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HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il)
Reynolds number: 50,000
Max Cl/Cd: 35.66 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq159b-il-50000-n5.txt
Download as CSV file: xf-hq159b-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/9 B 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5131   0.09372   0.08675  -0.0232   1.0000   0.0437
  -9.000  -0.5166   0.08859   0.08170  -0.0264   1.0000   0.0429
  -8.750  -0.5227   0.08282   0.07602  -0.0305   1.0000   0.0419
  -8.500  -0.5373   0.07600   0.06927  -0.0365   1.0000   0.0404
  -7.750  -0.5618   0.06094   0.05361  -0.0433   1.0000   0.0384
  -7.500  -0.5596   0.05678   0.04920  -0.0436   1.0000   0.0384
  -7.250  -0.5554   0.05253   0.04474  -0.0434   1.0000   0.0387
  -7.000  -0.5478   0.04894   0.04097  -0.0430   1.0000   0.0399
  -6.750  -0.5360   0.04657   0.03854  -0.0424   1.0000   0.0427
  -6.500  -0.5246   0.04352   0.03513  -0.0416   1.0000   0.0457
  -6.250  -0.5114   0.04021   0.03130  -0.0407   1.0000   0.0481
  -6.000  -0.4954   0.03690   0.02740  -0.0394   1.0000   0.0498
  -5.750  -0.4773   0.03393   0.02391  -0.0381   1.0000   0.0521
  -5.500  -0.4600   0.03222   0.02215  -0.0372   1.0000   0.0591
  -5.250  -0.4392   0.03012   0.01949  -0.0358   1.0000   0.0651
  -5.000  -0.4190   0.02811   0.01744  -0.0344   1.0000   0.0697
  -4.750  -0.3978   0.02658   0.01561  -0.0329   1.0000   0.0766
  -4.500  -0.3786   0.02539   0.01438  -0.0315   1.0000   0.0881
  -4.250  -0.3591   0.02413   0.01299  -0.0299   1.0000   0.0960
  -4.000  -0.3398   0.02308   0.01179  -0.0283   1.0000   0.1055
  -3.750  -0.3207   0.02208   0.01083  -0.0272   1.0000   0.1231
  -3.500  -0.3011   0.02110   0.00992  -0.0262   1.0000   0.1468
  -3.250  -0.2817   0.01986   0.00907  -0.0255   1.0000   0.2095
  -3.000  -0.2671   0.01804   0.00867  -0.0239   1.0000   0.4541
  -2.750  -0.2564   0.01772   0.00893  -0.0197   1.0000   0.6287
  -2.500  -0.2440   0.01770   0.00898  -0.0159   1.0000   0.7132
  -2.250  -0.2347   0.01768   0.00900  -0.0110   1.0000   0.7872
  -2.000  -0.2239   0.01759   0.00901  -0.0061   1.0000   0.8603
  -1.750  -0.1798   0.01758   0.00886  -0.0080   1.0000   0.9368
  -1.500  -0.1141   0.01758   0.00847  -0.0158   1.0000   0.9991
  -1.250  -0.0955   0.01755   0.00823  -0.0161   0.9953   1.0000
  -1.000  -0.0598   0.01769   0.00811  -0.0191   0.9858   1.0000
  -0.750  -0.0239   0.01788   0.00807  -0.0219   0.9766   1.0000
  -0.500   0.0152   0.01812   0.00808  -0.0252   0.9679   1.0000
  -0.250   0.0529   0.01833   0.00813  -0.0280   0.9574   1.0000
   0.000   0.0900   0.01854   0.00820  -0.0306   0.9465   1.0000
   0.250   0.1274   0.01876   0.00831  -0.0331   0.9364   1.0000
   0.500   0.1673   0.01899   0.00846  -0.0360   0.9277   1.0000
   0.750   0.1998   0.01921   0.00863  -0.0375   0.9170   1.0000
   1.000   0.2342   0.01945   0.00884  -0.0393   0.9071   1.0000
   1.250   0.2738   0.01967   0.00906  -0.0419   0.8988   1.0000
   1.500   0.3066   0.01989   0.00931  -0.0431   0.8879   1.0000
   1.750   0.3390   0.02009   0.00957  -0.0442   0.8762   1.0000
   2.000   0.3767   0.02011   0.00965  -0.0457   0.8612   1.0000
   2.250   0.4143   0.01997   0.00959  -0.0466   0.8426   1.0000
   2.500   0.4453   0.01988   0.00961  -0.0464   0.8216   1.0000
   2.750   0.4772   0.01982   0.00964  -0.0463   0.8041   1.0000
   3.000   0.5053   0.01986   0.00980  -0.0458   0.7868   1.0000
   3.250   0.5318   0.01994   0.01005  -0.0451   0.7678   1.0000
   3.500   0.5600   0.01994   0.01019  -0.0444   0.7490   1.0000
   3.750   0.5858   0.01999   0.01039  -0.0434   0.7274   1.0000
   4.000   0.6132   0.01996   0.01051  -0.0424   0.7048   1.0000
   4.250   0.6385   0.02001   0.01078  -0.0411   0.6781   1.0000
   4.500   0.6627   0.02008   0.01101  -0.0396   0.6464   1.0000
   4.750   0.6865   0.02010   0.01113  -0.0379   0.6055   1.0000
   5.000   0.7086   0.02016   0.01116  -0.0357   0.5505   1.0000
   5.250   0.7296   0.02046   0.01126  -0.0335   0.4869   1.0000
   5.500   0.7496   0.02110   0.01166  -0.0315   0.4276   1.0000
   5.750   0.7670   0.02198   0.01231  -0.0297   0.3617   1.0000
   6.000   0.7801   0.02322   0.01313  -0.0277   0.2836   1.0000
   6.250   0.7942   0.02460   0.01412  -0.0262   0.2114   1.0000
   6.500   0.8103   0.02603   0.01531  -0.0250   0.1482   1.0000
   6.750   0.8254   0.02784   0.01680  -0.0236   0.1127   1.0000
   7.000   0.8425   0.02968   0.01858  -0.0223   0.0949   1.0000
   7.250   0.8614   0.03151   0.02041  -0.0210   0.0848   1.0000
   7.500   0.8836   0.03325   0.02238  -0.0200   0.0762   1.0000
   7.750   0.9039   0.03499   0.02428  -0.0191   0.0665   1.0000
   8.000   0.9252   0.03707   0.02640  -0.0184   0.0609   1.0000
   8.250   0.9498   0.03976   0.02946  -0.0176   0.0571   1.0000
   8.500   0.9699   0.04253   0.03265  -0.0165   0.0528   1.0000
   8.750   0.9842   0.04488   0.03517  -0.0157   0.0477   1.0000
   9.000   0.9958   0.04817   0.03901  -0.0143   0.0447   1.0000
   9.250   1.0037   0.05196   0.04332  -0.0127   0.0432   1.0000
   9.500   1.0062   0.05586   0.04769  -0.0110   0.0420   1.0000
   9.750   1.0036   0.05967   0.05189  -0.0094   0.0409   1.0000
  10.000   0.9968   0.06337   0.05591  -0.0079   0.0399   1.0000
  10.250   0.9858   0.06681   0.05958  -0.0064   0.0391   1.0000
  10.500   0.9726   0.07049   0.06344  -0.0056   0.0384   1.0000
  10.750   0.9464   0.07603   0.06925  -0.0065   0.0394   1.0000
  11.000   0.9198   0.08259   0.07600  -0.0095   0.0404   1.0000
  11.250   0.8937   0.09047   0.08400  -0.0145   0.0418   1.0000
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