Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 140 (MVA H.17) AIRFOIL (goe140-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 140 (MVA H.17) AIRFOIL (goe140-il)
Reynolds number: 200,000
Max Cl/Cd: 74.78 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe140-il-200000.txt
Download as CSV file: xf-goe140-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 140 (MVA H.17) 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  -0.3034   0.10552   0.10219  -0.0217   1.0000   0.0328
  -8.250  -0.3048   0.10394   0.10068  -0.0248   1.0000   0.0330
  -8.000  -0.3037   0.10182   0.09862  -0.0284   1.0000   0.0331
  -7.750  -0.2980   0.09873   0.09559  -0.0324   1.0000   0.0332
  -7.500  -0.2889   0.09231   0.08916  -0.0243   1.0000   0.0342
  -7.250  -0.2834   0.08943   0.08633  -0.0237   1.0000   0.0349
  -7.000  -0.2809   0.08698   0.08393  -0.0237   1.0000   0.0355
  -6.750  -0.2808   0.08477   0.08180  -0.0236   1.0000   0.0362
  -6.500  -0.2856   0.08301   0.08011  -0.0227   1.0000   0.0368
  -6.000  -0.2424   0.07571   0.07279  -0.0353   0.9913   0.0404
  -5.750  -0.1762   0.06903   0.06587  -0.0577   0.9830   0.0424
  -5.500  -0.1596   0.06456   0.06146  -0.0569   0.9761   0.0434
  -5.250  -0.1254   0.06048   0.05735  -0.0619   0.9708   0.0451
  -5.000  -0.0881   0.05650   0.05328  -0.0684   0.9606   0.0477
  -4.750  -0.0143   0.05410   0.05023  -0.0835   0.9489   0.0523
  -4.500   0.0028   0.04770   0.04402  -0.0849   0.9379   0.0536
  -4.250   0.0271   0.04472   0.04102  -0.0862   0.9228   0.0552
  -4.000   0.0553   0.04207   0.03825  -0.0881   0.9054   0.0577
  -3.750   0.1025   0.04151   0.03708  -0.0920   0.8869   0.0638
  -3.500   0.1272   0.03734   0.03271  -0.0933   0.8689   0.0648
  -3.250   0.1469   0.03441   0.02978  -0.0933   0.8502   0.0662
  -3.000   0.1702   0.03252   0.02778  -0.0933   0.8296   0.0690
  -2.750   0.2065   0.03280   0.02746  -0.0936   0.8094   0.0762
  -2.500   0.2324   0.02944   0.02388  -0.0945   0.7909   0.0777
  -2.250   0.2567   0.02725   0.02162  -0.0948   0.7703   0.0797
  -2.000   0.2837   0.02585   0.02001  -0.0949   0.7494   0.0835
  -1.750   0.3154   0.02495   0.01861  -0.0951   0.7293   0.0911
  -1.500   0.3413   0.02320   0.01679  -0.0953   0.7073   0.0945
  -1.250   0.3703   0.02237   0.01558  -0.0954   0.6861   0.1061
  -1.000   0.3974   0.02129   0.01436  -0.0955   0.6643   0.1134
  -0.750   0.4254   0.02029   0.01310  -0.0957   0.6430   0.1250
  -0.500   0.4532   0.01951   0.01208  -0.0958   0.6228   0.1413
   0.000   0.5183   0.01753   0.00909  -0.0941   0.5858   0.0748
   0.250   0.5466   0.01664   0.00799  -0.0940   0.5684   0.0720
   0.500   0.5747   0.01629   0.00741  -0.0937   0.5518   0.0741
   0.750   0.6026   0.01580   0.00674  -0.0935   0.5361   0.0737
   1.000   0.6304   0.01540   0.00620  -0.0933   0.5213   0.0739
   1.250   0.6581   0.01514   0.00582  -0.0930   0.5072   0.0754
   1.500   0.6857   0.01475   0.00540  -0.0929   0.4943   0.0787
   1.750   0.7134   0.01457   0.00520  -0.0929   0.4825   0.0857
   2.000   0.7411   0.01442   0.00502  -0.0929   0.4718   0.1011
   2.250   0.7693   0.01425   0.00484  -0.0930   0.4613   0.1365
   2.500   0.7901   0.01255   0.00489  -0.0916   0.4521   1.0000
   2.750   0.8175   0.01287   0.00495  -0.0914   0.4434   1.0000
   3.000   0.8450   0.01311   0.00510  -0.0913   0.4339   1.0000
   3.250   0.8723   0.01343   0.00530  -0.0913   0.4258   1.0000
   3.500   0.8995   0.01370   0.00549  -0.0912   0.4175   1.0000
   3.750   0.9267   0.01403   0.00574  -0.0911   0.4101   1.0000
   4.000   0.9538   0.01431   0.00596  -0.0910   0.4026   1.0000
   4.250   0.9808   0.01465   0.00626  -0.0910   0.3957   1.0000
   4.500   1.0078   0.01493   0.00651  -0.0909   0.3886   1.0000
   4.750   1.0347   0.01531   0.00682  -0.0908   0.3829   1.0000
