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S6062 8% (s6062-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: S6062 8% (s6062-il)
Reynolds number: 1,000,000
Max Cl/Cd: 91.02 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s6062-il-1000000.txt
Download as CSV file: xf-s6062-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S6062 8%                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5322   0.08300   0.08144  -0.0128   1.0000   0.0086
  -8.500  -0.5377   0.07765   0.07612  -0.0160   1.0000   0.0079
  -8.250  -0.5410   0.07347   0.07197  -0.0189   1.0000   0.0088
  -8.000  -0.5505   0.06797   0.06649  -0.0244   1.0000   0.0085
  -7.750  -0.5521   0.06099   0.05946  -0.0309   1.0000   0.0087
  -7.500  -0.5507   0.05416   0.05253  -0.0353   1.0000   0.0087
  -7.250  -0.5455   0.04757   0.04576  -0.0379   1.0000   0.0092
  -7.000  -0.5328   0.04174   0.03972  -0.0387   1.0000   0.0100
  -6.750  -0.5180   0.03777   0.03549  -0.0386   1.0000   0.0103
  -6.500  -0.5079   0.03413   0.03156  -0.0377   1.0000   0.0104
  -6.250  -0.4983   0.02503   0.02176  -0.0394   0.9948   0.0111
  -6.000  -0.4691   0.02307   0.01966  -0.0411   0.9900   0.0117
  -5.750  -0.4376   0.02170   0.01816  -0.0428   0.9855   0.0126
  -5.500  -0.4075   0.02001   0.01624  -0.0438   0.9789   0.0138
  -5.250  -0.3746   0.02042   0.01653  -0.0443   0.9720   0.0154
  -4.750  -0.3302   0.01196   0.00708  -0.0414   0.9500   0.0087
  -4.500  -0.3069   0.01059   0.00550  -0.0401   0.9397   0.0080
  -4.250  -0.2830   0.00979   0.00457  -0.0391   0.9296   0.0080
  -4.000  -0.2579   0.00919   0.00388  -0.0384   0.9194   0.0084
  -3.750  -0.2322   0.00874   0.00334  -0.0379   0.9098   0.0091
  -3.500  -0.2062   0.00840   0.00290  -0.0373   0.9005   0.0099
  -3.250  -0.1795   0.00811   0.00255  -0.0370   0.8904   0.0103
  -3.000  -0.1528   0.00778   0.00212  -0.0366   0.8807   0.0110
  -2.750  -0.1259   0.00753   0.00179  -0.0362   0.8710   0.0132
  -2.500  -0.0990   0.00723   0.00150  -0.0359   0.8606   0.0305
  -2.250  -0.0730   0.00665   0.00128  -0.0358   0.8501   0.1280
  -2.000  -0.0465   0.00628   0.00114  -0.0357   0.8397   0.2079
  -1.750  -0.0202   0.00588   0.00103  -0.0355   0.8289   0.3061
  -1.500   0.0057   0.00541   0.00093  -0.0354   0.8175   0.4325
  -1.250   0.0315   0.00495   0.00086  -0.0351   0.8058   0.5645
  -1.000   0.0575   0.00463   0.00083  -0.0348   0.7937   0.6668
  -0.750   0.0832   0.00440   0.00082  -0.0342   0.7813   0.7504
  -0.500   0.1095   0.00431   0.00081  -0.0337   0.7681   0.7971
  -0.250   0.1357   0.00425   0.00081  -0.0332   0.7541   0.8357
   0.000   0.1615   0.00422   0.00081  -0.0325   0.7392   0.8699
   0.250   0.1871   0.00421   0.00082  -0.0318   0.7239   0.8979
   0.500   0.2124   0.00421   0.00082  -0.0310   0.7079   0.9226
   0.750   0.2382   0.00425   0.00083  -0.0304   0.6913   0.9458
   1.000   0.2677   0.00429   0.00084  -0.0305   0.6737   0.9661
