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HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/8 AIRFOIL (hq158-il)
Reynolds number: 200,000
Max Cl/Cd: 68.34 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq158-il-200000.txt
Download as CSV file: xf-hq158-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4984   0.08750   0.08408  -0.0205   1.0000   0.0385
  -8.250  -0.5024   0.08296   0.07961  -0.0251   1.0000   0.0395
  -8.000  -0.5096   0.07806   0.07477  -0.0310   1.0000   0.0399
  -7.750  -0.5128   0.07305   0.06968  -0.0367   1.0000   0.0402
  -7.500  -0.5138   0.06907   0.06553  -0.0399   1.0000   0.0406
  -7.250  -0.5117   0.06543   0.06170  -0.0413   1.0000   0.0408
  -7.000  -0.5158   0.05797   0.05418  -0.0426   1.0000   0.0417
  -6.750  -0.5069   0.05403   0.05032  -0.0421   1.0000   0.0428
  -6.500  -0.4977   0.05101   0.04728  -0.0416   1.0000   0.0442
  -6.250  -0.4882   0.04792   0.04410  -0.0411   1.0000   0.0461
  -6.000  -0.4778   0.04468   0.04065  -0.0406   1.0000   0.0489
  -5.750  -0.4696   0.04145   0.03671  -0.0395   1.0000   0.0546
  -5.500  -0.4583   0.03772   0.03307  -0.0384   1.0000   0.0564
  -5.250  -0.4459   0.03526   0.03054  -0.0369   1.0000   0.0587
  -5.000  -0.4242   0.02822   0.02225  -0.0329   1.0000   0.0293
  -4.750  -0.4117   0.02600   0.02010  -0.0322   1.0000   0.0350
  -4.500  -0.3906   0.02234   0.01566  -0.0298   1.0000   0.0303
  -4.250  -0.3698   0.02122   0.01422  -0.0281   1.0000   0.0291
  -4.000  -0.3490   0.01968   0.01245  -0.0267   1.0000   0.0288
  -3.750  -0.3268   0.01726   0.00977  -0.0255   1.0000   0.0294
  -3.500  -0.3051   0.01545   0.00789  -0.0244   1.0000   0.0316
  -3.250  -0.2690   0.01466   0.00705  -0.0262   0.9962   0.0395
  -3.000  -0.2334   0.01325   0.00566  -0.0279   0.9922   0.0566
  -2.750  -0.1979   0.01208   0.00481  -0.0300   0.9871   0.1276
  -2.500  -0.1707   0.00997   0.00470  -0.0311   0.9829   0.5923
  -2.250  -0.1389   0.00982   0.00475  -0.0316   0.9761   0.6920
  -2.000  -0.1033   0.00979   0.00476  -0.0328   0.9705   0.7481
  -1.750  -0.0707   0.00975   0.00474  -0.0334   0.9639   0.7909
  -1.500  -0.0365   0.00967   0.00466  -0.0343   0.9579   0.8264
  -1.250  -0.0032   0.00954   0.00458  -0.0347   0.9523   0.8675
  -1.000   0.0317   0.00937   0.00447  -0.0353   0.9458   0.9072
  -0.750   0.0817   0.00926   0.00434  -0.0393   0.9430   0.9392
  -0.500   0.1356   0.00916   0.00416  -0.0445   0.9399   0.9590
  -0.250   0.1851   0.00903   0.00398  -0.0490   0.9331   0.9740
   0.000   0.2364   0.00885   0.00376  -0.0538   0.9281   0.9860
   0.250   0.2797   0.00871   0.00359  -0.0572   0.9184   1.0000
   0.500   0.3029   0.00860   0.00342  -0.0564   0.9044   1.0000
   0.750   0.3255   0.00857   0.00334  -0.0554   0.8910   1.0000
   1.000   0.3489   0.00859   0.00330  -0.0544   0.8783   1.0000
   1.250   0.3730   0.00861   0.00331  -0.0535   0.8659   1.0000
   1.500   0.3974   0.00865   0.00332  -0.0526   0.8530   1.0000
   1.750   0.4221   0.00869   0.00333  -0.0517   0.8399   1.0000
   2.000   0.4468   0.00872   0.00335  -0.0507   0.8256   1.0000
   2.250   0.4715   0.00876   0.00338  -0.0498   0.8100   1.0000
   2.500   0.4966   0.00879   0.00344  -0.0488   0.7940   1.0000
   2.750   0.5216   0.00883   0.00348  -0.0479   0.7762   1.0000
   3.000   0.5463   0.00888   0.00353  -0.0469   0.7542   1.0000
   3.250   0.5709   0.00893   0.00356  -0.0458   0.7279   1.0000
   3.500   0.5951   0.00901   0.00358  -0.0446   0.6953   1.0000
   3.750   0.6188   0.00915   0.00368  -0.0434   0.6525   1.0000
   4.000   0.6417   0.00939   0.00376  -0.0421   0.5940   1.0000
   4.250   0.6628   0.00985   0.00390  -0.0405   0.5118   1.0000
   4.500   0.6830   0.01053   0.00420  -0.0391   0.4270   1.0000
   4.750   0.7040   0.01125   0.00461  -0.0380   0.3548   1.0000
   5.000   0.7246   0.01207   0.00515  -0.0370   0.2966   1.0000
   5.250   0.7473   0.01270   0.00562  -0.0363   0.2372   1.0000
   5.500   0.7672   0.01375   0.00625  -0.0353   0.1463   1.0000
   5.750   0.7853   0.01520   0.00727  -0.0340   0.0895   1.0000
   6.000   0.8062   0.01628   0.00833  -0.0329   0.0679   1.0000
   6.250   0.8270   0.01736   0.00937  -0.0320   0.0498   1.0000
   6.500   0.8480   0.01845   0.01047  -0.0310   0.0321   1.0000
   6.750   0.8658   0.02045   0.01256  -0.0291   0.0250   1.0000
   7.000   0.8866   0.02206   0.01431  -0.0278   0.0216   1.0000
   7.250   0.9070   0.02434   0.01670  -0.0266   0.0200   1.0000
   7.500   0.9274   0.02776   0.02040  -0.0254   0.0192   1.0000
   7.750   0.9470   0.03093   0.02393  -0.0240   0.0191   1.0000
   8.000   0.9647   0.03395   0.02733  -0.0225   0.0193   1.0000
   8.250   0.9782   0.03746   0.03128  -0.0207   0.0193   1.0000
   8.500   0.9847   0.04193   0.03620  -0.0187   0.0191   1.0000
   8.750   0.9818   0.04752   0.04230  -0.0164   0.0188   1.0000
  10.250   0.8964   0.07438   0.07103  -0.0088   0.0239   1.0000
  10.500   0.8732   0.08173   0.07849  -0.0138   0.0251   1.0000
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