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

GOE 738 AIRFOIL (goe738-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 738 AIRFOIL (goe738-il)
Reynolds number: 50,000
Max Cl/Cd: 10.11 at α=-1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe738-il-50000.txt
Download as CSV file: xf-goe738-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 738 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3810   0.11120   0.10437   0.0024   1.0000   0.3556
  -9.250  -0.3413   0.10528   0.09843   0.0019   1.0000   0.3633
  -9.000  -0.3594   0.10387   0.09715   0.0009   1.0000   0.3754
  -8.750  -0.3265   0.09944   0.09271   0.0005   1.0000   0.3884
  -8.500  -0.5882   0.06774   0.06097  -0.0326   1.0000   0.1926
  -8.250  -0.5974   0.06452   0.05776  -0.0312   1.0000   0.1917
  -8.000  -0.6206   0.06272   0.05599  -0.0268   1.0000   0.1908
  -7.750  -0.6452   0.06088   0.05411  -0.0223   1.0000   0.1898
  -7.500  -0.6674   0.05860   0.05172  -0.0182   1.0000   0.1889
  -7.250  -0.6856   0.05602   0.04896  -0.0146   1.0000   0.1882
  -7.000  -0.6994   0.05332   0.04601  -0.0112   1.0000   0.1878
  -6.750  -0.7090   0.05066   0.04304  -0.0081   1.0000   0.1883
  -6.500  -0.7151   0.04814   0.04014  -0.0052   1.0000   0.1899
  -6.250  -0.6959   0.04488   0.03632  -0.0066   0.9917   0.1939
  -6.000  -0.6576   0.04292   0.03438  -0.0098   0.9816   0.2007
  -5.750  -0.6321   0.04070   0.03157  -0.0111   0.9712   0.2085
  -5.500  -0.5987   0.03901   0.02985  -0.0132   0.9615   0.2174
  -5.250  -0.5633   0.03732   0.02777  -0.0156   0.9522   0.2297
  -5.000  -0.5365   0.03625   0.02660  -0.0162   0.9418   0.2423
  -4.750  -0.5035   0.03517   0.02547  -0.0178   0.9326   0.2585
  -4.500  -0.4715   0.03420   0.02441  -0.0191   0.9232   0.2788
  -4.250  -0.4455   0.03339   0.02358  -0.0193   0.9139   0.3016
  -4.000  -0.4090   0.03262   0.02296  -0.0210   0.9053   0.3353
  -3.750  -0.3884   0.03212   0.02268  -0.0202   0.8962   0.3660
  -3.500  -0.3521   0.03160   0.02234  -0.0217   0.8878   0.4133
  -3.250  -0.3376   0.03140   0.02238  -0.0197   0.8791   0.4516
  -3.000  -0.3048   0.03116   0.02257  -0.0202   0.8712   0.5141
  -2.750  -0.2927   0.03116   0.02294  -0.0172   0.8631   0.5762
  -2.500  -0.2669   0.03134   0.02368  -0.0150   0.8553   0.6720
  -2.250  -0.2278   0.03234   0.02504  -0.0136   0.8479   0.7775
  -2.000  -0.0349   0.03534   0.02761  -0.0348   0.8381   0.8785
  -1.750   0.1345   0.03606   0.02788  -0.0565   0.8286   0.9296
  -1.500   0.2716   0.03519   0.02670  -0.0752   0.8193   0.9754
  -1.250   0.3433   0.03457   0.02594  -0.0854   0.8092   1.0000
  -1.000   0.3537   0.03499   0.02626  -0.0844   0.8006   1.0000
  -0.750   0.3487   0.03603   0.02726  -0.0821   0.7913   1.0000
  -0.500   0.3585   0.03658   0.02773  -0.0807   0.7830   1.0000
  -0.250   0.3485   0.03777   0.02888  -0.0774   0.7739   1.0000
   0.000   0.3555   0.03847   0.02951  -0.0755   0.7655   1.0000
