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

GOE 703 AIRFOIL (goe703-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 703 AIRFOIL (goe703-il)
Reynolds number: 50,000
Max Cl/Cd: 22.31 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe703-il-50000-n5.txt
Download as CSV file: xf-goe703-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 703 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.4536   0.10032   0.09188  -0.0640   1.0000   0.0880
 -12.750  -0.5405   0.08249   0.07388  -0.0733   1.0000   0.0873
 -12.500  -0.6009   0.07336   0.06454  -0.0757   1.0000   0.0866
 -12.250  -0.6453   0.06773   0.05872  -0.0746   1.0000   0.0865
 -12.000  -0.6802   0.06378   0.05459  -0.0718   1.0000   0.0867
 -11.750  -0.7082   0.06077   0.05142  -0.0680   1.0000   0.0869
 -11.500  -0.7346   0.05832   0.04879  -0.0634   1.0000   0.0873
 -11.250  -0.7588   0.05632   0.04663  -0.0581   1.0000   0.0878
 -11.000  -0.7818   0.05469   0.04484  -0.0524   1.0000   0.0882
 -10.750  -0.8039   0.05333   0.04331  -0.0462   1.0000   0.0887
 -10.500  -0.8201   0.05227   0.04215  -0.0406   1.0000   0.0892
 -10.250  -0.8303   0.05146   0.04133  -0.0356   1.0000   0.0899
 -10.000  -0.8370   0.05058   0.04041  -0.0312   0.9992   0.0907
  -9.750  -0.8094   0.04883   0.03856  -0.0330   0.9905   0.0932
  -9.500  -0.7854   0.04695   0.03639  -0.0342   0.9814   0.0966
  -9.250  -0.7592   0.04539   0.03471  -0.0353   0.9720   0.1000
  -9.000  -0.7295   0.04405   0.03330  -0.0370   0.9634   0.1038
  -8.750  -0.6976   0.04240   0.03142  -0.0387   0.9547   0.1090
  -8.500  -0.6689   0.04132   0.03038  -0.0399   0.9451   0.1139
  -8.250  -0.6336   0.04001   0.02896  -0.0421   0.9374   0.1215
  -7.750  -0.5710   0.03790   0.02682  -0.0449   0.9198   0.1408
  -7.500  -0.5444   0.03702   0.02596  -0.0454   0.9096   0.1530
  -7.250  -0.5111   0.03607   0.02501  -0.0470   0.9020   0.1696
  -7.000  -0.4841   0.03532   0.02425  -0.0475   0.8927   0.1866
  -6.750  -0.4545   0.03458   0.02349  -0.0484   0.8839   0.2059
  -6.500  -0.4145   0.03372   0.02261  -0.0511   0.8786   0.2289
  -6.250  -0.3994   0.03339   0.02223  -0.0492   0.8658   0.2458
  -6.000  -0.3637   0.03261   0.02154  -0.0511   0.8595   0.2695
  -5.750  -0.3459   0.03225   0.02120  -0.0496   0.8482   0.2896
  -5.500  -0.3107   0.03152   0.02056  -0.0512   0.8410   0.3147
  -5.250  -0.2732   0.03078   0.01989  -0.0532   0.8344   0.3408
  -5.000  -0.2515   0.03041   0.01961  -0.0523   0.8237   0.3633
  -4.750  -0.2144   0.02963   0.01902  -0.0540   0.8177   0.3957
  -4.500  -0.1960   0.02932   0.01891  -0.0524   0.8075   0.4284
  -4.250  -0.1649   0.02882   0.01875  -0.0527   0.8000   0.4800
  -4.000  -0.0997   0.02865   0.01909  -0.0579   0.7959   0.5713
  -3.750  -0.0397   0.02968   0.02031  -0.0614   0.7882   0.6409
  -3.500   0.0031   0.03050   0.02105  -0.0624   0.7802   0.6857
  -3.250   0.0435   0.03102   0.02138  -0.0633   0.7745   0.7215
  -3.000   0.0583   0.03184   0.02212  -0.0603   0.7638   0.7488
  -2.750   0.0910   0.03232   0.02244  -0.0602   0.7568   0.7743
  -2.500   0.1248   0.03297   0.02295  -0.0604   0.7491   0.7951
  -2.250   0.1551   0.03358   0.02344  -0.0603   0.7407   0.8144
  -2.000   0.1978   0.03383   0.02352  -0.0623   0.7349   0.8324
  -1.750   0.2206   0.03444   0.02405  -0.0613   0.7257   0.8493
