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HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il)
Reynolds number: 200,000
Max Cl/Cd: 54.59 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq108-il-200000-n5.txt
Download as CSV file: xf-hq108-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5639   0.08665   0.08321  -0.0116   1.0000   0.0107
  -8.750  -0.5656   0.08196   0.07857  -0.0145   1.0000   0.0115
  -8.500  -0.5720   0.07610   0.07277  -0.0186   1.0000   0.0109
  -8.250  -0.5812   0.06991   0.06662  -0.0247   1.0000   0.0111
  -8.000  -0.5908   0.06337   0.06001  -0.0296   1.0000   0.0105
  -7.750  -0.5938   0.05773   0.05424  -0.0323   1.0000   0.0105
  -7.500  -0.5959   0.05108   0.04736  -0.0341   1.0000   0.0099
  -7.250  -0.5932   0.04514   0.04110  -0.0347   1.0000   0.0098
  -7.000  -0.5882   0.03852   0.03401  -0.0343   1.0000   0.0094
  -6.750  -0.5781   0.03334   0.02831  -0.0333   1.0000   0.0092
  -6.500  -0.5639   0.02932   0.02379  -0.0321   1.0000   0.0091
  -6.250  -0.5470   0.02610   0.02009  -0.0309   1.0000   0.0091
  -6.000  -0.5282   0.02354   0.01713  -0.0296   1.0000   0.0094
  -5.750  -0.5081   0.02142   0.01467  -0.0283   1.0000   0.0097
  -5.500  -0.4874   0.01967   0.01265  -0.0271   1.0000   0.0101
  -5.250  -0.4664   0.01822   0.01101  -0.0258   1.0000   0.0110
  -5.000  -0.4457   0.01699   0.00961  -0.0245   1.0000   0.0121
  -4.750  -0.4244   0.01634   0.00885  -0.0234   1.0000   0.0144
  -4.500  -0.4051   0.01509   0.00749  -0.0219   1.0000   0.0161
  -4.250  -0.3717   0.01395   0.00624  -0.0234   0.9941   0.0195
  -4.000  -0.3378   0.01327   0.00536  -0.0248   0.9880   0.0252
  -3.750  -0.3042   0.01249   0.00463  -0.0263   0.9821   0.0439
  -3.500  -0.2704   0.01193   0.00413  -0.0279   0.9758   0.0723
  -3.250  -0.2378   0.01122   0.00370  -0.0294   0.9691   0.1359
  -3.000  -0.2060   0.01051   0.00334  -0.0308   0.9617   0.2356
  -2.750  -0.1773   0.00955   0.00302  -0.0317   0.9534   0.4130
  -2.500  -0.1469   0.00900   0.00294  -0.0324   0.9460   0.5469
  -2.250  -0.1180   0.00878   0.00289  -0.0325   0.9363   0.6185
  -2.000  -0.0881   0.00865   0.00284  -0.0326   0.9275   0.6676
  -1.750  -0.0585   0.00854   0.00281  -0.0327   0.9190   0.7131
  -1.500  -0.0306   0.00847   0.00274  -0.0323   0.9089   0.7437
  -1.250  -0.0023   0.00842   0.00267  -0.0321   0.8990   0.7635
  -1.000   0.0250   0.00836   0.00262  -0.0317   0.8888   0.7861
  -0.500   0.0776   0.00827   0.00251  -0.0303   0.8664   0.8251
  -0.250   0.1040   0.00824   0.00246  -0.0298   0.8552   0.8392
   0.000   0.1305   0.00820   0.00242  -0.0292   0.8440   0.8537
   0.250   0.1568   0.00817   0.00239  -0.0285   0.8306   0.8693
   0.500   0.1839   0.00815   0.00235  -0.0280   0.8152   0.8866
   0.750   0.2125   0.00813   0.00231  -0.0278   0.7993   0.9055
   1.000   0.2444   0.00812   0.00228  -0.0284   0.7827   0.9253
   1.250   0.2787   0.00812   0.00227  -0.0296   0.7654   0.9480
   1.500   0.3146   0.00813   0.00230  -0.0312   0.7496   0.9716
   1.750   0.3509   0.00816   0.00231  -0.0330   0.7328   0.9969
   2.000   0.3771   0.00825   0.00236  -0.0327   0.7155   1.0000
   2.250   0.4021   0.00835   0.00243  -0.0321   0.6953   1.0000
   2.500   0.4271   0.00847   0.00250  -0.0315   0.6706   1.0000
   2.750   0.4519   0.00862   0.00261  -0.0307   0.6386   1.0000
   3.000   0.4762   0.00883   0.00270  -0.0299   0.5961   1.0000
   3.250   0.4995   0.00915   0.00279  -0.0289   0.5335   1.0000
   3.500   0.5220   0.00962   0.00296  -0.0279   0.4598   1.0000
   3.750   0.5446   0.01017   0.00322  -0.0271   0.3864   1.0000
   4.000   0.5671   0.01080   0.00360  -0.0264   0.3113   1.0000
   4.250   0.5903   0.01141   0.00398  -0.0259   0.2562   1.0000
   4.500   0.6145   0.01193   0.00439  -0.0254   0.2152   1.0000
   4.750   0.6382   0.01253   0.00481  -0.0250   0.1625   1.0000
   5.250   0.6843   0.01400   0.00589  -0.0239   0.0738   1.0000
   5.500   0.7077   0.01473   0.00653  -0.0234   0.0484   1.0000
   5.750   0.7308   0.01553   0.00730  -0.0228   0.0275   1.0000
   6.000   0.7532   0.01653   0.00835  -0.0220   0.0179   1.0000
   6.250   0.7757   0.01752   0.00946  -0.0211   0.0132   1.0000
   6.500   0.7969   0.01868   0.01073  -0.0202   0.0100   1.0000
   6.750   0.8184   0.01988   0.01210  -0.0192   0.0089   1.0000
   7.000   0.8392   0.02128   0.01368  -0.0182   0.0081   1.0000
   7.250   0.8599   0.02282   0.01539  -0.0171   0.0074   1.0000
   7.500   0.8801   0.02452   0.01731  -0.0161   0.0070   1.0000
   7.750   0.8997   0.02649   0.01953  -0.0150   0.0068   1.0000
   8.000   0.9177   0.02875   0.02210  -0.0138   0.0066   1.0000
   8.250   0.9335   0.03135   0.02505  -0.0125   0.0065   1.0000
   8.500   0.9463   0.03435   0.02854  -0.0109   0.0064   1.0000
   8.750   0.9550   0.03768   0.03230  -0.0093   0.0064   1.0000
   9.000   0.9589   0.04139   0.03644  -0.0074   0.0064   1.0000
   9.250   0.9571   0.04540   0.04087  -0.0054   0.0064   1.0000
   9.500   0.9489   0.04958   0.04542  -0.0035   0.0065   1.0000
   9.750   0.9330   0.05348   0.04959  -0.0012   0.0065   1.0000
  10.000   0.9132   0.05785   0.05419  -0.0005   0.0066   1.0000
  10.250   0.8921   0.06320   0.05974  -0.0023   0.0067   1.0000
  10.500   0.8690   0.07046   0.06717  -0.0070   0.0068   1.0000
  10.750   0.8471   0.08051   0.07729  -0.0152   0.0069   1.0000
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