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HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il)
Reynolds number: 100,000
Max Cl/Cd: 57.56 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq258-il-100000-n5.txt
Download as CSV file: xf-hq258-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/8 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3205   0.09195   0.08733  -0.0294   1.0000   0.0357
  -8.750  -0.3198   0.08812   0.08354  -0.0302   1.0000   0.0344
  -8.500  -0.3205   0.08415   0.07962  -0.0311   1.0000   0.0327
  -8.250  -0.3230   0.07960   0.07508  -0.0327   1.0000   0.0249
  -8.000  -0.3236   0.07577   0.07131  -0.0329   1.0000   0.0233
  -7.500  -0.4057   0.08227   0.07763  -0.0322   1.0000   0.0242
  -7.250  -0.4095   0.07913   0.07458  -0.0327   1.0000   0.0229
  -7.000  -0.4113   0.07544   0.07097  -0.0346   1.0000   0.0216
  -6.750  -0.4128   0.07126   0.06685  -0.0376   1.0000   0.0205
  -6.500  -0.4136   0.06682   0.06242  -0.0402   1.0000   0.0194
  -6.250  -0.4133   0.06154   0.05707  -0.0431   1.0000   0.0180
  -5.750  -0.3969   0.05394   0.04914  -0.0443   1.0000   0.0158
  -5.500  -0.3904   0.05015   0.04522  -0.0445   1.0000   0.0155
  -5.250  -0.3798   0.04629   0.04115  -0.0450   0.9999   0.0152
  -5.000  -0.3477   0.04107   0.03552  -0.0495   0.9951   0.0149
  -4.750  -0.3148   0.03647   0.03043  -0.0528   0.9903   0.0150
  -4.500  -0.2784   0.03248   0.02578  -0.0557   0.9861   0.0165
  -4.250  -0.2483   0.02917   0.02208  -0.0581   0.9814   0.0185
  -4.000  -0.2143   0.02653   0.01897  -0.0599   0.9771   0.0198
  -3.750  -0.1781   0.02385   0.01576  -0.0614   0.9740   0.0205
  -3.500  -0.1463   0.02178   0.01329  -0.0619   0.9691   0.0216
  -3.250  -0.1129   0.01999   0.01121  -0.0626   0.9649   0.0238
  -3.000  -0.0782   0.01861   0.00959  -0.0638   0.9614   0.0290
  -2.750  -0.0476   0.01765   0.00855  -0.0645   0.9557   0.0385
  -2.500  -0.0143   0.01656   0.00740  -0.0656   0.9509   0.0589
  -2.250   0.0218   0.01561   0.00667  -0.0677   0.9474   0.1235
  -2.000   0.0468   0.01387   0.00638  -0.0682   0.9409   0.4845
  -1.750   0.0752   0.01348   0.00647  -0.0676   0.9360   0.6813
  -1.500   0.0989   0.01332   0.00639  -0.0659   0.9291   0.7732
  -1.250   0.1208   0.01298   0.00627  -0.0631   0.9231   0.8710
  -1.000   0.1686   0.01274   0.00594  -0.0666   0.9208   0.9478
  -0.750   0.2066   0.01269   0.00571  -0.0691   0.9131   1.0000
  -0.500   0.2439   0.01268   0.00547  -0.0712   0.9077   1.0000
  -0.250   0.2737   0.01273   0.00537  -0.0718   0.8986   1.0000
   0.000   0.3073   0.01273   0.00524  -0.0731   0.8912   1.0000
   0.250   0.3391   0.01274   0.00515  -0.0739   0.8826   1.0000
   0.500   0.3686   0.01278   0.00512  -0.0742   0.8730   1.0000
   0.750   0.4002   0.01279   0.00505  -0.0749   0.8647   1.0000
   1.000   0.4301   0.01281   0.00503  -0.0752   0.8552   1.0000
   1.250   0.4586   0.01281   0.00500  -0.0751   0.8432   1.0000
   1.500   0.4872   0.01277   0.00492  -0.0749   0.8294   1.0000
   1.750   0.5153   0.01273   0.00489  -0.0745   0.8148   1.0000
   2.000   0.5429   0.01274   0.00488  -0.0742   0.8006   1.0000
   2.250   0.5702   0.01276   0.00490  -0.0738   0.7862   1.0000
   2.500   0.5965   0.01281   0.00498  -0.0732   0.7695   1.0000
   2.750   0.6231   0.01285   0.00510  -0.0726   0.7516   1.0000
   3.000   0.6502   0.01288   0.00515  -0.0721   0.7329   1.0000
   3.250   0.6760   0.01296   0.00527  -0.0714   0.7107   1.0000
   3.500   0.7020   0.01305   0.00539  -0.0707   0.6865   1.0000
   3.750   0.7276   0.01317   0.00551  -0.0699   0.6579   1.0000
   4.000   0.7527   0.01333   0.00566  -0.0689   0.6235   1.0000
   4.250   0.7772   0.01357   0.00595  -0.0679   0.5818   1.0000
   4.500   0.8007   0.01391   0.00618  -0.0668   0.5326   1.0000
   4.750   0.8230   0.01439   0.00649  -0.0655   0.4773   1.0000
   5.000   0.8443   0.01502   0.00692  -0.0643   0.4203   1.0000
   5.250   0.8650   0.01574   0.00746  -0.0631   0.3656   1.0000
   5.500   0.8859   0.01650   0.00808  -0.0620   0.3171   1.0000
   5.750   0.9059   0.01736   0.00877  -0.0609   0.2668   1.0000
   6.000   0.9247   0.01837   0.00948  -0.0597   0.2071   1.0000
   6.250   0.9418   0.01968   0.01039  -0.0586   0.1229   1.0000
   6.500   0.9550   0.02177   0.01169  -0.0570   0.0301   1.0000
   6.750   0.9725   0.02344   0.01333  -0.0553   0.0177   1.0000
   7.000   0.9917   0.02480   0.01497  -0.0539   0.0153   1.0000
   7.250   1.0094   0.02626   0.01670  -0.0523   0.0138   1.0000
   7.500   1.0256   0.02787   0.01856  -0.0506   0.0126   1.0000
   7.750   1.0409   0.02956   0.02044  -0.0489   0.0113   1.0000
   8.000   1.0509   0.03222   0.02332  -0.0469   0.0096   1.0000
   8.250   1.0669   0.03411   0.02542  -0.0454   0.0090   1.0000
   8.500   1.0822   0.03632   0.02792  -0.0438   0.0085   1.0000
   8.750   1.0961   0.03896   0.03087  -0.0422   0.0084   1.0000
   9.000   1.1075   0.04189   0.03416  -0.0404   0.0083   1.0000
   9.250   1.1154   0.04500   0.03765  -0.0384   0.0082   1.0000
   9.500   1.1186   0.04826   0.04129  -0.0362   0.0082   1.0000
   9.750   1.1158   0.05152   0.04492  -0.0335   0.0082   1.0000
  10.000   1.1084   0.05476   0.04859  -0.0309   0.0082   1.0000
  10.250   1.0975   0.05825   0.05238  -0.0288   0.0082   1.0000
  10.500   1.0843   0.06208   0.05648  -0.0275   0.0083   1.0000
  10.750   1.0693   0.06636   0.06102  -0.0273   0.0083   1.0000
  11.000   1.0533   0.07121   0.06609  -0.0282   0.0084   1.0000
  11.250   1.0358   0.07688   0.07197  -0.0305   0.0084   1.0000
  11.500   1.0182   0.08335   0.07862  -0.0342   0.0084   1.0000
  11.750   1.0000   0.09097   0.08629  -0.0392   0.0085   1.0000
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