  -1.500   0.2537   0.03472   0.02421  -0.0620   0.7186   0.8654
  -1.250   0.3083   0.03472   0.02403  -0.0664   0.7132   0.8792
  -1.000   0.3414   0.03528   0.02456  -0.0674   0.7036   0.8965
  -0.750   0.3953   0.03522   0.02438  -0.0719   0.6974   0.9137
  -0.500   0.4388   0.03514   0.02422  -0.0751   0.6900   0.9274
  -0.250   0.4679   0.03521   0.02424  -0.0759   0.6817   0.9383
   0.000   0.5070   0.03481   0.02372  -0.0782   0.6758   0.9459
   0.250   0.5297   0.03499   0.02389  -0.0781   0.6671   0.9543
   0.500   0.5613   0.03484   0.02368  -0.0794   0.6598   0.9616
   0.750   0.5989   0.03443   0.02316  -0.0814   0.6545   0.9679
   1.000   0.6193   0.03475   0.02352  -0.0813   0.6449   0.9756
   1.250   0.6523   0.03453   0.02325  -0.0827   0.6383   0.9817
   1.500   0.6876   0.03422   0.02287  -0.0846   0.6322   0.9873
   1.750   0.7093   0.03454   0.02324  -0.0847   0.6230   0.9940
   2.000   0.7454   0.03420   0.02285  -0.0867   0.6168   0.9991
   2.250   0.7615   0.03449   0.02313  -0.0852   0.6102   1.0000
   2.500   0.7653   0.03520   0.02389  -0.0818   0.6017   1.0000
   2.750   0.7858   0.03523   0.02387  -0.0807   0.5960   1.0000
   3.000   0.7944   0.03580   0.02446  -0.0779   0.5890   1.0000
   3.250   0.7958   0.03664   0.02535  -0.0741   0.5807   1.0000
   3.500   0.8175   0.03664   0.02532  -0.0731   0.5753   1.0000
   3.750   0.8192   0.03751   0.02623  -0.0694   0.5678   1.0000
   4.000   0.8187   0.03843   0.02720  -0.0653   0.5598   1.0000
   4.250   0.8423   0.03837   0.02711  -0.0645   0.5547   1.0000
   4.500   0.8307   0.03976   0.02857  -0.0590   0.5462   1.0000
   4.750   0.8315   0.04060   0.02944  -0.0551   0.5387   1.0000
   5.000   0.8597   0.04036   0.02918  -0.0548   0.5342   1.0000
   5.250   0.8201   0.04282   0.03173  -0.0457   0.5239   1.0000
   5.500   0.8306   0.04323   0.03215  -0.0431   0.5177   1.0000
   5.750   0.8678   0.04263   0.03153  -0.0439   0.5139   1.0000
   6.000   0.7779   0.04679   0.03576  -0.0287   0.5007   1.0000
   6.250   0.8090   0.04632   0.03529  -0.0285   0.4966   1.0000
   6.750   0.7582   0.05056   0.03955  -0.0162   0.4785   1.0000
   7.000   0.7773   0.05053   0.03952  -0.0145   0.4739   1.0000
   7.250   0.7270   0.05481   0.04384  -0.0078   0.4603   1.0000
   7.500   0.7401   0.05506   0.04409  -0.0056   0.4547   1.0000
   8.000   0.7233   0.05869   0.04773   0.0009   0.4360   1.0000
   8.500   0.7178   0.06158   0.05065   0.0064   0.4174   1.0000
   9.000   0.7437   0.06165   0.05074   0.0108   0.4036   1.0000
   9.500   0.7425   0.06450   0.05362   0.0155   0.3859   1.0000
  10.000   0.7421   0.06752   0.05668   0.0197   0.3684   1.0000
  10.500   0.7409   0.07095   0.06016   0.0234   0.3510   1.0000
  11.000   0.7412   0.07444   0.06370   0.0267   0.3337   1.0000
  11.500   0.7401   0.07839   0.06770   0.0296   0.3165   1.0000
  12.000   0.7382   0.08270   0.07207   0.0321   0.2994   1.0000
  12.250   0.7128   0.08844   0.07783   0.0324   0.2866   1.0000
  12.500   0.7313   0.08813   0.07754   0.0339   0.2826   1.0000
  12.750   0.7572   0.08687   0.07632   0.0356   0.2804   1.0000
  13.250   0.7186   0.09767   0.08717   0.0354   0.2609   1.0000
  13.750   0.7149   0.10370   0.09326   0.0361   0.2483   1.0000
<< Back to GOE 703 AIRFOIL (goe703-il)

Polar data table (+)

Polar graphs


<< Back to GOE 703 AIRFOIL (goe703-il)