   1.250   0.3019   0.00436   0.00084  -0.0319   0.6540   0.9810
   1.500   0.3386   0.00446   0.00086  -0.0338   0.6333   0.9905
   1.750   0.3752   0.00455   0.00088  -0.0358   0.6101   0.9977
   2.000   0.4047   0.00467   0.00091  -0.0362   0.5864   1.0000
   2.250   0.4298   0.00482   0.00096  -0.0357   0.5612   1.0000
   2.500   0.4551   0.00500   0.00103  -0.0352   0.5291   1.0000
   2.750   0.4798   0.00530   0.00111  -0.0347   0.4742   1.0000
   3.000   0.5054   0.00556   0.00122  -0.0344   0.4333   1.0000
   3.250   0.5308   0.00589   0.00135  -0.0340   0.3832   1.0000
   3.500   0.5558   0.00631   0.00152  -0.0337   0.3216   1.0000
   3.750   0.5808   0.00676   0.00172  -0.0334   0.2603   1.0000
   4.000   0.6052   0.00734   0.00198  -0.0331   0.1894   1.0000
   4.250   0.6297   0.00791   0.00226  -0.0327   0.1255   1.0000
   4.500   0.6528   0.00875   0.00270  -0.0322   0.0458   1.0000
   4.750   0.6779   0.00928   0.00308  -0.0319   0.0162   1.0000
   5.000   0.7043   0.00957   0.00340  -0.0316   0.0145   1.0000
   5.250   0.7303   0.00993   0.00379  -0.0313   0.0130   1.0000
   5.500   0.7561   0.01034   0.00427  -0.0309   0.0121   1.0000
   5.750   0.7809   0.01094   0.00497  -0.0304   0.0112   1.0000
   6.000   0.8047   0.01170   0.00585  -0.0298   0.0108   1.0000
   6.250   0.8291   0.01229   0.00651  -0.0293   0.0107   1.0000
   6.500   0.8536   0.01285   0.00712  -0.0288   0.0106   1.0000
   6.750   0.8775   0.01350   0.00786  -0.0282   0.0104   1.0000
   7.000   0.9011   0.01420   0.00863  -0.0276   0.0101   1.0000
   7.250   0.9241   0.01501   0.00952  -0.0269   0.0098   1.0000
   7.500   0.9463   0.01598   0.01058  -0.0261   0.0095   1.0000
   7.750   0.9678   0.01715   0.01185  -0.0251   0.0093   1.0000
   8.000   0.9896   0.01825   0.01305  -0.0243   0.0090   1.0000
   8.250   1.0113   0.01935   0.01425  -0.0235   0.0086   1.0000
   8.500   1.0317   0.02089   0.01595  -0.0225   0.0085   1.0000
   8.750   1.0515   0.02251   0.01773  -0.0215   0.0084   1.0000
   9.000   1.0699   0.02441   0.01983  -0.0204   0.0083   1.0000
   9.250   1.0861   0.02664   0.02229  -0.0190   0.0083   1.0000
   9.500   1.1000   0.02902   0.02493  -0.0176   0.0082   1.0000
   9.750   1.0960   0.03524   0.03176  -0.0141   0.0090   1.0000
  10.000   1.1007   0.03807   0.03484  -0.0121   0.0086   1.0000
  10.250   1.1351   0.03475   0.03112  -0.0134   0.0074   1.0000
  10.500   1.1360   0.03776   0.03436  -0.0112   0.0073   1.0000
  10.750   1.0680   0.03017   0.02718  -0.0047   0.0072   1.0000
  11.000   1.0494   0.03324   0.03042  -0.0021   0.0072   1.0000
  11.250   1.0313   0.03666   0.03403  -0.0007   0.0073   1.0000
  11.500   1.0117   0.04104   0.03857  -0.0006   0.0072   1.0000
  11.750   0.9917   0.04637   0.04407  -0.0018   0.0073   1.0000
  12.000   0.9696   0.05288   0.05071  -0.0041   0.0072   1.0000
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