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HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.5/8 AIRFOIL (hq158-il)
Reynolds number: 100,000
Max Cl/Cd: 49.01 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq158-il-100000-n5.txt
Download as CSV file: xf-hq158-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5065   0.07964   0.07491  -0.0266   1.0000   0.0192
  -8.000  -0.5097   0.07421   0.06954  -0.0320   1.0000   0.0194
  -7.750  -0.5128   0.06949   0.06481  -0.0346   1.0000   0.0190
  -7.500  -0.5133   0.06481   0.06007  -0.0371   1.0000   0.0188
  -7.250  -0.5124   0.06005   0.05521  -0.0391   1.0000   0.0187
  -7.000  -0.5092   0.05547   0.05048  -0.0404   1.0000   0.0186
  -6.750  -0.5032   0.05087   0.04563  -0.0412   1.0000   0.0189
  -6.500  -0.4957   0.04669   0.04120  -0.0413   1.0000   0.0187
  -6.250  -0.4864   0.04250   0.03669  -0.0409   1.0000   0.0185
  -6.000  -0.4748   0.03859   0.03241  -0.0402   1.0000   0.0184
  -5.750  -0.4610   0.03494   0.02831  -0.0391   1.0000   0.0185
  -5.500  -0.4451   0.03165   0.02456  -0.0379   1.0000   0.0186
  -5.250  -0.4274   0.02884   0.02119  -0.0365   1.0000   0.0189
  -5.000  -0.4083   0.02641   0.01830  -0.0350   1.0000   0.0193
  -4.750  -0.3887   0.02400   0.01553  -0.0336   1.0000   0.0199
  -4.500  -0.3697   0.02217   0.01356  -0.0325   1.0000   0.0226
  -4.250  -0.3494   0.02116   0.01237  -0.0313   1.0000   0.0267
  -4.000  -0.3284   0.01970   0.01072  -0.0298   1.0000   0.0289
  -3.750  -0.3080   0.01842   0.00929  -0.0284   1.0000   0.0315
  -3.500  -0.2837   0.01725   0.00807  -0.0280   0.9982   0.0372
  -3.250  -0.2498   0.01636   0.00711  -0.0296   0.9923   0.0570
  -3.000  -0.2163   0.01538   0.00622  -0.0312   0.9865   0.0984
  -2.750  -0.1851   0.01413   0.00558  -0.0328   0.9803   0.2232
  -2.500  -0.1569   0.01275   0.00544  -0.0335   0.9744   0.5265
  -2.250  -0.1283   0.01250   0.00549  -0.0332   0.9670   0.6590
  -2.000  -0.0968   0.01241   0.00547  -0.0333   0.9606   0.7313
  -1.750  -0.0695   0.01232   0.00544  -0.0324   0.9529   0.7930
  -1.500  -0.0390   0.01220   0.00530  -0.0321   0.9471   0.8462
  -1.250  -0.0067   0.01207   0.00515  -0.0325   0.9403   0.8851
  -1.000   0.0356   0.01199   0.00498  -0.0352   0.9352   0.9138
  -0.750   0.0818   0.01191   0.00477  -0.0390   0.9302   0.9369
  -0.500   0.1279   0.01184   0.00460  -0.0429   0.9232   0.9638
  -0.250   0.1747   0.01175   0.00442  -0.0470   0.9161   0.9954
   0.000   0.2092   0.01172   0.00431  -0.0485   0.9059   1.0000
   0.250   0.2385   0.01173   0.00423  -0.0490   0.8942   1.0000
   0.500   0.2675   0.01175   0.00419  -0.0493   0.8824   1.0000
   0.750   0.2961   0.01176   0.00416  -0.0494   0.8696   1.0000
   1.000   0.3241   0.01178   0.00414  -0.0493   0.8562   1.0000
   1.250   0.3514   0.01182   0.00417  -0.0491   0.8427   1.0000
   1.500   0.3782   0.01186   0.00420  -0.0487   0.8288   1.0000
   1.750   0.4047   0.01191   0.00424  -0.0482   0.8136   1.0000
   2.000   0.4311   0.01194   0.00427  -0.0475   0.7971   1.0000
   2.250   0.4570   0.01201   0.00440  -0.0469   0.7799   1.0000
   2.500   0.4826   0.01208   0.00451  -0.0462   0.7617   1.0000
   2.750   0.5086   0.01216   0.00461  -0.0455   0.7428   1.0000
   3.000   0.5340   0.01225   0.00474  -0.0447   0.7204   1.0000
   3.250   0.5593   0.01235   0.00488  -0.0438   0.6953   1.0000
   3.500   0.5842   0.01247   0.00508  -0.0429   0.6645   1.0000
   3.750   0.6085   0.01264   0.00522  -0.0417   0.6238   1.0000
   4.000   0.6318   0.01289   0.00535  -0.0404   0.5667   1.0000
   4.250   0.6535   0.01335   0.00552  -0.0388   0.4901   1.0000
   4.500   0.6739   0.01405   0.00585  -0.0374   0.4107   1.0000
   4.750   0.6949   0.01479   0.00631  -0.0363   0.3354   1.0000
   5.000   0.7153   0.01566   0.00697  -0.0352   0.2893   1.0000
   5.250   0.7376   0.01637   0.00757  -0.0345   0.2334   1.0000
   5.500   0.7591   0.01722   0.00823  -0.0337   0.1730   1.0000
   5.750   0.7798   0.01826   0.00902  -0.0328   0.1191   1.0000
   6.000   0.7999   0.01946   0.01000  -0.0319   0.0847   1.0000
   6.250   0.8206   0.02062   0.01121  -0.0308   0.0671   1.0000
   6.500   0.8403   0.02198   0.01267  -0.0295   0.0541   1.0000
   6.750   0.8601   0.02320   0.01382  -0.0288   0.0357   1.0000
   7.000   0.8792   0.02463   0.01530  -0.0277   0.0239   1.0000
   7.250   0.8983   0.02625   0.01728  -0.0263   0.0193   1.0000
   7.500   0.9164   0.02809   0.01930  -0.0249   0.0168   1.0000
   7.750   0.9330   0.03044   0.02183  -0.0236   0.0152   1.0000
   8.000   0.9494   0.03360   0.02526  -0.0221   0.0144   1.0000
   8.250   0.9663   0.03648   0.02856  -0.0207   0.0139   1.0000
   8.500   0.9796   0.03978   0.03232  -0.0190   0.0135   1.0000
   8.750   0.9887   0.04327   0.03631  -0.0172   0.0130   1.0000
   9.000   0.9932   0.04687   0.04039  -0.0153   0.0126   1.0000
   9.250   0.9928   0.05057   0.04452  -0.0134   0.0121   1.0000
   9.500   0.9868   0.05438   0.04870  -0.0114   0.0117   1.0000
   9.750   0.9746   0.05799   0.05260  -0.0093   0.0115   1.0000
  10.000   0.9582   0.06184   0.05668  -0.0080   0.0114   1.0000
  10.250   0.9393   0.06652   0.06155  -0.0083   0.0122   1.0000
  10.500   0.9196   0.07194   0.06716  -0.0107   0.0119   1.0000
  10.750   0.9003   0.07852   0.07387  -0.0149   0.0124   1.0000
  11.000   0.8823   0.08632   0.08175  -0.0206   0.0131   1.0000
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