0000
   9.250   0.7035   0.08433   0.07845  -0.0088   0.1236   1.0000
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