   0.250   0.3422   0.03971   0.03070  -0.0715   0.7566   1.0000
   0.500   0.3501   0.04039   0.03131  -0.0694   0.7473   1.0000
   0.750   0.3325   0.04163   0.03250  -0.0645   0.7380   1.0000
   1.000   0.3543   0.04206   0.03285  -0.0635   0.7276   1.0000
   1.250   0.3224   0.04350   0.03422  -0.0568   0.7180   1.0000
   1.500   0.3651   0.04367   0.03431  -0.0578   0.7069   1.0000
   1.750   0.3178   0.04535   0.03592  -0.0495   0.6974   1.0000
   2.000   0.3479   0.04594   0.03644  -0.0493   0.6863   1.0000
   2.250   0.3236   0.04725   0.03765  -0.0433   0.6764   1.0000
   2.500   0.3369   0.04822   0.03855  -0.0415   0.6656   1.0000
   2.750   0.3546   0.04906   0.03933  -0.0401   0.6541   1.0000
   3.000   0.3373   0.05056   0.04074  -0.0352   0.6437   1.0000
   3.250   0.3860   0.05107   0.04123  -0.0369   0.6319   1.0000
   3.500   0.3566   0.05287   0.04295  -0.0313   0.6206   1.0000
   3.750   0.3693   0.05412   0.04414  -0.0296   0.6093   1.0000
   4.000   0.3909   0.05518   0.04518  -0.0288   0.5984   1.0000
   4.250   0.3768   0.05725   0.04718  -0.0255   0.5881   1.0000
   4.500   0.4200   0.05783   0.04775  -0.0261   0.5770   1.0000
   4.750   0.3924   0.06044   0.05030  -0.0225   0.5673   1.0000
   5.000   0.4212   0.06160   0.05144  -0.0222   0.5579   1.0000
   5.250   0.4004   0.06456   0.05435  -0.0200   0.5522   1.0000
   5.500   0.4270   0.06590   0.05569  -0.0198   0.5437   1.0000
   5.750   0.4190   0.06857   0.05831  -0.0185   0.5379   1.0000
   6.000   0.4114   0.07140   0.06112  -0.0176   0.5349   1.0000
   6.250   0.4092   0.07443   0.06415  -0.0174   0.5359   1.0000
   6.500   0.4111   0.07746   0.06718  -0.0174   0.5379   1.0000
   6.750   0.4236   0.08062   0.07036  -0.0181   0.5408   1.0000
   7.000   0.3464   0.08692   0.07670  -0.0198   0.6202   1.0000
   7.250   0.3723   0.09047   0.08026  -0.0210   0.6133   1.0000
   7.500   0.3722   0.09145   0.08124  -0.0196   0.6005   1.0000
   7.750   0.3870   0.09440   0.08420  -0.0199   0.5937   1.0000
   8.000   0.4025   0.09648   0.08630  -0.0197   0.5802   1.0000
   8.250   0.4006   0.09826   0.08808  -0.0188   0.5715   1.0000
   8.500   0.4431   0.10291   0.09280  -0.0206   0.5598   1.0000
   8.750   0.4227   0.10291   0.09279  -0.0186   0.5488   1.0000
   9.000   0.4504   0.10704   0.09696  -0.0196   0.5408   1.0000
   9.250   0.4507   0.10832   0.09826  -0.0188   0.5272   1.0000
   9.500   0.4546   0.11101   0.10098  -0.0187   0.5197   1.0000
   9.750   0.4777   0.11430   0.10432  -0.0192   0.5076   1.0000
  10.000   0.4692   0.11597   0.10600  -0.0186   0.4980   1.0000
  10.250   0.5074   0.12124   0.11134  -0.0199   0.4881   1.0000
  10.500   0.4874   0.12152   0.11162  -0.0189   0.4776   1.0000
  10.750   0.5082   0.12568   0.11584  -0.0195   0.4691   1.0000
<< Back to GOE 738 AIRFOIL (goe738-il)

Polar data table (+)

Polar graphs


<< Back to GOE 738 AIRFOIL (goe738-il)