   5.000   1.0616   0.01559   0.00717  -0.0907   0.3767   1.0000
   5.250   1.0884   0.01593   0.00746  -0.0906   0.3713   1.0000
   5.500   1.1146   0.01620   0.00777  -0.0905   0.3638   1.0000
   5.750   1.1404   0.01642   0.00792  -0.0902   0.3553   1.0000
   6.000   1.1660   0.01659   0.00820  -0.0899   0.3465   1.0000
   6.250   1.1913   0.01682   0.00839  -0.0896   0.3381   1.0000
   6.500   1.2167   0.01697   0.00863  -0.0893   0.3294   1.0000
   6.750   1.2415   0.01714   0.00883  -0.0889   0.3193   1.0000
   7.000   1.2660   0.01735   0.00904  -0.0885   0.3097   1.0000
   7.250   1.2911   0.01751   0.00933  -0.0881   0.3001   1.0000
   7.500   1.3151   0.01767   0.00953  -0.0877   0.2860   1.0000
   7.750   1.3386   0.01790   0.00975  -0.0871   0.2691   1.0000
   8.000   1.3618   0.01823   0.01010  -0.0865   0.2522   1.0000
   8.250   1.3853   0.01861   0.01051  -0.0860   0.2329   1.0000
   8.500   1.4065   0.01921   0.01102  -0.0852   0.2031   1.0000
   8.750   1.4134   0.02132   0.01248  -0.0829   0.1123   1.0000
   9.000   1.4241   0.02303   0.01399  -0.0809   0.0685   1.0000
   9.250   1.4268   0.02527   0.01594  -0.0780   0.0290   1.0000
   9.500   1.4391   0.02646   0.01722  -0.0760   0.0263   1.0000
   9.750   1.4500   0.02764   0.01855  -0.0739   0.0249   1.0000
  10.000   1.4575   0.02887   0.01995  -0.0713   0.0243   1.0000
  10.250   1.4615   0.03025   0.02150  -0.0684   0.0239   1.0000
  10.500   1.4637   0.03192   0.02334  -0.0658   0.0235   1.0000
  10.750   1.4638   0.03392   0.02553  -0.0636   0.0233   1.0000
  11.000   1.4619   0.03632   0.02811  -0.0618   0.0231   1.0000
  11.250   1.4581   0.03921   0.03119  -0.0607   0.0229   1.0000
  11.500   1.4527   0.04256   0.03472  -0.0601   0.0227   1.0000
  11.750   1.4457   0.04640   0.03872  -0.0601   0.0225   1.0000
  12.000   1.4375   0.05062   0.04310  -0.0604   0.0222   1.0000
  12.250   1.4284   0.05513   0.04776  -0.0611   0.0220   1.0000
  12.500   1.4188   0.05979   0.05254  -0.0618   0.0218   1.0000
  12.750   1.4092   0.06447   0.05734  -0.0626   0.0215   1.0000
  13.000   1.3998   0.06907   0.06205  -0.0633   0.0213   1.0000
  13.250   1.3919   0.07343   0.06650  -0.0638   0.0211   1.0000
  13.500   1.3863   0.07740   0.07055  -0.0640   0.0210   1.0000
  13.750   1.3832   0.08093   0.07416  -0.0640   0.0210   1.0000
  14.000   1.3830   0.08400   0.07729  -0.0637   0.0209   1.0000
  14.250   1.3856   0.08655   0.07990  -0.0630   0.0209   1.0000
  14.500   1.3908   0.08865   0.08209  -0.0620   0.0210   1.0000
  14.750   1.3987   0.09030   0.08381  -0.0605   0.0211   1.0000
  15.000   1.4088   0.09163   0.08522  -0.0587   0.0214   1.0000
  15.250   1.4166   0.09357   0.08732  -0.0573   0.0218   1.0000
  15.500   1.4239   0.09571   0.08962  -0.0559   0.0223   1.0000
  15.750   1.4302   0.09813   0.09222  -0.0546   0.0230   1.0000
  16.000   1.4355   0.10086   0.09513  -0.0532   0.0237   1.0000
  16.250   1.4540   0.10250   0.09691  -0.0490   0.0246   1.0000
  16.500   1.4492   0.10666   0.10125  -0.0507   0.0248   1.0000
  16.750   1.4419   0.11136   0.10614  -0.0531   0.0250   1.0000
  17.000   1.4314   0.11682   0.11182  -0.0561   0.0253   1.0000
  17.250   1.4173   0.12330   0.11856  -0.0599   0.0257   1.0000
  17.500   1.3957   0.13181   0.12743  -0.0652   0.0265   1.0000
  17.750   1.3680   0.14213   0.13809  -0.0721   0.0273   1.0000
  18.000   1.3443   0.15203   0.14826  -0.0788   0.0280   1.0000
  18.250   1.3213   0.16228   0.15871  -0.0858   0.0286   1.0000
  18.500   1.2982   0.17323   0.16984  -0.0935   0.0292   1.0000
  18.750   1.2740   0.18532   0.18208  -0.1020   0.0299   1.0000
<< Back to GOE 140 (MVA H.17) AIRFOIL (goe140-il)

Polar data table (+)

Polar graphs


<< Back to GOE 140 (MVA H.17) AIRFOIL (goe140-il)