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

FX 63-143 AIRFOIL (fx63143-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 63-143 AIRFOIL (fx63143-il)
Reynolds number: 50,000
Max Cl/Cd: 10.7 at α=-1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx63143-il-50000.txt
Download as CSV file: xf-fx63143-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 63-143 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.3577   0.12694   0.12027  -0.0265   1.0000   0.2257
 -10.500  -0.3368   0.12188   0.11520  -0.0256   1.0000   0.2348
 -10.250  -0.3667   0.12210   0.11559  -0.0277   1.0000   0.2434
 -10.000  -0.3330   0.11614   0.10962  -0.0259   1.0000   0.2557
  -9.750  -0.3289   0.11278   0.10634  -0.0259   1.0000   0.2644
  -9.500  -0.3409   0.11092   0.10462  -0.0263   1.0000   0.2762
  -9.250  -0.3435   0.10872   0.10251  -0.0259   1.0000   0.2902
  -9.000  -0.3240   0.10454   0.09837  -0.0247   1.0000   0.3039
  -8.750  -0.3153   0.10120   0.09510  -0.0237   1.0000   0.3181
  -8.500  -0.3099   0.09798   0.09198  -0.0230   1.0000   0.3302
  -8.250  -0.3090   0.09522   0.08933  -0.0220   1.0000   0.3438
  -8.000  -0.3082   0.09252   0.08675  -0.0208   1.0000   0.3589
  -5.750  -0.5570   0.06097   0.05518  -0.0277   0.9828   0.1402
  -5.500  -0.5329   0.05442   0.04752  -0.0305   0.9727   0.1100
  -5.250  -0.5010   0.05011   0.04247  -0.0327   0.9632   0.0987
  -5.000  -0.4677   0.04642   0.03776  -0.0343   0.9535   0.0910
  -4.750  -0.4367   0.04358   0.03462  -0.0355   0.9442   0.0894
  -4.500  -0.3963   0.04088   0.03140  -0.0378   0.9361   0.0887
  -4.250  -0.3671   0.03906   0.02911  -0.0379   0.9272   0.0898
  -4.000  -0.3260   0.03711   0.02699  -0.0400   0.9198   0.0954
  -3.750  -0.2951   0.03584   0.02559  -0.0404   0.9115   0.1058
  -3.500  -0.2552   0.03452   0.02422  -0.0418   0.9037   0.1240
  -3.250  -0.2234   0.03362   0.02335  -0.0424   0.8959   0.1448
  -3.000  -0.1990   0.03258   0.02250  -0.0423   0.8880   0.1717
  -2.750   0.2033   0.03041   0.02213  -0.0912   0.9008   1.0000
  -2.500   0.2138   0.03068   0.02222  -0.0906   0.8887   1.0000
  -2.250   0.2398   0.03069   0.02202  -0.0923   0.8792   1.0000
  -2.000   0.2628   0.03079   0.02192  -0.0932   0.8689   1.0000
  -1.750   0.2727   0.03127   0.02223  -0.0918   0.8576   1.0000
  -1.500   0.3105   0.03113   0.02186  -0.0946   0.8492   1.0000
  -1.250   0.3025   0.03214   0.02280  -0.0902   0.8377   1.0000
  -1.000   0.3430   0.03206   0.02252  -0.0931   0.8308   1.0000
  -0.750   0.3148   0.03368   0.02412  -0.0858   0.8201   1.0000
  -0.500   0.3454   0.03400   0.02428  -0.0872   0.8141   1.0000
  -0.250   0.3059   0.03587   0.02616  -0.0783   0.8050   1.0000
   0.000   0.3283   0.03645   0.02659  -0.0784   0.7983   1.0000
   0.250   0.2966   0.03803   0.02814  -0.0704   0.7908   1.0000
   0.500   0.2848   0.03921   0.02926  -0.0655   0.7842   1.0000
   0.750   0.3215   0.03981   0.02973  -0.0676   0.7790   1.0000
   1.000   0.2714   0.04161   0.03152  -0.0575   0.7746   1.0000
   1.250   0.2595   0.04276   0.03260  -0.0528   0.7686   1.0000
   1.500   0.2960   0.04344   0.03316  -0.0545   0.7607   1.0000
   1.750   0.2661   0.04471   0.03439  -0.0473   0.7541   1.0000
   2.000   0.2830   0.04562   0.03522  -0.0464   0.7472   1.0000
   2.250   0.2876   0.04682   0.03635  -0.0441   0.7425   1.0000
   2.500   0.2713   0.04817   0.03766  -0.0394   0.7402   1.0000
   2.750   0.2634   0.04954   0.03897  -0.0360   0.7391   1.0000
   3.000   0.2600   0.05097   0.04035  -0.0334   0.7394   1.0000
   3.250   0.2598   0.05245   0.04179  -0.0314   0.7401   1.0000
   3.500   0.2629   0.05401   0.04330  -0.0299   0.7411   1.0000
   3.750   0.2731   0.05579   0.04504  -0.0296   0.7433   1.0000
   4.000   0.1261   0.05684   0.04625  -0.0152   0.9026   1.0000
   4.250   0.1498   0.05850   0.04786  -0.0166   0.8862   1.0000
   4.500   0.1758   0.06041   0.04971  -0.0182   0.8687   1.0000
   4.750   0.1958   0.06196   0.05123  -0.0189   0.8531   1.0000
   5.000   0.2142   0.06362   0.05286  -0.0194   0.8397   1.0000
   5.250   0.2449   0.06648   0.05569  -0.0219   0.8296   1.0000
   5.500   0.2633   0.06796   0.05717  -0.0223   0.8153   1.0000
   5.750   0.2726   0.06911   0.05831  -0.0216   0.8030   1.0000
   6.000   0.3013   0.07212   0.06132  -0.0238   0.7948   1.0000
   6.250   0.3177   0.07361   0.06283  -0.0240   0.7807   1.0000
   6.500   0.3296   0.07502   0.06425  -0.0236   0.7662   1.0000
   6.750   0.3448   0.07684   0.06609  -0.0238   0.7523   1.0000
   7.000   0.3663   0.07933   0.06863  -0.0248   0.7411   1.0000
   7.250   0.3995   0.08256   0.07191  -0.0271   0.7276   1.0000
   7.500   0.4037   0.08343   0.07281  -0.0259   0.7132   1.0000
   7.750   0.4116   0.08510   0.07452  -0.0255   0.7010   1.0000
   8.000   0.4382   0.08833   0.07784  -0.0272   0.6917   1.0000
   8.250   0.4525   0.09016   0.07975  -0.0273   0.6782   1.0000
   8.500   0.4560   0.09174   0.08137  -0.0266   0.6662   1.0000
   8.750   0.4811   0.09514   0.08485  -0.0282   0.6573   1.0000
   9.000   0.4973   0.09729   0.08711  -0.0286   0.6434   1.0000
   9.250   0.4992   0.09882   0.08871  -0.0279   0.6304   1.0000
   9.500   0.5111   0.10132   0.09130  -0.0282   0.6188   1.0000
   9.750   0.5424   0.10523   0.09535  -0.0301   0.6071   1.0000
  10.000   0.5545   0.10726   0.09748  -0.0301   0.5921   1.0000
  10.250   0.5556   0.10886   0.09916  -0.0296   0.5782   1.0000
  10.500   0.5644   0.11116   0.10156  -0.0297   0.5641   1.0000
<< Back to FX 63-143 AIRFOIL (fx63143-il)

Polar data table (+)

Polar graphs


<< Back to FX 63-143 AIRFOIL (fx63143-il)